Improving the Thermal Condition of the High-Pressure Turbine

Blade

V. M. Zubanov

, G. M. Popov

, S. A. Melnikov

, A. I. Sherban

and Liu Xin

Department of Aircraft Engine Theory, Samara National Research University,

34 Moskovskoe Highway, Samara, Russian Federation

Keywords: Aircraft Engine, Turbine Cooling, Cooled Blades, Cooling Efficiency Coefficient.

Abstract: An increase in the temperature of the gases forces the cooling system of the turbine blades to become more

and more complex. The presented article describes a complex and computer-intensive numerical model of a

working blade of a modern high-pressure turbine of a civil aviation aircraft gas turbine engine, which includes

a flow region, a blade body, internal cooling channels and coolant supply channels. Using this model, the

thermal state of the blade was determined and potential problem areas were found: hot gas leakage, coolant

stagnation and overheating. Based on the analysis, several options were proposed for changing the

configuration of the internal channels of the blade, which reduce the negative effects found. Although the

proposed design options did not fully achieve all the requirements for the blade, they made it possible to find

promising ways for further improvement. Also, the authors have practically worked out conjugate numerical

models to study the thermal state of the turbine.

1 INTRODUCTION

The operation of modern gas turbine engines (GTE)

is impossible without a cooling system for high-

temperature gas turbine components. Therefore, the

turbine cooling system includes cooling of the nozzle

blades (NB) and rotor wheels (RW). The cooling

schemes for the NB and RW in the first stage of a

high-pressure turbine (HPT) are the most complex.

They must ensure a component temperature at which

the turbine can operate effectively throughout its full

service life, bearing in mind that the total working

fluid temperature at the combustion chamber outlet

𝑇



∗

can exceed 1800 K in modern engines (Han

J.C.,2012, Inozemtsev A.A., 2006).

Different schemes of cooling air channels, which

contain heat exchange intensifiers, are used to cool

the HPT rotor blades (Kopelev S.Z., 1983, Nagoga

G.P.,1996, Vieser, W., 2002). The design of effective

blade cooling schemes is a time-consuming process,

https://orcid.org/0000-0003-0737-3048

https://orcid.org/0000-0003-4491-1845

https://orcid.org/0000-0002-0170-3846

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https://orcid.org/0000-0002-3137-8247

which is currently difficult to imagine without the use

of CFD tools.

The purpose of this work was to perform a

detailed analysis of the thermal condition of the

cooled HPT rotor blade (RB) using the Ansys CFD

Post software, and to develop and implement

measures to improve the thermal condition of the

blade and increase the efficiency of its cooling

scheme.

In the course of this work, alternative schemes of

internal cooling channels were developed based on

the analysis of the thermal condition of the HPT rotor

blades. Geometric models of the blade under study

were created with preservation of the external shape

and different schemes of internal cooling channels.

Using these models, several series of calculations

were carried out. The results of various cooling

schemes have been compared with each other in order

to assess the effect of changes in the HPT rotor blades

cooling scheme on its temperature condition.

234

Zubanov, V., Popov, G., Melnikov, S., Sherban, A. and Xin, L.

Improving the Thermal Condition of the High-Pressure Turbine Blade.

DOI: 10.5220/0012078400003546

In Proceedings of the 13th International Conference on Simulation and Modeling Methodologies, Technologies and Applications (SIMULTECH 2023), pages 234-242

ISBN: 978-989-758-668-2; ISSN: 2184-2841

 2023 by SCITEPRESS – Science and Technology Publications, Lda. Under CC license (CC BY-NC-ND 4.0)

2 COMPUTATIONAL MODEL

PREPARATION

As the initial data, the results of calculation of the

thermal condition of the initial variant of the HPT RB

cooling scheme were used. This cooling scheme is

referred to hereinafter as the "basic" one, since

modified cooling channels schemes were created

based on it. The geometric model of the "basic" blade

and its inner cavity is shown in Figure 1.

Figure 1: Shape of the basic rotor blade and its inner cavity.

Three-dimensional geometric models of the blade

and gas-air regions were created in the Siemens NX

software package. The RB inner cavity had a complex

system of channels and included different types of

convective cooling intensifiers, shown in Figure 2.

The blade has a total of three internal channels:

front, middle and rear. The rear part of the blade is

cooled by means of a vortex cooling matrix formed

by crossed ribs which are moulded on the inner

surfaces of the pressure and suction side. The coolant

exits from the vortex matrix into the flow path

through 8 slotted windows near the trailing edge. In

the front and middle inner channels there are ribs on

the pressure and suction side walls to intensify

convective cooling. The upper part of the vortex

matrix is fed with coolant by means of additional

openings from the middle channel. There are three

vertical rows of holes near the leading edge of the

blade to form a film cooling on the surface of the

blade. The first row has 16 holes, the second and third

rows have 17 holes each. All holes have a diameter of

0.55mm. The film cooling of the upper blade face is

formed by blowing out a part of the coolant through

four holes of 1 mm diameter on the peripheral end of

the HPT RB.

Figure 2: Blade cooling scheme, top view.

The computational model for the coupled

thermal-hydraulic calculation of the RB was created

in Ansys CFX (Ansys Workbench Product Release

Notes). The CFD model of the RB consisted of four

domains (Figure 3):

 the stationary domain of the NB blade passage

channel;

 the rotating domain of the RW blade passage

channel, also including the inter-disk cavity in

front of and after the RW;

 rotating domain of the front, middle and rear parts

of the inner cavity of the RW;

 the rotating domain of the cooling air inlet to the

blade.

 Inlet 1 is the cross-section at the main flow inlet

to the HPT;

Improving the Thermal Condition of the High-Pressure Turbine Blade

235

 Inlet 2 - cross-section at the inlet to the RB cooling

system;

 Inlet 3 - cross-section at the inlet to the cooling

area of the front surface of the RW disc;

 Inlet 4 - cross-section at the inlet to the cooling

area of the rear surface of the RW disc;

 Outlet - cross-section at the outlet of the main

flow from the HPT.

Figure 3: Computational model for coupled thermal-

hydraulic calculation of the rotor blade.

The numerical values of the initial data were taken

according to the results of the thermodynamic and

hydraulic calculation of the engine, as well as the 1D

calculation of the turbine.

The calculation took into account the fact that in

the model there are 2 working bodies (air and

combustion products), while taking into account that

they depend on the gas temperature (Dorofeev V.M.,

1973).

3 ANALYSIS OF BASIC ROTOR

BLADE COOLING SCHEME

Figures 4 and 5 show calculated temperature

distributions on external and internal surfaces of the

rotor blade at cooling air flow rate 𝐺



= 4.2% of gas

flow rate through throat of inter-blade channels of NB

𝐺

.

The maximum value of the calculated RB

temperature occurs in the periphery area closer to the

trailing edge and is 1136.8 °C.

Figure 6 shows the distribution of temperature

values and their average value in cross-sections

located at different heights of the RB.

The maximum value of the calculated RB

temperature occurs in the periphery area closer to the

trailing edge and is 1136.8 °C./

Figure 4: Temperature distribution on the external surfaces of

the rotor blade.

Figure 5: Temperature distribution on the internal surfaces of

the rotor blade.

Figure 6: Shape of the basic rotor blade and its inner cavity.

Figure 6 shows the distribution of temperature

values and their average value in cross-sections

located at different heights of the RB.

SIMULTECH 2023 - 13th International Conference on Simulation and Modeling Methodologies, Technologies and Applications

236

Based on the obtained cross-sectional temperature

distribution, the following problem areas were

identified for the studied RB.

Firstly, it is the overheated peripheral part of the

blade at the exit edge. And, secondly, there is an area

with temperature difference more than 150°С from

the middle of internal rib to the external surface of the

suction side at the blade bottom (see Fig. 6, section at

relative height from the hub equal to 5%). The

presence of this zone leads to stresses that reduce the

low-cycle fatigue margin.

Increased irregularity of the temperature field in

the blade is caused by vortex zones and stagnation

zones in the cooling channels. Examples are shown in

Figures 7 and 8.

On the basis of the problems described, proposals

have been developed for refining the internal cooling

channels of the RB.

Figure 7: Vortex flow zone in the film cooling channel.

In order to eliminate the vortex in the film cooling

channel, the shape of the channel was corrected,

namely the ratio of the maximum cross-sectional area

of the channel to the inlet cross-sectional area was

reduced by a factor of 1.3. The cross-sectional areas

of the film cooling channel have been modified as

shown in Figures 9 and 10.

To reduce the blade temperature near the trailing

edge, the shape of the internal channels was changed,

through which air entered the matrix and was then

discharged into the flow path of the turbine through

the holes in the trailing edge. This also reduced

hydraulic losses in the new channel configuration.

Figure 8: Stagnation zone in the film cooling channel.

Figure 9: Cross-sectional areas of the film cooling channel

of the basic scheme.

Connecting the film cooling channel to the vortex

matrix channel helped to get rid of the stagnation zone

and to additionally supply the overheated area around

the trailing edge at the blade periphery with cooling

air.

Improving the Thermal Condition of the High-Pressure Turbine Blade

237

Figure 10: Change in cross-sectional area of the film

cooling channel.

However, the film cooling operation was more

difficult to regulate with this arrangement, because

once all the channels were connected, the other two

channels began to influence the film cooling

parameters.

Also, to eliminate overheating near the trailing

edge, the number of holes connecting the middle

channel and the vortex matrix cavity was changed.

Instead of five holes, unevenly spaced across the

blade height, six holes were made without taking into

account the overflow between the film cooling

channel and the vortex matrix channel. In the new

schemes, the holes were positioned directly opposite

the openings in the trailing edge.

4 ANALYSIS OF NEW ROTOR

BLADE COOLING SCHEMES

Figures 11 and 12 show the new cooling scheme

variants, labelled "Variant 1" and "Variant 2".

For the "Variant 1" and "Variant 2" cooling

schemes, calculations are performed at the same

boundary conditions as for the blade with the basic

internal cooling channel scheme.

Figures 13 and 14 shows distribution of cooling

efficiency coefficient θ over blade height of all three

schemes at cooling air flow rate 𝐺



= 4.2% of

𝐺

.

Figure 11: Cooling scheme "Variant 1".

Figure 12: Cooling scheme "Variant 2".

Figure 13: Comparison of cooling efficiency with

different channel schemes.

SIMULTECH 2023 - 13th International Conference on Simulation and Modeling Methodologies, Technologies and Applications

238

Figure 14: Temperature at leading edge of blades with hot

air intake for "Variant 1" and "Variant 2" schemes.

Variant 1 performed better than the Variant 2. In

Variant 2, due to the geometry of the channels, large

vortex and stagnation zones were created, which were

not supplied with cooling air. The vortex inside the

cooling channels can be clearly seen in Figure 15. The

part of the channel that was not receiving cooling air

is also marked here. Therefore, for the subsequent

improvement of the cooling channel system, the

"Variant 1" scheme was selected.

The cross-sectional area of the channel connecting

the film cooling cavity and the vortex matrix was

halved to eliminate hot gas flowing into the blade's

inner cavity in order to increase the static pressure in

the film cooling cavity. The modified "Variant 1.2"

cooling channel scheme is shown in Figure 16.

As can be seen from Figure 20, it was possible to

significantly reduce the trailing edge temperature as a

result of redistributing the cooling air flow in the

vortex matrix.

Comparison of values of average cooling

efficiency coefficient of "Variant 1" and "Variant 1.2"

schemes by sections at different radii is shown in

Figure 17, and distribution of θ by sections at

different blade heights with "Variant 1.2" scheme is

shown in Figure 18. For each blade section the

cooling efficiency coefficient 𝛩



was

determined by the formula:

𝛩





𝑇

 

∗

𝑇

   

𝑇







∗

𝑇

 

∗

where 𝑇

 

∗

is the total gas temperature in relative

motion in front of the rotor blade at the corresponding

radius;

𝑇

 

∗

- total air temperature in relative motion at

the inlet to the lower end of the rotor blade;

𝑇

   

- average cross-sectional

temperature of the blade.

Figure 15: Area with insufficient cooling air.

Figure 16: Cooling scheme "Variant 1.2".

As can be seen from the graphs in Figure 17,

"Variant 1" has a higher cooling efficiency than

"Variant 1.2" according to preliminary calculations.

At the same time, the cooling efficiency of the two

schemes is lower than that of the basic scheme. One

of the reasons for this is that hot gas flows into the

film cooling holes. This can be seen from the

overheated leading edge in Figure 18.

Based on the data shown in the graphs in Figure

17, it can be concluded that the "Variant 1.2" scheme

Improving the Thermal Condition of the High-Pressure Turbine Blade

239

is almost identical to the base one, in terms of integral

values of section cooling efficiency at different radii.

Figure 19 shows a comparison of the blade

temperature distributions over the cross-section at 5%

of the airfoil height in the case of the base variant and

"Variant 1.2". It can be seen that in the "Variant 1.2"

scheme, the temperature gradient from the blade

internal surfaces to the external one was reduced by

changing the internal channel geometry.

Figure 17: Distribution of θ by blade height for basic

scheme and "Variant 1" and "Variant 1.2" schemes.

Figure 18: Distribution of θ across cross-sections at

different blade heights "Variant 1.2" at cooling air flow

rate 𝐺



= 4.2% of 𝐺

.

A comparison of the temperature distribution over

the blade surface with the basic scheme and with the

"Variant 1.2" cooling scheme is shown in Figure 20.

The "Variant 1.2" scheme has eliminated hot gas from

entering the blade cooling system through the film

cooling holes (see Figure 20, pressure side view of the

"Variant 1.2" blade).

Base scheme. "Variant 1.2" scheme.

Figure 19: Comparison of the temperature distribution

across the blade cross-section at 5% of blade height with

the basic cooling scheme and with the "Variant 1.2"

scheme.

Pressure side

view of blade with

‘basic’ cooling

scheme.

Pressure side view of

blade with ‘Variant

1.2’ cooling scheme.

Suction side view

of blade with

‘basic’ cooling

scheme.

Suction side view of

blade with ‘Variant

1.2’ cooling scheme.

Figure 20: Temperature distribution over the blade surface

with "basic" cooling scheme and "Variant 1.2" cooling

scheme.

The leading edge of the blade with cooling

scheme "Variant 1.2" is overheated in the periphery

SIMULTECH 2023 - 13th International Conference on Simulation and Modeling Methodologies, Technologies and Applications

240

area. This can be eliminated by changing the number

and geometry of the film cooling holes in the problem

area or by increasing the radius of the internal film

cooling channel at its 90⁰ turning point.

Also, as a result of changing the geometry of the

cooling channels (as evidenced by the CFD

calculation of the internal channels), an additional

positive effect was obtained by reducing the total

hydraulic resistance in the blade cavity by 116 kPa.

This means that the same cooling air flow rate can be

achieved at a lower pressure drop in the RB cooling

scheme.

For all blade cooling schemes the values of

cooling efficiency coefficient θ at different cooling

air flow rates have been calculated. The results of

these calculations are plotted on the general graph

from study of Shevchenko M. I. (2017), see Figure

21. To be able to compare the calculated data with the

graph, the values of θ have been recalculated using

the temperature 𝑇

 

∗

at the inlet to the cooling

cavity in the HPT disk, and the values of relative

cooling air flow rate were recalculated depending on

air flow rate at the inlet to the high pressure

compressor of the designed engine.

5 CONCLUSIONS

The results of the analysis of the thermal-hydraulic

calculation of the basic HPT rotor blade have been

evaluated and the main directions of its refinement

have been proposed, aimed at reducing the

temperature of the trailing edge and the temperature

difference between the external and internal surfaces

of the blade.

A number of new blade cooling schemes based on

the original geometric model have been developed.

By comparing the results of conjugate modelling of

the new blades with the initial data on the basic blade,

the advantages and disadvantages of alternative

schemes of internal cooling channels of the

investigated RB were revealed.

The result is the "Variant 1.2" scheme, which

eliminated the overheating of the trailing edge and

reduced the temperature difference between the

internal and external surfaces of the blade.

Therefore, the "Variant 1.2" scheme seems to be the

most promising for further development. At the same

time, the disadvantage of this scheme is difficulty of

regulation of its operation in comparison with "basic"

scheme, because all cavities inside the blade are

connected with each other by means of special

connection channels in the scheme "Variant 1.2".

In the future, with the help of the developed

mathematical model, searches will be made for the

shape of the internal channels of the blade, which will

make it possible to obtain its best thermal state. In this

case, rational convective cooling will be found at the

first stage. Then the film cooling will be also

modified.

Figure 21: Comparison of the results of the calculation of the cooling efficiency of the rotor blades in question with the

statistical data (Shevchenko M. I., 2017).

Improving the Thermal Condition of the High-Pressure Turbine Blade

241

REFERENCES

Kopelev S. Z. (1983). The tutorial, Cooled blades of gas

turbines. Thermal calculation and profiling. Nauka

Publisher.

Nagoga G. P. (1996). The tutorial, Effective ways of

cooling blades of high-temperature gas turbines.

Moscow: MAI Publishing House.

Shevchenko M. I. (2017). Dissertation, Design of cooled

GTE parts with advanced verification of thermal-

hydraulic models by the example of cooled gas turbine

blades.

Han J. C., Dutta S., Ekkad S. (2012). The paper, Gas Tubine

Heat Transfer and Cooling Technology. Second

Edition.

Vieser, W. (2002). The paper, Heat transfer prediction

using advanced two-equation turbulence models.

Inozemtsev A. A., Sandratsky V.L. (2006). The book, Gas

Turbine Engines.

Dorofeev V. M. (1973). The book, Thermogasodynamic

Calculation of Gas Turbine Power Plants. Moscow,

Mashinostroenie.

Ansys Workbench Product Release Notes. ANSYS, Inc.

and ANSYS Europe, Ltd. are UL registered ISO

9001:2000 Companies.

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