Field of the Invention
[0001] The present invention relates to two methods for operating and treating a turbine
assembly.
Background to the Invention
[0002] Modern turbines often operate at extremely high temperatures and are operated with
high speed. This causes high stresses that act on turbine components and may have
detrimental effects on selected parts or sections of the components especially prone
to fatigue. For example, the effect of temperature on the turbine blades and/or stator
vanes can be detrimental to the efficient operation of the turbine and can, in extreme
circumstances, lead to distortion and possible failure of the blade or vane. Hence,
turbines may include hollow blades or vanes incorporating cooling channels, inserts
and pedestals for cooling purposes to overcome this risk of failure. However, cooling
holes, especially at the disc of a turbine assembly that feed cooling medium to the
aerofoil, are regions that are exposed to high stress and are prone to cyclic life
limiting features. Moreover, structures connecting turbine assemblies, like turbine
blades or vanes, with a casing or with another components, e.g. a disc or platform,
tend to show fatigue, especially in regions of fastening structures, like, for example,
bolt holes.
[0003] Many turbine components exhibit a shorter than optimal operating life due to failure
mechanisms such as Low Cycle Fatigue (LCF). Hitherto the cyclic life of the component
is established analytically or via mechanical testing on bespoke test rigs. Depending
on the component material chemistry and metallurgical structure the design intent
may be either to remove the component from service prior to the onset of crack initiation
or to allow crack initiation and to predict crack growth through fracture mechanics
methods. It is normal and prudent practice for the Turbomachinery operator to maintain
a numerical count of operating cycles the machine and/or components has endured. It
is possible to modify the count with a multiplication factor to compensate for more
damaging operating regimes such as rapid starts to full power. Once the declared cyclic
life has been reached it is normal to declare the component unfit for further service
and the component is sentenced as scrap and sent for recycling. With these approaches
it can occur that components are discarded that may still have operating capacity
thus wasting resources and money.
[0004] It is a first objective of the invention to provide a method for operating and treating
a turbine assembly with which the above-mentioned shortcomings can be mitigated, and
especially, with which the benefit of environmental, energy and cost savings can be
provided. Moreover, turbine assemblies with longer cyclic life in comparison to turbine
assemblies operated and treated according to state of the art principles can be obtained.
[0005] It is a second objective of the present invention to provide an alternative method
for operating and treating a turbine assembly that results in the same advantages
as stated above.
Summary of the Invention
[0006] Accordingly, the present invention provides a method for operating and treating a
turbine assembly, wherein the method comprises at least the steps of: determining
at least a section of the turbine assembly that is prone to fatigue while operating
the turbine assembly, defining a specific operating time of the section of the turbine
assembly, operating the section up to the specific operating time, treating the section
that is prone to fatigue, especially low cycle fatigue, after operating the section
up to the specific operating time by material removal and thus expose a sub-section
of the section and operating the sub-section.
[0007] Due to the inventive matter an advantageous repair or refurbishment strategy can
be provided. It allows a rejuvenation of the section and thus the whole turbine assembly
by material removal of the affected areas of the turbine assembly. Typically, only
small aspects of the turbine assembly reach its cycle life; the remaining features
in the turbine assembly are often capable of many more cycles. This solution gives
an effective and economical method of salvaging components that have reached the end
of their initial cyclic life allowing them to be returned to engine service resulting
in significant cost savings to overhauled assemblies and supply chain. There is a
substantial environmental and energy saving benefit in not scrapping the turbine assembly
which necessitates re-melting, re-forging and re-machining which also carries the
attendant risk of being sentenced as scrap at any point of the process for reasons
such as material defect, incorrect dimensions during machining or handling damage.
The re-machining cost of the section and associated non-destructive testing is minimal
compared to the cost of the production of the turbine assembly or parts thereof, like
a disc forging, disc machining, curvic or hirth teeth and formation of blade roots.
Moreover, by treating the section further operating time of the sub-section of the
section can be gained.
[0008] Even if a term like section, material, sub-section, rim, layer or surface (also see
below) is used in the singular or in a specific numeral form in the claims and the
specification the scope of the patent (application) should not be restricted to the
singular or the specific numeral form. It should also lie in the scope of the invention
to have more than one or a plurality of the above mentioned structure(s).
[0009] A turbine assembly is intended to mean an assembly provided for a turbine, like a
gas turbine, wherein the assembly for example possesses or is at least a disc, a platform
or an aerofoil. Preferably, the turbine assembly is part of a turbine cascade and/or
turbine wheel comprising circumferential arranged aerofoils.
[0010] Moreover, it may further comprise an outer and an inner platform arranged at opponent
ends of the aerofoil(s) or a shroud and a root portion arranged at opponent ends of
the aerofoil(s). A "section of the turbine assembly that is prone to fatigue while
operating the turbine assembly" may be any structure or region of the turbine assembly
feasible for a person skilled in the art that shows fatigue while operated or under
selected operating conditions or after a specific operating time. This vulnerability
of the section to fatigue may be caused due to a specific geometry of the section
or its surroundings, a selected positioning at the turbine assembly e.g. relatively
to negatively influencing factors, like a flow path of hot flow medium, or a specific
function of the section, like a fastening function.
[0011] The term "specific operating time" should be understood as the time the turbine assembly
is operated until a pre-defined number of operating cycles are performed and thus
a certain degree of accumulated damage of the section caused by fatigue is to be expected.
Further, the specific operating time is the time the turbine assembly will be operated
beforehand of the (respective) refurbishment step. The specific operating time is
less than the maximum operating time of the turbine assembly, wherein the maximum
operating time is the time until the turbine assembly or a part thereof is damaged
to such a degree that it is unsafe or unfit to further operate it. The specific operating
time may also be referred to as pre-defined fatigue life and the maximum operating
time as nominal fatigue life.
[0012] The removal of material can be performed by any method feasible for a person skilled
in the art, like drilling, milling or machining away. It is, for example, possible
to use either machine tools and/or hand dressing techniques with or without the use
of guide fixtures. Furthermore, the material is removed from a nominal surface of
the section exposing a surface of the sub-section. In this context nominal surface
should be understood as the surface of the section as it is embodied at a start of
a service life of the turbine assembly. Further, a sub-section should be understood
as a part of the section e.g. being spatially positioned "beneath" the nominal surface
of the section.
[0013] It is further provided that the section is a section comprising a change in its contour.
Hence, an especially vulnerable embodied structure can be treated and refurbished.
The feature changing the contour may be any feature feasible for a person skilled
in the art, like an edge, a pedestal, a step, a counter bore or a hole. Preferably,
the section comprises a stress raiser or the feature causing the change in contour
is a stress raiser. In other words, the feature increases a stress concentration in
the section, wherein the stress is at the highest at the location of the feature or
adjacent to the feature.
[0014] Moreover, the section is preferably a hole with its surface or rim. Thus, a structure
widely used in Turbomachinery can be effectively rejuvenated. In this exemplary embodiment
the hole is the stress raiser, which increases the stress at the surface or rim of
the hole. Hence, the material to be removed is the material surrounding the hole or
stress raiser.
[0015] Furthermore, the section is preferably a hole, e.g. a cooling hole that has a breakout
that breaks out at its surface or rim at an angle building an acute corner, wherein
the acute corner is the life limiting area or feature. Further, the material of the
acute corner would be removed (e.g. by drilling, milling or be machined away) for
example perpendicular to a surface of the corner or at an angle to the surface.
[0016] Beneficially, the hole is a cooling hole or a bolt hole. Due to this, regions being
subjected to e.g. high connecting forces or high temperatures as well as to periodically
changes in temperature can be refurbished. The hole can be any cooling hole or bolt
hole feasible for a person skilled in the art, like an exit or entry hole for cooling
medium of an aerofoil or its root portion or of the aerofoil itself or a fastening
hole of e.g. the aerofoil or the disc. Preferably, the hole is a hole of a turbine
disc. The section that is prone to fatigue may be in these exemplary embodiments the
region at an acute corner of a cooling air feed hole at a bottom of a blade root slot,
a region around a bolt hole of a multi tie bolted rotor arrangement, a region around
an air transfer hole or a region around a bolt clearance hole.
[0017] Often the site of the cyclically aged material is localized at the feature changing
the contour of the section e.g. a hole and penetrates the sections structural material
surrounding the feature.
[0018] According to an advantageous embodiment of the invention the method comprises the
further step of: treating the section by material removal by enlarging a diameter
of the hole. Hence, material exposed to a high stress and being compromised during
beforehand operation of the turbine assembly is removed rejuvenating the hole. The
amount of material removed would vary from part to part and from service strategy
to service strategy. The amount of material to be removed should result in a balance
between an optimal life recovery for as few services as possible (or within acceptable/scheduled
levels of servicing) and acceptable levels of material removal without compromising
the overall integrity of the component. This amount could have any value between 1%
and 60% of an original diameter of the hole feasible for a person skilled in the art.
[0019] According to a preferred embodiment of the invention the method comprises the further
or alternative step of: treating the section by material removal from an acute corner
of a hole with a breakout that breaks out at a surface of the hole at an angle. Hence,
a life limiting area can advantageously be removed.
[0020] However, it was found, that by removing about 10% to 30% of the diameter of the hole
or about 10% to 30% of a diameter of the hole with the breakout at the acute corner
an advantageous refurbishment could be obtained. The proposed range has been shown
to have a satisfactory balance between a gain in cycle lives and a prevention of negative
effects on the overall integrity of the turbine assembly or its treated section. The
component refurbishment technique can rejuvenate the component fully or partly dependent
on the amount of material removed. Hence, it is possible to restart the cyclic "clock"
by fully removing all the aged material or partially "wind the cyclic clock back"
by removing a lesser amount of material if say another feature that could not be reworked
i.e. the blade root was the next feature to exhaust its cyclic life.
[0021] In case of the embodiment with the hole that breaks out at its surface at an angle
or that comprises the angled breakout the removal can be performed easily when the
acute corner is removed e.g. drilled/machined away.
[0022] Further, the geometry of the feature (hole) or the geometry around the feature such
as the hole break out angle affects or impacts the stress concentration in the material
at the feature or of the material surrounding the feature in different (radial) distances
from the feature or from a midpoint of the hole. Hence, the amount of material removed
is, for example, dependent of the geometry of the feature. Thus, also a function of
the hole can affect the amount of material to be removed. For example, at the surface
or rim of a cooling hole that, for example, breaks out at an angle the stress concentration
factor for the material is high and drops quickly with increasing distance from the
surface or rim of the hole. In contrast, at a surface or rim of a bolt hole the stress
concentration factor is normally lower. However, the decline of the stress concentration
factor is normally shallower in comparison with a cooling hole. Consequently, to gain
the same amount of cycle lives the removal of material may have to be higher for a
bolt hole in comparison with a cooling hole. Thus, it may be enough to remove 10%
to 25% of the original diameter of the hole for a hole with a diameter up to 16 mm,
mostly cooling holes, and 10% to 30% of the original diameter of a hole with a diameter
above 16 mm, normally bolt holes. For a cooling hole with a diameter of 5.6 millimeter
(mm) it might be enough to remove 0.7 mm (12.5%) and for a bolt hole with a diameter
of 42.8 mm, for example, 10.7 mm (25%) may have to be removed in order to sufficiently
rejuvenate the cyclic life.
[0023] Furthermore, other factors influence the amount of material that requires removal.
Such a factor may be any one feasible for a person skilled in the art, for example,
a characteristic of the section or turbine assembly (like material, material composition,
size) or an operating condition of the turbine engine (like temperature, temperature
change, speed, flow velocity of a working or cooling fluid). To determine these factors
empirical values, like in-service life experience or cycle life testing, can be included
into the consideration. Hence, the method comprises the further step of: determining
the amount of material to be removed by a consideration of at least one of the factors
consisting out of the group of: internal stress, contact stress, surface stress, temperature,
material capability, material thickness, operating speed, flow velocity of a flow
medium and desired future fatigue life.
[0024] The consideration can be done by any method feasible for a person skilled in the
art, like empirically or statistically. Beneficially, the method comprises the further
step of: determining the amount of material to be removed or an extent of the aged
or degraded material by at least one of the methods out of the group consisting of:
Finite Element Analysis, hand calculations, obtaining results from laboratory testing
or from examination of engine run components and/or applying statistics over the results.
Hence, results can be obtained easily. For example, during the procedures "laboratory
testing" or "examination of engine run components" components are cut-up (or sectioned)
and the material is tested or examined.
[0025] Preferably, the method comprises the further step of: determining the specific operating
time by a use of a cumulative damage technique. Thus, proven concepts can be employed.
This may be any technique feasible for a person skilled in the art, like the linear
damage rule, the Marko-Starky theory, damage theories based on endurance limit reduction,
the Corton-Dolon model, the Freudenthal-Heller approach, two-stage linear damage theories,
damage theories based on crack growth_concept, the damage curve approach, the double-damage
curve approach, the hybrid theory or the Palmgren-Miner rule.
[0026] In a further realisation of the invention the method comprises the further step of:
determining the specific operating time by a use of the Palmgren-Miner rule. By considering
the amount of material that would be compromised at different material depth away
from the feature (e.g. a hole) during a maximal operating time the specific operating
time can be determined easily. Moreover, the further operating time of the sub-section
can be also determined.
[0027] According to a further aspect of the invention the material removal from the section
is performed in one single step. Hence, the risk of damaging the turbine assembly
or the section during the refurbishment is minimised. Alternatively, the material
removal from the section can be performed in several steps. This provides the possibility
to perform a security check of the section before each subsequent refurbishment step.
[0028] Ideally, the refurbishment would be done such that the cyclic live recovered was
high and the times this would need was minimal, and also that an acceptable amount
of material was removed each time without compromising the turbine assembly or the
section thereof. In general, the more material removed, the more life will be recovered.
However, for further refurbishment cycles the rate reduces.
[0029] The number of reworking opportunities depends on how much structural material remains
after reworking to carry out the required component duty. This limited material removal
effectively moves back the LCF count closer to original design life allowing the turbine
assembly to be returned to service with a revised life limiting declaration or only
sufficient material to be removed to meet a pre-determined cyclic life. Thus, the
method comprises the further step of: repeating the steps of defining the specific
operating time, operating the section, treating the section and operating the sub-section
several times. Hence, a cyclical rejuvenation can be performed.
[0030] With this invention it may be possible to use a singular turbine assembly, e.g. a
disc, refurbished one or more times to achieve a full design life of the Turbomachinery.
The refurbished component life cycle cost approximates to:

[0031] The above formula is comparative to replacing Low Cycle Fatigue expired parts.
[0032] A number of the refurbishment steps may be planned beforehand of the first operating
of the turbine assembly. In this case the specific operating time and each subsequent
operating time are determined beforehand of the first operation by applying the Palmgren-Miner
rule using a number of summands, wherein the number of summands is equal to the number
of operating steps or the treatment/refurbishment steps is n-1, wherein n is the number
of positions examined to determine possible operating cycles. It may be also possible
to determine if a further refurbishment step is possible after the sub-section is
operated its further operating time. This would be done by determining the further
operating time after the beforehand operating time of the respective sub-section is
finished.
[0033] In addition, the further step of: defining the specific operating time, operating
the section, treating the section and operating the sub-section can be repeated until
an abort criterion is reached. This abort criterion may be any criterion feasible
for a person skilled in the art, like reaching a maximal operating time of the section,
reaching a rework limit, like a minimal material condition of the section (e.g. thickness,
instability) or an unsuitable geometry of the section.
[0034] A secure operation of the turbine assembly and thus the turbine engine can be provided
when the further step of: defining a specific operating time, operating the section,
treating the section and operating the sub-section are repeated till the defined specific
operating time in the step defining the specific operating time is less than a difference
between a maximum operating time of the section and the already performed operating
time(s).
[0035] In a further advantageous embodiment turbine assembly is a turbine disc.
[0036] The present invention also provides a method for operating and treating a turbine
assembly, that comprises at least the steps of: determining at least a section of
the turbine assembly that is prone to fatigue while operating the turbine assembly,
overlaying the section with an additional depth of material to form an additional
layer of material, defining a specific operating time of the layered section, operating
the turbine assembly up to the specific operating time, treating the layered section
by material removal from the additional layer of material and operating the section.
[0037] Due to the inventive matter an advantageous repair or refurbishment strategy can
be provided. It allows a rejuvenation of the layer and consequently of the section
and thus the whole turbine assembly by material removal of the affected areas of the
turbine assembly. This solution gives an effective and economical method of salvaging
components that have reached the end of their initial cyclic life allowing them to
be returned to engine service resulting in significant cost savings to overhauled
assemblies and supply chain. There is a substantial environmental and energy saving
benefit in not scrapping the turbine assembly which necessitates re-melting, re-forging
and re-machining which also carries the attendant risk of being sentenced as scrap
at any point of the process for reasons such as material defect, incorrect dimensions
during machining or handling damage. The re-machining cost of the section and associated
non-destructive testing is minimal compared to the cost of the production of the turbine
assembly or parts thereof, like a disc forging, disc machining, curvic or hirth teeth
and formation of blade roots. Moreover, by treating the layered section further operating
time of the section can be gained.
[0038] Definitions or explanations given in the text above should also apply to the embodiments
of the second aspect of the invention or should be understood analogously.
[0039] The amount of material being added by overlaying the section or a nominal surface
thereof, can be determined analogously to the determination of the amount of material
to be removed according to the method presented in the first aspect of the invention.
However, it may be also possible to add more or less material depending on the actual
properties of the turbine assembly or its section. Moreover, also the amount of material
being removed from the layered section can be determined with the above stated principles
or methods.
[0040] The material added might be the same material as the material of the section or it
may be another material or material combination, like a protective coating. Preferably,
the added material is the same material as the material of the section. Hence, the
same material properties apply and the same effects on the material of the layer can
be expected.
[0041] Advantageously, the method comprises the further step of: treating the layered section
by removing a whole amount of the added material of the additional layer. Hence, the
section can be refurbished effectively. Consequently, after the refurbishment the
section would be exposed. Additionally, the removed material is the material of the
additional layer. Further, also a removal of a depth of the section would be possible
and thus an exposure of a sub-section of the section.
[0042] According to a further aspect of the invention the material removal from the additional
layer is performed in one single step. Hence, the risk of damaging the turbine assembly
or the section during the refurbishment is minimised. Alternatively, the material
removal from the additional layer can be performed in several steps. This provides
the possibility to perform a security check of the section before each subsequent
refurbishment step. Hence, by removing the material in more than one step at least
the first material removal exposes a sub-layer of the layer or the section.
[0043] According to a preferred embodiment of the invention the method comprises the further
step of: repeating the steps of defining the specific operating time, overlaying the
section, treating the layered section and operating the section several times. Thus,
the turbine assembly can be used effectively and consequently saving resources and
money. Preferably, the steps are repeated till the defined specific operating time
in the step defining the specific operating time is less than a difference between
a maximum operating time of the layered section and the already performed operating
time(s).
[0044] In summary, both aspects of the invention are focused on the same features, namely,
the rejuvenation of a turbine assembly/component or part thereof by removing fatigued
material after a beforehand defined specific operating time to operate the rejuvenated
turbine assembly/component further.
Brief Description of the Drawings
[0045] The present invention will be described with reference to drawings in which:
- FIG 1:
- shows a schematically and sectional view of a gas turbine engine comprising several
inventive turbine assemblies,
- FIG 2:
- shows a perspective view of a part of a turbine assembly embodied as a turbine disc
of the gas turbine engine of FIG 1 comprising a cooling hole beforehand of a refurbishment,
- FIG 3:
- shows a perspective view of the part from FIG 2 after the refurbishment,
- FIG 4:
- shows a cross section through the turbine assembly along line IV-IV in FIG 2 comprising
a bolt hole,
- FIG 5:
- shows air transfer holes of the turbine assembly from FIG 2,
- FIG 6:
- shows a diagram depicting a stress magnitude of 980 MPa at a surface of a cooling
hole and of 600 MPa at a distance of 0.8 mm from the surface of the cooling hole during
operation,
- FIG 7:
- shows a diagram depicting crack initiation after 8086 cycles for a surface of a cooling
hole which was exposed to a stress of 980 MPa during operation,
- FIG 8:
- shows a diagram depicting crack initiation after 105143 cycles for a region being
at a distance of 0.8 mm from the surface of the cooling hole from FIG 7 which was
exposed to a stress of 600 MPa during operation,
- FIG 9:
- shows a diagram depicting the life gained in dependency of the amount of material
removed for different cooling holes,
- FIG 10:
- shows a diagram depicting the life gained in dependency of the amount of material
removed for different bolt holes and
- FIG 11:
- shows a flow-chart for a refurbishment strategy according to the inventive method.
Detailed Description of the Illustrated Embodiments
[0046] The present invention is described with reference to an exemplary turbine engine
26 having a single shaft 38 or spool connecting a single, multi-stage compressor section
30 and a single, one or more stage turbine section 34. However, it should be appreciated
that the present invention is equally applicable to two or three shaft engines and
which can be used for industrial, aero or marine applications.
[0047] The terms upstream and downstream refer to the flow direction of the airflow and/or
working gas flow through the engine 26 unless otherwise stated. If used, the terms
axial, radial and circumferential are made with reference to a rotational axis 36
of the engine 26.
[0048] FIG 1 shows an example of a gas turbine engine 26 in a sectional view. The gas turbine
engine 26 comprises, in flow series, an inlet 28, a compressor section 30, a combustion
section 32 and a turbine section 34, which are generally arranged in flow series and
generally in the direction of a longitudinal or rotational axis 36. The gas turbine
engine 26 further comprises a shaft 38 which is rotatable about the rotational axis
36 and which extends longitudinally through the gas turbine engine 26. The shaft 38
drivingly connects the turbine section 34 to the compressor section 30.
[0049] In operation of the gas turbine engine 26, air 40 which is taken in through the air
inlet 28 is compressed by the compressor section 30 and delivered to the combustion
section or burner section 32. The burner section 32 comprises a burner plenum 42,
one or more combustion chambers 44 defined by a double wall can 46 and at least one
burner 48 fixed to each combustion chamber 44. The combustion chambers 44 and the
burners 48 are located inside the burner plenum 42. The compressed air passing through
the compressor section 30 enters a diffuser 50 and is discharged from the diffuser
50 into the burner plenum 42 from where a portion of the air enters the burner 48
and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and
the combustion gas 52 or working gas from the combustion is channelled via a transition
duct 54 to the turbine section 34.
[0050] This exemplary gas turbine engine 26 has a annular combustor section arrangement
56, which is constituted by an annular array of combustor cans 46 each having the
burner 48 and the combustion chamber 44, the transition duct 54 has a generally circular
inlet that interfaces with the combustion chamber 44 and an outlet in the form of
an annular segment. An annular array of transition duct outlets form an annulus for
channelling the combustion gases to the turbine section 34.
[0051] The turbine section 34 comprises a number of turbine assemblies 10 embodied as blade
carrying discs 58 or turbine wheels 60 attached to the shaft 38. In the present example,
the turbine section 34 comprises two discs 58 each carry an annular array of aerofoils
62 each embodied as a turbine blade. However, the number of blade carrying discs 58
could be different, i.e. only one disc 58 or more than two discs 58. In addition,
turbine cascades 64 are disposed between the turbine blades. Each turbine cascade
64 carries an annular array of aerofoils 62 in the form of guiding vanes, which are
fixed to a stator 66 of the gas turbine engine 26. Between the exit of the combustion
chamber 44 and the leading turbine blades inlet guiding vanes or nozzle guide vanes
68 are provided and turn the flow of working gas 52 onto the turbine blades.
[0052] The combustion gas 52 from the combustion chamber 44 enters the turbine section 34
and drives the turbine blades which in turn rotate the shaft 38. The guiding vanes
68 serve to optimise the angle of the combustion or working gas 52 on to the turbine
blades. The turbine section 34 drives the compressor section 30. The compressor section
30 comprises an axial series of guide vane stages 70 and rotor blade stages 72. The
rotor blade stages 72 comprise a rotor disc 58 supporting aerofoils 62 as an annular
array of turbine blades.
[0053] The compressor section 30 also comprises a stationary casing 74 that surrounds the
rotor stages 72 in circumferential direction 76 and supports the vane stages 70. The
guide vane stages 70 include an annular array of radially extending aerofoils 62 embodied
as vanes that are mounted to the casing 74. The vanes are provided to present gas
flow at an optimal angle for the blades at a given engine operational point. Some
of the guide vane stages 70 have variable vanes, where the angle of the vanes, about
their own longitudinal axis, can be adjusted for angle according to air flow characteristics
that can occur at different engine operations conditions.
[0054] The casing 74 defines a radially outer surface 78 of a passage 80 of the compressor
section 30. A radially inner surface 82 of the passage 80 is at least partly defined
by a rotor drum 84 of the rotor which is partly defined by the annular array of blades.
[0055] During operation of the turbine engine 26 features of components of the engine 26,
like cooling holes 18 or bolt holes 20 of the turbine assemblies 10 or the discs 58
or of the aerofoils 62 are exposed to detrimental conditions inducing internal stresses
in the components. Thus, phenomena like low cycle fatigue (LCF) may occur which e.g.
results in the formation of cracks. Typically, only small aspects of the components
or the turbine assembly 10 reach its cycle life; the remaining features in the component
are often capable of many more cycles. However, these damaged regions deem the turbine
assembly 10 unfit for further operation. Hence, a repair strategy intended to rejuvenate
the turbine assembly 10 and consequently, a method for operating and treating the
turbine assembly 10 is proposed.
[0056] Hence, in a first step of this method a section 12 of the turbine assembly 10 that
is prone to fatigue while operating the turbine assembly 10 will be determined. This
may be done by inspecting the turbine assembly 10 after several operating cycles or
be based on live experience or be based on calculations.
[0057] As stated above, typical locations of cyclic life limiting features are cooling holes
18 or bolt holes 20 at the turbine disc 58. Thus, the section 12 comprises a change
in its contour 16 and is a hole 18, 20 with its surface 22 or rim. Changes in contour
16 or holes 18, 20 are stress raisers and increase the stress in the component at
their location or around their location. Hence, a stress concentration factor is especially
high for the material around the hole 18, 20. This is specifically the case for an
acute corner 23 of a hole 18 with a breakout 25 that breaks out at the surface 22
of the hole 18 at an angle α.
[0058] FIG 2 to 5 show exemplarily four sections 12, 12a that are prone to fatigue while
operating the turbine assembly 10, 10a. FIG 2 shows a perspective view of a part of
a turbine assembly 10 embodied as a turbine disc 58 of the gas turbine engine 26 beforehand
of a refurbishment. FIG 3 shows the part of the turbine disc 58 after the refurbishment.
[0059] The turbine assembly 10 comprises a cooling hole 18 that feeds cooling medium 86
to a blade root slot 88 of the disc 58. The cooling air feed hole 18 is positioned
at a bottom 90 of a blade root slot 88. After operating the turbine assembly 10 the
entire hole 18 or a part of the hole 18 will show fatigue. The part of the hole 18
may be, for example, a region or section 12 embodied as an inner surface 92 of the
hole 18 or the acute corner 23 of the hole 18, 20 with the breakout 25 that breaks
out at the surface 22 of the hole 18 at the angle α. In FIG 4 a bolt hole 20 of a
disc 58 of a multi tie bolted rotor arrangement is shown. The region or section 12
around the bolt hole 20 will show fatigue after operating the turbine assembly 10.
In FIG 5 air transfer holes 18 of a disc 58 are shown. The region or section 12a around
the air transfer hole 18 will show fatigue after operating the turbine assembly 10a.
[0060] To decide on a suitable refurbishment strategy several considerations are important.
The stress acting on different positions of the turbine assembly 10, especially, at
different distances from the stress raiser or from the surface 22 or rim of the hole
18, 20, needs to be considered. Moreover, the time until the cyclic life of the turbine
assembly 10 or the section 12 or different regions thereof is reached is taken into
account.
[0061] Hence, it is determined which maximum principal stress acts upon the material depending
on a distance from the stress raiser, in this case the surface 22 or rim of the hole
18, 20 or from the acute corner 23. The stress at a point varies in different directions.
These stresses are often expressed in terms of three "principle stresses" which occur
at a specific angle which depends on the part in question and at right angles to each
other. The largest of these, the "maximum principle stress" is often used to determine
the life. Alternatively, a mixture of all three stresses as well as shear stresses
may be used (this is known as the von Mises stress).
[0062] FIG 6 shows a diagram depicting the stress magnitude versus the perpendicular distance
from stress concentration or in other words the distance from the surface 22 or rim
of the cooling hole 18 having a diameter of 5.6 millimetre (mm). The y-axis refers
to the maximum principle stress in megapascal (MPa) and on the x-axis the distance
from the surface 22 or rim is plotted in millimetre (mm). As can be seen in the graph
at the surface 22 or rim of a cooling hole 18 a stress magnitude of 980 MPa acts upon
the material and at a distance of 0.8 mm from the surface 22 or rim of the cooling
hole 18 a stress magnitude of 600 MPa acts upon the material during operation.
[0063] To determine the time until the cyclic life or fatigue life of the selected locations
is reached strain-life curves are examined. Therefore, the maximum principle stress
is converted in to a total strain range. LCF life is calculated using a computer simulation,
finite element model. It models the stress versus time for 1 cycle, and then the location
with the worst combination of stress range and mean stress is found. This is used
to find a strain range, and then, using lifing data acquired from material specimen
testing which provides a strain range versus life curve, the life can be calculated,
after the curve is adjusted for the mean stress required. The time (LCF life) is represented
by the amount (N) of cycles, wherein one cycle is the time from when the engine 26
is first started to when it is stopped. Usually the number of cycles the parts in
question can endure is independent of the engine running time (provided the parts
do not creep, which means deform over time due to high temperature and stress). Therefore,
parts have a cyclic life, the higher the better. After the cyclic life is used up,
cracking occurs in the part, and then failure.
[0064] The life of a part is usually determined by the highest stressed point in that part.
In general, the higher the stress range at a point (i.e. the difference between the
highest and lowest stress at that point during 1 cycle), the lower the number of cycles
it can endure before it begins to crack. Sometimes the deformation during a cycle
i.e. the "strain range" is calculated (which is related to the "stress range"), and
similarly this relates to the life of the part.
[0065] The life is not only affected by the stress range, but also the average stress at
the life limiting point. The higher the average stress the worse it is. Lifing is
based on material testing which is done for different stress ranges. However, the
conditions in the test may not mirror the exact mean stress seen in a part. Therefore,
a mathematical mean stress adjustment is carried out on the test data to give life
data corresponding to the desired mean stress. The R value relates to the test data,
and it is the ratio of the test minimum stress/maximum stress, so R=-1 means the stress
in the test oscillates from a compressive (which is assigned a negative value), to
tensile (which is assigned a positive value), of equal magnitude.
[0066] FIG 7 and 8 show diagrams depicting strain versus the cycles to crack initiation.
The y-axis refers to the total strain range in percent (%) and on the x-axis N, cycles
for crack initiation is plotted (the graph with the broken line depicts the strain-life
curve, R=-1; the graph with the chain dotted line represents the strain-life curve,
mean stress adjusted, the horizontal solid lines the total strain range and the vertical
broken lines the respective LCF life).
[0067] As can be deviated from the diagram of FIG 7 cracking will occur at the surface 22
or rim after 8086 cycles so that the turbine assembly 10 reaches its LCF live after
8086 cycles or has an initial life of 8086 cycles. In FIG 8 is shown that for a region
being positioned 0.8 mm away from the surface 22 or rim of the hole 18 damage will
occur after 105143 cycles so that this region would reach its LCF live after 105143
cycles or has an initial life of 105143 cycles. Hence, the life of a region positioned
away from the hole is greater than the life at the surface 22 or rim of the hole 18.
[0068] The two lives (8086 and 105143) correspond to the different locations to the hole
18. The 8086 cycles is at the inner surface 92 of the hole 18, this highlights that
this section 12 has the lowest life. However, the region 0.8 mm from the inner surface
92 of the hole 18 has a life of 105143 cycles, so this part of the turbine assembly
10 is barely damaged after 8086 cycles (but some damage does occur). Therefore, a
cumulative damage approach will be used to assess the life of the hole 18 after refurbishment.
[0069] With these two values the specific operating time of the section 12 of the turbine
assembly 10 can be defined in the next step of the method for treating and operating
the turbine assembly 10. The specific operating time is the time or specifically is
the number of cycles the section 12 will be operated beforehand of the refurbishment.
[0070] The specific operating time is determined by a use of the Palmgren-Miner rule. Assuming
equal running intervals before and after a refurbishment service, and using Miner's
rule we get a service interval life (specific operating time),
ni of:

[0071] Hence, the section 12 is operated up to the specific operating time of 7508 cycles
in the respective step of the method for treating and operating the turbine assembly
10.
[0072] Additionally, a finite element analysis (FEA) of the turbine assembly 10 or the section
12 after as well as before the refurbishment could also be done, to more accurately
determine the stresses.
[0073] After this operating the section 12 is refurbished or treated by material removal
and depending on the damaged section 12 the acute corner 23 and/or a material layer
between the surfaces 92 and 94 is removed. Due to this a sub-section 14 of the section
12 with a surface 94 is exposed (see broken line in FIG 3). This sub-section 14 is
operated further. The further operating time was also gained out of the Palmgren-Miner
rule and is also 7508 cycles. Operating the turbine assembly 10 and thus the section
12 and the sub-section 14 for 7508 cycles before and after refurbishment gives the
same level of damage as running for 8086 cycles without refurbishment.
[0074] The amount of material to be removed is determined by a consideration of at least
one of the factors consisting out of the group of: internal stress, contact stress,
temperature, material capability, material thickness, operating speed, flow velocity
of a flow medium and desired future fatigue life as well as by at least one of the
methods out of the group consisting of: Finite Element Analysis, hand calculations,
obtaining results from laboratory testing or from examination of engine run components
and applying statistics over the results.
[0075] FIG 9 and 10 show the results of Finite Element Analysis of discs out of different
materials and having cooling or bolt holes 18, 20 with different diameters, geometries
or orientations. In each diagram the y-axis refers to the life gained as percent (%)
of initial life if this much material is removed and on the x-axis the distance from
the hole as percent (%) of the hole radius is plotted (Legend for the graphs in FIG
9 specifying the material of the disc, hole diameter and orientation of the respective
cooling hole 18: graph with broken line and diamond symbol: IN718, 12.5 mm; graph
with chain dotted line and rectangle symbol: steel, 16 mm; graph with chain dotted
line and triangle symbol: steel, 7 mm; graph with light solid line and cycle symbol:
steel, 5.6 mm; graph with solid line and asterisk symbol: IN718, 5.6 mm, vertical;
graph with bold solid line and cycle symbol: IN718, 5.6 mm, angled (peak stress on
obtuse corner); Note: for the first five graphs the peak stress is on an acute corner;
Legend for the graphs in FIG 10 specifying the material of the disc, hole diameter
and load factor increase of the respective bolt hole 20: graph with solid line and
cycle symbol: IN718, 34 mm; graph with chain dotted line and triangle symbol: IN718,
42.8 mm; graph with solid line and asterisk symbol: IN718, 34 mm, 1.45 load; graph
with broken line and diamond symbol: IN718, 42.8 mm, 1.2 load).
[0076] As could be seen in FIG 9 a life gain of 60% and even up to 70% can be obtained by
removing 10% of the distance from the radius of a cooling hole 18. By removing 25%
up to 95% life gain can be achieved. In the case of the bolt holes 20, as shown in
FIG 10, the removal of 10% results in a life gain of about 40% or even up to 85%.
By removing 30% a life gain of about 70% can be achieved.
[0077] Hence, it was found, that by treating the section 12 by material removal by enlarging
a diameter d of the hole 18, 20 or by removing material from or the "whole" acute
corner 23 of the hole 18 with the breakout 25 that breaks out at the surface 22 of
the hole 18 at the angle α, and especially, by removing about 10% to 30% of the diameter
d from the inner surface 92 or at the acute corner 23 a significant effect on the
life was gained without compromising the overall integrity of the section 12 and the
turbine assembly 10. These figures were determined as the range that would give optimal
life recovered for as little services as possible (or within acceptable/scheduled
levels of serving), with acceptable levels of material removal. The amount of material
removed would vary from part to part, and from service strategy to service strategy,
and will be chosen from the person skilled in the art according to its knowledge in
the field.
[0078] As can be seen by comparing the results for the cooling holes 18 (see FIG 9) and
the bolt holes 20 (FIG 10) the effects of the refurbishment is more effective for
the cooling holes 18, especially, angled holes with the peak stress at the acute corner
23 (see all graphs in FIG 9 except the graph with bold solid line and cycle symbol).
Thus, more material is generally needed to be removed for the bolt holes 20 to gain
the same amount of life as for a cooling hole 18. This is because the geometry at
a bolt hole 20 is usually slightly different from a cooling hole 18 resulting in a
lower stress "raiser" (a stress raiser - also known as a stress concentration factor
- is where the geometry causes the stress to increase as it nears the feature - hole
- in question). This is usually the case when the cooling hole 18 has the breakout
25 that breaks out at the angle α; and the peak stress usually occurs at the acute
corner 23 of the breakout 25. This refurbishment technique works better when the stress
is especially high close to the hole 18, which is more the case with the cooling holes
18 due to the higher stress concentration factor. Hence, the material removal of 10%
to 20% may be enough for a cooling hole 18 and should be up to 30% for a bolt hole
20.
[0079] The above stated findings will now be summarised exemplarily for a steel disc with
a hole having a diameter of 5.6 mm (see graph with light solid line and cycle symbol
in FIG 9). Assuming equal number of damage cycles for the 980 MPa and 600 MPa cycles
and using Miner's rule the life is 7508 joint cycles (where one joint cycle includes
one 980 MPa cycle and one 600 MPa cycle). The life of a cooling hole 18 is 7508 cycles,
then if it were refurbished by increasing the diameter of the hole from the original
diameter of 5.6 mm to the rework diameter of 7.2 mm (28.6% increase by enlarging the
diameter about 1.6 mm or by removing 0.8 mm from the acute corner 23 of the hole 18
at the region of the breakout 25, respectively) and the part of the disc that was
at 600 MPa, then was subjected to a stress of 980 MPa, it could do another 7508 cycles.
Therefore, in theory the life of the 5.6 mm cooling hole 18 would potentially increase
from 8086 engine cycles to 15016 engine cycles.
[0080] The method may comprise a single refurbishment step or more than one. Hence, the
method comprises the further step of: repeating the steps of defining the specific
operating time, operating the section 12, treating the section 12 and operating the
sub-section 14 several times. The refurbishment process could be repeated and more
life gained for the section 12 or the hole 18, 20, although the life gained after
each refurbishment would gradually drop.
[0081] As the stress versus distance from the hole curves basically offset after a refurbishment
step, i.e. it is like starting the stress versus distance from the hole curve again,
the amount of material removed in the further refurbishment step(s) and how many times
it was done will be selected so that a balance was reached between the gaining of
enough life to make the process worthwhile and permitting space and the overall integrity
of the part. Although analysis of the turbine assembly 10 or the section 12 before
or after the refurbishment can be done to accurately determine the stresses.
[0082] The repetition can be done until an abort criterion is reached and for example until
the defined operating time in the step defining the operating time is less than a
difference between a maximum operating time of the section 12 and the already performed
operating time(s).
[0083] In FIG 11 a flow chart of a possible component refurbishment strategy is shown. After
operating the component or turbine assembly 10 for a pre-defined time, like routinely
5000 cycles or routinely 1 to 4 years, operator or service records are interrogated
to determine if the assembly 10 has sufficient cyclic life to make it to the next
service. If yes, the assembly 10 is inspected out of security reasons and if no objections
are raised it is returned to service to operate a further time, like routinely 5000
cycles or routinely 2 to 4 years.
[0084] If the interrogation results in a negative result (decision no) e.g. when the selected
operating time has reached the value of the 7508 cycles disclosed above, it is checked
if the assembly 10 or its section 12 has reached its rework limit. Exceeding the rework
limit may compromise the overall integrity of the section 12 or assembly 10. If the
rework limit is reached (decision yes) the assembly 10 is scrapped. If the answer
is no the section 12 would be refurbished or reworked, respectively. Thereafter, the
assembly 10 is inspected out of security reasons and if no objections are raised it
is returned to service to operate a further time e.g. the further 7508 cycles disclosed
above.
[0085] In reference to FIG 5 an alternative method for operating and treating a turbine
assembly will be described. Components, features and functions that remain identical
are in principle substantially denoted by the same reference characters. To distinguish
between the embodiments, however, the letter "a" has been added to the different reference
characters of the embodiment in FIG 1 to 4. The following description is confined
substantially to the differences from the embodiment in FIG 1 to 4, wherein with regard
to components, features and functions that remain identical reference may be made
to the description of the embodiment in FIG 1 to 4.
[0086] FIG 5 shows a turbine assembly 10a being refurbished with an alternative method for
operating and treating the turbine assembly 10a. The method from FIG 5 differs in
regard to the method used in the embodiment according to FIG 1 to 4 in that after
determining a section 12a of the turbine assembly 10a that is prone to fatigue while
operating the turbine assembly 10a the section 12a or its surface 92 is overlain with
an additional depth of material to form an additional layer 24 of material providing
an outer surface 96 (see broken lines in FIG 5).
[0087] After a defining of a specific operating time of the layered section 12a and subsequently,
an operating of the turbine assembly 10a up to the specific operating time the layered
section 12a or layer 24 is treated by material removal from the additional layer 24
of material or its surface 96. Afterwards the section 12a is operated further.
[0088] Preferably, a whole amount of the added material of the additional layer 24 of the
layered section 12 is removed in the treatment step. Hence, the original surface 22
or rim or the surface 92 of the hole 18 is exposed. It is possible to either perform
the material removal from the additional layer 24 in one step or in several steps.
Hence, the method comprises the step of repeating the steps of defining the specific
operating time, overlaying the section 12a, treating the layered section 12a and operating
the section 12a several times. This may be done till an abort criterion or a work
limit is reached. This may be when the defined specific operating time in the step
defining the specific operating time is less than a difference between a maximum operating
time of the layered section 12a and the already performed operating time(s).
[0089] It should be noted that the term "comprising" does not exclude other elements or
steps and "a" or "an" does not exclude a plurality. Also elements described in association
with different embodiments may be combined. It should also be noted that reference
signs in the claims should not be construed as limiting the scope of the claims.
[0090] Although the invention is illustrated and described in detail by the preferred embodiments,
the invention is not limited by the examples disclosed, and other variations can be
derived therefrom by a person skilled in the art without departing from the scope
of the invention.
1. Method for operating and treating a turbine assembly (10), wherein the method comprises
at least the steps of:
- determining at least a section (12) of the turbine assembly (10) that is prone to
fatigue while operating the turbine assembly (10),
- defining a specific operating time of the section (12) of the turbine assembly (10),
- operating the section (12) up to the specific operating time,
- treating the section (12) that is prone to fatigue after operating the section (12)
up to the specific operating time by material removal and thus expose a sub-section
(14) of the section (12) and
- operating the sub-section (14).
2. Method according to claim 1,
wherein the method comprises the further step of:
- determining the amount of material to be removed by a consideration of at least
one of the factors consisting out of the group of: internal stress, contact stress,
surface stress, temperature, material capability, material thickness, operating speed,
flow velocity of a flow medium and desired future fatigue life.
3. Method according to claim 1 or 2,
wherein the method comprises the further step of:
- determining the amount of material to be removed by at least one of the methods
out of the group consisting of:
Finite Element Analysis, hand calculations, obtaining results from laboratory testing
or from examination of engine run components and applying statistics over the results.
4. Method according to any one of the preceding claims, wherein the method comprises
the further step of:
- determining the specific operating time by a use of the Palmgren-Miner rule.
5. Method according to any one of the preceding claims, wherein the section (12) is a
section (12) comprising a change in its contour (16).
6. Method according to any one of the preceding claims, wherein the section (12) is a
hole (18, 20) with its surface (22).
7. Method according to claim 6,
wherein the hole (18, 20) is a cooling hole (18) or a bolt hole (20).
8. Method according to claim 6,
wherein the method comprises the further step(s) of:
- treating the section (12) by material removal by enlarging a diameter (d) of the
hole (18, 20), especially, by removing about 10% to 30% of the diameter (d) of the
hole (18, 20) and/or
- treating the section (12) by material removal from an acute corner (23) of a hole
(18, 20) with a breakout (25) that breaks out at a surface (22) of the hole (18, 20)
at an angle (α), especially, by removing about 10% to 30% of a diameter (d) of the
hole (18, 20) with the breakout (25) at the acute corner (23).
9. Method according to any one of the preceding claims, wherein the method comprises
the further step of:
- repeating the steps of defining the specific operating time, operating the section
(12), treating the section (12) and operating the sub-section (14) several times.
10. Method according to any one of the preceding claims, wherein the method comprises
the further step of:
- repeating the steps of defining the specific operating time, operating the section
(12), treating the section (12) and operating the sub-section (14) several times until
the defined operating time in the step defining the operating time is less than a
difference between a maximum operating time of the section (12) and the already performed
operating time(s).
11. Method for operating and treating a turbine assembly (10a),
wherein the method comprises at least the steps of:
- determining at least a section (12a) of the turbine assembly (10a) that is prone
to fatigue while operating the turbine assembly (10a),
- overlaying the section (12a) with an additional depth of material to form an additional
layer (24) of material,
- defining a specific operating time of the layered section (12a),
- operating the turbine assembly (10a) up to the specific operating time,
- treating the layered section (12a) by material removal from the additional layer
(24) of material and
- operating the section (12a).
12. Method according to claim 11,
wherein the method comprises the further step of:
- treating the layered section (12a) by removing a whole amount of the added material
of the additional layer (24).
13. Method according to claim 11 or 12,
wherein the material removal from the additional layer (24) is performed in one step
or in several steps.
14. Method assembly to any one of claims 11 to 13,
wherein the method comprises the further step of:
- repeating the steps of defining the specific operating time, overlaying the section
(12a), treating the layered section (12a) and operating the section (12a) several
times and especially, until the defined specific operating time in the step defining
the specific operating time is less than a difference between a maximum operating
time of the layered section (12a) and the already performed operating time(s).
15. Method according to any one of claims 11 to 14,
wherein the method comprises the further steps of:
- determining the amount of material to be removed by a consideration of at least
one of the factors consisting out of the group of: internal stress, contact stress,
surface stress, temperature, material capability, material thickness, operating speed,
flow velocity of a flow medium and desired future fatigue life and/or - determining
the amount of material to be removed by at least one of the methods out of the group
consisting of: Finite Element Analysis, hand calculations, obtaining results from
laboratory testing or from examination of engine run components and applying statistics
over the results and/or - determining the specific operating time by a use of the
Palmgren-Miner rule and/or wherein the section (12a) is a section (12a) comprising
a change in its contour (16) and/or wherein the section (12a) is a hole (18, 20) with
its surface (22) and/or wherein the hole (18, 20) is a cooling hole (18) or a bolt
hole (20) and/or wherein the method comprises the further step(s) of: treating the
section (12a) by material removal by enlarging a diameter (d) of the hole (18, 20),
especially, by removing about 10% to 30% of the diameter (d) of the hole (18, 20)
and/or wherein the method comprises the further step of:
treating the section (12) by material removal from an acute corner (23) of a hole
(18, 20) with a breakout (25) that breaks out at a surface (22) of the hole (18, 20)
at an angle (α), especially, by removing about 10% to 30% of a diameter (d) of the
hole (18, 20) with the breakout (25) at the acute corner (23).