GT2014-26771

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Proceedings of ASME Turbo Expo 2014: Power for Land, Sea and Air
GT2014-26771
June 16-20, 2014, Düsseldorf, Germany
ENDWALL FILM COOLING USING THE STAGGERED COMBUSTOR-TURBINE GAP LEAKAGE FLOW
Yang Zhang, Xin Yuan*
Key Laboratory for Thermal Science and Power Engineering of Ministry of Education
Tsinghua University
Beijing 100084, P.R. China
*Email: yuanxin@mail.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 pressure side gill 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 upstream staggered slot, simulating
the combustor-turbine leakage gap flow. 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. The staggered slots were positioned upstream of the
cascades to simulate the combustor-turbine gap leakage. The
Pressure Sensitive Painting (PSP) technique was used to
detect the film cooling effectiveness distribution on the
endwall surface.
Through the investigation, the following results could be
achieved: 1）the film-cooling effectiveness on the endwall
surface downstream the slot and along the pitchwise direction
increased, with the highest parameter at Z/Pitch=0.6; 2) a
larger cooled region developed towards the suction side as the
blowing ratio increased; 3）the advantage of the staggered
slot was apparent on the endwall surface near the inlet area,
while the coolant film was obviously weakened along the
axial chord at a low blowing ratio. The influence of the
staggered slots could only be detected in the downstream area
of the endwall surface at the higher blowing ratio.
INTRODUCTION
The efficiency of a gas turbine increases with the
increase of the turbine inlet temperature. Modern gas turbines
are designed to operate at high turbine inlet temperature
which is above 1600o
C, placing high thermal loads on turbine
components. With adequate cooling, the lifetime of
components may be extended because of lower thermal
stresses on the turbine. The endwall region is considerably
more difficult to cool than the blade aerofoil surfaces due to
the complex secondary flow structure and strong pressure
gradient in the passage.
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
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

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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 degree at the design condition and ±5 degree
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 were strongly
influenced by the endwall pressure distribution. The leakage
flow and the film-cooling ejection would move towards the
low pressure region where high film-cooling effectiveness
was 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 degree
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.
Some recent research using PSP technology had
enlarged the horizon of endwall film cooling experiment.
Zhang and Moon [20] measured film-cooling effectiveness on
a turbine blade platform surface using the PSP technique.
They compared two kinds of axisymmetric platform profiles,
dolphin nose and shark nose. The results indicated that
although the pattern of the film effectiveness distribution was
quite different, the dolphin nose profile resulted in slightly
higher overall film effectiveness. Krueckels et al. [21]
measured film-cooling effectiveness on a turbine vanes
endwall surface using the liquid crystal measurement
technique. They considered the upstream gap (corresponding
to the gap between the combustor and turbine) and the purge
air exiting this gap in the investigations. The results had
shown that the effect of the purge air needed to be taken into
account for the layout of the cooling scheme. Zhang et al. [22]
compared film-cooling effectiveness of two models of
compound angle shaped holes on a turbine nozzle suction side
surface using PSP technique, one with a compound angle in
co-flow and the other in counter-flow direction to the cooling
supply. The effect of the direction of the compound angle was
found to be less significant compared to the compound angle
effect. Liu et al. [23] used three foreign gases N2 for low
density, CO2 for medium density, and a mixture of SF6 and
Argon for high density to study the effect of coolant density.
The film cooling increased as the density ratio increased. The
reason was that greater coolant to mainstream density ratio
resulted in lower coolant-to-mainstream momentum and
prevented coolant to lift off.
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 influence of complex slot geometry on endwall
film cooling. To help fill this gap, the current paper discusses
the effect of leakage flow from staggered slot on the film
cooling of a nozzle guide vane endwall. The factor of the
blowing ratio is also considered.

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EXPERIMENTAL METHODOLOGY
The film-cooling effectiveness was measured using the
PSP technique. PSP is a photo luminescent material that is
excited by visible light at 450 nm, emitting light that could be
detected by a high spectral sensitivity Charge-coupled Device
(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 and Figure 2.
Figure 1. THE TEST RIG WITH EXCITATION LIGHT
Figure 2. SCHEMATIC OF CASCADE TEST RIG
The image intensity obtained from PSP by the camera is
normalised with reference image intensity Iref taken without
main stream flow. Background noise in the optical setup is
eliminated by subtracting the nitrogen/air injection image
intensities with the image intensity obtained without main
stream flow and light excitation Iblk. The recorded light
intensity ratio can be converted to partial pressure ratio of
oxygen with the parameters obtained in the calibration, as
shown in Eq. (1):
 
 
 2
2
(1)
Oref b lk a ir
ra tio
w b lk O ref
PI I
f f P
I I P
 
  
 
 
Where I represent the intensity obtained at each pixel
and f (Pratio) is the parameter indicating the relationship
between the intensity ratio and the pressure ratio. The film-
cooling effectiveness can be determined by the correlation
between the PSP emitting intensity and the oxygen partial
pressure. Before the test, PSP should be calibrated to obtain
the curves representing the relationship between light
intensities and local partial pressure of oxygen. A PSP coated
copper coupon was used to simulate the experimental surface
with three thermocouples installed underneath the front
surface to measure the surface temperature during the
calibration. The sample coupon was located inside a sealed
chamber where a partial or total vacuum could be created.
The sample was heated by a heater at the rear of the coupon,
which could keep the sample at the desired temperature with
an accuracy of better than 0.5 K. Given the experiment
environment was at a pressure of approximately 1 atm and at
a temperature between 298 and 308 K, the PSP was calibrated
under two temperatures: 294.6 K and 302.5 K and at
pressures from vacuum to 1 atm. The calibration results are
presented in the curves indicating the relationship between
light intensity ratio and pressure ratio (Figure 3).
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
Pressure
I
T=302.5 K
T=294.6 K
Figure 3. CALIBRATION CURVE FOR PSP.
A temperature difference of less than 0.5 K between the
main stream and secondary flow should be guaranteed during
the tests. To obtain the film-cooling effectiveness, both air
and nitrogen are used as coolants. The molecular weight of
nitrogen is nearly the same as that of air, which makes the
change in local oxygen partial pressure at a fixed blowing
ratio possible. By comparing the difference in oxygen partial
pressure between the cases of air and nitrogen injection, the
light intensity can be obtained by using Eq. (2):
   
 
2 2
2
(2)
O Oair mix air mix
w
air O air
P PC C
I
C P

 

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The dimensionless temperature downstream of the
cooling holes could be obtained using the mass and heat
transfer analogy, as defined in Eq. (3):
   
   
   
 
2 2 2 2
2 2 2
2
(3)
O O O Omix air air mix
w
c O O ON air air
P P P PT T
I
T T P P P


 
  
 
The adiabatic wall temperature is reflected by the film-
cooling effectiveness, which is used as a dimensionless
parameter, defined as Eq. (4), for low speed and constant
property flows.
(4)a
w
c c
T T T T
I
T T T T
  
 
 
  
 
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 338 mm wide and 129 mm high. Not considering
the ununiform effect of the outlet flow field of the combustor,
the incidence angle was not 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. 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 plenums located below
the wind tunnel test section before being injected into the
main stream, as shown in the schematic diagrams in Figure 4.
Figure 4. THE TEST SECTION WITH CHANGEABLE SLOT MODULE AND THE
ASSEMBLY DRAWING OF THE TEST SECTION
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 cooling 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
plane. The four rows of fan-shaped holes are inclined 30
degree to the platform surface and held at an angle of 0, 30,
45 and 60 degree to axial direction respectively. The laidback
fan-shaped holes are featured with a lateral expansion of 10
degree from the hole-axis and forward expansion of 10 degree
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 used as leakage flow supply
in this experiment). The coolant supplied to each cavity is
independently controlled by a rotameter dedicated to that
cavity.
Figure 5. THE COOLANT SUPPLY PLENUM
Figure 6. THE STAGGERED SLOT AT THE ENDWALL INLET REGION

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Figure 7. DETAILS OF THE FAN-SHAPED ENDWALL FILM-COOLING HOLES
The staggered slots are positioned upstream of the
cascades to simulate the combustor-turbine gap leakage. The
slots are divided into two parts, PS (Pressure side) slot and SS
(Suction Side) slot. The PS slot was located upstream the
leading edge with a distance of 57.6% axial chord length and
the inclined angle was 60 degree to the anti-mainstream flow
direction. The distance of the SS slot was 57.6% axial chord
length with an inclined angle of 30 degree to the mainstream
flow direction. The two direction slots are connected directly
by sharp turning, which constructed the whole connection
part as “X” shape. The anti-mainstream slot, SS one, begins
upstream the blade leading edge and ends at 57.6%,
connected with the PS slot which covering the rest 49.1%
pitch. Then the SS slot began as a periodic structure. The
pitch-wise length of the SS and PS slots are 68mm and 42mm
respectively. The width of the staggered slot is 2mm and the
thickness of the endwall surface was 4mm. The PS and SS
slot share the same coolant plenum and the same coolant inlet.
To form a uniform inlet flow for the coolant supply, three or
four inlet tubes are introduced to the plenums. Several rows
of small staggered guide plates are located in the plenum to
separate the flow into small groups and then some main
vortex could be broken. The detail geometry of the guide
plates in the plenum are contained in Fig.6. The original
design of the slot is a simple inclined leakage with a 30
degree angle to the endwall surface. The inlet area of original
slot and the staggered slots is same.
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.
Figure 8. STAGGERED SLOTS CONFIGURATION AT ENDWALL INLET
REGION (WITH INNER STUCTURE OF COOLANT SUPPLY CHANNEL)
Table 1 Discrete film hole location and orientation
Hole
Name
Position
X/Cax
Number D
(mm)
Radial/
Compound
Angle to
Surface
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 Slot Film Cooling Endwall Film Cooling M
Air
(L/min)
N2
(L/min)
Air
(L/min)
N2
(L/min)
Film Cooling With Original Slot Injection
1 51 53 72 75 0.7
2 77 80 102 106 1.0
3 102 106 133 138 1.3
Film Cooling With Staggered Slot Injection
4 51 53 72 75 0.7
5 77 80 102 106 1.0
6 102 106 133 138 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

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Inlet and exit angles 0 & 72 °
Inlet/Outlet Mach number 0.1 & 0.3
Inlet mainstream velocity 35 m/s
Mainstream flow temperature 305.5 K
Injection flow temperature 305.0 K

Figure 9. TEST VANE WITH PRESSURE MEASUREMENT HOLES ON TOP
(MANUFACTURED WITH FAST PROTOTYPE)

Figure 10. FIVE-HOLE PROBE AND ITS TWO DIMENSIONAL
DISPLACEMENT WORKTABLE
Pitch
Span
0 0.5 1 1.5 2
0
0.25
0.5
0.75
1 0.965
0.97
0.975
0.98
0.985
0.99
0.995
1

Figure 11. SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITHOUT FILM COOLING

Though the phantom cooling effects have been
investigated through PSP measurements, the aerodynamic
performance is also an important factor for the NGVs. The
five-hole probe total pressure measurement section is added
to the cascades to probe the total pressure loss. A two
dimensional displacement worktable is used to make the
measurement on a single plain possible. The vanes with film
cooling holes are replaced by the pressure test vanes to
measure the static pressure distribution on airfoil surface. The
diameter of static pressure hole is 1mm with individual cavity
made through Stereolithography (SLA) fast prototype
technology. The static pressure is tested on the vane top where
some pressure sensors are located.

The pressure coefficient at outlet was measured by the
five-hole probe and the airfoil surface pressure distribution
was measured by the Kulite sensor. The aerodynamic results
show the periodicity of the cascades and structure of
secondary flow. The simple case without any film cooling
injection was measured as the baseline condition. The test
range covers two pitches and the near endwall area where
non-dimensional span height is between 0.75 and 1.0, as
shown in the black rectangular. The test time is 5 seconds at
every point for the five-hole probe, and the measurement
frequency is 100HZ. The airfoil surface pressure was probed
at 23 points on PS and SS respectively.
0 0.2 0.4 0.6 0.8
0.96
0.97
0.98
0.99
1
X/Cax
PressureCoefficient
VaneA PS
VaneB PS
VaneA SS
VaneB SS

Figure 12. NON-DIMENSIONAL PRESSURE COEFFICIENT DISTRIBUTIONS
ON AIRFOILS

The figure shows the pressure coefficient distribution at
the outlet plain with a distance of 0.5 axial chord length from
trailing edge. The high pressure loss area could be captured in
the map which demonstrates the position of wake and corner
vortex core. The low pressure coefficient area along the
spanwise direction shows the position and strength of the
trailing edge wake. The shape of wakes and vortex show that
the cascades have a reasonable vane to vane periodicity
quality. The secondary flow can be captured in the map
obviously According to the airfoil surface pressure coefficient
distribution shown in Figure 12, vane-to-vane comparisons of
the experimental measurement points demonstrate that a good
level of periodicity, too.
Pitch
Span
0 0.5 1 1.5 2
0.75
0.875
1
0.97
0.98
0.99
1

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PRESSURE LOSS WITH ENDWALL FILM COOLING, M=0.7
Pitch
Span
0 0.5 1 1.5 2
0.75
0.875
1
0.97
0.98
0.99
1

Pitch
Span
0 0.5 1 1.5 2
0.75
0.875
1
0.97
0.98
0.99
1

0 0.5 1 1.5 2
0.97
0.98
0.99
1
Pitch
MeanPressureCoefficient
M=0.7 M=1.0 M=1.3

Figure 16. COMPARISON OF PRESSURE LOSS NEAR ENDWALL SURFACE
WITH DIFFERENT FILM COOLING BLOWING RATIOS

The group of maps shows the pressure coefficient
distributions at outlet with different endwall cooling blowing
ratios. The distributions demonstrate that the total pressure
loss is hardly influenced by the endwall film cooling blowing
ratio. However, slight effects can be captured near the
endwall surface where the non-dimensional span is between
0.75 and 1.0. As the blowing ratio increases, the total pressure
loss near endwall surface is relatively higher. Simultaneously,
the strength of the corner vortex and wake is not weakened by
the endwall film cooling ejection. The Figure 16 shows the
spanwise averaged pressure coefficient distribution along
pitch direction. The curves demonstrate that with the blowing
ratio increasing, the total pressure loss near endwall surface is
slightly increasing. The huge amount of coolant ejection
strongly mixes with the main flow, which causes the total
pressure loss increases in this area.
RESULTS AND DISCUSSION
Though the cascade is 2-d linear, 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 and stagnation line to be hardly
cooled, while the different flow direction near the suction side
avoids this harmful interaction. According to the contours,
without the staggered slot leakage geometry the strong
pressure gradient will strongly bring the coolant to the suction
side, leaving an apparent uncooled area near the pressure side,
especially near the stagnation line.
In the current study, five coolant cavities are used for the
slot leakage gap flow 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 original and staggered slot
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 blowing ratios are shown in
Figures 17–22, 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 17–19 show the film-cooling
effectiveness distribution on the endwall surface with original
and staggered slot leakage flow 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 and stagnation line 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 near but not 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 holes outlet geometry on the flow field being
avoided in the experiment.
The left subplot in Figures 17–19 shows the film-cooling
effectiveness distributions on the endwall with original slot
injection film cooling when the blowing ratio on the endwall

8
is controlled to be M=0.7, M=1.0 and M=1.3 respectively.
The right subplot in Figures 17–19 shows the film-cooling
effectiveness distributions on the endwall with staggered slot
injection. When the blowing ratio is M=0.7, the cooled area
of staggered slot is slightly larger in the middle pitch area
downstream the slot, while the cooled area is restricted
outside 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 and
near stagnation line region for original slot case, while this
area is covered by staggered slot injection coolant at the
blowing ratio of M=1.0. The right subplot in Figure 19 shows
the film-cooling effectiveness distributions on the endwall
surface with staggered slot injection 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 and
stagnation line is obviously larger than the baseline case with
original slot.
Although valuable insight can be obtained from the
distribution maps (Figs. 17–19), the spanwise averaged plots
(Figs. 20–22) offer additional insight and provide clear
comparisons for large amounts of data. The effectiveness is
averaged from the SS to the PS (Figs. 17–19) 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 and the slot
location. Figures 20–22 indicate that, with the staggered slot
injection, the endwall film-cooling effectiveness increases in
the downstream area of the slot. The locally largest film-
cooling effectiveness difference appears at Cax=-0.16, where
the leakage gap injection appears. The average is significantly
higher because the coolant injected from the staggered slot
covers the endwall sufficiently, especially at middle pitch
region where the local pressure is relatively high. The
staggered slot injection effect is clearly seen on the upstream
half (axial chord position between 0 and 0.4) of the endwall.
original M=0.7 i= 0degZ/ZP
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
staggered M=0.7 i= 0degZ/ZP
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
Figure 17 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 0.7, WITH ORIGINAL AND STAGGERED SLOT
INJECTION)
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
INJECTION)
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
X/Cax
1
2
3
0 0.2 0.4 0.6 0.8 1
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1 0 0.1 0.2 0.3 0.4
INJECTION)
-0.2 0 0.2 0.4 0.6 0.8 1 1.2
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 original
i= 0deg M=0.7 staggered
Figure 20 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 0.7, WITH ORIGINAL AND
STAGGERED SLOT INJECTION)

9
-0.2 0 0.2 0.4 0.6 0.8 1 1.2
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
endwall
X/Cax
-0.2 0 0.2 0.4 0.6 0.8 1 1.2
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
endwall
X/Cax
With the staggered slot film cooling, the momentum of
the coolant is high enough to cover the endwall surface
though the better cooled area is limited to a small region near
the middle pitch region. A higher blowing ratio leads to more
coolant being injected from the slot near the PS and
stagnation line such that the better cooled area becomes wider
on the endwall (represented by red curves in Figs. 23–25, the
slot is located at cascades inlet). 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
staggered slot is enough to cool the high pressure area. With a
high blowing ratio the injection could cover the area near the
stagnation line and middle pitch region even overcool this
area with endwall cooling holes nearby, while the PS 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 PS corner region is satisfied.
Figures 17–19 and Figures 20–22 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 17 and 20, the main phenomenon
with original and staggered slot is that, at a low blowing ratio,
the injection area is both small. The coolant could hardly
inject from the slot near the pressure side (pitch between 0.8
and 0.1, axial chord between -0.12 and 0, the slot is located at
the cascades inlet with Cax=-0.12). This phenomenon shows
that the low blowing ratio could not overcome the high
pressure factor in this area, that the corner vortex and pressure
gradient weaken the film cooling near PS and stagnation line.
This condition is obviously changed at higher blowing ratios
as shown in Figures 19 and 22. When the blowing ratio is
M=1.3, the coolant could inject from the staggered slot, while
the film-cooling effectiveness is high not only near the middle
pitch region but also in the downstream area, especially near
the pressure side. This indicates that the staggered slot film
cooling is sensitive to the blowing ratio. Figure 19 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 staggered slot will move
along the passage vortex and then enter at the main passage
film cooled traces which causes the film-cooling effectiveness
in the upstream part of endwall to be obviously higher,
especially at the middle pitch part, as shown in Figure 19.
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4
Figure 23 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE INJECTION POSITION (THE BLOWING RATIO IS 0.7, WITH
ORIGINAL AND STAGGERED SLOT INJECTION)
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4

10
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4
X/C ax
0.1 0.2 0.3
0.6
0.7
0.8
0.9
1
0 0.1 0.2 0.3 0.4
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
Figure 26 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL AT STREAMWISE LOCATION 1 (THE BLOWING RATIO IS 0.7,
WITH ORIGINAL AND STAGGERED SLOT INJECTION, Cax=-0.12)
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
WITH ORIGINAL AND STAGGERED SLOT INJECTION, Cax=-0.12)
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 26 and 27 compare the local film-cooling
effectiveness distribution at streamwise location 1 with
different blowing ratios (M=0.7 and M=1.3). The position of
the computing area is indicated by the PS to SS white line
along the pitch direction in Figures 17–19 (Line 1 on the left
side). With the staggered slot injection, the local film-cooling
effectiveness apparently improves at middle pitch area, as
shown in Figures 26 and 27 where the curve representing the
staggered slot cooling condition is apparently higher at
upstream area. Meanwhile, the film-cooling effectiveness in
the main passage and near the SS is also obviously changed.
The well protected region is enlarged 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 staggered slot film cooling quickly, which makes the film-
cooling effectiveness of downstream part 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.
Figures 28 and 29 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.

11
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
WITH ORIGINAL AND STAGGERED SLOT INJECTION, Cax=0.13)
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
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
0 0.05 0.1
0
0.2
0.4
near SS
0.9 0.95 1
0
0.1
0.2
0.3
near PS
Figures 30 and 31 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 (Figure 30), no apparent unprotected
area could be found at the PS corner region (pitch is between
0.9 and 1.0), while the influence of the staggered slot film
cooling could not be probed in this area, the downstream part
of the endwall. This indicates that the effects of the staggered
slot film cooling are not apparent in the downstream corner
region of endwall surface when the blowing ratio is relatively
low. Figure 31 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, for the staggered slot case.
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 staggered slot film cooling on the endwall. The film-
cooling effectiveness curve representing the case of staggered
slot film cooling is obviously above the curves representing
the baseline case in the mid-pitch region (pitch is between 0.4
and 0.7 ) as shown in Figure 31. As the blowing ratio
increases, the influence of staggered slot injection is
apparently weakened by the secondary flow. The higher
momentum of the coolant injection flow could not effectively
overcome the mixing trend of the passage vortex and then
partially form a high film-cooling effectiveness area at the
mid-pitch.

12
CONCLUSIONS
In general, staggered slot injection apparently affects the
coolant distribution on the endwall surface, especially for the
near inlet area. The results show that with an increasing
blowing ratio, the film-cooling effectiveness increases on the
endwall surface, especially near the leakage gap 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 leakage flow film-cooling
on the endwall surface.
As the blowing ratio varies from M=0.7, to M=1.3, the
influence of staggered slot leakage flow on the endwall film
cooling increases near inlet area. Simultaneously, the area of
influence will move towards mid-pitch and the suction side.
In conclusion: 1 ） the film-cooling effectiveness on the
endwall surface downstream the slot and along the pitchwise
direction increased, with the highest parameter at Z/Pitch=0.6;
2) a larger cooled region developed towards the suction side
as the blowing ratio increased; 3 ） the advantage of the
staggered slot was apparent on the endwall surface near the
inlet area, while the coolant film was obviously weakened
along the axial chord at a low blowing ratio. The influence of
the staggered slots could hardly be detected on downstream
area of the enwall surface even at higher blowing ratio.
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
TE =Trailing Edge
V =velocity, m/s
X , Z =Cartesian coordinate system
 =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
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”.
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[4] Wright, L.M., Blake, S., Han, J.C., 2006. “Effectiveness
Distributions on Turbine Blade Cascade Platforms
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Th., 2009. “Experimental and Numerical Study of the
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[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
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[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.

13
[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,
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[15] Cardwell, N.D., Sundaram, N., Thole, K.A., 2006.
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Roughness on Endwall Film-Cooling”. Journal of
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[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.
[20] Zhang, L., Moon, H.K., 2011. “Comparison of Two
Axisymmetric Profiles on Blade Platform Film Cooling”.
In ASME Turbo Expo 2011, Vancouver, British
Columbia, Canada, ASME Paper No. GT2011-45102.
[21] Krueckels, J., Colban, W., Gritsch, M., Schnieder, M.,
2011. “Validation of a First Vane Platform Cooling
Design”. In ASME Turbo Expo 2011, Vancouver,
British Columbia, Canada, ASME Paper No. GT2011-
45252.
[22] Zhang, L., Yin, J., Moon, H.K., 2012. “The Effect of
Compound Angle on Nozzle Suction Side Film Cooling”.
In ASME Turbo Expo 2012, Copenhagen, Denmark,
ASME Paper No. GT2012-68357.
[23] Liu, K., Yang, S., Han, J., 2012. “Influence of Coolant
Density on Turbine Blade Film-Cooling With
Compound-Angle Shaped Holes”. In ASME Turbo Expo
2012, Copenhagen, Denmark, ASME Paper No.
GT2012-69117.

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