Numerical Analysis of Gas Turbine blade cooling by using internal
Cooling methods
In partial fulfilment of the degree of
MASTER OF TECHNOLOGY
In
COMPUTATIONAL FLUID DYNAMICS
Guided By:
Presented By:
Zozimus Labana
(Asst. Prof)
Aerospace Department
UPES, Bidholi.
Bhavesh N Bhatt
(R500215022)
M.Tech CFD,
Aerospace Department
UPES, Bidholi.
Contents
Introduction
• A turbine blade is the individual component which
makes up the turbine section of a gas turbine. The
blades are responsible for extracting energy from the
high temperature, high pressure gas produced by the
combustor.
• There are two types of cooling methods use:
1.internal cooling: It works by passing cooling air
through passages internal to the blade.
a. Convection cooling
b. impingement cooling
2.External Cooling: It works on technique of passing
cooling air through external surface of blade.
a. Film cooling
b. cooling effusion
c. Pin fin cooling
d. Transpiration cooling
Cooling methods
• Internal and external cooling
Blade cross-section
•For keeping blade secure in this harsh
condition, cooling high pressure air is injected
from the root of the vanes.
•The high pressure injected cooled air is
going through the channel which has shape
according to requirement and significance,
absorbs the heat from the inside wall and
exited from blade tip
•The discharge air which is used for blade
coolant material is when continuously
supplied; it created the thin layers of air with
wall of blade and is actually prevent the
contact of hot gases with directly the wall of
blade.
Objectives
Simulation of 3-D geometry.
Compute blade wall temperature
Static Pressure and temperature distribution around the
Blade wall
Convective heat transfer from Blade wall
Comparison of Blade wall temperature before and after
cooling
Comparison of CHT and FSI results
Numerical Methodology
K-Omega SST Model
•
•
•
•
This model is generally used to predict flows in the near wall region.
K-Omega mathematical model has blending between k-omega and k-epsilon equations.
The equation for inner layer is given as:
It is widely used most robust two-equation model based on eddy viscosity in CFD.
Geometry Details
•
•
•
•
•
•
From literature review the geometry was considered.
NASA C3X Turbine blade profile is used as a geometry.
By using co-ordinates, blade profile is designed in solid works software.
Geometry clean up and computational were done in Ansys design Modeller.
True chord length is 144.9 mm, height is 76.2 mm.
Blade material is Silicon carbide.
Cooling channel
Cylindrical cooling channel
Optimized cooling channel
Grid Generation
•Tetrahedral mesh is generated.
•8 layers of inflation are generated around airfoil wall.
•For the blade no. of elements are 73,011.It is due to for solid domain, no
requirement of large no. of elements.
Prism layers
Parameters
Values
Transition ratio
0.77
No. of layers
8
Growth rate
1.2
Prism layers detail
Domain
Blade
Fluid domain
No of element
73011
961881
No of nodes
19288
192888
Orthogonal quality
0.22
0.22
Mesh statistics
Boundary Details
Sl.
No
01
Description
Value
inlet
3.21 bar
Pressure
02
outlet
periodic
1.925 bar
pressure
03
Inlet Mach
0.17
Pressure
inlet
Number
04
Inlet
783 k
temperature
05
Inlet velocity
95.15 m/s
Pressure
outlet
Boundary details for cooling
channel
Cooling hole
no
Temperature
Diameter
Mass flow rate
1
387
6.30
0.0078
2
388
6.30
0.0066
3
371
6.30
0.0063
4
376
6.30
0.0067
5
355
6.30
0.0065
6
412
6.30
0.0067
7
367
6.30
0.0063
8
356
3.10
0.0023
9
406
3.10
0.0014
10
420
1.98
0.00068
Methodology
START
•The flow conditions which is
secondary data, have been
acquired from research papers
and academic references.
•Ansys 15v and Ansys 18v
student version.
Obtain data from Research
Journals
Theoretical Estimation of
various flow parameters
Modelling Geometries and
Meshing
Importing to Fluent and
carrying out simulation
activities.
Report & Conclusion
STOP
Solver validation
Comparison of vane static pressure
Comparison of vane temperature
Tref = 783 k
Comparison of Heat transfer coefficients
•
•
The present work aims to investigate effects of different
parameters on the blade life in a typical high pressure, high
temperature turbine blade with internal convection cooling.
The case study for heat transfer simulations is the first stage
rotor blade of a mechanical drive gas turbine, which has
complicated internal cooling passages.
Results of some parameters are shown here:
Static pressure
Static Pressure
•
•
•
Here pressure distribution on shroud surface is showing.
Since the gap on the blade, flow starts to accelerate and this
acceleration leads to a thin boundary layer and very high
heat transfer coefficients.
Another drawback of hot gas flowing through the gap is the
increase in losses, and as a consequence loss of efficiency.
Static pressures have their highest values close to the
stagnation point of the blade in leading edge. As flow travels
to downstream, static pressure decreases to finally reach to
the boundary condition value.
After carefully studying the velocity or Reynolds number
with pressure change we can specify the transition of flow
around the blade surface. This will give good insight to flow
characteristics of domain and we can study transition of flow
from laminar to turbulent.
Static Temperature
Static Temperature
•
•
•
•
•
Here at shroud surface temperature distribution around blade
is shown by using SST k-w model. Here maximum
temperature is 791k on leading edge at suction side.
The tip of a gas turbine goes under very large thermal loads
which could result in the tip failure and melt-down of this
section. These high thermal loads are due to the very large
temperatures of hot gases flowing over the gap between the
tip and shroud of turbine.
It is also found that peak temperature is occurs at stagnation
point and temperature is reduced as heat flux is increases.
On blade leading surface and on outer surface temperature is
lower due to conduction in thick blade wall.
There is around 80 k drop in temperature from inlet to exit
of flow domain. The leading edge and starting of suction
side are hottest place in the flow domain. Suction side is
100k hotter than pressure side as suction side is more
exposed to the hot gas.
Heat transfer coefficient
Heat transfer coefficient
•
•
•
From results of htc, we can see that it is in the range of
1200 W/ (m2K) on most area of vane surface but at some
place it is near to 1700 W/ (m2K) which is because high
temperature of hot gas flowing through the vane. Because of
convective heat transfer in blade wall, HTC of some place is
very high.
There is large fluctuation in heat transfer coefficient near the
side of the trailing due to vanes of the cooling effect. There
is increase in HTC with boundary layer transition but
boundary layer is not captured by SST-k-w model. For
capturing the transition in flow we need to use transition
model.
the flow developed along the suction side, there is
decrement in heat transfer because of little increase in
thickness of boundary layers. There is strong decrease in
convective heat transfer because of adverse pressure
gradient which increase the thickness of boundary layers
Temperature before cooling
Temperature before cooling
•
Figure demonstrate the consequence of surface divider
temperature before applying the cooling in gas turbine
vane. Hot gas is streaming around the sharp edge at free
stream temperature. Due convective warmth exchange
surface of vane is get warmed and temperature is about
the same as hot stream gas.
•
After convection heat transfer, heat is also transfer
through conduction in blade metal. So overall vane
temparature raises nearly hot flow gas temparature.
• To overcome the high temparature of vane surface we
need to apply effective cooling method to vanes which
internally cooled the blade reduce the temparature of
vane surface.
Temperature after cooling
Temperature before cooling
•
We can see from results that before cooling temparature of
blade is 790 k on blade tip and on blade suction side it is
around 700k.After cooling effects it is around 550k on
suction surface and on tip it is 650k Temparature is
maximum on blade edges as it is thin in thickness.
Temperature is driving through high temperature at outer
wall to low temperature at inner wall.
•
The employments of ten openings cooling geometry have
great impact on surface temperature of vane. Here because
of warmth sink impact, interior convection is obvious
around center and cooling openings.
Here warmth is expelled inside because of spiral
temperature angle in cooling gaps. This prompts
diminishment in surface temperature of vane driving surface
at weight side. We can reason that in the main edge at first
and second cooling openings cooling is exceptionally
compelling where as at raise edge cooling is not that much
successful because of thin surface of vane.
•
Temperature after cooling
Wall heat flux by CHT
heat flux by FSI
Wall heat flux
•
•
In Fig wall heat flux computed by using Finite element
method is shown. There is maximum heat flux computed
by CFD method is 7.31× 10^5 w/m^2 and by FEA
method is 7.53× 10^5 w/m^2.
In these regions, which can seen near rows III and IV,
higher thermal conductivity of sic actually causes the
wall temparature to increase. In general, the conjugate
conduction reduces wall temperatures where heat flux is
into the wall and increases them where heat flux is into
the fluid. Result of heat flux is shown.
Blade wall temperature by FEA
Blade wall temperature by FEA
•
•
•
•
It is seen from the results that blade metal temperature is
varies from 649 k to 500 k on whole blade cross section
area. This variation is from leading side to trailing side and
in the direction of the linear side.
We can see that temperature at shroud is more than mid
profile of blade surface. After co paring the heat flux and
temperature we can seen that heat flux is maximum where
temperature is maximum.
The uses of ten holes cooling geometry have good effect on
surface temperature of vane. Here due to heat sink effect,
internal convection is evident around hub and cooling holes.
Here heat is removed internally due to radial temperature
gradient in cooling holes.
Maximum temperature on blade wall is same in both by
CHT and FSI
Conclusion
•
•
•
•
•
we applied here conjugate heat transfer simulation for
numerical methodology. The advantages of the conjugate
approach are innately better exactness.
After studied results we can see that pressure is high at
suction side and low at pressure side. Lower pressure gradient
at pressure side indicate leakage flow. Pressure ration is high
at pressure side which indicates higher velocity in that area.
After computing numerical results we compare those results
with experimental results. As we can seen that results
matching good with experimental. Here results of vane
pressure, vane temparature and vane htc are compared and are
showing good agreement with experimental data.
This behavior also indicates that the thermal conduction is not
simply one-dimensional, but heat travels along the vane wall
toward low temperature regions.
These results show that it is too simplistic to assume that
increasing vane conductivity will monotonically drive all
surface temperatures toward the vane mean temperature,
especially for a two temperature mixing problem such as
turbine film cooling.
Future scope
• By optimizing shape and location of cooling
hole we can increase cooling rate.
• Use of different material to check thermal
strength and structure parameters.
• Use of different cooling method internal or
external or mix of both.
• By using ribs or slot in cooling channel we
can increase the heat transfer.
References
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Thank you