International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
Design and Analysis of Gas Turbine Blade
with Varying Pitch of Cooling Holes
Lalit Dhamecha #1, Shubham Gharde*2, Ganraj More#3 , M.J.Naidu*4
#1
Department of Mechanical Engineering Smt. Kashibai Navale College of Engineering, Pune, India.
Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering, Pune, India.
3
Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering, Pune, India.
4
Assistant Professor Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering,
Pune, India
2
Abstract — Gas turbine play a vital role in the
today’s industrialized society, and as the demand for
power increase, the power output and thermal
efficiency of gas turbine must also increase. One
method of increasing both the power output and
thermal efficiency of the engine is to increase the
temperature of the gas entering the turbine. In the
advanced gas turbine, the inlet temperature of
around 1500°C is used , however, this temperature
exceeds the melting temperature of the metal
airfoils. Therefore, along with high temperature
material development, a refined cooling system must
be developed for continuous safe operation of gas
turbines with high performance. Gas turbine blades
is cool internally and externally. This paper is
mainly focus on external cooling of the turbine i.e.
film cooling. In film cooling, relatively cool air is
injected from the inside of the blade which travels
through the entire blade length and form a
protective film around the blade-surface.
In present work attempt has been made to analyse
the failure of the gas turbine blade through
structural analysis. The analysis is conducted for
two different blade configuration one is the base line
configuration with film cooling cylindrical holes
along the entire length and in another configuration
is the holes along the leading edge are branched
together for anti-vortex considerations. The two
configurations is further study for two different pitch
to diameter ratio(y/d) of the cooling holes.
Keywords — Gas Turbine, Film Cooling, Failure
Analysis, Structural Analysis.
I. INTRODUCTION
Gas turbines is the main component in industries like
power generation; processing plant and aircraft
propulsion. In 1960s, by material property limited
the turbine blade and gas turbine firing temperature
around the 800C.But now a day’s gas turbine
engines operate at high temperatures (1200-1500°C)
to improving thermal efficiency and power output.
Due to increasing the gas temperature, the heat
transfer to the blades will also increase significantly
resulting in their thermal failure. With the existing
materials, it is not possible to go to higher
temperatures. Therefore a suitable cooling method
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must be developed for continuous safe operation of
gas turbines with high performance. In order to
employ high gas temperature in gas turbine stages, it
is necessary to cool the casing, nozzles, rotor blades
and discs. Cooling of these components can be
achieved either by air or liquid cooling. By using the
liquid cooling method to face the some
disadvantages of these processes i.e. the problems of
leakage, corrosion. Besides other side, air cooling
method, it’s allows to be discharged into the main
flow without any problem. The blade metal
temperatures can be reduced by around 200-300°C.
By using suitable blade materials (nickel-based
alloys) now available, an average blade temperature
of 800°C can be used. This gives the permit
maximum gas temperatures around 1200-1500°C.
Air cooling method is briefly described in two main
parts i.e. internal cooling & external cooling
Internal cooling of the blade can be achieved
by passing cooling air through internal cooling
passage from hub towards the blade tips. The
internal passages may be circular or elliptical and are
distributed near the entire surface of a blade. The
cooling of the blade is achieved by conduction and
convection. Relatively hot air after traversing the
entire blade length in the cooling passages escapes to
the main flow from the blade tips. A part of this air
can be usefully utilized to blow out thick boundary
layers from the suction surface of the blades. Hollow
blades can also be manufactured with a core and
internal cooling passage. Cooling air enters the
leading edge region in the form of jet and then turns
towards the trailing edge. Cooling of the blade takes
place due to both jet impingement (near the leading
edge) and convection heat transfer.
External cooling of the turbine blades is
achieved in two ways. The cooling air enters the
internal passage from hub towards the tips. On its
way upward it’s allow to flow over the blade surface
through a number of small orifices inclined to the
surface. A series of such holes is provide at various
sections of the blade along their lengths. The cooling
air flowing out of these small holes forms a film
over the blade surface. Besides cooling the blade
surface it decreases the heat transfer from the hot
gases to the blade metal. Another variation of this
method is the blade surface is made of a porous wall
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International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
which is equivalent to providing an infinite number
of orifice. Cooling air is force through this porous
wall which forms an envelope of a comparatively
cooler boundary layer or film. This film around the
blade prevents it from reaching very high
temperatures. Besides the effusion of the coolant
over the entire blade surface causes uniform cooling
of the blade.
The present paper attempts to study the effect of
variation in blade configuration and pitch to
diameter ratio(x/d) of the cooling holes on stress
distribution . The blade is analyses for two x/d ratio
of 3 and 4 . The structural analysis of turbine blade
is done using ansys 16 workbench Solver. The
material used for the analysis is inconel718.
II. LITERATURE SURVEY
Theju V et.al had mainly done the research work
on jet engines turbine blade; the study was done on
two different materials Inconel -718 and Titanium T6; to investigate the effect of temperature and
induced stresses on turbine blade. The study
concluded that the Titanium T-6 would have lesser
value of deformation and lower strength; however if
cost of material is the primary issue then inconal-718
could be selected it would have little higher
deformation and higher strength. Also Inconel have
good material properties at higher temperature than
titanium.
V. Veeragavanet.al had done their research on
aircraft turbine blades, his main focus was on
10C4/60C50 turbine blade models. Conventional
alloys Such as titanium, zirconium, molybdenum
were chosen for analysis .he studied the effect of
temperature on different material for the certain
interval of times and concluded that molybdenum
had better temperature resistance capability.
G Narasa Raju et.al had done research on
different types of cooling techniques which maintain
temperature of bade within allowable limits. Finite
Element Analysis is used to study steady state
thermal and structural performance for N155 and
Inconel-718 .four different models consisting of
solid blade and blades with varying number of
holes(5,9,13) were analysed to find out the number
of cooling holes. They had used two materials
Inconel 718 and Inconel 155 for their research and
found out Inconel 718 has better thermal properties
as blade temperature and stress induce is lesser.
V. NagaBhushana Rao et.al had done research on
turbine blade used in marine applications. The blade
was investigated for structural analysis at elevated
temperatures and under the action of large
centrifugal force, the material used was nickel based
super alloy, it was observed that maximum stresses
and strains were observed near to the root of the
turbine blade and upper surface along the blade root,
maximum temperature is observed at the blade tip
and minimum at the root of the blade; temperature
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distribution is decreasing from tip to root and
temperatures observed are below the melting point
of blade material.
Bhatti et al.performed transient state stress analysis
on an axial flow gas turbine blade and disk using
finite element techniques. They have chosen Inconel
718, a high heat resistant alloy of chromium, nickel
& niobium. The study was focused on centrifugal &
thermal stress arising in the disk.
Deepanraj et.al. had considered that the titanium –
aluminium alloy as the blade material and performed
structural and thermal analysis with varying number
of cooling passages. They further studied the effect
of varying the cooling air temperature on the
temperature distribution in the blades. It is
concluded that the blade configuration with 8 holes
gives an optimum blade temperature of 800C.
A.K.Mattaet.al.studied the stressanalysis for N –
155 & Inconel 718 material. On solid blades it is
reported that Inconel 718 show better results for high
temperature operation.
Ervanet. al.had suggested that high turbine
efficiency can be obtained by minimizing the air
flow required for cooling by effectively utilizing its
cooling potential. He suggested a cooling technology
which has three main parts:- a) The leading edge is
provided with impingement cooling; b) the middle
section of blade contains cooling pipes with
obstacles provided along the length to enhance
turbulence in the cooling air and c)the trailing edge
of the blade is provided with pin – fins for effective
cooling.
G. Narendranath et.al.examine the first stage rotor
blade off the gas turbine analyzed using ANSYS 9.0.
The material of the blade was specified as N155.
Thermal and structural analysis is done using
ANSYS 9.0 Finite element analysis software. The
temperature variations from leading edge the trailing
edge on the blade profile is varying from 839.531C
to 735.162C at the tip of the blade. It is observed
that the maximum thermal stress is 1217 and the
minimum thermal stress is the less than the yield
strength value i.e., 1450.
K haribrahmaiah et.al. Examined the thermal
analysis of gas turbine with four different models
consisting of blade with and without holes and
blades with varying number of holes(5,9&13) were
analyzed. Transfer rate and temperature distribution,
the blade with 13 holes is considered as optimum.
Steady state thermal and structural analysis is carried
out using ANSYS software with different blade
materials of Chromium steel and Inconel- 718.
While comparing these materials Inconel-718 is
better thermal properties and induced stresses are
lesser than the Chromium steel.
V.Vijaya Kumar et.al.examine the ―preliminary
design of a power turbine for maximization of an
existing turbojet engine‖. For proper understanding
of the combined mechanical and the thermal stresses
for the mechanical axial and centrifugal forces. The
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International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
peripheral speed of rotor and flows velocities is kept
in the reasonable range so to minimize losses. In
which the base profiles is analyzed later for flow
condition through any of the theoretical flow
analysis method such as ―potential flow approach.‖
III. Blade Design
Gas turbine is a power producing mechanical device.
Combustion gases from the combustor directly goes
to turbine blade. Blade is the main component of any
gas turbine that extract the power from high
temperature and high pressure gas.
Blade is made of material that can withstand high
temperature. Inconel alloy and titanium are generally
used. Blade cooling is studied by taking Inconel as
material as it is less costlier than titanium and give
performance in the similar way.
A. Material properties
Blade material:-Inconel (Ni-Cr alloy)
Properties
Unit
Inconel
718
Young’s
MPa
2E5
modulus
Density
kg/m3
8190
Poisson’s ratio
0.284
Tensile yield
MPa
1069
strength
Allowable stress
MPa
641.8
Allowable Shear
MPa
385.08
Stress
Specific heat
J/kg-K
586
Thermal
W/m-K
25
Conductivity
B. Steps In Blade Design
The Blade model profile is generated by using
CATIA V5 software. Key Points are created along
the profile in working plane. The points are joined
by drawing Spline curves to obtain a smooth
contour. The contour(2D) model is then converted
into area and then volume(3D) model was generated
by extrusion. The profile was extruded in y direction
up to 117mm i.e. the blade height. The hub is also
generated similar grounds. These two volumes are
then combined into single volume. The blade with
two pitch to diameter ratios (y/d) 3 and 4 are
modeled and analyzed for structural failure.
Where d= Diameter of cooling holes.
y= Distance between two cooling holes.
Fig. 1 Catia Model Of The Blade
C. List Of Selected Key points
SR.
NO
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
X
Z
0.07
80.761
15.055
21.565
28.282
25.188
42.251
49.43
56.672
63.921
71.118
78.211
85.16
91.33
98.51
104.883
104.883
97.195
89.384
81.528
73.687
65.913
58.274
50.886
43.888
37.151
30.187
92.782
15.054
7.216
0.18
-0.768
5.864
9.472
12.672
15.41
17.631
19.282
20.318
20.712
20.452
19.55
18.036
15.954
13.356
10.296
6.83
3.83
5.446
6.283
6.423
5.92
4.783
2.95
0.283
-3.284
7.331
-10.964
13.572
-14.95
-14.78
-11.169
IV.FINITE ELEMENT ANALYSIS OF A GAS
TURBINE BLADE
The turbine blade is analysed for its structural
performance under the action of various forces. The
structural analysis of a gas turbine is carried out
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International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
using ANSYS 16.0 software. Single blade is taken
into consideration for analysis as turbine blades and
is mounted on the periphery of hub symmetrically
along the axis of rotation of the blade. The cross
section of the blade is in the X-Z plane and the
length of the blade is along the Y axis. Centrifugal
forces generated during service by rotation of the
disc were also taken into consideration.
Static analysis was carried out to determine the
mechanical stresses and elongation experienced by
the gas turbine blade. In this analysis, the gas forces
are assumed to be distributed evenly, the tangential
and axial forces acts along the centroid of the blade.
The centrifugal force also acts through the centroid
of the blade in the radial direction.
Tangential Force (Ft) = 980 N
Axial Force (Fa)
= 250 N
Centrifugal Force (Fc) = 282661.8 N
V. RESULTS AND DISCUSSION
Blade failures can be caused by a number of
mechanisms under the turbine operating conditions
of high rotational speed at elevated temperature in
corrosive environments. To identify the causes of
this failures mechanical analysis of the turbine
blade was conducted. This work focuses on failure
analysis of gas turbine blades through structural
analysis. The present work deals with the modelling
and analysis of gas turbine blades. The structural
analysis was performed on the turbine blades using
ANSYS 16.0 software and results were discussed.
The force of high temperature gases impinging on
the turbine blade is divided in two components:
tangential force (Ft) and axial force (Fc). The
Centrifugal, Axial and Tangential forces acting on
the blade are considered as loads in structural
analysis. The maximum mechanical stresses and
elongations
induced in the turbine blade are
observed to be within the safe limit. Maximum
elongations are observed at the tip section of the
blade and minimum elongations at the root Section
of the blade.
Fig. 3 Total Deformation Of Blade With Y/d=4
Fig. 4 Von-Mises Stress Of Blade With Y/d=3
A. Blade With Cylindrical Holes
Fig-5: Von-Mises Stress Of Blade With Y/d=4
Fig. 2 Total Deformation Of Blade With Y/d=3
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International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
B. Blade With Branched Cylindrical
Holes At Leading Edge
Fig. 9 Von-Mises Stress Of Blade With Y/d=4
Fig. 6 Total Deformation Of Blade With Y/d=3
Fig. 7 Total Deformation Of Blade With Y/d=4
Fig. 8 Von-Mises Stress Of Blade With Y/d=3
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Above analysis Shows that the total deformation due
to action of all forces such as axial, tangential and
centrifugal forces is maximum at the tip section of
the turbine blade material and minimum deformation
is observed at the root section. The maximum
deformation observed is 0.22522 mm for blade
having branched holes at leading edge and having
y/d=4. The minimum deformation observed is
0.13485 mm for blade having cylindrical holes and
y/d=4.The Von-Mises due to action of all forces
such as axial, tangential and centrifugal forces is
maximum at the root section of the turbine blade
material and minimum stress is observed at the tip
section. The maximum Stress observed is 1878.5
Mpa for blade having branched holes at leading edge
and having y/d=4. The minimum stress observed is
1122.6 Mpa for blade having cylindrical holes and
y/d=3. It is Observed in all the cases that the
maximum stress occurring at the tip is at a very
small area and that the majority of the blade section
is having stress within the allowable limits. The high
stress zones can be avoided by reducing the number
of holes at the root section or by providing a fillet at
the root to avoid abrupt change in cross-section and
reduce stress concentration in that area.
VI. CONCLUSION
Gas turbine blade with film cooling holes is studied
for its structural performance considering Inconel
718 as the blade material. The analysis was
conducted on ANSYS 16.0 software. Maximum
elongations are observed at the tip section and
minimum at the root section of the blade. The
elongations observed are within the safe limit.
Maximum stresses are observed near the root section
and minimum at the tip section of the blade. The
maximum stresses are observed near the film
cooling holes due to high stress concentration along
the holes. The maximum stress of 1878.5 Mpa
occurs at the trailing edge nearer to the root of the
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International Journal of Engineering Trends and Technology (IJETT) – Volume 34 Number 8- April 2016
blade with y/d=4. This stress exceeds the yield
stress of the material and which might lead to the
failure of the turbine blade. At all other parts of
turbine blade, the stresses induced are within the
same limits.
[6]
Amjed
Ahmed
Jassim,A
L
Luhaibi,Mohammad
Tariq―Thermal Analysis Of Cooling Effect On Gas Turbine
Blade‖,International Journal Of Research in Engineering
And Technology, ISSN:2319-1163.
VII. REFERENCES
[1]
[2]
[3]
[4]
[5]
Theju V,Uday P S,PLV Gopinath Reddy,C.J. Manjunath, ―Design
And Analysis Of Gas Turbine Blade,‖ International Journal of
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2319-8753.
V.NagaBushana Rao,N. Niranjan Kumar,N. Madhulata,A.
Abhijeet, ―Mechanical Analysis Of 1st Stage Marine Gas
Turbine Blade‖,International Journal Of Advanced Science
And Technology vol.68(2014).pp.57-64.
R D V Prasad,G Narasa Raju,M S Srinivasa Rao,N
Vasudeva Rao, ―Steady State Thermal And Structural
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