1. 1
Proceedings of ASME Turbo Expo 2014: Power for Land, Sea and Air
GT2014-26824
June 16-20, 2014, Düsseldorf, Germany
EXPERIMENTAL INVESTIGATION OF NONPERIODIC ENDWALL FILM COOLING IN NEIGHBORING PASSAGES
WITH UPSTREAM ROTATING 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 distribution of film-cooling holes is considered to be
the same between neighbouring high pressure turbine
passages in most cascade experiments. Because of the
difference in accounts of combustors and vanes, however, the
flow fields of neighbouring passages are completely different.
The secondary flow, especially the passage vortex, is
dominated by the upstream inlet rotating flow whose relative
flow direction is the reverse between the neighbouring vane
passages. Specific rotating directions introduce new
challenges in film-cooling design. The present experiment
compares three groups of endwall film-cooling with
anticlockwise rotating flow inlets at different clocking
positions, and the film-cooling effect is analysed to
investigate the effects of inlet rotating flow. The inlet flow
condition of neighbouring passages is simulated by switching
the position of the swirler by means of which rotating inlet
flow conditions in different positions are achieved.
The GE-E3
airfoil is used in the cascades, with a scaled-
up factor of 2.2. The inlet Reynolds number is 3.5×105
and
the Mach number is 0.1. The effects of the blowing ratio and
relative positions of the swirler are investigated in the
experiment. Adiabatic film-cooling effectiveness is probed by
using pressure-sensitive painting (PSP). The coolant is
simulated by nitrogen by which a density ratio of around 1.0
can be achieved. Fan-shaped film-cooling holes are
introduced into the endwall surface as well as trailing edge
discharge holes. The cooling performance of the combustor-
turbine gap leakage flow is also considered simultaneously.
Conclusions are as follows: (1) the anticlockwise
direction of rotating flow influences the endwall film-cooling
effectiveness, especially for the upstream part of the endwall
which is mainly covered by the leakage flow; (2) the film-
cooling effectiveness in the neighbouring passages differs
depending on the position of the inlet rotating flow core; (3)
the film-cooling performance at the downstream part of the
endwall is partly influenced by the upstream rotating flow
inlet as well.
INTRODUCTION
As the inlet temperature in modern gas turbines increases,
higher thermal loading on the high-pressure turbine makes
heavy demands on guide-vane cooling. The introduction of
the film-cooling technique for nozzle guide vanes (NGV) of
advanced industrial gas turbines makes further improvement
in performance possible. With regard to the upstream
combustor outlet flow condition, research is necessary about
the effects of inlets' non-uniform boundary condition induced
by the swirler in the combustor. Given that the current
research in the primary stage concerns the basic mechanism,
an anticlockwise swirler embedded in a movable plate is used
to simulate the rotating flow with different core positions at
the turbine inlet.
In research on rotating flow inlet conditions, Khanal [1]
conducted a numerical investigation of transonic high-
pressure turbine behaviour under the combined influence of
swirl and hot-streak. The cases considered included both the
effects of the combustor-NGV clocking and the direction of
the swirl. The CFD result indicated that the swirl directions
strongly influenced the blade heat transfer characteristics and
their dependence on the clocking position. Qureshi [2]
investigated the influence of upstream swirling flow on high
pressure (HP) vane heat transfer and aerodynamics. The
vortex centre was aligned with the vane leading edge. This
paper presented measurements of HP vane surface and
endwall heat transfer for two vane positions. Qureshi [3]
compared the experimental measurements with and without
inlet swirl upstream the HP turbine stage. This paper
presented time-averaged experimental heat transfer
measurements performed on the rotor casing surface and on
the rotor blade surface at different span positions, as well as
time-averaged rotor casing static pressure measurements.
Khanal [4] compared two non-uniform inlet conditions of
practical interest using numerical simulation: a hot-streak
typical of a rich-burn combustion environment and a
combined swirl and hot-streak typical of a lean-burn
combustion environment. The results indicated that the
swirl/hot-streak combination gave much larger unsteady
fluctuations on both the rotor blade forcing and the heat load.

2. 2
In film-cooling research using pressure-sensitive
painting (PSP), Zhang and Jaiswal [5] measured film-cooling
effectiveness on a turbine vane endwall surface using the PSP
technique. 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 [6]
used the back-facing step to simulate the discontinuity of the
nozzle inlet with regard to the combustor exit cone. Nitrogen
gas was used to simulate cooling flow and a tracer gas to
indicate oxygen concentration so that the film’s effectiveness
could be obtained by mass transfer analogy. An experimental
study was performed by Wright and colleagues [7] to
investigate film-cooling effectiveness measurements of 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 captured with the PSP. Wright and
colleagues [8] used the PSP technique to measure the film-
cooling effectiveness on a turbine blade platform with three
different stator-rotor seals. Three slot configurations placed
upstream of the blades were used to model advanced seals
between the stator and rotor. Therefore, PSP was proven to be
a valuable tool for obtaining detailed film-cooling
effectiveness distributions. Gao and colleagues [9] 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. Detailed film-cooling effectiveness
distributions on the platform were also obtained with 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. Charbonnier and colleagues [10] applied the PSP
technique to measure coolant gas concentration. An engine
representative density ratio between the coolant and the
external hot gas flow was achieved by the injection of CO2.
Studies of the incidence angle effect on the flow field and
heat transfer have also been performed by researchers. Gao
and colleagues [11] studied the influence of the incidence
angle on the film-cooling effectiveness for a cutback squealer
blade tip. Three incidence angles were investigated: zero in
the design condition and ±5 in the off-design conditions.
Based on mass transfer analogy, film-cooling effectiveness
was 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 regards blade endwall platform film-cooling research,
Yang and colleagues [12] 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 were reported. Research by Kost and
Mullaert [13] indicates that both the leakage flow of endwall
upstream slots and film-cooling ejection are strongly
influenced by the endwall pressure distribution. The leakage
flow and film-cooling ejection 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 and colleagues [14]. 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
moved towards the suction side and then formed a triangular
cooled area.
The effects of rotation on platform film-cooling were
investigated by Suryanarayanan and colleagues [15] 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 flow
clearly towards the suction side surface. As regards
investigations into combustor-turbine leakage flow, Thole’s
group made significant contributions: the influences of slot
shape and position as well as width were studied in a series of
researches [16-18].
Oke and Simon [19] investigated the film-cooling flow
introduced through two successive rows of slots, a single row
of slots and slots with particular area distributions in the
pitchwise direction. Wright and colleagues [20] used a 30 °
inclined slot upstream of the blades to model the seal between
a stator and a rotor. Twelve discrete film holes were located
on the downstream half of the platform for additional cooling.
Rehder and Dannhauer [21] 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 [22] investigated
the leakage flow and misalignment effects on 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 can
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 endwall film-cooling, indicating
that the flow field parameter can change the injection flow
trace. Few studies, however, have considered the combined
effect of inlet rotating flow and endwall film-cooling. To help
fill this gap, the current paper discusses the effect of
anticlockwise inlet rotating flow clocking positions on the
film-cooling of a nozzle guide vane endwall. The blowing
ratio is also considered.

3. 3
EXPERIMENTAL METHODOLOGY
Pressure-sensitive paint (PSP) techniques are based on a
physical process called oxygen-quenched photoluminescence.
This process first involves excitation by a suitable light
source so that the active part of PSP emits light, a process that
can be interrupted by collisions of the PSP surface with
oxygen molecules. The result is that the PSP molecules may
relax back to their unexcited state without emitting visible
light if the local oxygen partial pressure is too high. Given
that the local oxygen partial pressure is related to the local
pressure of gas which contains oxygen, such as air, the
emitted light intensity is directly related to the local pressure
of surrounding air. A Charge-coupled Device (CCD) camera
with high spectral sensitivity and light-emitting diode (LED)
lights are used in the present study to receive the emitted light
and to excite the Ruthenium-based paint respectively. The
paint is excited at 450 nm and the camera is fitted with a 600
nm band pass filter. The paint used returns 600 nm light in the
energized state. In the current study, the main stream is air,
which contains approximately 21% oxygen, and the cooling
flow is pure nitrogen in which the partial pressure of oxygen
is 0%. The film effectiveness is then expressed in terms of
oxygen concentration, which is detected and measured using
the PSP. The resulting effectiveness equation is then given by
2
mix aw
N c
C C T T
C C T T
  
 
 
  
 
（1）
where
C and mixC represent the oxygen concentration of the
main stream and the air/nitrogen mixture (0% and 21%),
respectively. This approach gives film effectiveness values
between 0% (for far upstream and downstream locations) and
100% (inside film-cooling holes). The resulting equivalent
adiabatic effectiveness is then expressed using
   
 
2 2
2
O O
air mix air mix
air O air
P PC C
C P


  （2）
To measure the film-cooling effectiveness, four images
are taken at the same main stream temperature value for each
PSP film-cooling test. First, a dark image is taken without
LED light and the main stream flow. A reference image is
then taken without the main stream, but with the LED light on.
An air injection image and a nitrogen injection image are
subsequently obtained, both with the main stream flow and
LED light on, and the coolant gas is air and nitrogen,
respectively. The reference image distribution, divided by the
nitrogen-injection image and the air-injection image, is then
obtained from the four images for each test sequence. The
resulting distribution of reference ratio data (derived from the
air-injection image) then provides information on the change
in oxygen concentration owed to the change in pressure,
which does not contribute to the magnitude of local film-
cooling effectiveness. The other distribution of reference ratio
data, derived from the nitrogen-injection image, then yields
contributions to effectiveness values from absolute oxygen
concentrations. With these two groups of reference ratio
distributions, the local adiabatic film-cooling effectiveness
distributions are obtained in appropriate fashion by means of
mass transfer/heat transfer analogy.
Figure 1. CALIBRATION SYSTEM.AND CALIBRATION CURVES FOR PSP.
Before each experimental measurement, the PSP,
employed along a test surface, is calibrated to obtain the
relationships between light intensities and local partial
pressures of oxygen. Figure 1 shows a schematic diagram of
the calibration system employed for this purpose. Here, a
PSP-coated copper coupon is used to simulate the
experimental surface, and three thermocouples are installed
underneath the front surface to measure the time variation of
surface temperature as the calibration process is underway.
The sample coupon is located inside a sealed chamber, where
a partial or total vacuum is created. The temperature of the
sample is first increased by a heater located at the back of the
coupon, which maintains the sample at a desired temperature
with accuracy better than ±0.5 K. The camera faces the
sample coupon through a transparent window. Because
surface temperatures between 298 K and 308 K are possible
with this system, the PSP is subsequently calibrated for
surface temperatures of 276.5 K, 298 K, and 308 K, and for
pressures from vacuum (approximately 1/3 of an atmosphere)
to one atmosphere. The calibration is undertaken at these
three temperatures to provide complete information on the
influence of temperature. Examples of resulting calibration
data are presented in Figure 1, which shows the

4. 4
interdependence between light intensity ratio and pressure
ratio. Note that quantitative variations between the data sets
for the three temperature levels are quite small, and evidences
only weak dependence of calibration data on surface
temperature level.
The dimensionless temperature downstream of the
cooling holes is also determined from light intensity
distributions, and is defined using Eq. (3), which is expressed
using an equation of the form
c
T T
T T



 

（3）
Variations of surface adiabatic wall temperature are
presented using dimensionless film-cooling effectiveness,
which is defined for low speed and constant property flows
using Eq. (4), which is given by
aw
c
T T
T T
 




（4）
For the calibration case, based on a 95% confidence
level, the experimental uncertainties of the dimensionless
temperature and the adiabatic film-cooling effectiveness are
both approximately 3%, for an effectiveness representative
value of 0.5. These uncertainty values increase as the
effectiveness approaches zero, resulting in an uncertainty of
approximately 20%.
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 measuring 338 mm in width and 129 mm in height.
Not considering the non-uniform effect of the flow field at the
combustor outlet, the incidence angle is not selected to be the
variable in the experiment. The predominant vortex in the
combustor makes 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 is maintained at 35 m/s for all the inlet flow
conditions, corresponding to a Mach number of 0.1. A 2:1
scale model of the GE-E3
guide vanes with a blade span of
129 mm and an axial chord length of 79 mm is used. To
obtain an easier assembly method, the inlet rotating module is
movable and the swirler position can be changed with the
main blade part untouched. The inlet rotating flow module
has three specific pitchwise positions through which the
extreme position of leading edge-aligned and mid-pitch-
aligned can be achieved. During the test the main cascade
module is kept at the same relative position with the CCD so
that the film-cooling effectiveness distribution can be
compared accurately. 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 2 and Figure 3.
Figure 2. THE TEST SECTION WITH MOVABLE INLET SWIRLER AND THE
ASSEMBLY DRAWING OF THE TEST SECTION
Figure 3. SCHEMATIC OF CASCADE TEST RIG
Past studies in the open literature have shown that the
passage cross-flow sweeps the film coolant from endwall to
mid-span region because of the vortex in the passage. To
reflect this phenomenon more clearly, all 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 because of jet deflection. Four rows of radial cylindrical
film-cooling holes are arranged on the gill region to form a

5. 5
fully covered coolant film. Figures 4 to 8 show the hole
configurations and the geometric parameters of the blade.
Four rows of compound angle laidback fan-shaped holes
are arranged on the endwall to form a fully covered coolant
film. Figure 4 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 ° to
the platform surface and held at an angle of 0, 30, 45 and 60 °
respectively in the axial direction. The laidback fan-shaped
holes are featured with a lateral expansion of 10 ° from the
hole-axis and forward expansion of 10 ° into the endwall
surface, as shown in Figure 4. 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 5. (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 4. DETAILS OF THE FAN-SHAPED ENDWALL FILM-COOLING HOLES
The fixed fan upstream of the cascades is used to
simulate the real combustor swirler bringing a rotating flow to
the NGV inlet. The swirler is located 392mm upstream of the
cascade inlets to allow the rotating flow to fully develop in
the test section. The swirler is fixed on a movable plate which
makes position migration possible. The extreme position of
the swirler is limited by the left and right side of the plate.
The corresponding position of the swirler related to left side is
Vane A-aligned. The corresponding position of the middle
pitch and right side is middle-pitch aligned and Vane A/B-
aligned respectively. The Vane A-aligned and Vane B-aligned
positions are not technically the same condition when the
number of combustor cambers and nozzle guide vanes differ.
The Vane A-aligned and Vane B-aligned positions can
simulate the flow condition of ‘neighbouring’ passages. As
shown in Figure 5, the outer diameter of the swirler is 92mm
and its inner diameter is 31mm. The guide vanes of the
swirler turn the inlet flow into an anticlockwise rotating flow
boundary condition. The clockwise rotating flow condition is
not considered in the test because the rotating direction effect
is included in the following experiments. A slot is located at
the cascade inlet to simulate the combustor-turbine leakage
gap with a slot width of 2mm and an inclined angle of 30 ° to
the endwall surface. The leakage flow is supplied by an
individual coolant plenum where the blowing ratio can be
controlled independently.
Figure 5. THE POSITIONS OF SWIRLER AND THE CONFIGURATION OF
ENDWALL INLET SLOT (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 varies from 0.01 to 0.02 units.
Thus, relative uncertainties for very low effectiveness can be
very high, 100% at a 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 Slot Film-cooling Endwall Film-cooling M
Air
(L/min)
N2
(L/min)
Air
(L/min)
N2
(L/min)
Film-cooling Without Endwall Holes Injection (Vane A-aligned)

6. 6
1 77 80 0 0 1.0
Film-cooling Without Endwall Holes Injection (Mid-pitch-aligned)
2 77 80 0 0 1.0
Film-cooling Without Endwall Holes Injection (Vane B-aligned)
3 77 80 0 0 1.0
Combination of Slot and Endwall Holes Film-cooling
4 77 80 102 106 1.0
Film-cooling Without Endwall Holes Injection (Mid-pitch-aligned)
5 77 80 102 106 1.0
Film-cooling Without Endwall Holes Injection (Vane B-aligned)
6 77 80 102 106 1.0
Table 3 Geometric and flow conditions
Scaling factor 2.20
Scaled up chord length 135.50 mm
Scaled up axial chord length 79.00 mm
Pitch/chord 0.80
Span/chord 0.95
Reynolds number at inlet 3.5×105
Inlet and exit angles 0 & 72 °
Inlet Mach number 0.1 & 0.25
Inlet mainstream velocity 35 m/s
Mainstream flow temperature 305.5 K
Injection flow temperature 305.0 K
Figure 6. FIVE-HOLE PROBE AND ITS TWO DIMENSIONAL DISPLACEMENT
WORKTABLE
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.
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 7. SPATIALLY RESOLVED MAP OF NON-DIMENSIONAL TOTAL
PRESSURE LOSS WITHOUT FILM COOLING
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, as shown in Figure 7. 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 8. 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 8, Vane-to-vane comparisons of
the experimental measurement points demonstrate that a good
level of periodicity, too.

7. 7
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 differ from the endwall main flow direction. The
interaction between the endwall film-cooling coolant and the
secondary flow, especially the passage vortex, hardly cools
the endwall near PS and the stagnation line, and the different
flow direction near the suction side avoids this adverse
interaction. According to the contours, with endwall film-
cooling hole ejection the passage vortex pushes the coolant to
the suction side, leaving an apparently uncooled area near the
pressure side, especially near the stagnation line.
In the current study, five coolant cavities are used for the
slot leakage 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 in the same position so that the
different inlet rotating flow conditions can be compared
precisely. In this study, three different inlet rotating flow
positions were chosen for the typical design condition, Vane
A-aligned, mid-pitch-aligned and Vane B-aligned. The
blowing ratio of the coolant is fixed, so the film-cooling
effectiveness can be measured at a steady blowing ratio of
M=1.0, based on the mainstream flow inlet velocity.
The film-cooling effectiveness distributions and laterally
averaged values at different inlet rotating flow conditions are
shown in Figures 9 to 14: three different inlet rotating flow
positions are chosen for the typical design condition, Vane A-
aligned, mid-pitch-aligned and Vane B-aligned. The same
trend can be found in the contours so that the area coverage of
coolant film is larger with endwall film-cooling hole injection.
Figures 9 to 11 show the film-cooling effectiveness
distribution on the endwall surface with and without endwall
film-cooling hole injection, and the blowing ratio is
maintained at M=1.0. When the rotating flow position
changes, the area protected by the coolant changes. Though
the coolant can cover the main part of the endwall surface, the
unprotected area near the pressure side is still apparent
(shown by the red curve). This phenomenon shows the strong
pressure gradient in the turbine cascades, which dominates
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 zero and 0.3). A similar case can be
observed near the leading edge where the coolant can only
inject, apparently from the cooling holes at the leading edge.
The PS and SS leg of the horseshoe vortex can 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 are internally blocked,
which causes the slight effect of the hole outlet geometry on
the flow field being avoided in the experiment.
The left subplot in Figures 9 to 11 shows the film-
cooling effectiveness distributions on the endwall with main
passage film-cooling hole injection when the different swirler
position upstream of the cascades is controlled to be Vane A-
aligned, mid-pitch-aligned and Vane B-aligned respectively.
The right subplot in Figures 9 to11 shows the film-cooling
effectiveness distributions on the endwall with only an
upstream slot leakage gap flow. The blowing ratio in all of the
cases is fixed as M=1.0. When the swirler is in the Vane A-
aligned position, the cooled area is slightly larger in the red
curves of the contour, and the cooling effectiveness is uniform
on the endwall surface. For the mid-pitch case shown in
Figure 10, near the PS corner region, the cooled area is
smaller. When the swirler is in mid-pitch position, an
apparently unprotected area can be found near the PS corner
region, and this area is covered by the endwall cooling hole
injection in the other two cases where the swirler is facing the
vanes. The right subplot in Figure 11 shows the film-cooling
effectiveness distributions on the endwall with only leakage
gap flows when the swirler is located in the Vane B-aligned
position. Differently from the Vane A-aligned and mid-pitch-
aligned cases, the high film-cooling effectiveness area near
the PS stagnation region (Vane A) is obviously larger than the
other two cases, whereas the film-cooling effectiveness in the
mid-pitch and Vane B stagnation regions is very low.
Although valuable insight can be obtained from the
distribution maps (Figures 9 to 11), the spanwise averaged
plots (Figures 12 to 14) offer additional insight and provide
clear comparisons for large amounts of data. The
effectiveness is averaged from the SS to the PS (Figures 9 to
11) of the passage in the axial chord direction. The data
outside the airfoil was deleted from the averaged results. The
peaks in the plot correspond to the film-cooling holes’
location. Figures 12 to 14 indicate that, with different swirler
positions, the endwall film-cooling effectiveness changes
noticeably at the upstream half area. The largest film-cooling
effectiveness difference appears at Cax=-0.12 where the slot is
located. The averaged results of Vane A-aligned and mid-
pitch-aligned cases are significantly higher because the
coolant injected from the slot covers the endwall sufficiently,
especially near the Vane B corner region and the middle-pitch
area. The rotating flow effect is clearly seen on the upstream
half (axial chord position between -0.12 and 0.6) of the
endwall.
VaneA M=full 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 0.2 0.4 0.6 0.8
VaneA M=slot 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 0.2 0.4 0.6 0.8
Figure 9 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE SWIRLER POSITION IS VANE A-ALIGNED, WITH AND WITHOUT
Vane B
Vane A

8. 8
ENDWALL FILM-COOLING HOLE INJECTION, VANE A AND VANE B ARE
MARKED IN FIGURE 9)
MidPitch M=full 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 0.2 0.4 0.6 0.8
MidPitch M=slot 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 0.2 0.4 0.6 0.8
Figure 10 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE SWIRLER POSITION IS MID PITCH ALIGNED, WITH AND WITHOUT
ENDWALL FILM COOLING HOLES INJECTION)
VaneB M=full 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 0.2 0.4 0.6 0.8
VaneB M=slot 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 0.2 0.4 0.6 0.8
Figure 11 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(THE SWIRLER POSITION IS VANE B-ALIGNED, WITH AND WITHOUT
ENDWALL FILM-COOLING HOLE INJECTION)
Baseline M=full 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 0.2 0.4 0.6 0.8
Baseline M=slot 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 0.2 0.4 0.6 0.8
Figure 12 FILM-COOLING EFFECTIVENESS DISTRIBUTION ON ENDWALL
(BASELINE CASE WITHOUT SWIRLER, WITH AND WITHOUT ENDWALL
FILM COOLING HOLES 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=full VaneA
i= 0deg M=full MidPitch
i= 0deg M=full VaneB
i= 0deg M=full Baseline
Figure 13 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL WITH FILM COOLING HOLE INJECTION(THE BLOWING
RATIO IS 1.0, WITH ROTATING FLOW AT THREE RELATIVE POSITIONS,
VANE A AND VANE B ARE MARKED IN FIGURE 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
i= 0deg M=slot VaneA
i= 0deg M=slot MidPitch
i= 0deg M=slot VaneB
i= 0deg M=slot Baseline
Figure 14 LATERALLY AVERAGED FILM-COOLING EFFECTIVENESS ON
THE ENDWALL WITHOUT FILM COOLING HOLE INJECTION (THE
BLOWING RATIO IS 1.0, WITH ROTATING FLOW AT THREE RELATIVE
POSITIONS)
VaneA M=full i= 0degZ/ZP
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 15 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE GILL REGION WITH ENDWALL FILM-COOLING HOLE
INJECTION (THE BLOWING RATIO IS 1.0, WITH THE ROTATING FLOW
POSITION VANE A-ALIGNED)

9. 9
MidPitch M=full i= 0degZ/ZP
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 16 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE GILL REGION WITH ENDWALL FILM-COOLING HOLE
INJECTION (THE BLOWING RATIO IS 1.0, WITH THE ROTATING FLOW
POSITION MID PITCH-ALIGNED)
VaneB M=full i= 0degZ/ZP
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 17 FILM-COOLING EFFECTIVENESS DISTRIBUTION NEAR THE
PRESSURE SIDE GILL REGION WITH ENDWALL FILM-COOLING HOLE
INJECTION (THE BLOWING RATIO IS 1.0, WITH THE ROTATING FLOW
POSITION VANE B-ALIGNED)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation1
Z/Pitch
i= 0deg M=full VaneA
i= 0deg M=full MidPitch
i= 0deg M=full VaneB
i= 0deg M=full Baseline
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 18 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITH ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 1 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=-0.12, LOCATION 1 IS IN
FIGURES 9-11, VANE A AND VANE B ARE MARKED IN FIGURE 9)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation1
Z/Pitch
i= 0deg M=slot VaneA
i= 0deg M=slot MidPitch
i= 0deg M=slot VaneB
i= 0deg M=slot Baseline
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 19 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITHOUT ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 1 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=-0.12, LOCATION 1 IS IN
FIGURES 9-11)
With the endwall film-cooling, the momentum of the
coolant is high enough to cover the endwall surface though
the cooled area near PS is limited to a small region as shown
by the red curves. The inlet rotating flow position, however,
can significantly influence the passage vortex then influence
the local blowing ratio of near PS film-cooling hole. In the
Vane A-aligned and Vane B-aligned cases a higher local
blowing ratio leads to more coolant being injected from the
endwall cooling holes near the PS corner region such that the
cooled area becomes wider on the endwall (represented by
red curves in Figures 15 to 17; one film-cooling hole is
located near the pressure side). 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 the coolant to move along the corner vortex
towards the main passage. The film-cooling effectiveness
distributions indicate that the cooling performance of near the
PS hole is enough to cool the high pressure area. With a
proper swirler position such as leading edge-aligned (Vane A
or B), the injection can cover the area near the pressure side
combined with the leakage flow from the upstream slot,
whereas the corner region is still exposed to the hot
environment when the blowing ratio swirler is facing the
middle-pitch area. Vane A- or Vane B-aligned swirler
positions could obviously improve the cooling effectiveness
near PS, so the performance near the corner region is satisfied
by the proper design of the swirler position.
Figures 9 to 11 and Figures 12 to 14 indicate the
difference in film-cooling effectiveness distribution in the
upstream and downstream areas of the endwall. When the
swirler position is Vane A-aligned, as shown in Figures 9 and
15, the injection is uniform and can sufficiently cover the
endwall surface, especially at the inlet and near pressure side
region. Yet in the Vane B-aligned case, shown in Figures 11
and 17, the coolant can hardly inject near the Vane B

10. 10
stagnation line and middle-pitch region (pitch between 0.0
and 0.6, axial chord between -0.12 and zero; the leakage gap
is located upstream of the cascades). This phenomenon shows
that in the Vane B-aligned case, the leakage flow cannot
overcome the rotating inlet flow factor in this area where the
corner vortex and secondary flow weaken the Vane B
stagnation line and middle-pitch area film-cooling. This
condition obviously changes in the mid-pitch-aligned case, as
shown in Figures 10 and 16. When the swirler position is
mid-pitch, the coolant can inject near the Vane B leading edge
and cooling holes near the middle-pitch area, whereas the
film-cooling effectiveness is low near the Vane A pressure
side, especially at the leading edge. This indicates that the
endwall film-cooling is sensitive to the swirler position.
Figure 14 shows the behaviour of the injection flow near the
Vane A pressure side is apparently influenced by the inlet
rotating flow condition. The coolant from the near pressure
side cooling hole moves along the passage vortex and then is
mixed with the main passage cooling flow which causes the
film-cooling effectiveness near the pressure side to be lower,
especially in the mid-pitch-aligned case.
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation2
Z/Pitch
i= 0deg M=full VaneA
i= 0deg M=full MidPitch
i= 0deg M=full VaneB
i= 0deg M=full Baseline
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 20 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITH ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 2 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=0.13, LOCATION 2 IS IN
FIGURES 9-11, VANE A AND VANE B ARE MARKED IN FIGURE 9)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
SS
PS
StreLocation2
Z/Pitch
i= 0deg M=slot VaneA
i= 0deg M=slot MidPitch
i= 0deg M=slot VaneB
i= 0deg M=slot Baseline
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 21 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITHOUT ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 2 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=0.13, LOCATION 2 IS IN
FIGURES 9-11)
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
StreLocation3
Z/Pitch
i= 0deg M=full VaneA
i= 0deg M=full MidPitch
i= 0deg M=full VaneB
i= 0deg M=full Baseline
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 22 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITH ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 3 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=0.71, LOCATION 3 IS IN
FIGURES 9-11, VANE A AND VANE B ARE MARKED IN FIGURE 9)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2 SS
PS
StreLocation3
Z/Pitch
i= 0deg M=slot VaneA
i= 0deg M=slot MidPitch
i= 0deg M=slot VaneB
i= 0deg M=slot Baseline
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 23 LOCAL FILM-COOLING EFFECTIVENESS DISTRIBUTION ON THE
ENDWALL WITHOUT ENDWALL FILM COOLING HOLE INJECTION AT
STREAMWISE LOCATION 3 (THE BLOWING RATIO IS 1.0, WITH ROTATING
FLOW AT THREE RELATIVE POSITIONS , Cax=0.71, LOCATION 3 IS IN
FIGURES 9-11)
The phenomenon captured in this experiment has a close
relationship with the secondary flow field in the turbine
cascade. Previous literature provide important supporting
material. The research by Rehder and Dannhauer [18]
indicates that the coolant flow has an 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, can simultaneously force the coolant on the endwall
to move onto the airfoil suction side. Similar results were
found in the research report by Papa and colleagues [10].
They captured the phantom cooling phenomenon on the rotor

11. 11
blade suction side and the coolant was ejected from an
upstream slot. The paper indicates that the coolant from the
endwall moves towards the suction side and then forms 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. These comparable results provide a reasonable
explanation of the over-cooling phenomenon near the suction
side in this experiment.
Figures 18 and 19 compare the local film-cooling
effectiveness distribution at streamwise location 1 with
different swirler positions. The position of the computing area
is indicated by the PS to SS white line along the pitch
direction in Figures 9 to 11 (line marked as ‘1’). With the
mid-pitch-aligned swirler, the local film-cooling effectiveness
apparently improves near the suction side, as shown in
Figures 18 and 19 where the curve representing the mid-pitch
swirler cooling condition is apparently higher near the SS.
Meanwhile, the film-cooling effectiveness in the main
passage and near the PS is improved by the Vane B-aligned
swirler. The well-protected region is limited to the SS corner
region. After cooling the SS comer region, the coolant
strongly interacts with the secondary flows such as the
passage vortex and horseshoe vortex. The main flow
eliminates the momentum of the pressure side film-cooling
quickly, which makes the film-cooling effectiveness off the
SS lower. 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 difference to be smaller near the PS corner
region.
Figures 20 and 21 compare the local film-cooling
effectiveness distribution at streamwise location 2 with
different swirler positions. In the mid-pitch-aligned case, the
film-cooling effectiveness apparently improves near the
suction side. The higher effectiveness area is close to the
middle-pitch area. A well-protected region is that near the SS
area (pitch is between zero and 0.4). 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 unprotected area near PS. The main
flow still mixes the coolant and the hot gas in the passage,
which causes the injection flow to lift off the endwall surface,
which in turn causes the film-cooling effectiveness to barely
change at the PS corner region of the endwall.
Figures 22 and 23 show the local film-cooling
effectiveness distribution at streamwise location 3 where the
coolant is moved to the downstream part of the endwall, with
the swirler position controlled to be Vane A-aligned, mid-
pitch-aligned and Vane B-aligned. When the endwall film-
cooling holes are used (Figure 22), an apparently unprotected
area can be found at the PS corner region (pitch is between
0.9 and 1.0), and the influence of the inlet rotating flow can
still be observed in this area, the downstream part of the
endwall. The mid-pitch-aligned swirler can improve the
endwall film-cooling near SS at this streamline position. This
indicates that the effects of the inlet rotating flow are still
apparent in the downstream part of endwall surface when the
positions of the swirler differ. 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 inlet rotating flow
on the endwall. The film-cooling effectiveness curve
representing the case of the mid-pitch-aligned swirler is
obviously above the curves representing the vane leading
edge-aligned cases in the near SS region (pitch is between
zero and 0.4 ) as shown in Figure 22. As the swirler position
changes, the influence of the inlet rotating flow is apparently
weakened by the secondary flow. The difference between the
three curves is not as obvious as in the former comparisons at
the other stream locations. Figure 23 compares the local film-
cooling effectiveness distribution in the downstream area
when endwall film-cooling holes are blocked. The figure
shows that the upstream leakage flow can not arrive at
downstream part of the endwall.
CONCLUSIONS
In general, the swirler position affects the coolant
distribution on the endwall surface. The results show that
when the blowing ratio is fixed, the film-cooling effectiveness
changes slightly at the upstream part of the endwall with and
without main passage discrete hole film-cooling. No lift-off
and reattachment phenomenon is observed. The film-cooling
effectiveness increases as the axial chord increases, indicating
that the film-cooling ejections mix strongly with the main
flow near the leading edge, thus forming a region of low film-
cooling effectiveness. When the blowing ratio is fixed and the
swirler core changes, the difference is mainly obtained in the
upstream part on the endwall surface near the combustor
turbine leakage. The influence of the swirler migration could
also be detected in the main passage and downstream part but
was not as strong as in the upstream area.
As the clocking position varies from Vane A-aligned to
mid-pitch-aligned to Vane B-aligned at a fixed blowing ratio,
the film-cooling effectiveness first increases then decreases
near the inlet region on the endwall. Our conclusions are as
follows: (1) the anticlockwise direction of rotating flow
influences the endwall film-cooling effectiveness, especially
for the upstream part of the endwall which is mainly covered
by the leakage flow; (2) the film-cooling effectiveness in the
neighbouring passages is different because of the different
relative position of inlet rotating flow core; (3) the film-
cooling performance at the downstream part of the endwall is
partly influenced by the upstream rotating flow inlet as well.
NOMENCLATURE
C =concentration of gas / actual chord length of scaled-up
blade profile
D =film-hole diameter, mm

12. 12
i =incidence angle
I =light intensity
L =length of film hole, mm
M =blowing ratio, ρcVc/ρ∞V∞ & Measurement Condition
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 =background value
c =coolant fluid
in =inlet
mix =mixture condition
O2 =pure oxygen
ratio =partial pressure of oxygen
ref =reference value
sp =spanwise
 =free stream condition
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