U.S. patent application number 13/663398 was filed with the patent office on 2013-03-21 for turbine blades and methods of forming modified turbine blades and turbine rotors.
This patent application is currently assigned to HONEYWELL INTERNATIONAL INC.. The applicant listed for this patent is Honeywell International Inc.. Invention is credited to Steve Halfmann, Bob Mitlin, Mark C. Morris, Ardeshir Riahi.
Application Number | 20130071254 13/663398 |
Document ID | / |
Family ID | 41613630 |
Filed Date | 2013-03-21 |
United States Patent
Application |
20130071254 |
Kind Code |
A1 |
Mitlin; Bob ; et
al. |
March 21, 2013 |
TURBINE BLADES AND METHODS OF FORMING MODIFIED TURBINE BLADES AND
TURBINE ROTORS
Abstract
Turbine blades and methods of forming modified turbine blades
and turbine rotors for use in an engine are provided. In an
embodiment, by way of example only, a turbine blade includes a
platform and an airfoil. The platform includes a surface configured
to define a portion of a flowpath, and the surface includes an
initial contour configured to plastically deform into an intended
final contour after an initial exposure of the blade to an
operation of the engine. The airfoil extends from the platform.
Inventors: |
Mitlin; Bob; (Scottsdale,
AZ) ; Morris; Mark C.; (Phoenix, AZ) ;
Halfmann; Steve; (Chandler, AZ) ; Riahi;
Ardeshir; (Scottsdale, AZ) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Honeywell International Inc.; |
Morristown |
NJ |
US |
|
|
Assignee: |
HONEYWELL INTERNATIONAL
INC.
Morristown
NJ
|
Family ID: |
41613630 |
Appl. No.: |
13/663398 |
Filed: |
October 29, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
12273108 |
Nov 18, 2008 |
8297935 |
|
|
13663398 |
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Current U.S.
Class: |
416/90R ;
416/223A; 416/243; 416/244A |
Current CPC
Class: |
F01D 5/143 20130101;
Y10T 29/49336 20150115; Y02T 50/60 20130101; F01D 5/141 20130101;
F01D 5/18 20130101; F01D 5/30 20130101; Y10T 29/49764 20150115;
Y02T 50/673 20130101 |
Class at
Publication: |
416/90.R ;
416/244.A; 416/243; 416/223.A |
International
Class: |
F01D 5/30 20060101
F01D005/30; F01D 5/14 20060101 F01D005/14; F01D 5/18 20060101
F01D005/18 |
Goverment Interests
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
[0002] This invention was made with Government support under
W911W6-08-2-0001 awarded by the U.S. Army. The Government has
certain rights in this invention.
Claims
1. A turbine blade for use in an engine, the turbine blade
comprising: a platform including a surface configured to define a
portion of a flowpath, the surface including an initial contour
configured to plastically deform into an intended final contour
after an initial exposure of the blade to an operation of the
engine; and an airfoil extending from the platform.
2. The turbine blade of claim 1, wherein the operation of the
engine includes a predetermined maximum speed and a predetermined
maximum temperature.
3. The turbine blade of claim 1, wherein: the airfoil has an
airfoil shape defined by a convex suction side wall, a concave
pressure side wall, a leading edge, a trailing edge, a root and a
tip, the convex suction side wall, the concave pressure side wall,
and the tip each including an interior surface that defines an
interior with the root, the trailing edge including a plurality of
slots formed thereon.
4. The turbine blade of claim 1, wherein the initial contour is a
depression.
5. The turbine blade of claim 4, wherein the depression is located
adjacent to the concave pressure side wall.
6. The turbine blade of claim 4, wherein the depression is located
adjacent to the convex suction side wall.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present application is a divisional application of U.S.
Pat. No. 8,297,935 issued Oct. 30, 2012 (application Ser. No.
12/273,108, filed Nov. 18, 2008), the contents of which are herein
incorporated by reference in their entirety.
TECHNICAL FIELD
[0003] The inventive subject matter generally relates to turbines,
and more particularly relates to turbine blades and methods of
forming turbine blades and turbine rotors for use in turbines.
BACKGROUND
[0004] Gas turbine engines, such as turbofan gas turbine engines,
may be used to power various types of vehicles and systems, such as
aircraft. Typically, these engines include turbines that rotate at
a high speed when blades (or airfoils) extending therefrom are
impinged by high-energy compressed air. Consequently, the blades
are subjected to high heat and stress loadings which, over time,
may reduce their structural integrity.
[0005] To maintain blade structural integrity, a blade cooling
scheme is typically incorporated into the turbines. The blade
cooling scheme is included to maintain the blade temperatures
within acceptable limits. In some cases, the blade cooling scheme
directs cooling air through an internal cooling circuit formed in
the blade. The internal cooling circuit consists of a series of
connected, serpentine cooling passages, which incorporate raised or
depressed structures therein. The serpentine cooling passages
increase the cooling effectiveness by extending the length of the
air flow path. In this regard, the blade may have multiple internal
walls that form intricate passages through which the cooling air
flows to feed the serpentine cooling passages. The blade cooling
scheme may also include platform cooling, in some cases. For
example, openings may be formed through a turbine disk from which
the blades radiate, and the openings may direct cool air from a
cool air source onto a platform of the blade.
[0006] As the desire for increased engine efficiency continues to
rise, engine components are increasingly being subjected to higher
and higher operating temperatures. However, current engine
components, such as blades and blade platforms, may not be
adequately designed to withstand such temperatures over time. In
particular, current blade and blade platforms may be prone to
thermo-mechanical fatigue, plastic deformation, and/or platform
shingling, when exposed to the newer engine design operating
temperatures.
[0007] Accordingly, it is desirable to have an improved blade that
may operate with improved durability when exposed to high engine
operating temperatures. In addition, it is desirable to have an
improved blade that is relatively simple and inexpensive to
implement and that may be retrofitted into existing engines.
Furthermore, other desirable features and characteristics of the
inventive subject matter will become apparent from the subsequent
detailed description of the inventive subject matter and the
appended claims, taken in conjunction with the accompanying
drawings and this background of the inventive subject matter.
BRIEF SUMMARY
[0008] Turbine blades and methods of forming modified turbine
blades and turbine rotors for use in an engine are provided.
[0009] In an embodiment, by way of example only, a turbine blade
includes a platform and an airfoil. The platform includes a surface
configured to define a portion of a flowpath, and the surface
includes an initial contour configured to plastically deform into
an intended final contour after an initial exposure of the blade to
an operation of the engine. The airfoil extends from the
platform.
[0010] In another embodiment, by way of example only, a method of
forming a modified turbine blade includes providing an unmodified
blade configuration including a platform and an airfoil extending
from the platform, the platform including a surface defining a
portion of a flowpath of the engine, and subjecting the unmodified
blade configuration to one or more maximum engine operating
conditions, wherein the one or more maximum engine operating
conditions include rotating a turbine rotor including the
unmodified blade configuration thereon at a predetermined maximum
speed and exposing the turbine rotor to a predetermined maximum
temperature. The method also may include performing a plastic
analysis of the unmodified blade configuration, while the
unmodified blade configuration is subjected to the one or more
maximum engine operating conditions, calculating a residual plastic
deformation of the platform surface, based on the plastic analysis,
determining whether the residual plastic deformation of the
platform surface is within a predetermined acceptable range,
determining an inverse of the residual plastic deformation of the
platform surface, if a determination is made that the residual
plastic deformation of the platform surface is within the
predetermined acceptable range, and incorporating the inverse of
the residual plastic deformation of the platform surface into the
unmodified blade configuration to form the modified blade
configuration.
[0011] In still another embodiment, by way of example only, a
method of forming a turbine rotor includes providing an unmodified
blade configuration including a platform and an airfoil extending
from the platform, the platform including a surface defining a
portion of a flowpath of the engine. The unmodified blade
configuration is subjected to one or more maximum engine operating
conditions, wherein the one or more maximum engine operating
conditions include rotating a turbine rotor including the
unmodified blade configuration thereon at a predetermined maximum
speed and exposing the turbine rotor to a predetermined maximum
temperature. A plastic analysis of the unmodified blade
configuration is performed, while the unmodified blade
configuration is subjected to the one or more maximum engine
operating conditions. A residual plastic deformation of the
platform surface is calculated, based on the plastic analysis. A
determination is made as to whether the residual plastic
deformation of the platform surface is within a predetermined
acceptable range. An inverse of the residual plastic deformation of
the platform surface is determined, if a determination is made that
the residual plastic deformation of the platform surface is within
the predetermined acceptable range. The inverse of the residual
plastic deformation of the platform surface is incorporated into
the unmodified blade configuration to form the modified blade
configuration. A turbine blade is cast from a mold including the
modified blade configuration. The turbine blade is attached to a
hub to form the turbine rotor. The turbine rotor is then rotated at
the maximum engine operating conditions to impart a final contour
to the turbine blade.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] The inventive subject matter will hereinafter be described
in conjunction with the following drawing figures, wherein like
numerals denote like elements, and
[0013] FIG. 1 is a cross-sectional side view of a turbine section
of an engine, according to an embodiment;
[0014] FIG. 2 is a perspective view of a turbine blade, according
to an embodiment;
[0015] FIG. 3 is a flow diagram of a method of forming a turbine
blade, according to an embodiment; and
[0016] FIG. 4 is a flow diagram of a method of manufacturing a
turbine rotor, according to an embodiment.
DETAILED DESCRIPTION
[0017] The following detailed description is merely exemplary in
nature and is not intended to limit the inventive subject matter or
the application and uses of the inventive subject matter.
Furthermore, there is no intention to be bound by any theory
presented in the preceding background or the following detailed
description.
[0018] FIG. 1 is a cross-sectional side view of a portion of a
turbine section 100 of an engine, according to an embodiment. The
turbine section 100 receives high temperature (e.g, temperature
greater than 1100.degree. C.) gases from an upstream engine
combustor (not shown) to produce energy for the engine and/or
components coupled to the engine. In an embodiment, the turbine
section 100 includes a turbine nozzle 104 that has a plurality of
static vanes 106 mounted circumferentially around a ring 108. The
static vanes 106 direct the gases from the combustor to a turbine
rotor 110. According to an embodiment, the turbine rotor 110
includes a plurality of blades 112 (only one of which is shown)
that are attached to a hub 114 and retained in axial position by a
retention plate 116. When the blades 112 are impinged upon by the
gases, the gases cause the turbine rotor 110 to spin. According to
an embodiment, an outer circumferential wall 118 surrounds the
static vanes 106 and the plurality of blades 112 to define a
flowpath 122. The circumferential wall 118 also defines a portion
of a compressor plenum 120 that is disposed radially outwardly
relative to the flowpath 122. The compressor plenum 120 receives
bleed air from a compressor section (not shown), which may be
directed through one or more openings in the outer circumferential
wall 118 towards the plurality of blades 112 to cool the blades
112.
[0019] FIG. 2 is a perspective view of a blade 200, according to an
embodiment. The blade 200 is configured to transform from a first,
initial shape to a second, permanent shape, where the first,
initial shape includes a contour imparted to the blade 200 during
manufacture and the second, permanent shape includes a contour
imparted to the blade 200 as a result of being subjected to maximum
engine operating conditions. In an embodiment, the maximum engine
operating conditions may include maximum conditions at which an
engine may be designed to operate. For example, a maximum condition
may include a predetermined maximum speed at which a turbine may be
designed to rotate within the engine. In an embodiment, the
predetermined maximum speed may be a rotational speed in a range of
from about 10,000 rotations per minute (rpm) to about 50,000 rpm.
In other embodiments, the predetermined maximum speed may be
greater than or less than the aforementioned range. In another
example, a maximum condition may include a predetermined maximum
temperature to which the turbine is designed to be exposed. For
example, the predetermined maximum temperature may be in a range of
from about 1100.degree. C. to about 1700.degree. C. In other
embodiments, the predetermined maximum temperature may be greater
than or less than the aforementioned range. In still other
embodiments, the maximum condition may include other parameters in
accordance with which an engine is designed to operate.
[0020] In any case, the blade 200 may comprise a single crystal
blade including a nickel-based superalloy, in an embodiment.
Suitable nickel-based superalloys include, but are not limited to,
CMSX3, or SC180. In other embodiments, the blade 200 may comprise a
polycrystalline structure. Suitable materials include, but are not
limited to, MAR-M-247EA or MAR-M-247DS.
[0021] The blade 200 includes a blade attachment section 202, an
airfoil 204, and a platform 206. The blade attachment section 202
provides an area in which a shape 208 is machined. In an
embodiment, the shape 208 corresponds with a shape formed in a
respective blade attachment slot (not shown) of the turbine hub
(e.g., hub 114 in FIG. 1). For example, in some embodiments, the
shape 208 may be a firtree shape. In other embodiments, the shape
208 may be a beveled shape. However, in other embodiments, any one
of numerous other shapes suitable for attaching the blade 200 to
the turbine may be alternatively machined therein.
[0022] The airfoil 204 has two outer walls 210, 212, each having
outer surfaces that define an airfoil shape. The airfoil shape
includes a leading edge 214, a trailing edge 216, a pressure side
218 along the first outer wall 210, a suction side 220 along the
second outer wall 212, a blade tip 222, a pressure side discharge
trailing edge slot 224, and an airfoil platform fillet 226. Though
not shown, the blade 200 may have an internal cooling circuit
formed therein, which may extend from an opening in the platform
206 through the blade 200 and may include various passages that
eventually communicate with the trailing edge slot 224 or other
openings (not shown) that may be formed in the blade 200.
[0023] The platform 206 has a surface 228 that is adapted to define
a portion of a flowpath that extends through the engine (e.g.,
flowpath 122 of FIG. 1). In an embodiment, the surface 228 may be
adapted to include the first, initial shape imparted thereto during
manufacture of the blade 200, and to transform to the second,
permanent shape after being subjected to the maximum engine
operating conditions. According to an embodiment, the first initial
shape of the surface 228 may include a specially-designed
depression 230. The depression 230 may be located adjacent to the
concave, pressure side wall 210, in an embodiment. In another
embodiment, the depression 230 may be located adjacent to the
convex, suction side wall 212. Although the depression 230 is shown
in FIG. 2 as having a particular shape, the depression 230 may have
a different shape, in other embodiments. For example, though the
depression 230 is defined by a generally smooth surface in FIG. 2,
the surface may include protrusions, bumps, or other features in
other embodiments. In any case, after the blade 200 is exposed to
the maximum engine operating conditions, the depression 230
transforms into an intended shape 232 (shown in phantom), which
provides the platform 206 with a desired final contour for defining
a portion of the flowpath (e.g., flowpath 122 in FIG. 1), when the
blade 200 is attached to a turbine that is disposed in a turbine
engine.
[0024] To manufacture the blade 200 described above, method 300 may
be employed. FIG. 3 is a flow diagram of method 300 of forming a
turbine blade, according to an embodiment. As used herein, a blade
(such as blade 200 in FIG. 2) including an intentionally
incorporated depression (such as specially-designed depression 230
in FIG. 2) may be referred to hereinafter as a "modified blade". A
blade that does not have and has never included an intentionally
incorporated depression may be referred to hereinafter as an
"unmodified blade".
[0025] Method 300 includes providing a blade configuration of an
unmodified blade, step 302. In an embodiment, the unmodified blade
configuration includes a platform and an airfoil extending from the
platform, where the platform includes a surface defining a portion
of a flowpath of the engine. According to another embodiment, the
unmodified blade configuration includes other features of a blade,
such as a blade attachment section, suction side walls, pressure
side walls, and the like, which may have particular shapes. In any
case, the unmodified blade configuration may be provided as a
computer-generated (e.g., virtual) model or as an actual blade.
[0026] In accordance with an embodiment, the unmodified blade
configuration is subjected to one or more maximum engine operating
conditions, step 304. The maximum engine operating conditions may
include operating conditions described above in conjunction with
the blade 200. The unmodified blade configuration, which may be
provided as a virtual or an actual blade, may be virtually
subjected to the maximum engine operating conditions, in an
embodiment. For example, the maximum engine operating conditions
may be provided as computer-generated data which is used during the
simulated operation of a computer-generated model of the unmodified
turbine blade configuration. In accordance with another embodiment
in which the unmodified blade configuration is implemented as an
actual blade, the actual blade is disposed on a turbine rotor, and
an engine within which the turbine rotor is disposed is powered on
to simulate the maximum engine operating conditions.
[0027] Next, a plastic analysis of the unmodified blade
configuration is performed, while the unmodified blade
configuration is subjected to the maximum engine operating
conditions, step 306. According to an embodiment, the plastic
analysis includes determining a flow structure of the flowpath
defined by the platform surface (e.g., surface 228 in FIG. 2) of
the unmodified blade configuration, while the unmodified blade
configuration is subjected to the maximum engine operating
conditions. For example, the flow structure of the flowpath may be
determined by performing a computational fluid dynamics analysis,
which may include performing calculations to identify algorithms
that simulate an interaction of a gas flowing along the flowpath
relative to platform surface. The analysis may be based purely on
computer-generated data, or may be based at least partially on data
that is collected from an actual blade operated in a controlled
physical environment.
[0028] In another embodiment, the plastic analysis may additionally
or alternatively include collecting temperature data across an area
of the platform surface, while the unmodified blade configuration
is subjected to the maximum engine operating conditions. In one
example, a thermal analysis may be performed on the platform
surface to determine platform temperature data. According to
another embodiment, the temperature data may be obtained across an
entirety of the platform surface. In other embodiments, the
temperature data may be obtained across a portion of the platform
surface. The temperature data may be physically collected by
sensors that may be embedded on an actual blade, in an embodiment.
In another embodiment in which simulated operation of a
computer-generated turbine blade configuration is performed, the
temperature data may be computer-generated, using data from the
computational fluid dynamics analysis.
[0029] Residual plastic deformation of the platform is calculated
from the plastic analysis, step 308. For example, analyses
identifying stresses across the area of the platform surface may be
performed in order to locate "critical areas" of residual plastic
deformation. The "critical areas" may be locations on the platform
at which the residual plastic deformation may fall outside of a
predetermined acceptable range. The predetermined acceptable range
may be a range that provides a maximum desired residual plastic
deformation, when the unmodified blade configuration is subjected
to the maximum engine operating conditions. It will be appreciated
that the predetermined acceptable range of residual plastic
deformation may depend on various factors, such as a particular
desired design of a resultant blade, a particular design of engine
components surrounding the resultant blade, engine operating
parameters, and the like. A residual plastic deformation may be a
permanent deformation in a material resulting from an application
of stress to the material. In an embodiment, calculations are
performed by identifying stresses and residual plastic deformation
across the area of the platform surface to obtain identified
critical areas. In accordance with an embodiment, calculations may
be performed by identifying a maximum stress that a particular
location of the platform surface can withstand without experiencing
plastic deformation at the area. Calculations of the maximum stress
may be obtained over an entirety of the platform surface, in an
embodiment, or over a portion of the platform surface in another
embodiment. The maximum stress calculations are used in conjunction
with the temperature data collected and with data related to the
strength of the particular material from which the blade may be
comprised to determine a contour of the residual plastic
deformation.
[0030] A determination is made as to whether the residual plastic
deformation on the platform is within a predetermined acceptable
range, step 310. In an embodiment, if a determination is made that
the residual plastic deformation is not within the predetermined
acceptable range, material may be added onto the blade to place the
residual plastic deformation in the critical areas of the blade
within the predetermined acceptable range, step 312. In one
embodiment, the material may be substantially identical in
formulation to the material of which the blade comprises. In
another embodiment, the material may be different material. In
still another embodiment, the material may be added to the blade in
the form of a tapering of the platform surface or as an additional
flange or rib that protrudes from the platform surface. In
accordance with other embodiments, the material may be formed in
another shape and added to the blade. When an actual blade is being
operated in a controlled physical environment, the material may be
added to the blade by a suitable conventional brazing or welding
process. When simulated operation of a computer-generated turbine
blade is performed, the material may be added by increasing the
dimensions of the computer-generated turbine blade.
[0031] In accordance with an embodiment, to ensure that an
appropriate amount of material has been added to the blade, steps
304, 306, 308, 310, and 312 may be repeated. If after step 310, a
determination is made that the residual plastic deformation is
within the predetermined acceptable range, an inverse of the
residual plastic deformation is then determined, step 314. For
example, the residual plastic deformation may be a protrusion and
the inverse of the residual plastic deformation may be a depression
that is sunken into an opposite side of the platform, where the
depression may have a shape that is a mirror-image of the
protrusion. To define the residual plastic deformation in
mathematical terms, the contour of the residual plastic deformation
(e.g., the protrusion) may be mapped as a plurality of data points,
and an inverse of each data point is obtained to define the contour
of the depression. In an embodiment, the inverse of the residual
plastic deformation is then incorporated into the unmodified blade
configuration to form the modified blade configuration, step
316.
[0032] The modified blade configuration may be used to manufacture
a turbine rotor, in an embodiment. FIG. 4 is a flow diagram of a
method 400 of manufacturing a turbine rotor, according to an
embodiment. In an embodiment, blades are cast using molds having
the modified blade configuration, step 402. The blades are then
attached to a hub to form a turbine rotor, step 404. In an
embodiment, the blades and hub are cast using a conventional
casting process. In an example, the blades and hub are cast
together into an equiaxed nickel-based superalloy component. In
another example, the blades may be cast from a nickel base
superalloy, and the hub may be manufactured from a powder metal
nickel superalloy. In any case, the hub and the blades may include
substantially identical materials, in an embodiment. In other
embodiments, the hub and the blades may be made of different
materials. The number of blades included in the turbine rotor
depends on a size of the hub and a number of blade attachment slots
that may be formed in the hub. In one example, the hub may include
thirty-eight blade attachment slots, and thirty-eight blades are
included. In other embodiments, more or fewer slots and blades may
be included. Still in other embodiments, the cast blades or blade
ring may be metallurgically bonded to the hub to form the turbine
rotor.
[0033] After the turbine rotor is manufactured, the turbine rotor
is rotated and subjected to the maximum engine operating
conditions, step 406. In an embodiment, the turbine rotor is
incorporated into a turbine engine, and the turbine engine is
operated at the maximum engine operation conditions. In another
embodiment, the turbine rotor is disposed in an enclosure, maximum
engine operation conditions including the predetermined maximum
temperature, are simulated in the enclosure, and the turbine rotor
is rotated at the predetermined maximum speed. In any case, as a
result of the rotation and exposure to the maximum engine operating
conditions, depressions in the platform surfaces of the blades may
deform into the intended final contour so that the intended final
contour is permanently imparted to the platform surface.
[0034] By forming the blades in the manner described above, the
blades may be capable of operating at higher temperatures and at
higher rotational speeds than conventional blades. In particular,
the improved blades may operate with minimal incidence of undesired
deformation, because including the specially-designed depression in
the blade and subsequently exposing the blade to maximum engine
operating conditions permanently plastically deforms the blade to
the intended final contour. This permanent plastic deformation may
impart favorable residual stresses in areas adjacent to the plastic
deformation. Thus, when the blade is subjected to the maximum
engine operating conditions again at a subsequent time, further
plastic deformation may be minimized The result is a blade which
has improved durability and maintains performance in a high
temperature and high speed environment over conventional
blades.
[0035] While at least one exemplary embodiment has been presented
in the foregoing detailed description of the inventive subject
matter, it should be appreciated that a vast number of variations
exist. It should also be appreciated that the exemplary embodiment
or exemplary embodiments are only examples, and are not intended to
limit the scope, applicability, or configuration of the inventive
subject matter in any way. Rather, the foregoing detailed
description will provide those skilled in the art with a convenient
road map for implementing an exemplary embodiment of the inventive
subject matter. It being understood that various changes may be
made in the function and arrangement of elements described in an
exemplary embodiment without departing from the scope of the
inventive subject matter as set forth in the appended claims.
* * * * *