Evolution of Turbine Blade, Single Crystal Blade, Super Alloys, and Blade Cooling
Turbine blade
from Turbo-Union RB199 jet
engine.
This is a blade with an outer
shroud which prevents gas leaking round the blade tip in which case it wouldn't
contribute to the force on the aerofoil. The platform at the base of the
aerofoil forms a continuous annulus ring which, together with cooling-air
cavity purge flow prevents hot gas leakage onto the turbine discs. The short
extension, or shank, between the platform and fir-tree fixing in the disc
allows space for cooling-air entry to blade, may control blade vibration modes
and heat transfer to disc rim. The
turbine blades have a golden colour in this engine cutaway.
A turbine blade is a
radial aerofoil mounted in the rim of a turbine disc and which produces a
tangential force which rotates a turbine rotor. Each turbine disc has many
blades. As
such they are used in gas turbine engines and steam turbines.
The blades are responsible for extracting energy from the high temperature,
high pressure gas produced by the combustor. The turbine blades are often
the limiting component of gas turbines. To survive in this difficult
environment, turbine blades often use exotic materials like superalloys and
many different methods of cooling that can be categorized as internal and
external cooling, and thermal barrier coatings.
Blade fatigue is a
major source of failure in steam turbines and gas turbines. Fatigue is caused
by the stress induced by vibration and resonance within the operating range of
machinery. To protect blades from these high dynamic stresses, friction dampers
are used. Blades of wind turbines and water
turbines are designed to operate in different conditions, which typically
involve lower rotational speeds and temperatures.
Introduction
Diagram of a twin spool jet
engine. The high-pressure turbine is connected by a shaft to the high-pressure
compressor to form one spool, or complete rotating assembly(purple)- and the
low-pressure turbine is connected to the low-pressure compressor to form the
other spool (green).
In a gas turbine engine, a
single turbine stage is made up of a rotating disk that holds many turbine
blades and a stationary ring of nozzle guide vanes in front of the blades. The
turbine is connected to a compressor using a shaft (the complete rotating
assembly sometimes called a "spool"). Air is compressed, raising the
pressure and temperature, as it passes through the compressor. The temperature
is then increased by combustion of fuel inside the combustor which is located
between the compressor and the turbine. The high-temperature, high-pressure gas
then passes through the turbine. The turbine stages extract energy from this
flow, lowering the pressure and temperature of the gas and transfer the kinetic
energy to the compressor. The way the turbine works is similar to how the
compressor works, only in reverse, in so far as energy exchange between the gas
and the machine is concerned, for example. There is a direct relationship
between how much the gas temperature changes (increase in compressor, decrease
in turbine) and the shaft power input (compressor) or output (turbine).
For a turbofan engine the number
of turbine stages required to drive the fan increases with the bypass-ratio unless
the turbine speed can be increased by adding a gearbox between the turbine and
fan in which case fewer stages are required. The number of turbine stages
can have a great effect on how the turbine blades are designed for each stage.
Many gas turbine engines are twin-spool designs, meaning that there is a
high-pressure spool and a low-pressure spool. Other gas turbines use three
spools, adding an intermediate-pressure spool between the high- and
low-pressure spool. The high-pressure turbine is exposed to the hottest,
highest-pressure air, and the low-pressure turbine is subjected to cooler,
lower-pressure air.
The difference in conditions
leads to the design of high-pressure and low-pressure turbine blades that are
significantly different in material and cooling choices, though aerodynamic and thermodynamic principles
are the same. Under these severe operating conditions inside the gas and
steam turbines, the blades face high temperature, high stresses, and
potentially high vibrations. Steam turbine blades are critical components in
power plants which convert the linear motion of high-temperature and
high-pressure steam flowing down a pressure gradient into a rotary motion of
the turbine shaft.
Environment and
failure modes
Turbine blades are subjected to
very strenuous environments inside a gas turbine. They face high temperatures,
high stresses, and a potential environment of high vibration. All three of
these factors can lead to blade failures, potentially destroying the engine,
therefore turbine blades are carefully designed to resist these conditions.
Turbine blades are subjected to
stress from centrifugal force (turbine stages can rotate at tens of
thousands of revolutions per minute (RPM)) and fluid forces that can
cause fracture, yielding, or creep failures. Additionally,
the first stage (the stage directly following the combustor) of a modern gas
turbine faces temperatures around 2,500 °F (1,370 °C), up from
temperatures around 1,500 °F (820 °C) in early gas turbines. Modern
military jet engines, like the Snecma M88, can see turbine temperatures of
2,900 °F (1,590 °C). Those high temperatures can weaken the
blades and make them more susceptible to creep failures. The high temperatures
can also make the blades susceptible to corrosion failures. Finally,
vibrations from the engine and the turbine itself can cause fatigue failures.
Higher temperatures lead to
better cycle efficiency in a turbine, calling for a need for materials
innovation that continually pushes the previous benchmark. As the most common
modes of failure in SC and polycrystalline blades are intrinsically different,
SC blades, thanks to their lack of grain boundaries, are ideal for first and
second stage turbine blade applications. Turbine blades are typically made of
Ni-Cr superalloys, which present a unique microstructure.
Single crystal (SC) superalloys are
typically cast with the ⟨001⟩ crystallographic direction along the turbine
blade's axis, and often with a controlled secondary orientation, which has been
found to be ideal for reducing localized stresses. Their microstructure
consists of cuboidal γ′ precipitates aligned along ⟨001⟩ within
a γ matrix. These alloys exhibit anisotropic mechanical properties due to the
FCC structures of γ and γ′, also known as orthotropic properties. These include
tensile and creep properties, which are also influenced by temperature alloy
composition. This anisotropy strongly affects deformation and damage during
thermo-mechanical fatigue (TMF).
Phase Angle
- For an in-phase (IP) TMF cycle, maximum tensile
loading occurs at the maximum temperature of the cycle. The associated
deformation modes are thus high-temperature creep at this tensile loading
state, which can be intensified by internal stress and centrifugal force,
even at moderate temperatures. Conversely, low temperatures correspond to
a compressive state characterized by low-temperature plasticity, typically
occurring near or within cooling channels designed into the blade.
- In certain areas, such as the blade platform,
with insufficient cooling, they can become “hot spots”. In areas like
this, out-of-phase (OP) TMF cycle can be observed. It is characterized
by creep relaxation caused by compressive stress at high
temperatures, and plastic deformation by tensile stress at low
temperatures. Thus, while running and hot, the area will be under
compressive stress and will experience creep relaxation, that will
translate into tension when returned to room temperature. Oxide formation
at high temperatures causes low ductility in SC alloys, which can lead to
cracking in tensile stress, so areas experiencing OP oftentimes have
shorter lifespans, despite experiencing lower average stress and inelastic
strain than IP.
- In comparing the effects of IP vs OP, it is
found that maximum tensile stress is the largest contributor to lifetime,
resulting in OP having a shorter lifespan.
- TMG is a very damaging process, and its fatigue
lifespan for a SC superalloy at any given orientation is 70-90% lower than
isothermal cyclic loading conditions at peak temperature.
Deformation and
Cracking
- Deformation in SC alloys tends to be localized
due to the lack of grain boundaries and anisotropy of the crystal. Deformation
twinning is a common occurrence in SC alloys, which causes crack
growth localized to along the twins, leading to failure. It is a leading
cause of TMF failure, especially of second-generation SC alloys that
contain Re. However, under conditions of temperature or stress gradients
along these deformation bands, it can be difficult to interpret mechanics.
Fatigue crack paths in SC
initiate with Mode I tensile opening and are subsequently propagated
by crystallographic shearing along slip planes. This is in opposition to the
behavior of conventional crack growth that follows the Stage I/Stage II pattern
of crack nucleation (crystal shear) and steady growth (Mode I).
- In OP-TMF, oxidation-fatigue interactions, and
γ′ coarsening at high temperatures are the main causes of crack
nucleation. Cracks tend to form in spots with oxide growth, the process of
which is exacerbated by high temperature loading conditions, leading to
well-defined, planar cracks-a key feature of brittle materials. It has
been observed that localized clusters of twinned plates form ahead of the
crack tip on the {1 1 1} plane, also known as deformation twinning.
This is caused by partial dislocation movement in the corresponding slip
system ({1 1 1}<112>). When big enough, they enable crack
propagation, preferentially localized to along the twins.
- Without the temperature cycling, in low
cycle fatigue (LCF), the deformation mode is dominated by diffusion-based
mechanisms.
- While most propagation takes place at the low
temperature part of the OP-TMF cycle, the high temperature and compressive
state is responsible for damage caused by dissolution and
recrystallization of the γ′ phase at the crack tip. Study of crack closure
shows that crack growth rate is independent of maximum cycle temperature
and dwell time as a function of change in effective stress intensity
factor (ΔKeff), indicating that the driving force for
crack propagation is mitigated by stress relief at the crack tip as a
result of closure.
- For IP-TMF, cracks tend to form at weak
interfaces. There is no observable difference between crack growth rate as
a function of ΔKeff, showing a similar outcome to the
corresponding LCF conditions. At prolonged hold times a reduction in crack
growth closure and increase in K result in crack growth. Thus, IP-TMF and
LCF tests are also described by crack growth rate vs ΔKeff.
- Under stress-controlled conditions, IP has the
shortest fatigue life, followed by LCF, then OP.
Dwell Time
The addition of a compressive
hold time impacts the direction of crack propagation as observed in OP-TMF.
Without a dwell time, cracks initiate at the surface (at areas of oxidation
spiking), and grow perpendicular to the direction of applied stress, before the
growth along twin plates. Upon extension of dwell time beyond 10 minutes, the
crack directly grows along the twin plates until failure. Application of dwell
time decreases fatigue lifespan, even at high temperatures, likely as a result
of stress relaxation, failure mode, increased inelastic strain, and high
tensile mean stress. For IP-TMF, the effects of dwell time on crack
initiation and propagation are accurately predicted by a combined model of
fatigue and creep.
The yield strength anomaly in the
pure γ′ phase describes a state in which the material has a higher yield
strength at higher temperatures, so the impressive mechanical properties of SC
superalloys is a result of the interaction between the γ and γ′ phases’
cuboidal structure. However, this state is unstable, and rafting formed as a
result of high temperature and long dwell times in TMF tests will weaken the
material, leading to a lower yield stress. This effect is thus exacerbated by
lower values for minimum temperature in TMF cycles. As the yield stress
decreases, the inelastic strain range increases, and cracks form sooner.
Crystal
Orientation
Stiffness
TMF life for a given strain range
is determined by the elastic modulus with respect to strain
orientation in a strain controlled TMF test. For Ni alloys, the [001] direction
exhibits lower stiffness, suppressing stress, and corresponding to a longer
life. Typically fatigue lifespan of the different crystal orientations in the
controlled strain test is longest in <001>, which displays elastic
cycling behavior, followed by <011> then <111>, which both display
plastic deformation. Creep damage is the leading cause of fracture,
characterized by octahedral slip planes in <001> and <011>, and
cubic in <111>, which can be expected for FCC. Additionally, due to
the <001> direction’s ability to be exposed to TMF, there is opportunity
for rafting to develop, in which the γ’ strengthening phase has the opportunity
to coalesce into plates. For loading in <001>, OP-TMF leads to
P-type (parallel to loading direction) rafting, and IP-TMF leads to N-type
(perpendicular to loading direction).
Slip planes
Fatigue life is also impacted by
the number of slip planes. For crystallographic orientations with a higher
number of active slip planes, the deformation is less localized, improving
fatigue life in OP-TMF.
Twinning
Twinning is seen less in
<011>-oriented samples in OP-TMF than in <001>.
Low Cycle
Fatigue (LCF)
LCF is characterized by high
strain applied over a low number of cycles at constant temperature. For
<001> and <111> -oriented specimens, cycle softening, or a
reduction in failure stress, is observed at high temperatures, due to the
dissolution of the γ’ phase, dislocation removal, and precipitates, the
creation of dislocations at phase interfaces to relieve stress, and the removal
of dislocations via thermally activated mechanisms. In low-strain conditions,
in <111>-oriented SC specimens experience cyclic hardening, or the
increase in failure stress, as a result of the formation of aligned dislocation
arrays in the matrix preventing interactions between dislocations from other
slip systems. This reduces the fatigue life of the <111> oriented SC
turbine blades.
Fatigue resistance improves with
a uniform distribution of secondary γ′ precipitates in the γ matrix. Cracks
typically initiate and grow along persistent slip bands (PSBs). However, at
high temperatures and strain levels, cracks form earlier due to dislocation
entanglement and stress-driven coarsening of γ′ precipitates.
High Cycle
Fatigue
High cycle fatigue (HCF) is
the leading cause of SC turbine blade failure, characterized by a high cycle
number and a low amplitude stress field, that leads to elastic deformation.
There is no difference between fatigue life between <001> or <111>
orientations in HCF testing, however, features that concentrate stress, such as
holes or notches, can increase the number of endured loading cycles before
failure. These stress concentrators cause samples to not display cleavage
planes upon fracture, similarly to polycrystalline specimens.
A rough zone with multiple planar
facets forms at crack initiation sites due to stress concentrations. Slip bands
tend to develop near stress concentrators, inducing formation of slip bands,
while casting defects create localized high shear stress but have a lower K
than the overall specimen threshold. Intense shear forces around these defects
lead to recrystallization and cavitation at high temperatures by creation of
nucleation sites. Additionally, some superalloys precipitate topologically
close packed structures and carbide precipitations in the rough zone, causing
slip band formation away from the defect area. When the rough zone’s K reaches
the global threshold, macroscopic HCF cracks propagate.
Fretting Fatigue
Fretting fatigue is typically a
result of intense vibrations experienced by the joints and discs of turbine
blades during operation, due to centrifugal and aerodynamic forces. This causes
stress concentrations that induce fatigue and eventually, failure via increased
crack nucleation. Dislocations can move through any of 12 slip planes in a
coherent FCC crystal. As temperature increases, slip line density decreases,
with respect to crystallographic orientation. Thus, fracture is the primary
mode of deformation.
Materials
A limiting factor in early jet
engines was the performance of the materials available for the hot section
(combustor and turbine) of the engine. The need for better materials spurred
much research in the field of alloys and manufacturing techniques, and that
research resulted in a long list of new materials and methods that make modern
gas turbines possible. One of the earliest of these was Nimonic, used
in the British Whittle engines.
The development of superalloys in
the 1940s and new processing methods such as vacuum induction melting in the
1950s greatly increased the temperature capability of turbine blades. Further
processing methods like hot isostatic pressing improved the alloys used for
turbine blades and increased turbine blade performance. Modern turbine blades
often use nickel-based superalloys that incorporate chromium, cobalt, and
rhenium.
Aside from alloy improvements, a
major breakthrough was the development of directional solidification (DS) and
single crystal (SC) production methods. These methods help greatly increase
strength against fatigue and creep by aligning grain boundaries in one
direction (DS) or by eliminating grain boundaries altogether (SC). SC research
began in the 1960s with Pratt and Whitney and took about 10 years to be
implemented. One of the first implementations of DS was with the J58 engines of
the SR-71.
A turbine blade with thermal
barrier coating. This blade has no tip shroud so tip leakage is controlled by
the clearance between the tip and a stationary shroud ring attached to the
turbine case.
Another major improvement to
turbine blade material technology was the development of Thermal Barrier
Coatings (TBC). Where DS and SC developments improved creep and fatigue
resistance, TBCs improved corrosion and oxidation resistance, both of which
became greater concerns as temperatures increased. The first TBCs, applied in
the 1970s, were aluminide coatings. Improved ceramic coatings became
available in the 1980s. These coatings increased turbine blade temperature
capability by about 200 °F (90 °C). The coatings also improve blade
life, almost doubling the life of turbine blades in some cases.
Most turbine blades are
manufactured by investment casting (or lost-wax processing). This
process involves making a precise negative die of the blade shape that is
filled with wax to form the blade shape. If the blade is hollow (i.e., it has
internal cooling passages), a ceramic core in the shape of the passage is
inserted into the middle. The wax blade is coated with a heat-resistant
material to make a shell, and then that shell is filled with the blade alloy.
This step can be more complicated for DS or SC materials, but the process is
similar. If there is a ceramic core in the middle of the blade, it is dissolved
in a solution that leaves the blade hollow. The blades are coated with a TBC,
and then any cooling holes are machined.
Ceramic Matrix Composites (CMC),
where fibers are embedded in a matrix of polymer derived ceramics, are
being developed for use in turbine blades. The main advantage of CMCs over
conventional superalloys is their light weight and high temperature
capability. SiC/SiC composites consisting of a silicon carbide matrix
reinforced by silicon carbide fibers have been shown to withstand
operating temperatures 200°-300 °F higher than nickel superalloys. GE
Aviation successfully demonstrated the use of such SiC/SiC composite
blades for the low-pressure turbine of its F414 jet engine.
List of turbine
blade materials
Note: This list is not inclusive
of all alloys used in turbine blades.
- U-500 This
material was used as a first stage (the most demanding stage) material in
the 1960s, and is now used in later, less demanding, stages.
- Rene 77
- Rene N5
- Rene N6
- PWA1484
- CMSX-4
- CMSX-10
- Inconel
- IN-738 – GE
used IN-738 as a first stage blade material from 1971 until 1984, when it
was replaced by GTD-111. It is now used as a second stage material. It
was specifically designed for land-based turbines rather than aircraft
gas turbines.
- GTD-111 Blades
made from directionally solidified GTD-111 are being used in many GE
Energy gas turbines in the first stage. Blades made from equiaxed
GTD-111 are being used in later stages.
- EPM-102 (MX4 (GE), PWA
1497 (P&W)) is a single crystal superalloy jointly developed
by NASA, GE Aviation, and Pratt & Whitney for the High Speed
Civil Transport (HSCT). While the HSCT program was cancelled, the
alloy is still being considered for use by GE and P&W.
- Nimonic 80a was
used for the turbine blades on the Rolls-Royce Nene and de Havilland Ghost
- Nimonic 90 was
used on the Bristol Proteus.
- Nimonic 105 was
used on the Rolls-Royce Spey.
- Nimonic 263 was
used in the combustion chambers of the Bristol Olympus used
on the Concorde supersonic airliner.
- 3D printed thermoplastic resin to
create wind turbine blades is in development in a partnership
between ORNL, NREL, and GE Renewable Energy.
Cooling
At a constant pressure
ratio, thermal efficiency of the engine increases as the turbine
entry temperature (TET) increases. However, high temperatures can damage the
turbine, as the blades are under large centrifugal stresses and materials are
weaker at high temperature. So, turbine blade cooling is essential for the
first stages but since the gas temperature drops through each stage it is not
required for later stages such as in the low-pressure turbine or a power
turbine. Current modern turbine designs are operating with inlet
temperatures higher than 1900 kelvins which is achieved by actively cooling the
turbine components.
Methods of
cooling
Laser-drilled holes permit film
cooling in this first-stage V2500 nozzle guide vane.
Turbine blades are cooled using
air, except for limited use of steam cooling in a combined cycle power plant.
Water cooling has been extensively tested but has never been introduced. The
General Electric "H" class gas turbine has cooled rotating blades and
static vanes using steam from a combined cycle steam turbine although GE was
reported in 2012 to be going back to air-cooling for its
"FlexEfficiency" units. Liquid cooling seems to be more
attractive because of high specific heat capacity and chances of evaporative
cooling but there can be leakage, corrosion, choking and other problems which
work against this method. On the other hand, air cooling allows the
discharged air into main flow without any problem. Quantity of air required for
this purpose is 1–3% of main flow and blade temperature can be reduced by
200–300 °C. There are many techniques of cooling used in gas turbine
blades; convection, film, transpiration cooling, cooling effusion, pin fin
cooling etc. which fall under the categories of internal and external cooling.
While all methods have their differences, they all work by using cooler air
taken from the compressor to remove heat from the turbine blades.
Internal cooling
Convection
cooling
Blade cooling by
convection
It works by passing cooling air
through passages internal to the blade. Heat is transferred by conduction through
the blade, and then by convection into the air flowing inside of the blade. A
large internal surface area is desirable for this method, so the cooling paths
tend to be serpentine and full of small fins. The internal passages in the
blade may be circular or elliptical in shape. Cooling is achieved by passing
the air through these passages from hub towards the blade tip. This cooling air
comes from an air compressor. In case of gas turbine the fluid outside is
relatively hot which passes through the cooling passage and mixes with the main
stream at the blade tip.
Impingement
cooling
Impingement Cooling
A variation of convection
cooling, impingement cooling, works by hitting the inner surface of the blade
with high velocity air. This allows more heat to be transferred by convection
than regular convection cooling does. Impingement cooling is used in the
regions of greatest heat loads. In case of turbine blades, the leading edge has
maximum temperature and thus heat load. Impingement cooling is also used in mid
chord of the vane. Blades are hollow with a core. There are internal cooling
passages. Cooling air enters from the leading edge region and turns towards the
trailing edge.
External cooling
Film cooling
Rendering of a turbine blade with
cooling holes for film cooling
Film Cooling
Film cooling (also called thin film
cooling), a widely used type, allows for higher cooling effectiveness than
either convection and impingement cooling. This technique consists of
pumping the cooling air out of the blade through multiple small holes or slots
in the structure. A thin layer (the film) of cooling air is then created on the
external surface of the blade, reducing the heat transfer from main flow, whose
temperature (1300–1800 kelvins) can exceed the melting point of
the blade material (1300–1400 kelvins). The ability of the film cooling
system to cool the surface is typically evaluated using a parameter called
cooling effectiveness. Higher cooling effectiveness (with maximum value of one)
indicates that the blade material temperature is closer to the coolant
temperature. In locations where the blade temperature approaches the hot gas
temperature, the cooling effectiveness approaches to zero. The cooling
effectiveness is mainly affected by the coolant flow parameters and the
injection geometry. Coolant flow parameters include the velocity, density,
blowing and momentum ratios which are calculated using the coolant and
mainstream flow characteristics. Injection geometry parameters consist of hole
or slot geometry (i.e. cylindrical, shaped holes or slots) and injections
angle. A United States Air Force program in the early 1970s funded the
development of a turbine blade that was both film and convection cooled, and
that method has become common in modern turbine blades.
Injecting the cooler bleed into
the flow reduces turbine isentropic efficiency; the compression of the cooling
air (which does not contribute power to the engine) incurs an energetic
penalty; and the cooling circuit adds considerable complexity to the engine. All
of these factors have to be compensated by the increase in overall performance
(power and efficiency) allowed by the increase in turbine temperature.
In recent years, researchers have
suggested using plasma actuator for film cooling. The film cooling of
turbine blades by using a dielectric barrier discharge plasma
actuator was first proposed by Roy and Wang. A horseshoe-shaped plasma
actuator, which is set in the vicinity of holes for gas flow, has been shown to
improve the film cooling effectiveness significantly. Following the previous
research, recent reports using both experimental and numerical methods
demonstrated the effect of cooling enhancement by 15% using a plasma actuator.
Cooling effusion
Cooling by
effusion
The blade surface is made of
porous material which means having a large number of small orifices on the
surface. Cooling air is forced through these porous holes which forms a film or
cooler boundary layer. Besides this uniform cooling is caused by effusion of
the coolant over the entire blade surface.
Pin fin cooling
In the narrow trailing edge film
cooling is used to enhance heat transfer from the blade. There is an array of
pin fins on the blade surface. Heat transfer takes place from this array and
through the side walls. As the coolant flows across the fins with high
velocity, the flow separates and wakes are formed. Many factors contribute
towards heat transfer rate among which the type of pin fin and the spacing
between fins are the most significant.
Transpiration
cooling
This is similar to film
cooling in that it creates a thin film of cooling air on the blade, but it is
different in that air is "leaked" through a porous shell rather than
injected through holes. This type of cooling is effective at high temperatures
as it uniformly covers the entire blade with cool air. Transpiration-cooled
blades generally consist of a rigid strut with a porous shell. Air flows
through internal channels of the strut and then passes through the porous shell
to cool the blade. As with film cooling, increased cooling air decreases
turbine efficiency, therefore that decrease has to be balanced with improved
temperature performance.
Evolution of Turbine Blade, Single Crystal Blade, Super Alloys, and Blade Cooling
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