1. 1
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: yangzhang09@mails.tsinghua.edu.cn
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

2. 2
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

3. 3
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 air
ratio
blk O ref
PI I
f 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 1.2
0.1
0.3
0.5
0.7
0.9
1.1
1.2
P/Pref
I/Iref
T=302.5 K
T=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

4. 4
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

5. 5
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
Position
X/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

6. 6
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= 0degZ/ZP
X/Cax
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= 0degZ/ZP
X/Cax
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)
Baseline M=1.0 i= 0degZ/ZP
X/Cax
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=1.0 i= 0degZ/ZP
X/Cax
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 10 FILM COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 1.0, WITH AND WITHOUT PRESSURE SIDE
INJECTION)

7. 7
Baseline M=1.3 i= 0degZ/ZP
X/Cax
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=1.3 i= 0degZ/ZP
X/Cax
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 11 FILM COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 1.3, WITH AND WITHOUT PRESSURE SIDE
INJECTION)
-0.2 0 0.2 0.4 0.6 0.8 1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
endwall
X/Cax
i= 0deg M=0.7 Baseline
i= 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 1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
endwall
X/Cax
i= 0deg M=1.0 Baseline
i= 0deg M=1.0 PsCooling
Figure 13 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 1.0, WITH AND WITHOUT
PRESSURE SIDE INJECTION)
-0.2 0 0.2 0.4 0.6 0.8 1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
endwall
X/Cax
i= 0deg M=1.3 Baseline
i= 0deg M=1.3 PsCooling
Figure 14 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 1.3, WITH AND WITHOUT
PRESSURE SIDE INJECTION)
Baseline M=0.7 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 0.2 0.4 0.6 0.8
PsCooling M=0.7 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 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)
Baseline M=1.0 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 0.2 0.4 0.6 0.8
PsCooling M=1.0 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 0.2 0.4 0.6 0.8
Figure 16 FILM COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE INJECTION POSITION (THE BLOWING RATIO IS 1.0, WITH
AND WITHOUT PRESSURE SIDE INJECTION)
Baseline M=1.3 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 0.2 0.4 0.6 0.8
PsCooling M=1.3 i= 0degZ/ZP
X/C ax
1
2
0.1 0.2 0.3 0.4
0.5
0.6
0.7
0.8
0 0.2 0.4 0.6 0.8
Figure 17 FILM COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE INJECTION POSITION (THE BLOWING RATIO IS 1.3, WITH
AND WITHOUT PRESSURE SIDE INJECTION)

8. 8
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation1
Z/Pitch
i= 0deg M=0.7 Baseline
i= 0deg M=0.7 PsCooling
0 0.05 0.1
0.2
0.3
0.4
near SS
0.9 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 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation1
Z/Pitch
i= 0deg M=1.3 Baseline
i= 0deg M=1.3 PsCooling
0 0.05 0.1
0.2
0.3
0.4
near SS
0.9 0.95 1
0.1
0.2
0.3
0.4
near PS
Figure 19 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 1 (THE BLOWING RATIO IS 1.3,
WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.18)
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

9. 9
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 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation2
Z/Pitch
i= 0deg M=0.7 Baseline
i= 0deg M=0.7 PsCooling
0 0.05 0.1
0.2
0.3
0.4
near SS
0.9 0.95 1
0.1
0.2
0.3
0.4
near PS
Figure 20 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 2 (THE BLOWING RATIO IS 0.7,
WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.3)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation2
Z/Pitch
i= 0deg M=1.3 Baseline
i= 0deg M=1.3 PsCooling
0 0.05 0.1
0.2
0.3
0.4
near SS
0.9 0.95 1
0.1
0.2
0.3
0.4
near PS
Figure 21 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 2 (THE BLOWING RATIO IS 1.3,
WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.3)
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 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation3
Z/Pitch
i= 0deg M=0.7 Baseline
i= 0deg M=0.7 PsCooling
0 0.05 0.1
0.4
0.5
0.6
near SS
0.9 0.95 1
0.2
0.4
0.6
near PS
Figure 22 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 3 (THE BLOWING RATIO IS 0.7,
WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.78)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation3
Z/Pitch
i= 0deg M=1.3 Baseline
i= 0deg M=1.3 PsCooling
0 0.05 0.1
0.4
0.5
0.6
near SS
0.9 0.95 1
0.2
0.4
0.6
near PS
Figure 23 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 3 (THE BLOWING RATIO IS 1.3,
WITH AND WITHOUT PRESSURE SIDE INJECTION, X/Cax=0.78)
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

10. 10
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

11. 11
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.
[19] Piggush, J.D., Simon, T.W., 2007. “Heat Transfer
Measurements in a First-Stage Nozzle Cascade Having
Endwall Contouring: Misalignment and Leakage
Studies”. Journal of Turbomachinery, 129, 782–790.