HANOI UNIVERSITY OF SCIENCE AND TECHNOLOGY
THESIS
REDUCE TIP LEAKAGE FLOW USING
SQUEALER TIP IN AN AXIAL
TURBINE
DO DINH CHINH
ID: 20202603M
CLASS: 20BKTHK
Advisors:
PhD. Dinh Cong Truong
Department:
Faculty:
Department of Aerospace Engineering
School of Transportation Engineering
Hanoi, 05/2021
Advisor’s sign
CỘNG HÒA XÃ HỘI CHỦ NGHĨA VIỆT NAM
Độc lập – Tự do – Hạnh phúc
BẢN XÁC NHẬN CHỈNH SỬA LUẬN VĂN THẠC SĨ
Họ và tên tác giả luận văn: Đỗ Đình Chinh
Đề tài luận văn: Giảm xoáy đầu mút cánh turbin dọc trục sử dụng đầu
mút lõm.
Chuyên ngành: Kỹ thuật Hàng không
Mã số SV: 20202603M
Tác giả, Người hướng dẫn khoa học và Hội đồng chấm luận văn
xác nhận tác giả đã sửa chữa, bổ sung luận văn theo biên bản họp Hội đồng
ngày 06/09/2021 với các nội dung sau:
1. Chỉnh sửa hình thức, lỗi chế bản (lỗi đánh máy, tại phần mục lục
thêm phần kết luận và tài liệu tham khảo)
2. Bổ sung danh mục từ viết tắt, kí hiệu
3. Đánh số các phương trình và dẫn nguồn các phương trình
4. Đổi tên tiêu đề mục 2.1
5. Làm rõ thêm ý nghĩa của giá trị y +
6. Thêm biểu diễn w trong hình 33
Ngày 10 tháng 9 năm 2021
Giáo viên hướng dẫn
Tác giả luận văn
CHỦ TỊCH HỘI ĐỒNG
1. Student’s information:
Full name: Do Dinh Chinh
ID: 20202603M
Email:
[email protected]
Class: 20BKTHK
The project was done at Department of Aerospace Engineering
Date of assignment:
14/12/2020
Date of completion:
15/07/2021
2. Missions of the thesis:
Investigate the effect of tip clearance, squealer tip on aerodynamic
performance in an axial turbine.
3. Student’s statement:
I assure that this thesis was my independent research under the instructions
of my advisor PhD. Dinh Cong Truong. This research is not a copy from any
previous research paper.
Hanoi, 15 July 2021
Do Dinh Chinh
Acknowledgements
This research is conducted as a fulfillment to pursue a Master degree at the Hanoi
University of Science and Technology, Department of Aerospace Engineering.
First I would like to thank PhD. Dinh Cong Truong for his enthusiastic and
dedicated assistance through each stage of the process. Without his dedicated
instructions, I could have not completed this project. I would like to express my
sincere thanks to my friends and my colleges for advises and encouragements in
this time. I am grateful to my parents who taught me to cherish excellence.
Without their support, this work would have not been completed. This is the first
time I did a project in this field of study so it is inevitable that there are some
shortcomings. Finally, I would like to thank the commitee for take the time in
reading this research work and I look forward to receiving the comments and
corrections to complete this study.
Student
Do Dinh Chinh
TABLE OF CONTENTS
CHAPTER 1. INTRODUCTION ....................................................................... 1
1.1
Introduction ............................................................................................... 1
1.2
Previous research ....................................................................................... 2
1.3
Tip clearance of rotor blade ....................................................................... 4
1.4
Tip leakage flow ......................................................................................... 5
1.5
Flat tip and squealer tip ............................................................................. 8
CHAPTER 2. NUMERICAL ANALYSIS ....................................................... 10
2.1
Turbine model .......................................... Error! Bookmark not defined.
2.1.1
Turbine “LISA”..................................................................................... 10
2.1.2
Stator and rotor blade geometry ............................................................ 11
2.2
Numerical method .................................................................................... 16
2.2.1
Turbine performance curves ................................................................. 16
2.2.2
The fundamental equations of fluid dynamics ...................................... 17
2.2.3
Simulation procedure ........................................................................... 21
2.3
Meshing .................................................................................................... 22
2.4
Boundary conditions ................................................................................ 25
2.5
Convergence criteria ................................................................................ 27
CHAPTER 3. RESULTS AND DISCUSSIONS.............................................. 28
3.1
Grid dependency test and validation ........................................................ 28
3.2
Pressure, velocity and temperature contours............................................ 32
3.3
Effect of tip clearance .............................................................................. 36
3.4
Reduce tip leakage flow using squealer tip .............................................. 38
3.5
Effects of the squealer tip on aerothermal performance …………….. 38
CONCLUSION AND FUTURE WORK
REFERENCES
TABLE OF FIGURES
Figure 1 Location of turbine in aircraft engine ..................................................... 1
Figure 2 High pressure shrouded (left) and unsrhouded (right) turbine rotor
blades ..................................................................................................................... 5
Figure 3 Illustration of tip leakage flow over a flat tip ......................................... 6
Figure 4 Outline of the flow in the region of an unshrouded turbine rotor blade . 6
Figure 5: Rotor tip with flat and squealer show clearly the cavity squealer tip
with a small figure at the tip region ....................................................................... 8
Figure 6: Schematic view of LEC’s LISA research axial turbine ........................ 10
Figure 7: Sketch of the turbine first stage with the relevant dimensions............. 11
Figure 8: Rotor blade (left) and stator blade (right) ........................................... 12
Figure 9: Stator blade geometric parameters and profile pressure distribution 12
Figure 10: Rotor blade geometric parameters and profile pressure distribution 12
Figure 11: Stator and rotor blade design in Ansys Design Modeler ................... 14
Figure 12: Conceptual view of rotor blade without squealer tip (WST) and with
cavity squealer tip (CST) ...................................................................................... 15
Figure 13: Rotor blade with cavity squealer tip and fillet radius at the hub ...... 15
Figure 14: 3D mesh of the stator blade ............................................................... 22
Figure 15: 3D mesh of rotor blade ...................................................................... 23
Figure 16: Diffuser computational domain ......................................................... 23
Figure 17: The computational domain of turbine with WST .............................. 24
Figure 18: 3D mesh of the computational domain with CST .............................. 25
Figure 19: Complete computational domain when mirrored around the
rotational axis ...................................................................................................... 25
Figure 20: Mesh dependency test results ............................................................. 29
Figure 21: 𝑦 + contours on stator and rotor blade for Mesh 1 to Mesh 4 .......... 29
Figure 22: Measured point and computed total pressure ratio compare two
interface cases and Blanco’s computed results ................................................... 30
Figure 23: Measured point and computed adiabatic efficiency compare two
interface cases ...................................................................................................... 32
Figure 24: Static pressure contour of the stator and rotor at the mid-span plane
.............................................................................................................................. 33
Figure 25: Relative Mach number of the stator and rotor at the mid-span plane
.............................................................................................................................. 33
Figure 26: Total pressure and Static entropy at the outlet plane of the stator ... 34
Figure 27: Total pressure and Static entropy at the outlet plane of the rotor..... 34
Figure 28: Temperature contour on the surface of stator and rotor blades ....... 35
Figure 29: Static pressure contour at the end wall of the rotor .......................... 35
Figure 30: Flow visualization of the recirculation bubble over the rotor tip
surface .................................................................................................................. 36
Figure 31: Peak efficiency at different rotor blade tip clearance ....................... 37
Figure 32: Efficiency at mass flow rate of 11.7 kg/s ........................................... 37
Figure 33: Squealer tip parameters ..................................................................... 38
Figure 34: Pressure and Static entropy contour at the rotor outlet plane .......... 39
Figure 35: Aerothermal performance of LISA turbine with cavity squealer tip.. 42
Figure 36 Distribution of temperature [K] on the stator blade ........................... 43
Figure 37: Pressure [Pa] contours on the shroud casing of rotor blade without
squealer and with w/τ = 100% ............................................................................. 43
Figure 38: Streamline through the tip clearance of rotor blade with w/τ =100%
.............................................................................................................................. 45
Figure 39: Static entropy contours on the blade with w/τ = 100% ..................... 45
Figure 40: Nu contours on the blade with h/τ= 150% ........................................ 47
Figure 41: Nu contours on the blade with w/τ = 200% ....................................... 47
LIST OF TABLES
Table 1 LISA research turbine facility controlling parameters ........................... 11
Table 2 Design parameters of the first stage blades ............................................ 13
Table 3: Design specifications of cavity squealer tip .......................................... 16
Table 4 Measured operating condition at turbine design .................................... 16
Table 5 Thermodynamic properties of the gas used in the CFD analysis ........... 26
Table 6 Boundary conditions in CFD analysis .................................................... 26
Table 7: Mesh dependency test results ................................................................ 28
Table 8 Pressure ratio compared to Blanco’s results and measured point ......... 31
Table 9 Maximum efficiency and stall point of turbine stage .............................. 32
Table 10 Flow angle at the inlet and outlet locations.......................................... 36
Table 11 Values of tip clearance investigated and computed adiabatic efficiency
.............................................................................................................................. 37
Table 12 Aerodynamic performance at different cases ....................................... 39
Table 13: Effect of cavity squealer on aerodynamic and aerothermal
performances for LISA turbine............................................................................. 41
CHAPTER 1. INTRODUCTION
1.1 Introduction
In the aviation industry, increasing the performance of aircraft is the most
important thing to improve aircraft operating cost and reduce emissions. In
addition to improvements in aerodynamics and materials of the structure, engine
improvement is the top concern of many studies. Turbines are always mentioned
as the essential part in the engine, directly affecting the performance of the
engine. Research to improve turbine efficiency plays an important role in
increasing overall engine performance. Turbine in general and turbine blades in
particular are the parts operating under extreme conditions: continuous high
temperatures, aerodynamic loads and large centrifugal forces. Therefore,
experimenting with turbines in particular and aviation engines in general faces
many difficulties in terms of cost and equipment. The methods of numerical
simulation have been created to solve this difficulty thanks to the use of
calculation models based on complex solving equations. In this project, I focus
on analyzing the performance of a turbine stage using CFD simulation. The
location of turbine is shown in the following Fig. 1.
Figure 1 Location of turbine in aircraft engine
Improving turbine performance is an issue of great concern in the jet engine
and power sector. Researches in turbine blade technology can be categorized as
reducing the tip clearance, casing grooves, airflow injection and so on. One of
these methods is the use of squealer tip to reduce tip leakage losses. This paper
presents an analysis of the squealer tip configuration, in which the leakage flow
through the tip gap was extensively investigated using computational fluid
dynamics (CFD) methods. The effects on aerothermal performance of the axial
1
turbine were evaluated based on efficiency and Nusselt number. The turbine
studied in this investigation is an axial annular turbine named “LISA”, which was
experimentally tested at the Laboratory for Energy Conversion (LEC), Institute
of the ETH Zürich, Switzerland. Numerical calculations have been performed
using 3-D Reynolds Averaged Navier-Stokes (RANS) equations with the shear
stress transport (SST) turbulence model and “total energy” option with “mixingplane” option between rotor and stator interfaces. The impact on aerothermal
performance and leakage loss of various geometric parameters related to the
height and width of cavity on tip are also discussed. The numerical results
showed that the created vortex directly affects the turbine’s aerothermal
performance and most of the different sizes of the cavity gave higher
performance than the original case without squealer tip with a maximum of
0.88% and 9.64% increase in efficiency and averaged Nusselt number. This
research work used the CFD simulation to investigate the effect of tip clearance
on aerodynamic performance of an axial turbine. Then apply two methods which
are using squealer tip to improve the performance of the turbine.
1.2 Previous research
Flow structure in turbomachinery passages is extremely complex. In turbine,
rotor is a rotational part therefore there is always a small space between the rotor
tip and casing called tip clearance or tip gap. Some research declared that this tip
clearance produces a lot of losses and vortices, so it reduces the performance of
the turbine. As we can see the curved passages, the clearance between the blades
and the end walls give rise to non-uniform velocity profiles, pressure gradients
and temperature gradients. These unsteady flows generate leakage flows
therefore reduce efficiency of the turbine.
Currently, there are many studies on aerodynamic enhancing methods to limit
the influence of the tip leakage flow. One of the most popular methods is to
design the squealer tip for the blade. In one study, Heyes et al. [19] showed that
blade tip geometry had a positive effect on the aerodynamic performance of axial
turbine cascades by limiting the undesirable effects of the tip leakage flow. In
terms of thermodynamics, Ameri et al. [20] showed that a squealer tip directly
slowed down the leakage flow, while also increasing the total heat transfer
2
coefficient compared to the original design. An experimental study by Camci et
al. [21] showed that the suction side squealer offered a better aerodynamic
performance with respect to cavity squealer in a single-stage, low speed, rotating
axial flow turbine, which was studied at Pennsylvania State University. A
numerical simulation was performed by Kavurmacioglu et al. [22] and the
authors pointed out a reduction in aerodynamic loss on a suction side squealer
when compared to the conventional flat tip. Key and Arts [23] compared the flat
tip with the cavity squealer tip, which was based on the flow characteristics at
both low and high-speed conditions in a linear cascade. They discovered that
squealer tips would lower aerodynamic loss in the case of the flat tip under
specified conditions. An experiment by Newton et al. [24] measured the heat
transfer coefficients and the pressure coefficients in a linear cascade with flat tip,
suction side squealer and cavity squealer. Their results indicated a fall in heat
transfer when using the squealers. Krishnababu et al. [25] investigated the effects
of the blade tip’s geometry and concluded that cavity tip increased the
aerodynamic performance and heat transfer. Lee and Kim [26] studied the
influences of the tip gap’s height on aerodynamic performance when using a
cavity squealer tip in a linear cascade turbine. Schabowski and Hodson [27]
investigated the aerodynamic effects of various tip designs in a low-speed linear
turbine cascade and found that the cavity squealer tip led to a lower aerodynamic
loss.
At present, the number of studies on the simultaneous effects of aerodynamics
and heat of the tip leakage flow is very limited. When Lee and Chae [26] studied
the effects of squealer rim height on aerodynamic losses downstream of a highturning turbine rotor, they came to the conclusion that by increasing the rim’s
height, the aerodynamic loss height reduced until the squealer rim’s height-tochord ratio reached 2.75%. Zhou and Hodson [27] conducted the experimental
and numerical works of the squealer geometry’s effect on the aerothermal
performance of the tip leakage flow of cavity tips. They reported that squealer
height affected aerodynamic loss complexly, and the heat transfer coefficient
reduced with increasing the height and reducing the width while reducing the
width mitigated aerodynamic loss. Kang and Lee [28] investigated the effects of
3
squealer rim’s height on the heat transfer on the floor of cavity squealer tip in a
high turning turbine blade cascade and found that the average heat transfer rate
decreased with an increase of height. Recently, Senel et al. [29] studied the
influence of the squealer’s width and height on the aerothermal performance of a
high-pressure turbine blade with four different squealer height and seven
squealer width values being investigated. The results indicated that proper
squealer’s width and height selection played an important role in improving the
aerothermal performance.
In this study, squealer tip configurations with varied squealer’s width and
height were studied to find the effect of squealer tip on the aerodynamic
efficiency, thermodynamic performance, and leakage mass flow rate of the axial
turbine in comparison with the case without the squealer tip.
1.3 Tip clearance of rotor blade
In normal, the tip clearance is not too big to prevent losses, but it could not
be too small, because the expansion of rotor blade because thermal expansion
and inertial force during operation could damage the casing.
One practiced method of mitigating the over the tip leakage flow is achieved
by introducing a shroud to the rotor blade. In Fig. 2 [11], two high pressure
turbine rotor blades are depicted: one shrouded and the other free-tip. Both rotor
blades present orifices from which fluid (air bled-off of the HP compressor) is
ejected to create a boundary layer of cooled air (~700 K) that protects the metal
(through film cooling) from burning and from the combustion products.
4
Figure 2 High pressure shrouded (left) and unsrhouded (right) turbine rotor
blades
Fig. 2 shows the structure of a shrouded turbine blade (left) and an unshroud
turbine blade (right). Even though the shroud over the rotor increases the
aerodynamic efficiency of the turbine stage, the added weight at the tip of the
blade creates considerable mechanical (mainly centrifugal) stresses at the root of
the blade and to the disc itself. Therefore, the rotational speed of a shrouded
blade will have a lower limit compared to an unshrouded one. Since the work
output is proportional to square of the blade rotational speed (via Euler turbine
equation), an advantage of using unshrouded blades becomes apparent.
Nonetheless, the shroud damps out the blade vibrations which is an advantage as
compared to unshrouded rotor.
1.4 Tip leakage flow
In order to optimize the performance of a turbine stage, it is essential to
minimize aerodynamic losses that occur within it. One of the most important
sources of losses is due to over the tip leakage flow (OTL) in rotor blade.
5
Figure 3 Illustration of tip leakage flow over a flat tip
The over the tip leakage (OTL) flow has its origins on the static pressure
difference that occurs at each side of the rotor airfoil at the tip. In this gap the
fluid is not deflected by the blade and hence does not contribute to the work
output of the stage. The fluid enters the gap on the pressure side of the rotor
blade and continues to the other side where it mixes with the core flow and rolls
up into a vortex. An additional vortex due to the endwall boundary layer of the
casing (outer passage vortex) interacts with the over the tip leakage vortex [15].
These sequences can be observed from the sketch of Fig. 4.
Figure 4 Outline of the flow in the region of an unshrouded turbine rotor blade
Researches have been published recently on the tip leakage flows in turbines
from theoretical study. Study of Rain [1] has found the models of flow through
the tip gap of an axial compressor. Rain has gigured out the structure of flow on
6
the tip gap surface of a compressor. Moore et al. has presented the effect of
Reynolds number to flow on a tip gap. Research showed that there is a large
separated flow ay the sharp edge of blade with the high Reynolds number over
10000. Moore et al [2] also calculated turbulence model with a high Reynolds
number from 100 to 10,000 [2]. Bindon [3] showed the development of vortex
along the leading edge to trailing edge of the blade. The detailed development of
tip clearance loss from the leading to trailing edge of a linear turbine cascade was
measured and the contributions made by mixing, internal gap shear flow, and
endwall secondary flow were identified, separated, and quantified for the first
time. Only 13 percent of the overall loss arises from endwall secondary flow and
of the remaining 87 percent, 48 percent is due to mixing and 39 percent is due to
internal gap shear. All loss formation appears to be dominated by phenomena
connected with the gap separation bubble [3] Yamamoto [4] has found that the
clearance gaps size and the cascade incidences were chosen as the most
important variables affecting the mechanisms. Flows close to the endwall and
inside the clearance were surveyed in great detail using a micro five-hole pitot
tube of 0.6 mm head size [4]. Tallman and Lakshminarayana [5, 6])
demonstrated that reduced tip clearance results in less mass flow through the gap,
a smaller leakage vortex, and less aerothermal losses in both the gap and the
vortex. The structure of the aerothermal losses in the passage changed
dramatically when the outer casing motion was incorporated, but the total losses
in the passage remained very similar [5]. And finally to full turbine test rigs
(Prasad and Wagner [7], Stephan et al. [8], Xiao et al. [9], McCarter et al. [10],
and Blanco [11]). Sjolander [12] presented an overview of the tip leakage flow,
summarized its effect on the performance of axial turbine stage. Storer and
Cumpsty [13] carried out experimental and numerical investigation to explore the
characteristics of tip leakage flow. The results revealed tip leakage flow and
vortex are major source of loss in compressor. Gao et al. [14] studied effect of
three types of casing contouring on aerodynamic performance of an unshrouded
turbine rotor. The curved passages represents secondary flows by the deflection
of vortex tubes in a flow with an initial normal vorticity distribution results in a
streamwise component of vorticity at the exit of the passage [15, 16]. One of the
7
most important sources of losses is due to over the tip leakage (OTL) flow in
rotor blades since it accounts for over one third (>1/3) of the overall turbine stage
losses [17]. At the trailing edge of each turbine blade there is a momentum deficit
of the flow field called wake. Meyer [18] defined a wake as a negative jet
directed at the blade profile trailing edge.
The present work investigated the performance of turbine “LISA” and the
variation of rotor blade tip, using three-dimensional (3D) Reynolds-averages
Navier-Stockes (RANS) equations to find its effect on the aerodynamic
performances.
1.5 Flat tip and squealer tip
Usually in the manufacture of rotor blades, the tip is often made with the
shape flat (Flat tip). Therefore, a number of studies have been conducted to
evaluate the effect of the tip with other profiles on turbine performance. One of
the design improvements that have proven to be effective is the use of a squealer
tip. By piercing down the tip in a specific profile, the losses are reduced. In this
article we will use the same profile as the flat tip. Fig. 5 illustrates the flat tip
(left) and squealer tip (right):
Figure 5: Rotor tip with flat and squealer show clearly the cavity squealer tip
with a small figure at the tip region
In fact the squealer tip method has been studied on compressor blades to
increase aerodynamic performance. And for turbines, researchers have applied it
to improve the cooling capacity of the blades. With that idea, applied to a
specific object, in this project, we conducted a study on changing the end tip
configuration from a flat shape to a squealer shape and evaluate its influence on
the vortex in the turbine passage, and also to improve turbine performance in this
8
case. We will also evaluate the impact of this method on thermal performance in
the later part of the study.
9
CHAPTER 2. NUMERICAL ANALYSIS
2.1 Turbine model
2.1.1 Turbine “LISA”
The turbine studied in this investigation is an axial annular turbine named
“LISA”. This turbine was tested at the Laboratory for Energy Conversion (LEC)
Institute of the ETH Zürich, Switzerland. LISA is a continuously operating
scaled subsonic turbine test rig where the generated power is released to a
generator that ensures stable operating conditions. Thus, not only a steady state
operation of the turbine is accomplished but also lower temperatures and flow
velocities are achieved and therefore intrusive measurement techniques can be
accurately used [9]. Fig. 6 depicted the schematic view of LEC’s LISA turbine
[9]:
Figure 6: Schematic view of LEC’s LISA research axial turbine
The air circulates in a quasi-closed loop; an opening to the atmosphere exists
at the exit of the turbine. The mass flow rate through the compressor is altered by
adjustable inlet guide vanes and is measured by a calibrated venturi nozzle. To
control the turbine inlet temperature the air passes through a water-cooled heat
exchanger. The control of the turbine rotational speed is done with a DC
generator to an accuracy of ±0.1 rpm. Some characteristics of the LISA turbine
are described below [9]:
10
Table 1 LISA research turbine facility controlling parameters
Compressor power
750 kW
Turbine speed (max.)
3000 rpm
Compressor mass
flow rate
Generator power
6 to 13 kg/s
Turbine inlet temperature
33 to 55 ⁰C
400 kW
Turbine exit pressure
Atmospheric
Working fluid
Air
Turbine tip diameter
800 mm
Therefore the measurement planes positions are defined as follows:
𝐿1
𝐶𝑎𝑥,𝑠𝑡𝑎𝑡𝑜𝑟
= 0.5
;
𝐿2
𝐶𝑎𝑥,𝑠𝑡𝑎𝑡𝑜𝑟
= 0.15
;
𝐿3
𝐶𝑎𝑥,𝑟𝑜𝑡𝑜𝑟
= 0.15
Where 𝐶𝑎𝑥 is the axial chord distance of the respective blade rows as
demonstrated in Fig. 7:
Figure 7: Sketch of the turbine first stage with the relevant dimensions
At the position, there are measurement probes to mersure some
characteristics needed.
2.1.2 Stator and rotor blade geometry
Every blade has a different parameters, based on these parameters we design
the turbine using design module ANSYS Design Modeler 19.1. Stator and rotor
blade row is shown in Fig. 8.
11
Figure 8: Rotor blade (left) and stator blade (right)
The stator 1 and rotor profiles at three span sections are depicted in Fig. 9
and Fig. 10 respectively, along with their pressure distribution at designed
operation. Relevant design parameters of both blades are given in Tab. 2 [10].
Figure 9: Stator blade geometric parameters and profile pressure distribution
Figure 10: Rotor blade geometric parameters and profile pressure distribution
Tab. 2 presents some design parameters of the rotor and stator blade:
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