Heat transfer engineering an international journal, tập 31, số 7, 2010

Heat Transfer Engineering, 31(7):527–554, 2010
Copyright C Taylor and Francis Group, LLC
ISSN: 0145-7632 print / 1521-0537 online
DOI: 10.1080/01457630903425320

Gas Turbine Blade Tip Heat Transfer
and Cooling: A Literature Survey
BENGT SUNDEN and GONGNAN XIE
Department of Energy Sciences, Lund University, Lund, Sweden

Gas turbines are widely used for aircraft propulsion, land-base power generation, and other industrial applications like
trains, marines, automobiles, etc. To satisfy the fast development of advanced gas turbines, the operating temperature must
be increased to improve the thermal efficiency and output work of the gas turbine engine. However, the heat transferred to the
turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important
to cool the turbine blades for a long durability and safe operation. Cooling the blade must include cooling of the key regions
being exposed to the hot gas. The blade tip region is such a critical area and is indeed difficult to cool. This results from the
tip clearance gap where the complex tip leakage flow occurs and thereby local high heat loads prevail. This paper presents a
literature survey of blade tip leakage flow and heat transfer, as well as research of external and internal cooling technologies.
The present paper does not intend to review all published results in this field, nor review all papers from the past to now. This
paper is limited to a review of recently available published works by several researchers, especially from 2001 to present,
concerning blade tip leakage flow associated with heat transfer, and external or/and internal tip cooling technologies.

INTRODUCTION
A gas turbine is an engine designed to convert the energy
of a fuel into some form of useful power, such as shaft power
or thrust. Today, gas turbines (GTs) are widely used in aircraft
propulsion, land-based power generation, and other industrial
applications. For example, GTs are used to power commercial
airplanes, marines, trains, electric power generators, automobiles, and gas pipeline compressor drivers. Figure 1 illustrates
a commercial gas turbine engine. The reasons that gas turbine

engines are widely used for aircraft propulsion include that they
are light, compact, and have a high power-to-weight ratio. As
shown in Figure 1, there are three main components of a gas
turbine engine: compressor, combustor, and turbine. The compressor is used to compress the intake air to a specific high pressure, the combustor is used to burn the input fuel and produce
the high temperature gas, and the turbine extracts the energy of
the gas and converts it into power work. A number of components sometimes occurs in the gas turbine system to improve the
The authors acknowledge financial support from the TURBO POWER consortium funded by the Swedish Energy Agency (STEM), SIEMENS Industrial
Turbomachinery, and VOLVO AERO Corporation.
Address correspondence to Professor Bengt Sunden, Division of Heat Transfer, Department of Energy Sciences, Lund University, PO Box 118, S-22100,
Lund, Sweden. E-mail:

network output or thermodynamic efficiency, e.g., intercooler,
recuperator, regenerator, and combustion reheater. However, the
balance of additional power, efficiency, cost, complexity, durability, compactness, etc. must be carefully evaluated.
The temperature–entropy diagram for the basic cycle of a
gas turbine engine with friction is shown in Figure 2. The ideal
standard cycle is assumed to be adiabatic, reversible, and frictionless. The overall thermodynamic efficiency depends on the
efficiencies of all components, such as compressor efficiency,
turbine efficiency, and combustion efficiency. Clearly, the turbine efficiency will affect the cycle efficiency at some degree.
Thus, improving the turbine efficiency will help to improve the
overall performance of a gas turbine engine, while losses in
turbine efficiency and/or output work will reduce the overall
performance of the system. Apart from the component efficiencies, the operating temperature of gas turbine system affects the
overall performance.
It is well recognized that one way to increase the power output and thermodynamic efficiency of a gas turbine engine is to
increase the turbine inlet temperature (TIT). From the principles
of engineering thermophysics [1, 2], the reason is that at a fixed
pressure ratio the net work output of a gas turbine increases
with increasing turbine blade (also called rotor) inlet temperature. Figure 3 shows recent development of TIT from 1950
to 2010. Current advanced gas turbine engines are operating at

TIT of about 1200–1500◦ C. To pursue higher power, the inlet

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Turbine Inlet Temoperature, K

2600
New cooling concept

2400

Projected trend
new material

2000

Film,Impingement
Sophisticated
cooling system

1600
Simple cooling

1200
Allowable metal temperature

Uncooled turbines

1000
Figure 1 Gas turbine illustration. From />epower/gallery/GasTurbines.htm.

1950

1960

1970

1980

1990

2010

Year
temperature should be raised increasingly to higher certain targets. For example, to double the power of the aircraft, the TIT
should be increased from 1500 to 2000◦ C.
However, the heat transferred to the blade increases with the
increase of the blade inlet temperature, and the allowable melting temperature of materials increases at a slower rate. This
means that the turbine blade inlet temperature may exceed the
material melting temperature by more than 500◦ C. Thus, it is
critical to cool turbine blades for a safe and long-lasting operation. The blades can only survive if effective cooling methods
are used. Various internal and external cooling techniques are
employed to decrease the blade material temperature below its
melting point. Figure 4 depicts the typical cooling technology
for internal and external zones. The leading edge is cooled by jet
impingement with film cooling, the middle portion is cooled by

internal serpentine ribbed-turbulators passages, and the trailing
edge is cooled by pin-fins with ejection. In internal cooling, the
relatively cold air, bypassed/discharged from the compressor,
is directed into the hollow coolant passages inside the turbine
blade. In external cooling, the bypassed air is ejected through
those small holes, which are located in the turbine blade discretely. The commonly used cooling technique for the highpressure turbine blade is a combination of internal and film
cooling. Most recent developments in TIT increase have been

Figure 3 Developments of gas turbine inlet temperature over recent years.
Reproduction from Rolls Royce plc.

achieved by better cooling of the turbine blade and have improved the understanding of the heat transfer mechanisms in the
turbine passages. Several recent publications reviewing the gas
turbine heat transfer and cooling technology investigations are
available. These include a relevant book [3], edited volumes [4,
5], and journal papers [6–9].
Internal convective
cooling
Film
cooling

Hot gas

Tip cap cooling

Trailing edge
ejection

Rib turbulated
cooling

Impingement
cooling

Pin-fin cooling

Cooling air
Figure 2 Temperature–entropy diagram for a basic gas turbine cycle.

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Figure 4 Typical cooling techniques for a blade.

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Shroud
Shroud
Blade
Leakage flow

Clearance Gap

Blade Tip
Pressure side

Moving rotor

Suction side

Rotation

Figure 5 Clearance gap and leakage flow of unshrouded turbine blade.

Cooling of the blade should include the cooling of all regions
exposed to high-temperature gas and thermal load. Among such
regions, particularly for high-pressure turbines, is the blade tip
area. Gas turbine blades usually have a clearance gap between
the blade tip and stationary casing or the shroud (as schematically shown in Figure 5). The clearance gap is necessary to
allow for the blade rotation and for its mechanical and thermal
expansion. However, due to the pressure difference between the
pressure side and suction side, the hot gas leaks through the
gap. This is known as the tip leakage flow. The leakage flow
is undesirable, because it is associated with the generation of
a secondary flow resulting in reduction of the work done and
hence of the overall efficiency, and results in higher heating at
the high pressure tip corner from mid-chord to trailing edge. The
hot leakage flow increases the thermal loads on the blade tip,
leading to high local temperature. Thus, it is essential to cool the
turbine blade tip and near the tip regions. However, it is difficult
to cool such regions and to seal against the hot leakage flow.
The blade tip operates in an environment between the rotating
blade and the stationary casing, and experiences the extremes of
the fluid-thermal conditions within the turbine [10–12]. A more
detailed discussion of the blade tip can be found in [13].
Because the blade lifetime may be reduced by a factor of 2
if the blade metal temperature prediction is off by only 30◦ C, it

is very critical to predict accurately the local heat transfer and
local blade temperature to prevent hot spots and thus increase
the turbine blade life. It is important for the gas turbine designers
to know the effects of increased heat load in the area exposed to
hot gas and be able to design efficient cooling schemes to protect
the blade. Therefore, fundamental and detailed studies of heat
transfer and flow relating to the blade tip or near blade tip regions
are needed to provide better understanding and prediction of the
heat loads on such regions accurately.
Besides conventional techniques of experimental measurements with advanced apparatus, computational fluid dynamics
(CFD) plays an increasingly important role in design and research studies of gas turbines. During the past two decades,
CFD has been developed so rapidly that many advanced
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computational codes and commercial softwares have continuously appeared for solving the heat transfer and flow field of
complex geometries like gas turbine passages. By validating the
codes with experimental data, many computational results based
on CFD are accurate and reliable. This will contribute to the
prediction and design of turbomachinery components, without
doubt, including the turbine blade and its tip. The highly accurate
computational results can contribute to the design and manufacture of gas turbine blades and improve the durability and safe
operation.
This paper does not and cannot review all the interesting
and important progress related to gas turbine heat transfer and
cooling (some may be found in [1–5]), but tries to summarize
the recently published results in the concerned field of blade tip
heat transfer and development of cooling technology. The first
studies on blade tip heat transfer were reviewed earlier [11–13].
In this paper, published literature from 1995 to 2008 and on,
especially the recent years 2001–2008, are reviewed.

This paper is organized as follows. Gas turbine heat transfer
and the need for cooling techniques are introduced first. Then the
blade tip leakage complicated flow associated with heat transfer
on tips or near the tip regions is reviewed. Next the development of external tip cooling methodology is reviewed, while
the last section reviews the development of internal tip cooling
methodology. A summary is presented in the final section.

BLADE TIP LEAKAGE FLOW AND HEAT TRANSFER
Generic Flow Pattern Associated With Tip Leakage Flow
The flow field in a turbine is very complex. It is strongly
three-dimensional, unsteady, and viscous, with several types of
secondary flows, endwall flows, and vortices (passage vortex,
counter vortex, horseshoe vortex, leakage flow vortex, etc.).
Transition flow and high turbulence intensity result in additional
complexities. Figure 6 depicts the complex flow phenomena
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Figure 6
[14].

Complex flow and heat transfer phenomena in the turbine gas path

in a turbine blade gas path [14]. The understanding of such
complex flow fields and heat transfer characteristics is necessary
to improve the blade design and prediction in terms of efficiency

as well as the evaluation of mechanical and thermal fatigue. Tip
leakage flow is a dominant source of unsteadiness and threedimensionality of the flow in turbomachineries. As depicted in
Figure 7, the tip leakage flow passes through the tip clearance
driven by the pressure gradient between the pressure side and
suction side. Also, the leakage flow tends to roll up into a vortex
and interacts with the secondary flow. Thus, the leakage flow
and its interaction with other flow features show very complex
phenomena.
A perfect blade tip will not allow any leakage flow, and no
secondary flows to reduce stage efficiency will be generated,
nor losses for downstream stages created, and cooling is not
required. Thereby no thermodynamic losses occur [11]. Thus,
the two main objectives of blade design are to reduce the leakage

Figure 8 Different kinds of blade tips [11].

Figure 7 Schematics of blade tip leakage flow characteristics [11].

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flow as much as possible and to cool the blade tips using small
quantities of extracted cooling air. However, all the blade tips in
modern gas turbines do allow some leakage flow and secondary
flows are generated. Today, there are several major types of
blade tips: (a) flat tip, (b) recessed tip with peripheral squealer
sealing rims, and (c) attached tip shrouds [11], as shown in
Figure 8. Each blade tip has its advantages and disadvantages.
Although it is easy to design a flat tip and its cooling scheme,
very few turbines use flat tips. High leakages lead to bad tip
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531

Figure 9 Streamlines of blade tip flow pattern [15].

aerodynamics and results in higher heat loads on the tip. A
recessed tip with sealing rims is the most common design in
practice today for high-pressure turbine blades. A recessed tip
with rim reduces the risk of blade damage if the tip rubs against
the shroud; however, the design of a recessed tip is more complex
because of the cooling of the rim and the need to prevent losses
by oxidation and erosion. Blades with attached tip shrouds are
mostly used in low-pressure turbine blades. This tip has the
lowest aerodynamic loss when properly installed, but it requires
greater attention to stresses because of the heavier weight and
requires a more complex cooling system.
Ameri et al. [15] performed calculations on flow and heat
transfer of a GE-E3 rotor tip considering three types: plane tip,
2% recess tip, and 3% recess tip. A two-dimensional (2D) cavity
flow problem was used to validate the k-ω turbulence model.
These authors found two dominant flow structures in the recess
region, which strongly affect the heat transfer rate, as shown
in Figure 9, but no significant effect on the adiabatic efficiency
was observed for these three tips. Also, Ameri et al. [16] studied
the effects of tip clearance and casing recess on heat transfer
and stage efficiency in axial turbines. Their numerical study
reconfirmed a linear relationship between the efficiency and the

tip gap height. Introduction of a recessed casing resulted in a
drop in the rate of heat transfer on the pressure side, and a
marked reduction of the heat load and peak values on the blade
tip. Ameri et al. observed that the recessed casing has a small
effect on the efficiency but can have a moderating effect on the
flow underturning at smaller tip clearances.

Experimental Measurements for Tip Region Flowfield
and Heat Transfer
Bunker et al. [17], and Ameri and Bunker [18] reported
results of a combined experimental and simulation study designed to investigate the detailed distribution of the convective
heat transfer coefficient on the flat tip surface with both sharp
and rounded edges for a large power generation turbine. This
study showed good agreement between experiments and comheat transfer engineering

Figure 10 Sharp and rounded edge tip heat transfer coefficients [18].

putations. Figure 10 presents a sample of these experimental
and computed tip heat transfer coefficients for the sharp and
rounded edge tips. Ameri [19] also conducted experimental and
numerical studies of detailed heat transfer coefficient distribution on the rounded blade tip of a gas turbine equipped with
a mean-camberline strip. Generally good agreement between
experimental data and computations was achieved, as shown
in Figure 11. Results showed that the mean-camberline strip
could reduce the tip leakage flow but the total pressure loss was
not reduced comparatively, and the sharp edge tip was better in
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Figure 11 Heat transfer coefficient and flow pattern for blade tip with meancamberline strip [19].

reduction of the tip leakage flow and tip heat transfer compared
to the rounded edge tip.
Thorpe et al. [20, 21] reported experimental measurements of
time-mean/time-resolved heat transfer and static pressure on the
over-tip casing of a transonic axial flow turbine. They presented
axial and circumferential distributions of the heat transfer rate
as well as adiabatic wall temperature, Nusselt number, and static
pressure. They found that the rate of heat transfer to casing wall
and the wall temperature varied strongly with axial position
through the rotor, and the effects of the vane exit flow features
were small. Through assessments of the relative importance
of different time varying phenomena to the casing heat load
distribution, they concluded that up to half of the casing heat
load was associated with the tip leakage flow. Also, discussion
about shroudless turbine design accounting for the high heat flux
was addressed. Thorpe et al. [22] also experimentally studied the
blade tip heat transfer and aerodynamics in a transonic turbine
stage. They observed high heat transfer rates near the nose of the
blade tip and also in the region of high blade lift near the midaxial chord, and proposed three primary mechanisms: vane–
shock interaction, relative total temperature fluctuations, and
fluctuations in tip leakage flow speed and direction driving the
unsteady heat transfer.
Chana and Jones [23] presented detailed experimental measurements of heat transfer and static pressure distributions on
the shroudless rotor blade tip and casing with and without inlet
nonuniform temperatures. Also, a simple 2D model was developed to estimate the heat transfer rate to tip and casing as a

function of Reynolds number. Results showed that the overall
heat load was reduced with inlet nonuniformity, that the highest
heat transfer rate was on the pressure side of the blade where the
highest random unsteadiness was marked, and that the average
static pressures did not show significant difference between the
two cases. Camci et al. [24] investigated experimentally aerodynamic characteristics of full and partial length squealer rims in
an axial turbine. Figure 12 shows a schematic picture of partial
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Figure 12 Geometries of partial squealer rims [24].

squealer rims studied. Results showed that the partial squealer
rim could seal the tip effectively, and a mid-size partial rim was
most effective in reducing the tip leakage flow. Compared to
the two studied channel arrangements having partial rims near
the corners of the suction and pressure sides, the sealing performance of the mid-size rim on the suction side was even better.
This indicated that the partial squealer rims on the suction side
were capable of reducing the exit total pressure loss by the tip
leakage flow to a significant degree. Key and Arts [25] studied
the tip leakage flow characteristics for flat and squealer turbine
tips. The experiments were conducted at different Reynolds
number and Mach number conditions for a fixed value of the
tip gap in a nonrotating, linear cascade arrangement. Oil flow
visualization was used, as shown in Figure 13, and the static
pressure and aerodynamic loss were measured. These authors
found that the squealer tip showed a significant decrease in velocity through the tip gap, and for the flat tip the increase of
Reynolds number would cause an increase in the tip velocity
level whereas for the squealer tip the sensitivity was not much.
Their data are valuable for validation of CFD computations,
and in turn CFD can provide insight to some details of the flow

physics in the tip region.
Azad et al. [26, 27] and Teng et al. [28] measured the heat
transfer coefficient and static pressure distributions on gas turbine tips in a five-bladed stationary linear cascade. Various regions of high and low heat transfer coefficients at the tip surface
were observed. The heat transfer coefficients increase with an
increase of the inlet turbulence intensity. Compared to the flat
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Figure 13 Flow visualization of squealer and flat tip [25].

tip, the squealer tip showed a lower overall heat transfer coefficient. Also, a reduced tip gap clearance resulted in a weaker
unsteady wake effect on the blade tip heat transfer and a reduction in the heat transfer coefficient over the blade tip surface.
Azad et al. [29] also studied the effect of the squealer geometry
arrangement on a blade tip. The detailed heat transfer coefficient distributions of six tip geometry cases were obtained. It
was shown that the suction-sided squealer could provide a better benefit compared to other cases, and the mid-chamber lined
squealer behaved better than the pressure-sided squealer. Also,
a single squealer provided better performance in reducing the
overall heat transfer than a double squealer.
Dhadwal and Kurkov [30] used a dual-laser probe integrated
fiber optic system to measure the blade tip clearance in a rotating turbomachinery. A symmetric configuration of the probe
installation could offer better resolution. The time-of-flight measurements were robust and reliable. Saxena et al. [31] presented
a comprehensive investigation of the effect of various tip sealing
geometries on the blade tip leakage flow and heat transfer of a
scaled up high-pressure turbine. They found that compared to
other geometries, the tripped strips placed against the leakage
flow (as shown in Figure 14a) led to the lowest heat transfer

on the tips with a reduction of 10–15%. The use of strips and
pin-fins did not decrease the tip surface heat transfer coefficients. Saxena and Ekkad [32] also experimentally investigated
the effect of squealer tip geometries on the blade tip leakage
and associated heat transfer in the same facility. It was found
that the suction-sided squealer rim might be favorable for reducing the heat transfer coefficients on the tip surface, whereas
the pressure-sided squealer did not reduce the heat transfer and
behaved like the plane tip. Nasir et al. [33] also investigated
the effect of tip gap and squealer geometry on the detailed
heat transfer engineering

Figure 14 Blade tip geometries for test [31].

heat transfer over a high pressure turbine rotor blade tip. The
squealer studied altered the tip gap flow significantly and hence
resulted in lower heat transfer coefficient. Also, experimental
results showed that some partial burning of the squealers might
be good for overall reduction in the heat transfer coefficient.
Rhee and Cho [34, 35] experimentally measured local
heat/mass transfer characteristics on tip, shroud, and near-tip
surface of a rotating blade in a low-speed annular cascade. The
effects of rotation and incoming flow incidence angle were examined. Results showed that the heat transfer was complex with
complicated flow patterns such as flow acceleration, laminarization, transition, separation, and tip leakage flow, and the blade
rotation caused increased incoming flow turbulence intensity
while the tip leakage flow was reduced. Also, they found that
the heat/mass transfer coefficients were about 1.7 times than
those on the blade surface and shroud, and due to the reduced
tip leakage flow under rotation the heat/mass transfer coefficients on the tip slightly decreased while they remained similar
on the shroud. With a positive incidence angle, more uniform
and higher heat transfer rate were found on the tip because of
the increased tip gap flow and high flow angle. Rhee and Cho

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Figure 15 Blade geometries for test [40].

[36, 37] experimentally studied the effect of vane/blade position
on heat transfer in a stationary blade and shroud in a low-speed
wind tunnel. They presented detailed mass transfer measurements, and the results showed that the mass transfer coefficients
in the upstream region varied up to 25% due to the blockage effect as the blade position changed. The size and level of the peak
region were affected strongly. Also, distinctly different patterns
near the blade tip were observed due to the variation in the tip
leakage flow.
Matsunuma [38] observed the effect of Reynolds number
and freestream turbulence on turbine tip clearance flow. Threedimensional flow fields at the exit of the turbine with and without
tip clearance were measured. Results indicated that variations
heat transfer engineering

in Reynolds number and freestream turbulence intensity did not
affect the mass-averaged tip clearance loss. Due to the strong
interaction between the leakage vortex and tip-side passage vortex, the decrease in flow angle at lower Reynolds numbers was
larger than that at higher Reynolds numbers. Kwak and Han
[39] and Kwak et al. [40] conducted a series of measurements
on the tip and near-tip region heat transfer coefficients of a
turbine blade with flat or squealer tip, and the effects of rim location and height as well as tip clearance on heat transfer were
measured. The geometry is shown in Figure 15. The blade tip
clearance was 1.0%, 1.5%, and 2.5% and the rim height was

2.1%, 4.2%, and 6.3% of the blade span, as shown in Figure 15.
Experimental results showed that the heat transfer coefficients
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on the tip surface were higher than those on the shroud and on
the near-tip region of the pressure and suction sides, and with an
increase of the tip clearance the heat transfer on the tip surface
increased whereas heat transfer on the shroud and the suction
side first increased and then decreased. On the blade pressure
side the heat transfer coefficient was kept constant. They also
found that higher rims could reduce the heat transfer coefficient
on the tip and shroud, while on the pressure side and suction side
the reduction was not significant. The suction-sided rim could
provide lower heat transfer coefficient on the tip and near-tip
region than the double-sided rim case. Kwak and co-workers
[41, 42] also performed measurements on detailed heat transfer
coefficients on the squealer tip and near-tip region of a turbine
blade. Results showed that the overall heat transfer coefficients
on the squealer tip were higher than those on the shroud surface and the near-tip region of the pressure side and suction
side. Near the tip region the heat transfer coefficient showed
no significant reduction. Also, the suction-sided squealer tip revealed the lowest heat transfer coefficients on the blade tip and
near tip.
Papa et al. [43] investigated experimentally the effects of
squealer or winglet-squealer tip and tip clearance on the average and local mass transfer coefficients for a large-scale gas
turbine blade, and used the heat–mass analogy to obtain heat
transfer coefficients. Flow visualization on the tip surface was
presented. Compared to the winglet-squealer tip, the squealer

tip provided a higher average mass/heat transfer coefficient.
Rehder and Dannhauer [44] studied the effect of the tip leakage
flow on the three-dimensional (3D) flow field and end-wall heat
transfer. Results showed that when the leakage mass flow rate
increased from 1% to 2%, significant changes in the secondary
and end-wall heat transfer occurred. The secondary flow was
amplified as the leakage flow was ejected perpendicular to the
main flow direction, whereas it was reduced significantly as
the leakage flow was ejected tangentially. Govardhan et al. [45]
investigated the 3D flow in a large deflection turbine cascade
with tip clearance 0.08%, 1.5%, and 3.0% of the chord. They
found that there was a strong horseshoe vortex in front of the
leading edge for 0.08% clearance, while for 3% clearance there
was no vortex. A small tip separation vortex was also observed
on the tip surface, which made the flow from the pressure side
to be accelerated. The passage vortex did not diminish as the
tip clearance increased. Also, Govardhan et al. [46] investigated
the effect of endwall and tip clearance on the flow in a twodimensional turbine rotor blade cascade. Five incidence angles
were chosen: −10, −5, 0, 5, and 10◦ . Results showed that as
the tip clearance was increased the adverse pressure gradient
upstream the leading edge was reduced, and with the increase
of incidence angle the blade loading due to the static pressure
gradient also increased.
Porreca et al. [47] conducted experimental and numerical investigation on flow dynamics and performance of partially and
fully shrouded axial turbines, as shown in Figure 16. Experimental results showed that for the partial shroud case a strong
tip leakage vortex was developed from the first rotor and transheat transfer engineering

535

Figure 16 Shroud configuration and probe planes [41].

ported through the downstream blade row. CFD computational
results showed a good agreement with the measured data at the
midspan for the first stage. The overall second stage efficiency
for the full shroud case could be improved by 1%. Newton et al.
[48] measured the heat transfer coefficient and pressure coefficient on the tip and near-tip region of a generic turbine blade
in a five-blade linear cascade. Two tip clearances of 1.6% and
2.8% of chord were considered and three tip geometries were
studied: plane tip, suction-sided squealer, and cavity squealer.
They found that the flow separation at the pressure side edge
dominated the flow through the plain gap, that the highest heat
transfer was located in such a region that the flow reattached
on the tip, and that the suction-sided and cavity squealers could
reduce the heat transfer in the gap. The suction-sided squealer
provided an overall net heat flux reduction of 15%, while the
cavity squealer revealed no net heat flux reduction. Palafox
et al. [49] measured new detailed flow fields for a very large
low-speed, high-pressure turbine rotor blade using particle image velocimetry (PIV). The interaction between the tip leakage
vortex and passage vortex was clearly characterized, and the effect on the tip leakage vortex was examined. Results showed
that a separation bubble under the tip significantly affected
the leakage flow, and the end-wall movement influenced the
shape and size of the bubble distinctly, while the relative blade
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casing movement distorted the shape of the tip leakage vortex

and shifted it closer to the suction side.
Jin and Goldstein [50, 51] simulated and measured local
mass and heat transfer on a turbine tip and near-tip regions. They
concluded that for the smallest tip clearance the mass transfer on
the tip was significant along the pressure side. At the largest tip
clearance the separation bubble on the tip could cover the whole
width of the tip on the second half of the tip surface. A high
mainstream turbulence could reduce the average mass transfer
rate on the tip, whereas a higher mainstream Reynolds number
provided higher local and average values on the tip and near-tip
surfaces. Stephens et al. [52] and Douville et al. [53] conducted
experiments to study the effects of thickness-to-gap and gap-tochord ratios on the tip-gap flows. They also performed surface
flow visualization on the blade tip for better understanding of the
gap flow behavior. The partial squealer tip or plasma actuators
were used to control the tip leakage flow. Results showed that
the squealer tip could effectively reduce the pressure loss, and
by the use of a plasma actuator the effect depended strongly on
the unsteady frequency. Srinivasan and Goldstein [54] measured
the local mass transfer on the tip of a turbine blade in a fiveblade linear cascade with a blade-centered configuration, and
used a moving end wall mounted on the top of a wind tunnel
to observe the effect of relative motion between the casing and
the tip. Results showed that at a clearance of 0.6% there was
a small but definite reduction of 9% in the heat/mass transfer,
and at 0.86% clearance only a small effect of the wall motion
on the Sherwood number occurred. At all higher clearances
no measurable effect of the relative motion on the Sherwood
number was observed.

Computational Tip Leakage Flow and Heat Transfer
Multiple numerical studies have been carried out on blade tip

leakage flow associated with heat transfer. Numerical prediction
and analysis can provide details of the flow field and thermal
distribution that sometimes are difficult to obtain by experimental measurements. Dorney et al. [55] performed a parallelized
unsteady analysis of the effects of tip clearance on the transient
and time-averaged flow fields in a supersonic turbine. Results
indicated that improved performance could be traced by a reduction in the strength of the shock system in the vane or rotor,
and the reduction in losses was greater than the losses generated by increasing the tip clearance. Green et al. [56] conducted
computations and experiments on averaged and time-dependent
aerodynamics of a single-stage high-pressure turbine tip cavity
and stationary shroud. The computational results showed good
correlation with the time-resolved data. This in turn provided
confidence of the CFD modeling ability to predict turbine passages, blade tip, and shroud. They found the largest amount of
unsteady surface pressure activity at the 15% span location, especially on the suction surface near the leading edge. Past the
leading edge unsteady pressure amplitudes with respect to vane
passing frequency dropped off rapidly, and unsteady pressure
heat transfer engineering

Figure 17 Schematics of blade tips with winglet [57].

amplitudes were much larger for all shroud locations than at the
blade tip locations, Also, the results suggested that the blade
tip configuration had very little impact on the time-accurate
behavior for the stationary shroud.
Saha et al. [57] performed calculations to observe the effect
of a winglet on the flow and heat transfer for both a flat tip and
a squealer tip, as shown in Figure 17. All the winglets were
located on the pressure side only. They found that for a flat tip
the winglet resulted in approximately 30% reduction in the local
heat transfer coefficient on the tip, and a significant reduction
in the strength of the leakage flow and vorticity, whereas for

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a two-sided squealer tip the winglet produced only marginal
improvements. They concluded that the suction-side squealer
with constant winglet width offered better performance with
lower heat transfer coefficient and pressure loss than the others.
Lampart et al. [58] simulated numerically the effect of interaction of the main flow with rotor and tip leakage flows in a
high pressure axial turbine stage. The proposed method could
trace and evaluate the process of mixing of the tip leakage and
windage flows with the main stream, and the interaction with
secondary flows and separation.
Starodubtsev et al. [59] proposed a 3D numerical model for
simulation of the viscous turbulent flow in a one-stage gas turbine and validated the results with experimental measurements.
They stated that their method could be applied for a variety of
turbine studies and design task. Han et al. [60] analyzed numerically 3D flow fields near the tip region in an annular cascade
with tip clearance and rotation and in a linear cascade with the
validation of numerical results by flow visualization. Results
showed that rotation could weaken the leakage flow, which decreased the size of the separation bubble on the tip surface, and
the tip vortex became larger and moved to the suction side as
the tip leakage flow was increased by an enlarged tip clearance.
Intaratep et al. [61] studied the interaction between the rotor
blade tip leakage flow and inflow disturbances. They found that
the passage flow consisted of shear layers shed from the suction
side tip gap and a high velocity deficit region extending from
the suction side to the pressure side tip gap, and the local perturbations near the blade tip induced the streamwise mean velocity
perturbations in the tip leakage vortex. Yang et al. [62] numerically simulated the leakage flow and heat transfer on a flat tip,
a double squealer tip, and a single suction side squealer tip of a

scaled up GE-E3 blade. The rotational effect was observed under high pressure ratio and high temperature. It was found that
the heat transfer coefficient decreased by increasing the squealer
cavity depth, while the shallow squealer cavity was the most effective in reducing the overall heat load. Although the rotation
changed significantly the tip leakage flow pattern and local heat
transfer coefficient distribution on the tip, the area-averaged heat
transfer coefficient was affected only slightly.
Mumic et al. [63, 64] numerically studied the tip leakage flow
and heat transfer on the first stage of a high pressure turbine.
A flat tip and a squealer tip with tip clearance of 1.0%, 1.5%,
and 2.5% blade span were considered. Three turbulence models
were used to assess the prediction of the heat transfer. It was
found that the three models could provide similar results in
reasonable agreement with the experimental data. The low-Re
k-ω model could yield better prediction of blade tip heat transfer
compared to the other two models. As shown in Figure 18, the
leakage flow increased and moved toward the trailing edge side
as the tip gap was increased. The high heat transfer coefficients
on the rim were increased due to acceleration of the flow going
into the cavity and from the cavity into the rim region, and the
heat transfer coefficient near the leading edge cavity increased
and extended toward the trailing edge. The flat tip heat transfer
was higher than the squealer tip heat transfer. Mischo et al. [65]
heat transfer engineering

Figure 18
[63].

537

Comparison of heat transfer coefficient and simulated flow field

numerically studied the flow field near the blade tip for different
shapes of the recessed cavities. An improved design of the blade
tip was presented. It was found by an appropriate profiling of the
recessed shape, the total tip heat transfer Nusselt number was
significantly reduced by 15% and 7% compared to the flat tip
and baseline recessed shape, respectively, as shown in Figure 19.
The CFD analysis predicted a 0.38% total efficiency increase
for the rotor equipped with the new recess design compared to
the flat tip.
Hamik and Willinger [66] introduced a new concept for passive turbine tip leakage control: A jet was injected roughly
perpendicular to the tip gap flow, as shown in Figure 20. They
also presented an analytical model to describe the reduction of
the tip gap discharged coefficient due to the tip injection. They
stated that the blade tip injection could increase the turbine efficiency. Prakash et al. [67] proposed an improved tip having
a pressure-side inclined squealer shelf and used CFD to study
different tip geometries, as shown in Figure 21. It was found that
the inclined shelf could reduce the leakage flow and improve
the efficiency, indicating that it was superior to a vertical shelf.
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Figure 20 Blade tip with internal injection [66].

temperature, while too much coolant flow results in reduced
turbine efficiency and power. Therefore, it is very important to

design a turbine cooling system considering the balance of the
minimum coolant air flow and maximum benefit of a high inlet

Figure 19 3D CFD flow for different tip shapes [65].

BLADE TIP EXTERNAL COOLING TECHNOLOGY
Needs of Cooling Technology for Blade Tip
Without doubt, cooling is required for the turbine blades,
including all regions being exposed to the high-temperature hot
gas. Due to an unavoidable gap clearance between the blade tip
and casing or shroud, the hot gas flowing through the gap results
in a large thermal load on the blade tip. The potential damage
due to the large heat load will lead to blade oxidation, as shown
in Figure 22. Hence, the blade tip is a key region that needs
cooling.
The turbine blades are cooled by the use of extracted/
bypassed air from the compressor of the gas turbine. This extraction results in a reduction of the thermodynamic efficiency
and power output. Too little coolant flow results in high blade
heat transfer engineering

Figure 21 Blade tip with squealer shelf [67].

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539

Figure 23 Schematics of film cooling.

conditions (Reynolds number, Mach number, free-stream turbulence, and unsteadiness). Obviously, a high and uniform cooling
effectiveness will ensure overall performance of the blade surface cooling. In general, a higher blowing ratio at a specific
temperature ratio gives a higher film cooling performance, and
thereby the heat is transferred to the blade surface and hence the
protection of surface is improved. However, too high a blowing
ratio leads to jet penetration into the mainstream resulting in a
reduced cooling performance, while too small a blowing ratio
does not force enough coolant to cover the hot surface. Thus, it
is important to optimize the amount of coolant for film cooling
at the engine operating conditions. For a better cooling performance, it is necessary to study the film cooling hole pattern,
e.g., shape, angle, location, and distribution, which affect the
film cooling performance.
Additional review papers related to film cooling of gas turbines are exemplified by refs. [68]–[71]. This paper is limited
to a review of recent publications on blade tip cooling, and thus
does not include all the results of external film cooling on turbine
blades.

Blade Tip External Cooling
Figure 22 Material loss due to oxidation [11].

temperature. If a proper cooling system is designed, the gain
from high firing temperature is so significant that it can outweigh the losses in the efficiency and power output, and offset
the complexity and cost of the cooling technology.
The turbine blade tip and near-tip regions are difficult to cool
and are subjected to potential damage because of the high heat
load caused by tip leakage flow. A common way to cool the tip
is to extract the cooling air from the internal coolant passages
through some film holes that are located on the blade surface
discretely. This cooling is known as film cooling. The relatively

cool air passes these holes and forms a thin protective layer/film
to protect the tip surface from the highly hot mainstream.
Figure 23 depicts the film cooling concept. The performance
of the film cooling depends on the coolant-to-hot mainstream
pressure ratio (blowing ratio), temperature ratio, and the hole
location, configuration (hole size, spacing, shape, angle and
number), distribution (leading-edge, trailing-edge, pressure and
suction side, endwall, tip), and on the representative flow
heat transfer engineering

A summary of Professor D. E. Metzger’s blade tip cooling
studies on blade tip cooling was published by Kim et al. [72].
Comparison of various tip cooling configurations and their effects on film effectiveness and heat transfer coefficients were
presented. Figure 24 shows the clearance gap and tip film cooling configuration and Figure 25 shows the cross sections. Four
film cooling configurations were tested: (1) discrete slot injection, (2) round hole injection, (3) pressure side flared hole injection, and (4) grooved-tip cavity injection. It was found that for
case 4 the overall film cooling performance varied significantly
with injection locations and that among the plane-tip injections
the discrete slot injection provided better performance than the
others.
Yang et al. [73] numerically studied various film hole configurations on plane and squealer tips of a turbine blade. Three
configurations were tested: (1) the camber arrangement, (2) the
upstream arrangement, and (3) the two rows arrangement, as
schematically shown in Figure 26. The effects of rotation were
observed. It was found that at high blowing ratios the latter two
cases provided better film cooling performance on the plane
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Figure 26 Various film cooling hole arrangements [73].
Figure 24 Schematics of clearance gap and film cooling configuration.

and squealer tips than the former one. Higher blowing ratios
resulted in a higher cooling effectiveness on the shroud for all
cases. They also found that rotation decreased the plane-tip film
cooling effectiveness while it slightly affected the squealer-tip
film cooling due to the large cavity depth.
Mhetras et al. [74] observed the effects of shaped holes on
the tip pressure side, coolant jet impingement on the pressure
side squealer rim from tip holes, and varying blowing ratios
for a squealer tip. The film cooling effectiveness distributions
on the blade tip, near-tip pressure side rim, and the inner pressure side rim were measured using a pressure-sensitive paint
(PSP) technique. Numerical simulations were also performed
for prediction of the film cooling. It was found that a higher
blowing ratio provided higher effectiveness on the tip rim, cavity flow, and inner rim walls, and the presence of serpentine
passages could supply coolant to the holes so that a significant
impact on film cooling performance was achieved. Good agreement between the experiments and simulation was achieved.
Mhetras et al. [75] also measured the film cooling effectiveness of shaped holes near the tip pressure side and cylindrical

holes on the squealer cavity floor using PSP. The pressure side
squealer rim wall was cut near the trailing edge. It was found
that the cutback squealer rim provided high film cooling effectiveness in the trailing edge of the blade tip compared to a full
squealer. Due to the combined effect of tip- and pressure-side
coolant injection, high and uniform effectiveness was found on
the tip rim and inner and outer squealer rim walls.
Ameri and Gigby [76] performed computations to predict the
heat transfer coefficient distribution on a blade tip with cooling

holes. The simulation model for prediction of the tip heat transfer
and cooling effectiveness based on a 3D Reynolds-averaged NS
solver, was assessed by the data of Kim and Metzger [77].
Through the numerical flow visualization it was shown that the
distance from the pressure side to the edge of the film cooling
hole might be an important parameter. Christophel et al. [78–
80] experimentally investigated the adiabatic effectiveness and
heat transfer coefficients along and near the blade tip using
pressure side film cooling holes. Results showed that the cooling
effectiveness of the holes was better for a small tip gap than for
a large tip gap. With blowing, the tip heat transfer coefficients
were increased above those without blowing, and increased with
increasing blowing ratio. The area-averaged net flux reduction

8/3d
W

d
3d

1.5W

a
1/3W

b

Figure 25 Cross section of film cooling configuration.

heat transfer engineering

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suggested a small dependence on the coolant flow rate and higher
cooling benefit for a small tip gap.
Kwak and Han [81] measured heat transfer coefficient and
film cooling effectiveness on the squealer tip of a gas turbine
blade in a five-bladed linear cascade. The blade model was
equipped with a single row of film cooling holes on the pressure
side near the tip region and the tip surface along the camberline. They found that the overall film cooling effectiveness was
increased but heat transfer coefficients were decreased for the
squealer tip compared to the plane tip at the same tip gap and
blowing ratio. High film cooling effectiveness occurred near the
trailing edge cavity because of the coolant accumulation. Kwak
and Han [82] also measured the distributions of heat transfer
coefficient and film cooling effectiveness on a turbine blade
tip. Three tip gas clearances, i.e., 1.0%, 1.5%, and 2.5%, and
three blowing ratios, i.e., 0.5, 1, and 2, were tested. Results
showed that with the increasing blowing ratio, the film cooling
effectiveness increased but the heat transfer coefficient on the tip
slightly decreased. The static pressure on the shroud increased,
and with the increase of gap clearance the heat transfer coefficient and film effectiveness increased. By addition of pressure
side injection the film cooling effectiveness could be increased.
Ahn et al. [83] also observed the effects of the presence of the
squealer tip, the locations of film cooling holes, and the tip gap

clearance on the film cooling effectiveness compared to a plane
tip. It was found for the squealer tip with tip and pressure-side
injection that the film cooling effectiveness was higher than that
with only tip injection or with only pressure-side injection. For
the plane tip the film cooling effectiveness was significant but
negligible for squealer tip.
Gao et al. [84, 85] studied the effect of incidence angle on
film cooling effectiveness for a cutback squealer blade tip in
a five-blade linear cascade. The film cooling effectiveness was
measured based on mass transfer analogy using PSP techniques.
One row of shaped holes was located along the pressure side
just below the tip and two rows of cylindrical holes were located on the tip. It was found that the film cooling effectiveness
distribution was altered, and the peak of laterally averaged effectiveness was shifted to upstream or downstream depending on
the incidence angle, but the overall area-averaged film cooling
effectiveness was not changed significantly. Also, the coolant
jet spread more on the cavity floor at positive incidence angles,
resulting in relatively high and uniform film coverage on the
cavity floor.
Other detailed studies related to blade tip heat transfer and
cooling topics have been summarized in many theses. These can
be found in refs. [86–97].

541

Figure 27 A typical serpentine passage inside a blade.

through the blade but are not limited to a simple straight channel. A common serpentine passage may consist of a first pass,
a sharp 180◦ turn/bend, and a second pass. A typical serpentine
passage is schematically shown in Figure 27. The coolant flows
radially outward from the hub and then turns 180◦ and travels

radially inward from the tip to the hub. Also, rib turbulators
might be mounted on the leading or/and trailing walls to enhance the heat transfer between the blade wall and coolant. The
flow field in the turn/bend is very complex, and so is the heat
transfer, because the channel configuration, its aspect ratio, the
turn geometry, and the rib configuration and location will affect
the flow and heat transfer. Because the rotation alters the flow
and hence the heat transfer coefficient distribution, the rotation
effect should be considered.
This paper does not review all research works on heat transfer
enhancement in single-pass ribbed channels, but reviews mainly
the findings of flow and heat transfer in two-pass or U-bend
channels with/without rib turbulators, especially in the turn/bend
region.

BLADE TIP INTERNAL COOLING TECHNOLOGY
Experimentally Internal Cooling for Blade Tip
Apart from external film cooling the blade tip region, a number of serpentine passages can be used as channels for internal coolant air to cool the blade. These cooling passages wind
heat transfer engineering

Park and Lau [98], Park et al. [99–102], Kukreja et al.
[103], and Lee et al. [104] conducted a series of naphthalene
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sublimation experiments on local heat/mass transfer distributions on the leading and trailing walls of rotating smooth and
ribbed two-pass channels. The effects of channel orientation,

channel shape, rotation, sharp turn, and angled ribs were observed. It was found that rotation did not lower the spanwise
average heat transfer on the leading wall, and the sharp turn
reduced the heat/mass transfer on the leading and trailing walls.
Due to the complex flow field with secondary flows, separated
and re-attached flows, and flow recirculation in the turn and
near the ribs on the walls, there existed large variation of local
heat/mass transfer in the turn and immediately downstream the
turn, as shown in Figure 28.
Mochizuki et al. [105] performed detailed measurements of
the local heat transfer coefficients in turbulent flow through
smooth and rib-roughened serpentine passages with 180◦ sharp
bend, and performed flow visualization to reveal the generation
of secondary flows. Results showed that for a smooth channel
the heat transfer downstream from the bend was controlled by
secondary flows and the heat transfer coefficients on the wall
surfaces differed from one another. For ribbed channel, due to
the interaction of two secondary flows by the ribs and bend, the
ribs could affect strongly heat transfer in the bend and second
pass. Chen et al. [106] presented 3D detailed mass (heat) transfer
distributions along four active walls of a square duct with a 180◦
bend and ribs in the first pass. Results showed that the effect
of the bend was clearly visible in the ribbed duct following the
bend. Due to the high velocity resulting from the bend, local
acceleration and turbulence production generated by ribs, the
higher mass transfer rates occurred near the corners of the outer
wall. Astarita and co-workers [107, 108] measured the detailed
heat transfer distribution near a 180◦ sharp turn of a square
channel with and without rib turbulators. It was observed that
for the smooth channels there were three high heat transfer zones
in the turn, while the only high heat transfer zone left was placed

after the second outer corner and exhibited a smaller extension.
The averaged normalized Nusselt number slightly increased for
the both side heating condition compared to that for one-side
heating conditions.
Ekkad et al. [109] measured the detailed heat transfer distributions inside straight and tapered two-pass channels with and
without rib turbulators. It was found that the tapered channel
with ribs provided 1.5–2.0 times higher Nusselt number ratios
over the tapered smooth channel in the first pass, while in the
after-turn region of the second pass the ribbed and smooth channels provided similar Nusselt number ratios. Ekkad et al. [110]
and Pamula et al. [111] also measured the detailed heat transfer
distribution inside a two-pass square channel connected by two
rows of holes on the divider walls, shown in Figure 29. It was
found that the proposed feed system, from first pass to second
pass using crossflow injection holes, produced higher Nusselt
numbers on the second-pass walls with the enhancement factor
as high as two to three times than that obtained in the second
pass for a channel with a conventional 180◦ turn. Son et al.
[112] carried out particle image velocimetry (PIV) experiments
to study the correlation between the high Reynolds number
heat transfer engineering

Figure 28 A typical result inside a two-pass channel [98].

(Re = 30,000) turbulent flow and wall heat transfer characteristics in a two-pass square channel with a smooth wall and a 90◦
rib-roughened wall. Compared with the heat transfer experimental data of Ekkad and Han [113], the PIV measurement results
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543

Figure 29 The geometry of a two-pass channel having injection holes.

showed that the flow impingement is the primary factor for the
two-pass square channel heat transfer enhancement rather than
the flow turbulence level itself. Besides, the secondary flow characteristics are correlated with the wall heat transfer enhancement
for smooth and ribbed wall two-pass square channels.
Chanteloup et al. [114] measured flow characteristic effects
on the wall heat transfer distribution of a two-pass internal
coolant ribbed passage of gas turbine airfoils. Results showed
that ribs at 45◦ increased the average heat transfer gradients,
and the ratio of high to low Nusselt numbers was up to 6 in
the U-shaped heat transfer distribution downstream the ribs.
Chanteloup and Bolcs [115] also measured flow characteristics
in a 180◦ bend region and downstream of the bend of two-leg
internal coolant passages of gas turbine airfoils with film cooling hole ejection. 45◦ angled ribs were located on the bottom
and top walls of both legs. Results showed that adding bleeding
holes having high ratio between the channel inlet mass flow and
the extracted mass flow would affect significantly the flow in the
two-legged cooling channel. Due to the high variations in the
streamwise velocity, the large variations in the heat transfer occurred near the upstream part of the bend. Iacovides et al. [116]
reported flow and heat transfer in a U-bend with 45◦ ribs rotating channel. It was found that the Nusselt number in the ribbed
channel was twice that for a smooth channel [117], and the flow
and average Nusselt numbers were relatively unaffected by rotation but led to local hot or cold spots resulting in significant
implications for the level of thermal stresses induced.
Hsieh and Liao [118] and Hsieh and co-workers [119–121]
measured the effects of rotation and uneven heating conditions
as well as passage aspect ratio on the local heat transfer and pressure drop in a rotating two-pass ribbed rectangular or smooth
square channel. Results showed that higher heat transfer on both

the leading and trailing walls was caused by a complicated 3D
accelerated flow and secondary flow in the U-bend region. For
a ribbed channel, steamwise-periodic fully developed flow was
achieved after a sufficient distance. The intensity of the shear
layer was greater in the vicinity of the ribs compared to a smooth
surface. However, the size of the separation region was smaller
than that of a stationary duct as the rotation number increased.
Hsieh et al. also found that the rotation makes the turbulent
intensity and shear stress distribution more random in the transheat transfer engineering

Figure 30 Different cross sections of profiled ribs [122].

verse direction. For a smooth channel, they found no separation
in the first and second channels except for a certain size pocket
of separation on the inner wall in the U-bend region. The influence of the U-bend and rotation on the mean velocity field
was apparent, and the rotation may alter the development of the
mean and fluctuating motion.
Acharya et al. [122] and Nikitopoulos et al. [123] investigated
experimentally the effects of rib with different cross-stream profiles on the surface mass (heat) transfer distribution along four
active walls of a square duct having a sharp 180◦ bend. The cross
sections of the profiled ribs are shown in Figure 30. These authors found that the profiled ribs enhanced the heat transfer due
to the generation of secondary and longitudinal vorticity that
interacts with Coriolis-induced secondary flows in the channel.
It was suggested that the use of profiled ribs might be a viable
and effective solution to local heat transfer enhancement and/or
spatial redistribution in actual rotating, ribbed multipass cooling channels for gas turbine applications. Liou and co-workers
[124–134] conducted a series of experiments on flow and heat
transfer in rotating two-pass smooth and various angled ribbed
channels using LCT or/and LDV (see Figure 31). The effects
of the divider thickness, rib arrangement, channel cross-section

shape, channel orientation, and rotation conditions were observed in detail.
Al-hadhrami and Han [135] tested the effect of various 45◦
angled rib turbulators on Nu ratio in a rotating two-pass square
channel, as shown in Figure 32. It was found that the Nu ratio
in the 180◦ turn region and the differences among different
angled rib orientations were increased with increasing rotation
number. Al-hadhrami et al. [136] also studied the heat transfer
in two-pass rotating rectangular channels with five different
orientations of 45◦ V-shaped ribs, as shown in Figure 33. Results
showed that there was relatively low heat transfer enhancement
in the 180◦ turn region due to suppression of the vortices induced
from the V-shaped ribs by the turn and no ribs placed at the
turn. Parallel 45◦ rib arrangements provided better heat transfer
compared to the other cases.
Prabhu and Vedula [137] investigated the local pressure drop
characteristics in a square-cross-sectioned smooth channel with
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Figure 31 LCT and LDV test facility [124].

a sharp 180◦ bend rotating about an axis normal to the freestream direction. They found that the local pressure drop characteristics in the bend region are affected by a change in the
rotation number but the influence of the Reynolds number was
weak. Another finding is that the friction factor was less sensitive to rotation for a bend with a hydraulic diameter ratio of 0.24
compared to bends with ratios of 0.37 and 0.73, respectively.
Ratana-Rao and Prabuhu [138] and Ratana-Rao et al. [139] experimentally studied the effect of several turn treatments on the

pressure drop distribution in smooth and ribbed squared channels with a sharp 180◦ bend. Results showed that short and long
guide vanes placed at the center of the bend in a smooth channel
resulted in a reduction of about 28% in overall pressure drop,
and for the ribbed channel a maximum decrease of 15% to 16%
in overall pressure drop was achieved in the case of the long
guide vane located at the center of the bend and multiple 180◦
extended guide vanes.
Azad et al. [140] measured the heat transfer in a two-pass
rectangular rotating channel with 45◦ angled rib turbulators.
Results showed that the heat transfer from the first pass trailing
and second pass leading surfaces was enhanced by rotation.
45◦ parallel ribs provided a better heat transfer augmentation
than 45◦ cross ribs. Fu et al. [141, 142], and Liu et al. [143]
reported heat transfer coefficients and friction factors in twopass rectangular channels with rib turbulators placed on the
leading and trailing surfaces. Five kinds of ribs were considered:
45◦ angled, V-shaped, discrete 45◦ angled, discrete V-shaped,
heat transfer engineering

Figure 32 Two-pass channel with various 45◦ angled ribs [135].

and crossed V-shaped. It was found that due to the turn effect
the rotation effect was greater on heat transfer in the first pass
than in the second pass. The discrete V-shaped ribs showed the
best overall thermal performance, as shown in Figure 34.
Nakayama et al. [144] measured flow and heat transfer in
stationary two-pass channels with a sharp 180◦ turn. Three turn
clearances were considered. It was found that flow recirculation
appeared in the upstream corner in the turn section as well as
along the divider wall after the turn, and the local maxima of
the Sherwood number on the short-side walls inside and after

the turn were mainly caused by the velocity component normal
to each wall.
Zhou et al. [145] measured the heat transfer and pressure drop
in a rotating smooth two-pass coolant passage. It was found that
rotational effects were important in the bend region at lower
Reynolds number with significant enhancement along the bendtrailing surface, and a higher density ratio enhanced the heat
transfer on both the leading and trailing walls of the inlet, bend,
and outlet. In the bend region the enhancement was significant
on the leading surface.
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545

Figure 34 Two-pass channel with various ribs [141].
Figure 33 Two-pass channel with various 45◦ V-shaped ribs [136].

Kim et al. [146–149] measured the detailed heat/mass transfer and pressure drop in a rotating two-pass duct with transverse
ribs. It was found that due to the rotation of the duct, the Sherwood number ratios and pressure coefficients were high on the
trailing surface in the first pass and on the leading surface in
the second pass. In the turn region of the stationary duct two
Dean vortices were transformed into one large asymmetric vortex cell, which changed the heat/mass transfer and pressure drop
characteristics. Cho et al. [150] measured the effect of cross ribs
on heat/mass transfer in a two-pass duct under rotating conditions. Results showed that for the stationary case the turning
effect dominated the secondary flow at the end of the turn, and
for the rotating case in the first pass the Sherwood numbers
heat transfer engineering

on the trailing surface were higher than those on the leading
surface, while in the second pass the Sherwood numbers were
higher on the leading surface. Cho et al. [151] also measured
the heat/mass transfer and flow characteristics in a two-pass rotating rectangular duct with and without 70◦ angled ribs, and
conducted numerical simulations to analyze the flow pattern.
Results showed that large overall heat transfer on the leading
and trailing surfaces for the first and second passes depended on
the rotating speed and turn geometry, but the local heat transfer
was affected mainly by the rib arrangement.
Bunker [152] presented a method to provide substantially
increased convective heat flux on the internal cooled tip cap
of a turbine blade, where arrays of discrete shaped pins were
fabricated and placed, as shown in Figure 35. The detailed heat
transfer distribution over the internal tip cap was obtained based
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Figure 36 Secondary flow and temperature contours in a rotating smooth
channel [157].

Figure 35 Geometries of tip-cap with pin arrays [152].

on a large-scale model of a sharp 180◦ tip turn. Five tip cap
surfaces were tested. It was found that the effective heat transfer
coefficient could be increased by up to a factor of 2.5 due to
the combination of impingement and cross-flow convection on

the pins, as shown in Figure 35. The tip turn pressure drop was
negligible compared to that of a smooth surface.

Computationally Internal Cooling for Blade Tip
With the fast development of computer resources, the increase of computational power makes it economical to simulate
flow and heat transfer inside turbine blade passages. The thermal
and cooling performance then can be optimized and designed
based on numerical analysis. Chen et al. [153, 154], Jang et al.
[155, 156], and Al-Qahtani et al. [157, 158] calculated the 3D
flow and heat transfer in rotating two-pass square channel with
smooth walls or 45◦ /60◦ angled ribs by a second-moment closure
model and a two-layer k-ε isotropic eddy viscosity model. Good
agreement with experimental data of Ekkad and Han [113] was
achieved. The comparison of the results showed that the nearwall second-moment closure model provided accurate predictions of the complex 3D flow and heat transfer resulting from
the rotation and strong wall curvatures. Also, it was observed
that angled ribs with high blockage ratio and a 180◦ sharp turn
produced strong nonisotropic turbulence and heat flux, which
heat transfer engineering

affected significantly the flow field and heat transfer coefficient,
as shown in Figure 36.
Lin et al. [159] performed computations of 3D flow and heat
transfer in a U-shaped square duct for rotating and nonrotating conditions. The flow streamlines, velocity vector fields, and
contours showed how the fluid flow in a U-duct evolved from a
unidirectional one to one with convoluted secondary flows due to
the Coriolis force, centrifugal buoyancy, staggered inclined ribs,
and a 180◦ bend, and also how the nature of the fluid flow affected the surface heat transfer. Suga [160] predicted turbulence
and heat transfer in two types of square sectioned U-bend duct
flows with mild and strong curvature by recent second moment
closures. Suga and Abe [161] applied a higher order version

of the generalized gradient diffusion hypothesis along with the
TCL (two-component-limit) model. They found that the second
moment closure was good enough for predicting flow and heat
transfer in the case of mild curvature, but only the TCL model
was reliable for the strong curvature case.
Iacovides [162] carried out computations of turbulent flows
through stationary and rotating rib-roughened U-bends to explore both numerical and turbulence modeling. Because of using
body-fitted grids and higher order schemes for the discretization
of the convective transport of all flow variables, grid requirements could be reduced. Concerning the turbulence modeling,
the comparisons suggested that a low-Re second-moment closure becomes necessary, but the second-moment closure cannot
account for the effects of negative rotation. Moreover, the finer
mesh computations would examine the predicted turbulence
fields closely. Nikas et al. [163] presented computations of heat
and fluid flow through a square-ended U-bend that rotates about
an axis normal to both the main flow direction and also the axis
of curvature. The main flow features were well reproduced by
all models, but the mean flow within and after the bend was
better reproduced by the low-Re models. On the other hand, turbulence levels within the rotating U-bend were underpredicted,
but low-Re DSM models produced a more realistic distribution.
Along the leading side, all models overpredicted the heat transfer just after the bend, and for the trailing side, the heat transfer
vol. 31 no. 7 2010


B. SUNDEN AND G. XIE

predictions of the low-Re DSM with a differential length-scale
correction term were close to the measurements. Raisee et al.
[164] considered the application of low-Re linear and nonlinear
eddy-viscosity models for the numerical prediction of the velocity and pressure field in flow through two 90◦ curved ducts: one
of a square cross section and one of a rectangular cross section.

The results indicated that for the bend of square cross section
the curvature induced a strong secondary flow, while for the
rectangular cross section the secondary motion was modified at
the corner regions. For both curved ducts, the secondary motion persisted downstream of the bend and disappeared slowly.
Another aspect was that, for the bend of square cross section,
comparisons indicated that both turbulence models could get
reasonable predictions. A wider range of data was available for
the bend of rectangular cross section, and it was found the nonlinear k-ε model showed superior predictions of the turbulence
field and the pressure and friction coefficients.
Murata and Mochizuki [165] numerically studied the centrifugal buoyancy effect on turbulent heat transfer in a rotating
two-pass square channel with 180◦ sharp turns by the large eddy
simulation (LES). It was found that with increasing buoyancy,
the pressure loss coefficient of the sharp turn was decreased and
that of the straight pass was increased in the first pass and decreased in the second pass, and due to the aiding and opposing
buoyancy contributions to the main flow the variation caused
by the buoyancy was larger for the heat transfer on the pressure
surface than on the suction surface. For the studied buoyancy
range the Colburn j factor was kept almost constant.
Sleiti and Kapat [166] predicted numerically the flow field
and heat transfer of high rotation numbers and density ratio flow
in a square internal cooling channel with U-turn. They found that
the four-side-averaged Nusselt number increased linearly with
increasing rotation number but slightly decreased with increasing density ratio. At the center of the U-bend the corner vortices
were suppressed with increasing rotation number, while an increased density ratio resulted in a decrease in all surfaces of the
U-turn. Sleiti and Kapat [167, 168] also predicted the 3D flow
field and heat transfer in a two-pass rib-roughened square channel. Results showed that in the U-turn high shear stresses were
found near the leading and trailing surface, and were increased
by increasing density ratio.
Etemad and Sunden [169, 170], and Etemad et al. [171]
used turbulence models with linear and nonlinear expressions

for the Reynolds stresses to investigate turbulent flow and heat
transfer in a square-sectioned U-bend. Five turbulence models
were evaluated: Suga’s quadratic and cubic low-Re k-ε, V2F
k-ε, RSM-EVH, and RSM GGDH. These models predicted
the stress-induced secondary motion in the straight inlet duct,
and this secondary motion had an impact on the flow in the
bend. It was found that Suga’s model performed slightly better
and offered a higher degree of robustness. Guleren and Turan
[172] used large-eddy simulation (LES) to carry out numerical predictions of developing turbulent flow through stationary
and rotating U-ducts with strong curvature. Their aim was to
validate LES in a strongly curved U-duct for three different
heat transfer engineering

547

cases: stationary, positive, and negative rotational cases. They
found that grid resolution had some effect on the profiles of
the Reynolds stresses. The wall function was responsible for
the excessive turbulent intensities, and LES was superior to the
two-component-limit turbulence model with the predictions of
mean velocities. The primary and secondary flow behavior can
give a better understanding of the origin and development of the
flow separation. Viswanathan and Tafti [173] predicted turbulent
flow field in a two-pass internal cooling duct with normal ribs by
detached eddy simulation (DES) and the unsteady Reynolds averaged Navier–Stokes equations (URANS). Results showed that
DES predicted a slower flow development than LES, whereas
URANS predicted it much earlier than LES computations and
experiments. DES could accurately predict the flow both in the
fully developed region as well as in the 180◦ bend of the duct.
Other related research works about turbulent heat transfer in

serpentine passages are available in research theses [174–186].
Valuable review articles have been presented [8, 9, 187, 188].

SUMMARY
As the turbine inlet temperature is continuously increased
for fast development of current gas turbine engines, the heat
transferred to the blade is increased. To satisfy the even increasingly high inlet temperature, turbine blade cooling becomes an
important issue for new designs. Such cooling includes blade
end-wall cooling, leading-edge cooling, trailing-edge cooling,
and tip cooling. The blade tip is one of the critical regions to be
cooled due to the high thermal load over the tip surface. Therefore, highly accurate and highly detailed local heat transfer and
flow data related to such regions are needed for analysis, and
cooling schemes must be designed to prevent the failure due to
the local hot spots.
As reviewed in the preceding sections, more available data
from experimental measurements and numerical simulations are
for blade tip clearance leakage flow associated with heat transfer
and for near-tip region flow field and heat transfer. Even with
sophisticated clearance control methods to employ, the gap is
never eliminated, and thereby the leakage flow occurs due to the
pressure difference between the pressure side and suction side.
The leakage flow has a pronounced influence on local heat/mass
coefficient distribution and hence the heat load. Thus, whatever
the tip geometry is and whatever the clearance control strategy
is, to develop novel and optimal techniques will require more
research on the detailed and accurate leakage flow and heat/
mass transfer characteristics over the blade tip and near-tip
region.
The blade tip is the most susceptible region subjected to the
large thermal load and is difficult to cool sufficiently. For external cooling, a common technique is to add film cooling through

the tip and near-tip region. The cooling performance is affected
significantly by most conditions, such as film cooling hole configuration, location, and distribution, and the representative flow
vol. 31 no. 7 2010


548

B. SUNDEN AND G. XIE

conditions. For internal cooling, serpentine cooling passages are
designed inside blades, so that the heat from the pressure side
and suction side is picked up by the turning coolant extracted
from compressors. The serpentine channel configuration, aspect
ratio and orientation, rib configuration and location, and rotation
and bend/turn geometry affect significantly the internal cooling
efficiency. Several studies have contributed to the cooling issues.
However, it is not enough to observe the parametric effects of
film cooling for the blade tip. More studies related to combined
film cooling and internal convective cooling are needed. Also,
although a large number of research works have concerned turbulent heat transfer and cooling issues inside serpentine (twopass, multipass) channels, studies concerning internal blade tip
cooling concept and research are still limited. Thus more studies related to these issues are required. Especially, the detailed
flow and heat transfer distribution characteristics and cooling
performance on the tip-cap walls or near-tip region need to be
investigated.
The CFD techniques act as an important role in research and
design of gas turbine components and can provide useful data
related to the detailed flow field and heat transfer coefficient
distribution along the gas turbine blades. With the fast development of CFD techniques as prediction tools, highly accurate
CFD computations are encouraged to provide insight into complex flow and heat/mass transfer as well as the cooling process
on the blade tip, and various turbulence models should be tested

and validated by available experimental data.

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vol. 31 no. 7 2010