1. 1
Proceedings of ASME Turbo Expo 2012: Power for Land, Sea and Air
GT2012
June 11-15, 2012, Copenhagen, Denmark
GT2012- 69298
EFFECT OF INCIDENCE ANGLE ON GAS TURBINE FIRST-STAGE NOZZLE GUIDE VANE LEADING EDGE AND GILL
REGION FILM COOLING
Yang Zhang, Xin Yuan
Key Laboratory for Thermal Science and Power Engineering of Ministry of Education
Tsinghua University
Beijing 100084, P.R. China
Email: zhangyange436@yahoo.com.cn
ABSTRACT
The nonuniformity of the Hp turbine inlet flow field put
forward higher requirements for NGV (Nozzle Guide Vanes)
leading edge and gill region film cooling. The assumption of
design condition in most of the experiments couldn’t reflect
the true operation environment in the Hp turbine NGV. The
factor of off-design condition was incorporated into the
experiment in this research. The GE-E3
Hp turbine nozzle
guide vanes were used in the experiment to investigate the
cooling performance of injection from leading edge and gill
region with inlet Reynolds numbers of Re=3.5×105
and inlet
Mach number of Ma=0.1. The compound angle fan-shaped
film cooling hole configuration was applied. The cooling
characteristics at off-design condition were analyzed and
compared in the paper. The leading edge and gill region film
cooling performance was assessed with the incidence angle
varying from i=-10deg to i=+10deg. The blowing ratio
varying from M=0.7 to M=1.3, was also selected as an
experimental variable. Film cooling effectiveness distribution
was measured using PSP (Pressure Sensitive Paint) technique.
The film cooling performance of the compound angle
fan-shaped holes was assessed at both design and off-design
conditions. The object of this research is to change the
concept that NGV leading edge film cooling experiment only
needs the data at design condition. Through the comparative
analysis of experimental results at different inlet flow angle,
the influence of off-design condition on NGV leading edge
and gill region film cooling could be illustrated at a
reasonable level.
INTRODUCTION
Higher performance of future gas turbines desires an
improvement of efficiencies which is usually achieved by
increasing the turbine inlet temperatures. However, the
turbine inlet temperatures (about 1600 ºC) are generally
above the material failure limit of turbine components (about
1300 ºC), which drives 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 PSP, an
experimental study has been performed by Wright et al. [1] to
investigate the film cooling effectiveness measurements by
three different steady state techniques: pressure sensitive paint,
temperature sensitive paint, and infrared thermograph. They
found that detailed distributions could be obtained in the
critical area around the holes, and the true jet separation and
reattachment behavior is captured with the PSP. Wright et al.
[2] used the PSP (pressure sensitive paint) 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 to obtain detailed film cooling
effectiveness distributions. Gao et al. [3] 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 PSP technique. 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. [4] 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. Zhang et al. [5] 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 film

2. 2
effectiveness by the mass transfer analogy could be obtained.
The studies of incidence angle effect on flow field and
heat transfer were also preformed by researchers. Gao et al. [6]
studied the influence of incidence angle on film cooling
effectiveness for a cutback squealer blade tip. Three incidence
angles were investigated 0 deg at design condition and ±5 deg
at off-design conditions. Based on 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. Montomoli et al. [7] studied the effect of wake
passing on a film cooled leading edge with different incidence
and injection angles. The separated region near leading edge
changed its extension when the incidence angle is different. In
particular, the separation interested an higher region of
pressure side at negative incidence.
Lee et al. [8] studied the effects of incidence angle on
the endwall convective transport within a high-turning turbine
rotor passage. Surface flow visualizations and heat/mass
transfer measurements at off-design conditions were carried
out at a fixed inlet Reynolds number for the incidence angles
of -10deg, -5 deg, 0 deg, +5 deg, and +10 deg. The results
showed that the incidence angle had considerable influences
on the endwall local transport phenomena and on the
behaviors of various endwall vortices. In the negative
incidence case, convective transport was less influenced by
the leading edge horseshoe vortex. In the case of positive
incidence, however, convective transport was augmented
remarkably along the leading edge horseshoe vortex, and is
much influenced by the suction-side corner vortex. Benabed
et al. [9] numerically investigated the influence of incidence
angle on asymmetrical turbine blade model showerhead film
cooling effectiveness. The results indicated that variation of
operating incidence angle from that fixed by design
conditions could strongly affect the thermal protection of the
blade especially for the low blowing ratios. The stagnation
point moved from the suction side to the pressure side as the
incidence angle changing from the negative values to the
positive ones. For the lowest value of the incidence angle, the
trajectory path of the suction side injection is inversed when
the blowing ratio increased, dangerously exposing the suction
side to hot gases.
Duikeren et al. [10] studied the over-tip casing heat
transfer at off-design condition. At part load conditions, the
fluid might hit the front section of the suction side of the
blade resulting in increased turbulence and heat transfer levels.
In severe cases, the boundary layer flowed around the leading
edge will separate, which improved the heat transfer even
more. Wagner et al. [11] investigated the performance of
different film cooling hole configurations at designed and
off-designed mainstream incidence angles. At +5deg
incidence on the pressure side, a beneficial interaction
between the jets of the pressure side row appeared. For
middle and high blowing ratio, the film cooling performances
were also better than 0deg incidence. The study of Camci, C.
and Arts, T. [12]indicated that the change of the position of
the stagnation point strongly altered the aerodynamic
behavior and convective heat transfer to the blade. The
free-stream mass flow rate was kept constant for each
experiment at different incidence levels in the experiment.
The results showed that the heat transfer both on the suction
surface and on the pressure surface was significantly
influenced by the changes in approaching flow direction,
especially near the regions where small separation bubbles
were located.
As for the leading edge and showerhead film cooling
research at off-design condition, Ahn, J. et al.[13] conducted
the film cooling experiments on the first stage rotor of a
3-stage axial turbine with off-design condition at 2400 rpm.
The blowing ratio was controlled to be 0.5, 1.0, and 2.0 while
the density ratios of 1.0 and 1.5 were obtained using nitrogen
and CO2 as coolant gases, respectively. Wolff, S.[14] the
noticed the effects of upstream passing wakes had close
relationship with changes in incidence angle. The periodic
impinging wakes were generated by a wake generator
consisting of moving bars upstream of the cascade inlet plane.
Lin, Y.-L. and Shih, T.I.-P. [15] used computational results to
show the interactions between the mainstream hot gas and the
cooling jets, and how those interactions affect surface
adiabatic effectiveness. Karni, J. and Goldstein, R.J.[16] used
the naphthalene sublimation technique to study the effect of
surface injection on the mass (heat) transfer from a circular
cylinder in crossflow. Streamwise and spanwise injection
inclinations were studied separately, and the effects of
blowing rate and injection location relative to the cylinder
front stagnation line were investigated. The effect of
turbulence on leading edge film cooling was another
important research direction. Ou. S., Mehendale, A. B.and
Han, J. C. [17-19] experimentally investigated the effects of
film opening shape and mainstream turbulence on the leading
edge heat transfer coefficient and film effectiveness.
Past research has shown that incidence angles can result
in changes to the local heat transfer on the leading edge and
gill region region. Many studies have investigated the effects
of inlet flow angle on the blade showerhead film cooling,
indicating the off-design condition could apparently change
the injection flow trace. Few studies, however, have
considered the combining effect of compound angle and
incidence angle on the leading edge as well as gill region film
cooling. To help fill this gap, the current paper discusses the
effect of incidence angle on the film cooling of a nozzle guide
vane leading edge and gill region. The factor of the blowing
ratio is also considered.
EXPERIMENTAL METHODOLOGY
The film cooling effectiveness was measured using the
PSP technique. PSP is a photo luminescent material that
excited by visible light at 450 nm, emitting light that could be
detected by a high spectral sensitivity CCD camera (PCO
Sensicam Qe high performance cooled digital 12 bit CCD
camera) fitted with a 600 nm band pass filter. The light
intensity is inversely proportional to the local partial pressure

3. 3
of oxygen. The layout of the optical system is shown in Fig.1.
The image intensity obtained from PSP by the camera is
normalized with a reference image intensity ( refI ) 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 ( blkI ). The recorded
light intensity ratio can be converted to partial pressure ratio
of oxygen with the parameters obtained in calibration, as
shown in Eq.(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 intensity ratio and pressure ratio.
Figure 1. THE TEST RIG WITH EXCITATION LIGHT
0 0.2 0.4 0.6 0.8 1 1.2
0.1
0.3
0.5
0.7
0.9
1.1
1.2
Pressure
I
T=302.5 K
Figure 2. CALIBRATION CURVE FOR PSP.
The film cooling effectiveness can be determined by the
correlation between the PSP emitting intensity and the oxygen
partial pressure. Calibration of the PSP was performed in a
vacuum chamber by varying the pressure from 0 atm to 1.0
atm at three different temperatures. A PSP coated test coupon
was placed in the vacuum chamber with transparent windows
through which the camera can detect the light intensity on the
coupon surface. The calibration curve is shown in Fig.2. A
temperature difference less than 0.5K between main stream
and secondary flow should be guaranteed during the tests. To
obtain film cooling effectiveness, both air and nitrogen are
used as coolant. 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 air and nitrogen injection cases, the film cooling
effectiveness can be obtained using Eq.(2).
EXPERIMENTAL FACILITY
The test section consists of an inlet duct, a linear turbine
cascade, and an exhaust section. The inlet duct has a cross
section of 318 mm (width) and 129 mm (height). Considering
the nonuniformity of the outlet flow field of combustor,
incidence angle was selected to be the variable in the
experiment. The predominant vortex in the combustor made
the velocity direction in the outlet section difficult to predict.
The position of the stagnation point is strongly affected by the
indefinite inlet flow angle, and then in turn changed the
leading edge and gill region film cooling effectiveness
distribution. To study different mainstream inlet angles, the
guide vanes are placed on a rotatable half-circular plate,
which serves as part of the endwall, as shown in Fig.3. By
turning the half-circular plate, incidence angles at design and
off-design conditions are achieved. During the test, the tail
boards, and the CCD camera were moved with the rotatable
plate to the same relative position as that at incidence angle of
0deg. In this study, three different positions were chosen for
the incidence angles of i = -10 deg, 0deg and +10 deg. During
the test, the cascade inlet air velocity was maintained at 35
m/s for all the inlet flow conditions, which corresponds to a
Mach number of 0.1. A two times scaled model of the GE-E3
guide vanes with a blade span of 129 mm and an axial chord
length of 78.8 mm was used. [20] For coolant air supply,
compressed air is delivered to a plenum located below the
wind tunnel test section before being injected into the
mainstream, as shown in the schematic diagrams in Fig. 4.
Past studies in the open literature have shown that the
passage cross flow sweeps the film coolant from endwall to
mid-span region due to the vortex in the passage. To reflect
this phenomenon more apparently, all of the film cooling
holes are positioned in straight lines. Studies on the flat plates
show that coolant from compound angle holes covers wider
area due to jet deflection. Six rows of compound angle
laidback fan-shaped holes are arranged on the leading edge
and gill region to form full covered coolant film. Fig. 5-Fig. 7
show the holes configurations and the blade geometric
parameters.

4. 4
Figure 3. THE TEST SECTION WITH ROTATABLE CASCADE AND THE
ASSEMBLY DRAWING OF TEST SECTION
Figure 4. SCHEMATIC OF CASCADE TEST RIG
Figure 5. SCHEMATIC OF THE COMPOUND ANGLE FILM COOLING HOLES
AT LEADING EDGE AND GILL REGION
Three rows are arranged on the LE at axial locations of
6.2 mm (LE1, 16 holes), 2.8 mm (LE2, 17 holes), and 3.0 mm
(PS3, 16 holes). The first row is located on the leading edge
suction side that the film cooling effectiveness of this row is
difficult to be accessed for camera position limitation. The
following five rows are positioned on the showerhead and gill
region, with the last one located at 20% axial chord
downstream of the leading edge. Three rows were provided
on the PS at axial locations of 5.9 mm (PS1, 17 holes), 10.6
mm (PS2, 16 holes) and 15.6 mm (PS3, 17 holes). The
leading edge hole diameter of metering part d was 1.0 mm
and the total length of a hole was 6d. The hole expansion
started at 4d. The gill region hole diameter of metering part d
was 1.0 mm and the total length of a hole was 8.5d. The hole
expansion started at 4.5d. The holes were staggered; therefore,
LE1, LE3, and PS2 had one hole less than LE2, PS1, and PS3.
Due to the large pressure gradient on the endwall, it is
difficult to control the local blowing ratios for every single
hole with one common coolant plenum chamber. In the
current study, two coolant cavities are used for the
showerhead and gill region respectively, as shown in Fig.6
(The other rows of cooling holes are designed to research the
pressure side film cooling. They are not used in this
experiment, though shown in the figure). The coolant
supplied to the two cavities is controlled by a shared
rotameter. As shown in Fig.6 and Fig.7, the six rows of
fan-shaped holes are inclined 33 deg to the airfoil surface and
held an angle of 30 deg to radial direction. The laidback
fan-shaped holes are featured with a lateral expansion of 5
deg from the hole axis and forward expansion of 5 deg into
the airfoil surface, as shown in Fig.7.
The uncertainties of the dimensionless temperature and
the film cooling effectiveness are estimated as 3% at a typical
value of 0.5 based on 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 6. COMPOUND ANGLE FILM COOLING HOLES CONFIGURATION ON
LEADING EDGE AND GILL REGION (COMPARED WITH RADIAL HOLES)

5. 5
Figure 7. DETAILS OF THE FAN-SHAPED FILM COOLING HOLES
RESULTS AND DISCUSSION
The film cooling effectiveness distributions and laterally
averaged values at different incidence angles are shown in
Fig.8- Fig.19., 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 that the area coverage of coolant film is larger at
higher blowing ratio. Although valuable insight can be
obtained from the distribution maps (Fig.8-Fig.10,
Fig.12-Fig.14 and Fig.16-Fig.18), the spanwise averaged
plots (Fig.11, Fig.15 and Fig.19) offer additional insight and
provide clear comparisons for large amounts of data. The
effectiveness is averaged from the hub to the tip (Fig.8-Fig.10,
Fig.12-Fig.14 and Fig.16-Fig.18) 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
fan-shaped holes location. Fig.11, Fig.15 and Fig.19 indicate
that increasing the injection rate increases the film cooling
effectiveness. The lowest film cooling effectiveness appears
at M=0.7. The average is significantly lower because the
coolant does not cover the leading edge and gill region
sufficiently, especially near the leading edge corner region
where the local pressure is relatively high. The blowing ratio
effect is clearly seen on the downstream half part (axial chord
position larger than 0.2) of the pressure side, with the
effectiveness being proportional to the mass flow ratio.
Though the cascade is 2-d symmetric, the relative
ejection direction of the coolant is different at the tip corner
and the root corner. The radial angle of 33 deg causes the
ejection direction to opposite relative to the endwall surface at
tip and root. The interaction between the coolant and the main
flow, especially the horse shoe vortex, makes the root corner
region to be hardly cooled. While the different flow direction
near the tip corner avoid this harmful interaction. The
aerodynamic relationship between leading edge coolant
ejection and main flow has been discussed by Michael
Cutbirth, J. and Bogard, D. G. [21] in their paper about flow
visualization of leading edge film cooling.
Fig.8-Fig.10 show the film cooling effectiveness
distribution on the leading edge and gill region surface at
different blowing ratio, while the incidence angle is controlled
to be i=0deg. With the blowing ratio increasing, the area
protected by the coolant is getting larger. Though the coolant
could cover the main part of the airfoil pressure side, the
unprotected area near the corner region is still apparent.
(shown with the white curve) This phenomenon represents
that the strong pressure gradient in the turbine cascades
dominant 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 corner region.
(axial chord position larger than 0.2) The similar case could
be observed near the leading edge junction region where the
coolant could only inject apparently from the cooling holes at
mid-span height. Though the other pressure side cooling holes
were not used in the experiment, the inner surface of the holes
were painted by with the PSP thus forming light reflection in
the cooling holes on the pressure side downstream part. All of
the cooling hoes unused were internally blocked, which cause
the slight effect of the hole outlet geometry on flow filed not
to be avoided in the experiment. But the influence is restricted
to the pressure side downstream part which is not discussed.
compound M=0.7 i= 0deg
Z/Z
sp
X/Xax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 8 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 0.7 AND THE
INCIDENCE ANGLE IS 0DEG)
compound M=1.0 i= 0deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 9 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.0 AND THE
INCIDENCE ANGLE IS 0DEG)

6. 6
compound M=1.3 i= 0deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 10 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.3 AND THE
INCIDENCE ANGLE IS 0DEG)
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/X
ax

compound i=0deg M=1.3
compound i=0deg M=1
compound i=0deg M=0.7
Figure 11 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS ON THE
PRESSURE SIDE (INCIDENCE ANGLE IS 0DEG)
compound M=0.7 i= -10deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 12 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 0.7 AND THE
INCIDENCE ANGLE IS -10DEG)
compound M=1.0 i= -10deg
Z/Zsp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 13 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.0 AND THE
INCIDENCE ANGLE IS -10DEG)
compound M=1.3 i= -10deg
Z/Z
sp
X/Xax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 14 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.3 AND THE
INCIDENCE ANGLE IS -10DEG)
With the blowing ratio increasing, the original
momentum of the coolant increases. Higher blowing ratio
leads the coolant injecting from the cooling holes near corner
region that the extreme position of coolant injection becomes
nearer to the endwall. (span height between 0.8 and 1.0, axial
chord between 0 and 1.0, two film cooling holes are located
near endwall) As the coolant leaves the cooling holes, the
trace of the injection flow is led by the corner vortex
developing near the leading edge pressure side. The vortex is
strong at the junction region, which causes the boundary of
coolant to move along the corner vortex and towards the
mid-span height. The film cooling effectiveness distributions
indicate that the cooling performance of leading edge and gill
region holes is not enough. The injection could cover the
mid-span part and even over cool this area, while the corner
region is still exposed to the hot environment. To increase the
blowing ratio could partly improve the cooling effectiveness,
but the performance near the corner region is not satisfied.
Fig.12-Fig.14 and Fig.16-Fig.18 indicate the film

7. 7
cooling effectiveness distributions at off-design conditions.
When the inlet flow angle is i=-10deg, as shown in
Fig.12-Fig.14, the main difference from the design condition
is that, at low blowing ratio the injection area is larger. The
coolant could inject from the cooling holes in the corner
region. (span height between 0.8 and 1.0, axial chord between
0 and 0.1, two film cooling holes are located near
endwall)This phenomenon shows that the incidence angle of
i=-10deg makes the high pressure area to move towards the
endwall slightly, that the corner vortex is weaker at the
pressure side. This increase is also obvious at higher blowing
ratio cases. When the blowing ratio is M=1.0, the coolant
could inject from the corner region cooling holes near the
endwall when the incidence angle is i=-10deg, while the film
cooling effectiveness is low when the incidence angle is
i=0deg. This indicates that the leading edge film cooling is
sensitive to the incidence angle even at higher blowing ratio.
Fig.16-Fig.18 show the similar trend that the behaviour of the
injection flow could be apparently influenced by the
incidence angle even at high blowing ratio. For instance,
compared with the obvious coolant injection in the corner
region (axial chord is between 0 and 0.1, span position is
between 0.8 and 1.0, two fan-shaped cooling holes are located
in this region) at leading edge in Fig.14, very low film
cooling effectiveness is probed when the incidence angle is
i=+10deg, as shown in Fig.18.
Fig.11, Fig.15 and Fig.19 compare the laterally average
film cooling effectiveness at different blowing ratio. As the
blowing increasing, the averaged effectiveness apparently
improves. Meanwhile the average effectiveness increases with
increasing axial chord. The well protected region is the
mid-span part of the airfoil (span height is between 0.2 and
0.8). In the corner region of the airfoil the coolant strongly
interacts with the secondary flows such as the corner vortex
and transversal flow. The main flow pushes the coolant
towards mid-span region, which cause the protected area to be
limited. On the other hand, the main flow fatherly mixes the
coolant and the hot gas in the passage, which leads the
injection flow to lift off the airfoil surface. These two factors
cause the average film cooling effectiveness decrease a lot in
the corner region of the airfoil.
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/X
ax

compound i=-10deg M=1.3
compound i=-10deg M=1
compound i=-10deg M=0.7
Figure 15 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS ON THE
PRESSURE SIDE (INCIDENCE ANGLE IS -10DEG)
compound M=0.7 i= +10deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 16 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 0.7 AND THE
INCIDENCE ANGLE IS +10DEG)
compound M=1.0 i= +10deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 17 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.0 AND THE
INCIDENCE ANGLE IS +10DEG)
compound M=1.3 i= +10deg
Z/Z
sp
X/X
ax
0 0.1 0.2 0.3 0.4
0
0.2
0.4
0.6
0.8
1
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
Figure 18 FILM COOLING EFFECTIVENESS DISTRIBUTION ON LEADING
EDGE AND PRESSURE SIDE (THE BLOWING RATIO IS 1.3 AND THE
INCIDENCE ANGLE IS +10DEG)

8. 8
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/X
ax

compound i=+10deg M=1.3
compound i=+10deg M=1
compound i=+10deg M=0.7
Figure 19 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS ON THE
PRESSURE SIDE (INCIDENCE ANGLE IS +10DEG)
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/Xax

compound M=0.7 i= -10deg
compound M=0.7 i= 0deg
compound M=0.7 i=+10deg
Figure 20 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS AT
DIFFERENT INCIDENCE ANGLE (THE BLOWING RATIO IS 0.7 )
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/Xax

compound M=1.0 i= -10deg
compound M=1.0 i= 0deg
compound M=1.0 i=+10deg
Figure 21 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS AT
DIFFERENT INCIDENCE ANGLE (THE BLOWING RATIO IS 1.0 )
-0.1 0 0.1 0.2 0.3 0.4 0.5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
X/X
ax

compound M=1.3 i= -10deg
compound M=1.3 i= 0deg
compound M=1.3 i=+10deg
Figure 22 LATERALLY AVERAGED FILM COOLING EFFECTIVENESS AT
DIFFERENT INCIDENCE ANGLE (THE BLOWING RATIO IS 1.3 )
Fig.9, Fig.13 and Fig.17 show the film cooling
effectiveness distributions on the airfoil at different incidence
angle when the blowing ratio is controlled to be M=1.0. When
the incidence angle is i=0deg (Fig.9) and i=+10deg (Fig.17),
an apparent unprotected area could be found in the leading
edge corner region (axial chord is between 0 and 0.1, span
position is between 0.8 and 1.0, two fan-shaped cooling holes
are located in this region), while this area is covered by the
coolant better at the incidence angle of i=-10deg, as shown in
Fig.13. This indicates that though the effects of incidence
angle is not apparent in the downstream corner region of
airfoil surface (axial chord is larger than 0.2, the triangular
region shown with the white curve), the influence could not
be ignored in the leading edge corner region. Fig.20-Fig.22
compare the laterally average film cooling effectiveness on
the airfoil at different incidence angle. Fig.20, Fig.21 and
Fig.22 show that the increasing in incidence angle (from
i=-10deg to i=+10deg) decreases the average effectiveness on
the leading edge and gill region when the blowing ratio is
M=0.7, M=1.0 and M=1.3 respectively. The lower film
cooling effectiveness on the leading edge and gill region
indicates that the coolant injection is influenced by the
incidence angle of i=+10deg. In this area, the main flow is
dominated by the pressure side leg of horseshoe vortex and
the corner vortex. Lower effectiveness means stronger
influence of the vortex, which shows that the incidence angle
could change the pressure side leg of the horseshoe vortex as
well as the corner vortex, and then lead the coolant injection
flow to move away from the corner region. The average
effectiveness curve representing the case of i=-10deg is
obviously above the curves representing i=0deg case and
i=+10deg case in Fig.20, Fig.21 and Fig.22.As the blowing
ratio increasing, the disadvantage of positive incidence angle
(i=+10deg) is not weakened. The higher momentum of the
coolant injection flow could not effectively overcome the
influence of horseshoe vortex, and the corner vortex.
CONCLUSIONS
In general, the incidence angle affects the coolant

9. 9
distribution on the leading edge and gill region surface
apparently. The results show that with blowing ratio
increasing, the film cooling effectiveness increases on the
airfoil surface. No lift off and reattachment phenomenon is
observed. The film cooling effectiveness increase with the
axial chord increasing, indicating that the film cooling
ejections mix with main flow strongly near leading edge thus
forming low film cooling effectiveness region. With blowing
ratio increasing, the improvement is mainly captured at the
downstream part on the pressure side gill region. The
influence of the blowing ratio is not apparent at leading edge.
As the incidence angle varies from +10deg, 0deg to
-10deg, at a fixed blowing ratio the film cooling effectiveness
increases on the leading edge and gill region. Simultaneously,
the pressure distribution at the hole outlet will apparently
change the coolant ejection position. The pressure distribution
at positive incidence angle could make the ejection position
relatively upstream. The peak point in the plots moves
slightly downstream, which cause the peak point axial chord
position of the film cooling effectiveness curves different.
The incidence angle could apparently change the position of
the stagnation line relative to the leading edge, thus changing
the pressure distribution near leading edge obviously. But the
pressure distribution at downstream part of the cascade is
hardly influenced by the incidence angle, which make the
film cooling effectiveness distribution not sensitive to
incidence apparently at downstream part. The influence of the
incidence angle is obvious at the leading edge region while its
influence is weakened as the axial chord increases.
NOMENCLATURE
C =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
LE =leading edge
M =blowing ratio, ρcVc/ρ∞V∞
Ma =Mach number
PS =pressure side
PSP =pressure sensitive paint
Re =Reynolds number
SS =suction side
V =velocity, m/s
X,Z =Cartesian coordinate system
 =film cooling effectiveness
Subscripts
aw =adiabatic
ax =axial chord
c =coolant fluid
in =inlet
mix =mixture condition
ref =reference value
sp =span wise
 =free stream condition
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