1. 1
Proceedings of ASME Turbo Expo 2014: Power for Land, Sea and Air
GT2014-26781
June 16-20, 2014, Düsseldorf, Germany
EXPERIMENTAL INVESTIGATION OF TURBINE PHANTOM COOLING ON ENDWALL WITH TRAILING EDGE
DISCHARGE 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
The film cooling ejection on High Pressure (Hp) turbine
component surface is strongly affected by the complex flow
structure in the nozzle guide vane or rotor blade passages. The
action of secondary flow in the main passage could dominate
the film cooling effectiveness distribution on the component
surfaces. The film cooling ejections from endwall and airfoil
trailing edge are mixed by the secondary flow. Considering a
small part of the coolant ejection from trailing edge discharge
flow will move from the airfoil trailing edge pressure side to
endwall downstream and then cover some area, the
interaction between the coolants injected from endwall and
airfoil trailing edge is worth investigating. Though the
temperature of coolant discharge flow from trailing edge
increases after the mixing process in the internal cooling
procedure, the ejections moving from airfoil to endwall still
have the potential of second order cooling. This part of the
coolant is called “Phantom cooling flow” in the paper. A
typical scale-up model of Hp turbine NGV is used in the
experiment to investigate the cooling performance of ejection
from trailing edge. Instead of the airfoil trailing edge platform
itself, the film cooling effectiveness is measured on the
downstream part of the endwall. This paper is focused on the
trailing edge discharge flow with compound angle effects and
the coolant from discharge holes moving from trailing edge to
endwall surface. The coolant flow is injected from the straight
discharge holes with a compound angle of 15deg and 45deg
respectively. The film cooling holes on the endwall are used
simultaneously to investigate the combined effects. The
blowing ratio and different configurations of compound angle
holes are selected to be the changing parameters in the paper.
The experiment is completed with the blowing ratio changing
from M=0.7 to M=1.3 and the compound angle is introduced
to the entire row of trailing edge discharge holes (full span),
with inlet Reynolds numbers of Re=3.5×105
and an inlet
Mach number of Ma=0.1
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] utilised the back-facing step to simulate the
discontinuity of the nozzle inlet to the combustor exit cone.
Nitrogen gas was used to simulate the 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] employed 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

2. 2
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
ejection 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.
In the trailing edge film cooling research field, Murata et
al. [8] compared four different cutback geometries at trailing
edge: two smooth cutback surfaces with constant-width and
converging lands and two roughened cutback surfaces with
transverse ribs and spherical dimples. The dimple surface was
proven to be a favourable cutback-surface geometry, because
it gave enhanced heat transfer without deterioration of the
high film cooling effectiveness. Benson et al. [9] used
magnetic resonance imaging experiments to provide the
three-dimensional mean concentration and three component
mean velocity field for typical trailing edge film-cooling
cutback geometry. They designed new trailing edge
geometries to modify the large scale mean flow structures
responsible for surface effectiveness degradation. Martini et
al. [10] conducted the experiments covering a broad variety of
internal cooling designs with trailing edge discharge. The
results clearly demonstrated the strong influence of the
internal cooling design and the relatively thick pressure side
structure on film cooling performance downstream of the
trailing edge ejection slot. Dannhauer [11] used a
combination of IR-thermography and thermocouples to
investigate two different trailing edge geometries with coolant
ejection. The first configuration was pressure side cutback
while the other was a row of cylindrical holes. Telisinghe et al
[12] studied the aero performance differences between a
conventional turbine blade trailing edge and a trailing edge
with a sharp cut-back. With regard to the discharge
coefficient, the new configuration had a larger discharge
coefficient compared with the conventional configuration.
As for blade endwall/platform film-cooling research,
Yang et al. [13] 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 has been reported. The research by Kost and
Mullaert [14] 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. [15]. 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 moves towards the
suction side and then forms a triangular cooled area.
The effects of rotation on platform film cooling has been
investigated by Suryanarayanan et al. [16], 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 regards investigations into
combustor-turbine leakage flow, Thole’s group made
significant contributions. With detailed investigations, the
influence of slot shape and position as well as width has been
analysed in a series of studies [17–19].
Oke and Simon [20] 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. [21] 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 [22] 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 visualisation. Piggush and Simon [23] 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 change the ejection
flow trace. Few studies, however, have considered the
combined effect of trailing edge discharge flow film cooling
and endwall film cooling. To help fill this gap, the current
paper discusses the effect of trailing edge discharge flow
phantom cooling Nozzle Guide Vane (NGV) endwall. The
factor of the blowing ratio and compound angle are
considered.

3. 3
EXPERIMENTAL METHODOLOGY
The film-cooling effectiveness is 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 Charge-coupled Device
(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 image intensity obtained from
the PSP by the camera is normalised with a reference image
intensity ( refI ) taken without mainstream flow. Background
noise in the optical setup is eliminated by subtracting the
nitrogen/air ejection image intensities with the image
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. 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
1. A temperature difference of less than 0.5K between the
mainstream and the secondary flow should be guaranteed
during the tests. To obtain film-cooling effectiveness, both air
and nitrogen are used as the coolant. The molecular weight of
nitrogen is almost the same as that of air, which makes the
density ratio close to 1.0. By comparing the difference in
oxygen partial pressure between the air and nitrogen ejection
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 had a cross
section 338 mm wide and 129 mm high. Not considering the
ununiformed 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, which in turn changed 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. [24]
Figure 2. THE TEST SECTION WITH MOVABLE TRAILING EDGE MODULE
AND THE ASSEMBLY DRAWING OF THE TEST SECTION
To obtain easier assemble method, the trailing edge
module was movable and could be changed with the main
blade part untouched. The trailing edge module had its own
coolant supply plenum which could make the trailing edge
discharge flow blowing ratio controlled independently.
During the test the trailing edge module was kept at the same
relative position with the Charge-coupled Device (CCD) so
that the film cooling effectiveness distribution could be
compared accurately. For coolant air supply, compressed air
was delivered to a plenum located below the wind tunnel test
section before being injected into the mainstream, as shown in
the schematic diagrams in Figure 2. Four vanes and three

4. 4
passages are included in the cascade. Two of the vanes have
surface film cooling and only one passage has endwall film
cooling.
Figure 3. SCHEMATIC OF CASCADE TEST RIG
Figure 4. THE TEST CASCADES WITH PSP ON ENDWALL SURFACE
Figure 5. DETAILS OF TRAILING EDGE DISCHARGE HOLES
Figure 6. THETRAILING EDGE DISCHARGE HOLES ON THE PRESSURE SIDE
AND THE INNER PIN FIN STRUCTURE
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 clearly, all of the film-cooling holes
were 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 were arranged on the gill region to form full
covered coolant film. Figures 3–8 show the test cascades
configurations and the geometric parameters of the blade.
Four rows of compound angle laidback fan-shaped holes
were 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 was located upstream of
the leading edge plane. The following three rows were 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 were 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 were 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 was 1 mm, and the
expansion starts at 3D. Four coolant cavities were used for the
four rows of holes respectively, as shown in Figure 7. (The
extra coolant plenum chamber was designed to simulate the
purge flow which was used as leakage flow supply in this
experiment). The coolant supplied to each cavity was
independently controlled by a rotameter dedicated to that
cavity.
Figure 7. DETAILS OF THE FAN-SHAPED ENDWALL FILM-COOLING HOLES
One row of trailing edge discharge hole was arranged on
the trailing edge surface at axial location of 77 mm (TE, 20

5. 5
holes) as shown in figure 7. The trailing edge discharge holes
were located on the pressure side trailing edge region so that
the film-cooling effectiveness of the airfoil surface was
difficult to access due to camera position limitation when the
endwall film-cooling effectiveness was investigated. The
trailing edge discharge flow was injected from the coolant
plenum with inner pin fin structure, which consisted of five
rows of pin fins and two connected cavities with a width of
4mm respectively. The spanwise height of the trailing edge
discharge hole was 4mm and the axial length was 9 mm. The
discharge holes were connected to the coolant plenum by
cylindrical channel with a diameter of 2mm and a length of
15mm. The discharge holes were located along the spanwise
direction; therefore, the distance between the last discharge
hole and enwall surface should be the same (3mm). Due to
the strong secondary flow in the trailing edge wake, it was
difficult to control the local blowing ratios for every hole with
one common coolant plenum chamber. One coolant cavity
was used for the trailing edge discharge flow, as shown in
Figure 8. (The coolant plenum had two inlet tubes as shown
in Figure 8. However, only the bottom one was used in the
test, while the top one was blocked in this investigation). The
coolant supplied to the cavity was controlled by a rotameter.
As shown in Figure 8, the trailing edge discharge holes were
inclined 15° and 45° to the endwall surface respectively to
study the effect of compound angle on phantom cooling.
Figure 8. TRAILING EDGE DISCHARGE HOLE CONFIGURATION (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
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 TE Film Cooling Endwall Film Cooling M
Air
(L/min)
N2
(L/min)
Air
(L/min)
N2
(L/min)
Film Cooling With 15deg TE Ejection
1 52 54 72 75 0.7
2 75 78 102 106 1.0
3 97 100 133 138 1.3
Film Cooling With 45deg TE Ejection
4 52 54 72 75 0.7
5 75 78 102 106 1.0
6 97 100 133 138 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
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
Ejection flow temperature 305.0 K
Figure 9. SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITHOUT FILM COOLING
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

6. 6
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
ejection 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, Figure 9. 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.
Figure 10. 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 10, Vane-to-vane comparisons of
the experimental measurement points demonstrate that a good
level of periodicity, too.
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 wakes, means
that the endwall near the trailing edge hardly cooled, while
the different flow direction in the main passage avoids this
harmful interaction. According to the contours, without the
trailing edge ejection the pressure gradient will strongly keep
the coolant outside the inner flow of the wakes, leaving an
apparent uncooled area near the trailing edge, especially near
the discharge holes. Five coolant cavities are used for the
trailing edge discharge 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 different trailing
edge film cooling could be compared precisely. In this study,
three different blowing ratios are 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 11–16, 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 11–13 show the film-
cooling effectiveness distribution on the endwall surface with
15deg and 45deg trailing edge 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 trailing edge is still apparent (shown
with the red curve). This phenomenon represents the strong
pressure gradient in the turbine cascades, dominating the
moving direction of the coolant traces. The momentum of the
coolant ejection is not strong enough to take the cool air into
the high pressure area near the corner region (axial chord
position between 1 and 1.1). 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.
Figure 11 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 0.7, WITH 15DEG AND 45DEG TRAILING EDGE
EJECTION)
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
trailing15 M=0.7 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6
trailing45 M=0.7 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6

7. 7
Figure 12 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 1.0, WITH 15DEG AND 45DEG TRAILING EDGE
EJECTION)
Figure 13 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE BLOWING RATIO IS 1.3, WITH 15DEG AND 45DEG TRAILING EDGE
EJECTION)
The left subplot in Figures 11–13 shows the film-cooling
effectiveness distributions on the endwall with 15deg trailing
edge discharge hole 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 11–13 shows the
film-cooling effectiveness distributions on the endwall with
45deg trailing edge discharge hole film cooling. When the
blowing ratio is M=0.7, the cooled area is slightly larger in
the red curves of the contour for the 45deg case, while the
cooled area is restricted to the Trailing Edge (TE) corner
region (red lines). At higher blowing ratios, near the TE
corner region, the cooled area is relatively larger. When the
blowing ratio is M=0.7, an apparent unprotected area can be
found at the edge of TE corner region for 15deg case, while
this area is covered by the TE ejection coolant at the blowing
ratio of M=1.0. The right subplot in Figure 13 shows the film-
cooling effectiveness distributions in the corner region with
45deg trailing edge discharge hole 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 TE is obviously larger than the 15deg
case.
Although valuable insight can be obtained from the
distribution maps (Figs. 11–13), the spanwise averaged plots
(Figs. 14–16) offer additional insight and provide clear
comparisons for large amounts of data. The effectiveness is
averaged from the SS to the PS (Figs. 11–13) of the passage
in the axial chord direction. The data outside the airfoil were
deleted from the averaged results. The peaks in the plot
correspond to the film-cooling holes’ location. Figures 14–16
indicate that, with the 45deg TE film cooling ejection, the end
wall film-cooling effectiveness increases in the downstream
area. Locally, the largest film-cooling effectiveness difference
near the trailing edge appears at Cax=1.02. The average for
45deg case is significantly higher because the coolant injected
from the trailing edge covers the endwall sufficiently,
especially near the corner region where the local pressure is
relatively high in the wake. The 45deg trailing edge ejection
effect is clearly seen near the trailing edge part (axial chord
position between 1 and 1.05) of the endwall.
Figure 14 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 0.7, WITH 15DEG AND 45DEG
TRAILING EDGE EJECTION)
Figure 15 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 1.0, WITH 15DEG AND 45DEG
TRAILING EDGE EJECTION)
trailing15 M=1.0 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6
trailing45 M=1.0 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6
trailing15 M=1.3 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6
trailing45 M=1.3 i= 0degZ/ZP
X/Cax
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 0.2 0.4 0.6
-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 trailing15
i= 0deg M=0.7 trailing45
-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=1.0 trailing15
i= 0deg M=1.0 trailing45

8. 8
Figure 16 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL (THE BLOWING RATIO IS 1.3, WITH 15DEG AND 45DEG
TRAILING EDGE EJECTION)
Figure 17 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
TRAILING EDGE EJECTION POSITION (THE BLOWING RATIO IS 0.7, WITH
15DEG AND 45DEG TRAILING EDGE EJECTION)
Figure 18 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
TRAILING EDGE EJECTION POSITION (THE BLOWING RATIO IS 1.0, WITH
15DEG AND 45DEG TRAILING EDGE EJECTION)
Figure 19 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
TRAILING EDGE EJECTION POSITION (THE BLOWING RATIO IS 1.3, WITH
15DEG AND 45DEG TRAILING EDGE EJECTION)
Figure 20 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 1 (THE BLOWING RATIO IS 0.7, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 1 IS IN FIGURES 17-19)
Figure 21 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 1 (THE BLOWING RATIO IS 1.3, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 1 IS IN FIGURES 17-19, THE
MEANING OF TRAILING, FAR, LEFT AND RIGHT IS IN FIGURE 15)
With the trailing edge 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
trailing edge. A higher blowing ratio leads to more coolant
being injected from the trailing edge cooling holes near the
corner region such that the cooled peak area becomes wider
on the endwall (represented by red curves in Figs. 17–19,
15deg and 45deg TE discharge holes are located near the
endwall). As the coolant leaves the cooling holes, the trace of
the ejection flow is led by the wake vortex developing
downstream the trailing edge PS and SS. The vortex is strong
at the junction region, which causes the boundary of coolant
to move along the wake central line and towards the outlet.
The film-cooling effectiveness distributions indicate that the
cooling performance of the TE discharge holes is enough to
cool the endwall surface covered by wakes. With a high
blowing ratio the ejection could cover the area near the
trailing edge, while the near TE 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 trailing edge corner
-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=1.3 trailing15
i= 0deg M=1.3 trailing45
trailing15 M=0.7 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
trailing45 M=0.7 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
trailing15 M=1.0 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
trailing45 M=1.0 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
trailing15 M=1.3 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
trailing45 M=1.3 i= 0degZ/ZP
X/C ax
1
2
3
0.9 1 1.1
-0.2
-0.1
0
0 0.2 0.4 0.6
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
TRAILING
FAR
Line1
Z/Length
i= 0deg M=0.7 trailing15
i= 0deg M=0.7 trailing45
0 0.05 0.1
0
0.1
0.2
Trailing
0.9 0.95 1
0
0.1
0.2
0.3
Far
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
TRAILING
FAR
Line1
Z/Length
i= 0deg M=1.3 trailing15
i= 0deg M=1.3 trailing45
0 0.05 0.1
0
0.1
0.2
Trailing
0.9 0.95 1
0
0.1
0.2
0.3
Far

9. 9
region is satisfied even for 15deg case.
Figures 11–13 and Figures 14–16 indicate the difference
in film-cooling effectiveness distribution along the wake
central line and along the vertical direction versus central line.
When the blowing ratio is M=0.7, as shown in Figures 11 and
14, the main difference with 15eg and 45deg TE discharge
film cooling is that, at a low blowing ratio, the ejection area
of 15deg case is small. The coolant could hardly inject from
the cooling holes on the trailing edge (pitch between -0.05
and -0.06, axial chord between 0.98 and 1.0, one discharge
hole is located near the endwall surface with a distance of
3mm). This phenomenon shows that the low blowing ratio
ejection could not overcome the high pressure factor in the
wake, that the wake vortex and secondary flow weaken the
trailing edge film cooling, even for the 45deg case. This
condition is obviously changed at higher blowing ratios as
shown in Figures 13 and 16. When the blowing ratio is
M=1.3, the coolant could inject from the trailing edge
discharge holes near the endwall for the 45deg case, while the
film-cooling effectiveness is high not only near the trailing
edge but also in the downstream area, especially in the wake
trace. Figure 13 shows the trend that the behaviour of the
ejection flow could apparently influence the downstream
effectiveness distribution at high blowing ratios. The coolant
from the trailing edge cooling holes will move along the wake
vortex and then arrive at the downstream part which causes
the film-cooling effectiveness at the outlet to be higher,
especially in the wake region, 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
visualisation 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 20 and 21 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 TRAILING to FAR white line along the
wake central line direction in Figures 11–13. With the 45deg
trailing edge ejection, the local film-cooling effectiveness
apparently improves near the trailing edge, as shown in
Figures 20 and 21 where the curve representing the 45deg
ejection condition is apparently higher near the TE.
Meanwhile, the film-cooling effectiveness in the wake and
near the outlet is hardly changed. The well protected region is
limited to the TE corner region. After cooling the TE comer
region, the coolant strongly interacts with the secondary flows
such as the passage vortex and wakes. The main flow
eliminates the momentum of the trailing edge film cooling
quickly, which makes the film-cooling effectiveness of the TE
to be the same. On the other hand, the main flow further
mixes the coolant and the hot gas on the endwall, which leads
the ejection flow to lift off the endwall surface and then move
to the main flow. These two factors hardly cause the film-
cooling effectiveness to change at the far side of the trailing
edge which is close to the outlet.
Figure 22 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 2 (THE BLOWING RATIO IS 0.7, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 2 IS IN FIGURES 17-19, THE
MEANING OF TRAILING, FAR, LEFT AND RIGHT IS IN FIGURE 15)
Figure 23 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 2 (THE BLOWING RATIO IS 1.3, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 2 IS IN FIGURES 17-19)
Figures 22 and 23 compare the local film-cooling
effectiveness distribution at vertical location 2 with different
blowing ratios. As the blowing increases, the film-cooling
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
LEFT
RIGHT
Line2
Z/Length
i= 0deg M=0.7 trailing15
i= 0deg M=0.7 trailing45
0 0.05 0.1
0
0.1
0.2
Left
0.9 0.95 1
0
0.1
0.2
0.3
Right
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
LEFT
RIGHT
Line2
Z/Length
i= 0deg M=1.3 trailing15
i= 0deg M=1.3 trailing45
0 0.05 0.1
0
0.1
0.2
Left
0.9 0.95 1
0
0.1
0.2
0.3
Right

10. 10
effectiveness apparently improves near the trailing edge.
Meanwhile, the higher effectiveness area approaches the two
sides of the wake central line, LEFT and RIGHT respectively.
The well protected region is at the central line and the near
region by the central line (non-dimensional length Z/Length is
between 0.3 and 0.7). In the TE corner region of the passage,
the coolant strongly interacts with the secondary flows such
as the passage vortex and wakes. The main flow pushes the
coolant towards the mid-passage region, which causes the
protected area to be larger on the left side. However, the main
flow still mixes the coolant and the hot gas in the passage,
which leads the ejection flow to lift off the endwall surface,
which hardly causes the film-cooling effectiveness to change
at the off central line region where the endwall surface is not
covered by the TE ejection.
Figure 24 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 3 (THE BLOWING RATIO IS 0.7, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 3 IS IN FIGURES 17-19, THE
MEANING OF TRAILING, FAR, LEFT AND RIGHT IS IN FIGURE 15)
Figure 25 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL ALONG LINE 3 (THE BLOWING RATIO IS 1.3, WITH 15DEG AND
45DEG TRAILING EDGE EJECTION, LINE 3 IS IN FIGURES 17-19)
Figures 24 and 25 show the local film-cooling
effectiveness distribution at vertical location 3 where the
coolant is moved to the downstream part of the endwall near
outlet, 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 high
effectiveness area could be found along the vertical line (non-
dimensional length Z/Length is between 0.3 and 0.7), while
the influence of the trailing edge film cooling could not be
probed in this area, the downstream part at the wake central
line. This indicates that the effects of the trailing edge film
cooling are not apparent far away from the TE corner region
of endwall surface when the blowing ratio is relatively low.
Figure 25 compares the local film-cooling effectiveness
distribution in the downstream area when the blowing ratio is
M=1.3. The figure shows that the increase in the blowing
ratio improves the local film-cooling effectiveness of the TE
corner region where the ejection is fully dominated by the
wakes. The lower film-cooling effectiveness near the LEFT
and RIGHT side indicates that the coolant ejection is
influenced by the main wake flow field, especially the wake
vortex which causes strong mixing flow in the wake trace. In
this area, the main flow is dominated by the trailing edge
wakes. Lower effectiveness means stronger influence of the
vortex, which shows that the distance from the TE could
change the influence of the trailing edge film cooling on the
endwall. The film-cooling effectiveness curve representing
the case of 45deg TE film cooling is almost the same as the
curves representing the 15deg case in the mid-pitch region
(non-dimensional length Z/Length is between 0.3 and 0.7) as
shown in Figure 25. At the downstream part dominated by the
TE wake, the influence of trailing edge ejection is apparently
weakened by the secondary flow. The higher momentum of
the coolant ejection flow could not effectively overcome the
mixing trend of the wake vortex and then form a low film-
cooling effectiveness area in the wake.
Figure 26 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (45DEG), M=0.7
Figure 27 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (15DEG), M=0.7
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
LEFT
RIGHT
Line3
Z/Length
i= 0deg M=0.7 trailing15
i= 0deg M=0.7 trailing45
0 0.05 0.1
0
0.1
0.2
Left
0.9 0.95 1
0
0.1
0.2
0.3
Right
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
LEFT
RIGHT
Line3
Z/Length
i= 0deg M=1.3 trailing15
i= 0deg M=1.3 trailing45
0 0.05 0.1
0
0.1
0.2
Left
0.9 0.95 1
0
0.1
0.2
0.3
Right
Pitch
Span
0 0.5 1
0.75
0.875
1
0.97
0.98
0.99
1
Pitch
Span
0 0.5 1
0.75
0.875
1
0.97
0.98
0.99
1

11. 11
Figure 28 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH ENDWALL FILM COOLING, M=0.7 (BASELINE)
Figure 29 COMPARISON OF SPANWISE AVERAGED PRESSURE LOSS ALONG
PITCH WITH DIFFERENT TRAILING EDGE DISCHARGE GEOMETRY (M=0.7)
When the trailing edge blowing ratio is controlled to be
M=0.7, the geometry of trailing edge has strong influence on
the outlet total pressure loss. The 45deg discharge flow can
obviously increase the pressure loss in the trailing edge wakes
and corner vortex, as shown in Figures 26-28. The 45deg
discharge flow will lead another high total pressure loss
region above the corner vortex to form. However, the total
pressure loss does not increase apparently when the discharge
angle is 15deg. And the width of the wake is less than that of
the baseline case without any trailing edge discharge. This
result shows the 15deg discharge has better aerodynamic
performance than the 45deg design, though its secondary
cooling performance on endwall surface is worse. The Figure
29 shows the spanwise averaged total pressure coefficient
distribution along pitch direction. Compared with 45deg case,
the 15deg design has less total pressure loss in the wake area.
Figure 30 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (45DEG), M=1.0
Figure 31 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (15DEG), M=1.0
Figure 32 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH ENDWALL FILM COOLING, M=1.0 (BASELINE)
Figure 33 COMPARISON OF SPANWISE AVERAGED PRESSURE LOSS ALONG
PITCH WITH DIFFERENT TRAILING EDGE DISCHARGE GEOMETRY (M=1.0)
The high pressure loss area near corner region represents
the strength and position of secondary flow in the main
passage. When the trailing edge discharge angle is 45deg, a
new vortex forms above the original one at a non-dimensional
span of 0.875, as shown in Figure 26, 30 and 34. It is the
result of interaction between main passage secondary flow
and the discharge flow. The higher total pressure loss has
negative influence on the rotating blades downstream. The
mechanism behind this phenomenon could be illustrated
though formation of shear flows. The big angle between the
discharge flow and main flow causes strong shear effects on
the turning air, which makes a new vortex possible. Then the
new vortex moves up with the lift-off flow on suction side.
The final double vortex structure is the result of interaction
between 45deg coolant ejection and the secondary flow in
upstream passage.
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
0.98
0.99
1
Pitch
MeanPressureCoefficient
LE 45deg M=0.7 TE 15deg M=0.7
Pitch
Span
0 0.5 1
0.75
0.875
1
0.97
0.98
0.99
1
Pitch
Span
0 0.5 1
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
0.98
0.99
1
Pitch
MeanPressureCoefficient
LE 45deg M=1.0 TE 15deg M=1.0

12. 12
Figure 34 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (45DEG), M=1.3
Figure 35 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH TRAILING EDGE DISCHARGE (15DEG), M=1.3
Figure 36 SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITH ENDWALL FILM COOLING, M=1.3 （BASELINE）
Figure 37 COMPARISON OF SPANWISE AVERAGED PRESSURE LOSS ALONG
PITCH WITH DIFFERENT TRAILING EDGE DISCHARGE GEOMETRY (M=1.3)
As the blowing ratio increases, the advantage of 15deg
design is still apparent. The high total pressure loss regions of
45deg case exist in the wake and corner vortex range. And the
spanwise averaged pressure coefficient distributions in
Figures 33 and 37 show the worse aerodynamic performance
of 45deg design. Come to the conclusion, the 15deg railing
edge discharge design has smaller aerodynamic loss and the
position of the corner vortex core does not change. However,
the 45deg design causes extra aerodynamic loss and the high
total pressure loss region interacts with the corner vortex.
Though the secondary cooling function of 45deg case is
better, its disadvantage of aerodynamic loss cannot be
ignored.
CONCLUSIONS
In general, the trailing edge discharge flow affects the
coolant distribution on the downstream part of endwall
surface. The results show that with an increasing blowing
ratio, the film-cooling effectiveness increases on the endwall
surface, especially near the TE corner region. The film-
cooling effectiveness difference is weakened with the axial
chord increase, indicating that the trailing edge phantom film-
cooling ejection mixes with the main flow strongly at the
passage outlet, thus forming a low influence region in the area
further downstream. With increasing blowing ratios, the
improvement is also captured at the part further downstream
near the trailing edge region. The influence of the blowing
ratio is apparent for the trailing edge phantom coiling on the
endwall surface.
As the blowing ratio varies from M=0.7 to M=1.3, the
influence of the trailing edge discharge flow on the endwall
film cooling increases near the TE corner region.
Simultaneously, the area of influence will be limited at the
trailing edge region. In conclusion: 1）the film-cooling
effectiveness increases on the endwall surface near the
trailing edge region, with the highest parameter at the
impingement point: 2) a one-peak cooled region develops
along the mean line downstream the trailing edge as the
blowing ratio increases; 3）the advantage of the trailing edge
discharge flow phantom cooling was apparent when the
compound angle was 45deg, while the coolant film was
obviously weakened along the axial chord at smaller
compound angle 15deg. The influence of the trailing edge
discharge flow phantom cooling could only be detected at the
downstream area of the trailing edge, while it has little
influence on the main passage endwall film cooling.
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
Pitch
Span
0 0.5 1
0.75
0.875
1
0.97
0.98
0.99
1
Pitch
Span
0 0.5 1
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
0.98
0.99
1
Pitch
MeanPressureCoefficient
LE 45deg M=1.3 TE 15deg M=1.3

13. 13
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
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Film Cooling Study Using Pressure-Sensitive Paint”.
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[2] Zhang, L., Moon, H.K., 2003. “Turbine Nozzle Endwall
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[3] Wright, L.M., Gao, Z., Varvel, T.A., and Han, J.C., 2005.
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196–205.
[11] Dannhauer, A., 2009. “Investigation of Trailing Edge
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ASME Paper No. GT2009-59343.
[12] Telisinghe, J., Ireland, P., Jones, T., Barrett, D., Son, C.,
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Paper No. GT2006-91207.
[13] Yang, H., Gao, Z., Chen, H.C., Han, J.C., Schobeiri,
M.T., 2009. “Prediction of Film Cooling and Heat
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Stage”. Journal of Turbomachinery, Transactions of the
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[14] Kost F., Mullaert, A., 2006. “Migration of Film-Coolant
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