U.S. patent application number 15/264098 was filed with the patent office on 2018-03-15 for additively deposited gas turbine engine cooling component.
The applicant listed for this patent is Rolls-Royce Corporation. Invention is credited to Bruce Varney.
Application Number | 20180073390 15/264098 |
Document ID | / |
Family ID | 59626505 |
Filed Date | 2018-03-15 |
United States Patent
Application |
20180073390 |
Kind Code |
A1 |
Varney; Bruce |
March 15, 2018 |
ADDITIVELY DEPOSITED GAS TURBINE ENGINE COOLING COMPONENT
Abstract
An example gas turbine engine component includes a component
configured to separate a cooling air plenum from a heated gas
environment. The component includes a substrate defining a surface,
and a unitary structure. The unitary structure includes a cooling
region and a cover layer. The cover layer defines a hot wall
surface configured to face the heated gas environment. The cooling
region is disposed between the cover surface and the substrate and
includes a plurality of support structures extending between the
cover layer and the surface of the substrate. At least some of the
support structures define a respective bond surface bonded to the
substrate at the surface of the substrate. An example technique for
fabricating the gas turbine engine component includes additively
depositing the unitary structure on the surface of the
substrate.
Inventors: |
Varney; Bruce; (Greenwood,
IN) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Corporation |
Indianapolis |
IN |
US |
|
|
Family ID: |
59626505 |
Appl. No.: |
15/264098 |
Filed: |
September 13, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2260/202 20130101;
F01D 25/12 20130101; F23R 3/002 20130101; F23R 3/005 20130101; B23P
15/04 20130101; F05D 2220/32 20130101; F05D 2260/203 20130101; F05D
2230/31 20130101; F01D 5/186 20130101; F05D 2240/30 20130101; F05D
2230/20 20130101; F01D 5/182 20130101; F05D 2240/35 20130101; F01D
9/041 20130101; F23R 3/06 20130101; F05D 2260/94 20130101; F05D
2240/12 20130101; F05D 2260/201 20130101 |
International
Class: |
F01D 25/12 20060101
F01D025/12; F01D 5/18 20060101 F01D005/18; F01D 9/04 20060101
F01D009/04; F23R 3/00 20060101 F23R003/00; F23R 3/06 20060101
F23R003/06; B23P 15/04 20060101 B23P015/04 |
Claims
1. A gas turbine engine component configured to separate a cooling
air plenum from a heated gas environment, the gas turbine engine
component comprising: a substrate defining a surface; and a unitary
structure comprising a cooling region and a cover layer, wherein:
the cover layer defines a hot wall surface configured to face the
heated gas environment, the cooling region is disposed between the
cover layer and the substrate and comprises a plurality of support
structures extending between the cover layer and the surface of the
substrate, and at least some of the support structures define a
respective bond surface bonded to the substrate at the surface of
the substrate.
2. The gas turbine engine component of claim 1, wherein the
plurality of support structures defines a plurality of cooling
channels, respective cooling channels of the plurality of cooling
channels being between adjacent support structures of the plurality
of support structures.
3. The gas turbine engine component of claim 2, wherein the hot
wall surface defines a plurality of cooling apertures, wherein
respective cooling apertures of the plurality of cooling apertures
are fluidly connected to respective cooling channels of the
plurality of cooling channels for fluidly connecting the heated gas
environment to the cooling region.
4. The gas turbine engine component of claim 1, wherein the unitary
component comprises a turbine vane defining an exterior surface
comprising the hot wall surface and defining an internal hollow
chamber comprising the cooling air plenum, wherein the cover layer
comprises a coversheet, and wherein the substrate comprises a
spar.
5. The gas turbine engine component of claim 1, wherein the unitary
component comprises a combustor component configured to separate
the cooling air plenum from a combustion chamber comprising the
heated gas environment.
6. The gas turbine engine component of claim 1, wherein the unitary
component comprises a dual wall structure configured to separate
the cooling air plenum from the heated gas environment, wherein the
dual wall structure defines a hot section wall comprising the hot
wall surface, wherein the substrate defines a cold section wall,
wherein the cold section wall defines a plurality of impingement
apertures that extend through a thickness of the cold section wall,
wherein the plurality of support structures comprises a plurality
of pedestals that connect the cold section wall to the hot section
wall to define a plurality of cooling channels between the cold
section wall and the hot section wall, and wherein the plurality of
impingement apertures, the plurality of cooling channels, and the
cooling apertures fluidly connect the cooling air plenum to the
heated gas environment.
7. The gas turbine engine component of claim 1, wherein the unitary
component comprises a flame tube, a combustion ring, a combustor
casing, a combustor guide vane, a turbine vane, a turbine disc, or
a turbine blade.
8. A method of fabricating a gas turbine engine component
configured to separate a cooling air plenum from a heated gas
environment, the method comprising: additively depositing a unitary
structure on a surface of a substrate, wherein the unitary
structure comprises a cooling region and a cover layer, wherein the
cover layer defines a hot wall surface configured to face the
heated gas environment, wherein the cooling region is disposed
between the cover layer and the substrate and comprises a plurality
of support structures extending between the cover layer and the
surface of the substrate, and wherein at least some of the support
structures define a respective bond surface bonded to the substrate
at the surface of the substrate.
9. The method of claim 8, wherein the plurality of support
structures defines a plurality of cooling channels, respective
cooling channels of the plurality of cooling channels being between
adjacent support structures of the plurality of support
structures.
10. The method of claim 8, further comprising installing the
unitary structure in a gas turbine engine.
11. The method of claim 10, wherein installing the unitary
structure comprises bonding the gas turbine engine component to a
gas turbine engine surface.
12. The method of claim 11, wherein the bonding comprises diffusion
bonding.
13. The method of claim 10, wherein the installing the unitary
structure comprises connecting the gas turbine engine component to
an air-cooling system of the gas turbine engine.
14. The method of claim 8, wherein the unitary structure comprises
a dual wall component.
15. The method of claim 8, wherein the unitary structure comprises
a flame tube, a combustion ring, a combustor casing, a combustor
guide vane, a turbine vane, a turbine disc, or a turbine blade.
16. The method of claim 9, wherein the hot wall surface defines a
plurality of cooling apertures fluidly connected to the plurality
of support structures, wherein respective cooling apertures of the
plurality of cooling apertures are fluidly connected to respective
cooling channels of the plurality of cooling channels for fluidly
connecting the heated gas environment to the cooling region.
17. The method of claim 8, wherein additively depositing the
unitary structure comprises: directing a material stream and an
energy stream at a focal region on the surface of the substrate,
and moving the focal region along a predetermined path.
18. The method of claim 17, wherein the energy beam comprises a
laser beam and wherein the material stream comprises one or more of
metal or alloy powder, wire, or ribbon.
Description
TECHNICAL FIELD
[0001] The present disclosure relates to forming a cooled component
in a gas turbine engine.
BACKGROUND
[0002] Hot section components of a gas turbine engine may be
operated in high temperature environments that may approach or
exceed the softening or melting points of the materials of the
components. Such components may include air foils including, for
example turbine blades or vanes which may have one or more surfaces
exposed high temperature combustion or exhaust gases flowing across
the surface of the competent. Different techniques have been
developed to assist with cooling of such components including, for
example, application of a thermal barrier coating to the component,
construction the component as single or dual wall structure, and
passing a cooling fluid, such as air, across or through a portion
of the component to aid in cooling of the component. Maintaining
the efficiency and operation of such cooling systems is useful to
facilitate engine performance and prevent over heating of the
engine.
SUMMARY
[0003] In some examples, the disclosure describes an example gas
turbine engine component configured to separate a cooling air
plenum from a heated gas environment. The gas turbine engine
component includes a substrate defining a surface, and a unitary
structure. The unitary structure includes a cooling region and a
cover layer. The cover layer defines a hot wall surface configured
to face the heated gas environment. The cooling region is disposed
between the cover surface and the substrate. The cooling region
includes a plurality of support structures extending between the
cover layer and the surface of the substrate. At least some of the
support structures define a respective bond surface bonded to the
substrate at the surface of the substrate.
[0004] In some examples, the disclosure describes an example
technique for fabricating a gas turbine engine component. The
technique includes additively depositing a unitary structure on a
surface of a substrate. The unitary structure includes a cooling
region and a cover layer. The cover layer defines a hot wall
surface configured to face the heated gas environment. The cooling
region is disposed between the cover layer and the substrate and
includes a plurality of support structures extending between the
cover layer and the surface of the substrate. At least some of the
support structures define respective bond surface bonded to the
substrate at the surface.
[0005] The details of one or more examples are set forth in the
accompanying drawings and the description below. Other features,
objects, and advantages will be apparent from the description and
drawings, and from the claims.
BRIEF DESCRIPTION OF THE FIGURES
[0006] FIG. 1 is a conceptual cross-sectional view of an example
component for a gas turbine engine that includes a unitary
structure bonded to a surface of a substrate.
[0007] FIG. 2 is conceptual cross-sectional view of an example
component for a gas turbine engine that includes a unitary
structure bonded to a surface of a substrate, where a cover layer
of the unitary structure defines cooling apertures.
[0008] FIG. 3 is a conceptual diagram of an example turbine blade
for use in a gas turbine engine.
[0009] FIG. 4 is a conceptual cross-sectional view of an example
dual walled turbine blade for use in a gas turbine engine that
includes a unitary structure bonded to a surface of a
substrate.
[0010] FIG. 5 is a cross-sectional view of an example combustor
that includes a flame tube with a sidewall.
[0011] FIG. 6 is a schematic and conceptual block diagram of an
example system for fabricating a gas turbine engine component that
includes a unitary structure bonded to a substrate.
[0012] FIG. 7 is a conceptual flow chart of an example technique
for fabricating a gas turbine engine component that includes a
unitary structure bonded to a substrate.
DETAILED DESCRIPTION
[0013] The disclosure generally describes gas turbine engine
components configured to separate a cooling air plenum from a
heated gas environment, in which the gas turbine engine component
includes a substrate defining a surface and a unitary structure
comprising a cooling region and a cover layer. The cover layer may
define a hot wall surface configured to face the heated gas
environment. The cooling region may be disposed between the cover
layer and the substrate and includes a plurality of support
structures extending between the cover layer and the surface of the
substrate. At least some of the support structures define a
respective bond surface bonded to the substrate at the surface of
the substrate. The surface of the substrate may include a
two-dimensional collection of points (flat surface), a
three-dimensional collection of points whose cross section is a
curve (curved surface), the boundary of any three-dimensional
solid, or a continuous boundary dividing a three-dimensional space
into two regions. By disposing the bond between the unitary
structure and the substrate at the interface between the support
structures and the substrate, the gas turbine engine components may
possess an increased working life and better mechanical stability
compared to components in which the bond interface is between the
cover layer and the support structures. While not wishing to be
bound by any specific theory, it is believed that this is due to
the bond interface potentially introducing thermal resistivity. By
moving the bond interface further from the hot wall surface, more
of the unitary structure is exposed to cooling air, which may lead
to increased heat transfer from the unitary structure to the
cooling air.
[0014] Hot section components, such as turbine surfaces, air foils,
and flame tubes of a combustor of a gas turbine engine may be
operated in high temperature gaseous environments. In some
examples, the temperature of the gaseous environments may approach
or exceed the melting point or softening point of a material from
which at least a portion of the component is formed. For example,
operating temperatures in a high pressure turbine section of a gas
turbine engine may exceed melting or softening points of superalloy
materials used in the high pressure turbine section, e.g., to form
substrates of blades or vanes.
[0015] In some examples, to reduce or substantially prevent melting
or softening of the engine components, the components may include a
dual wall structure having a hot wall (e.g., coversheet), a cold
wall (e.g., substrate), and a cooling region between the hot wall
and the cold wall. The cooling region may include support
structures between the hot wall and the cold wall. In some
examples, the cooling system may function by flowing relatively
cold air from the compressor section of the gas turbine engine
through cooling channels in the cooling region of the dual wall
structure. These cooling channels may exhaust some or all of the
cooling air through cooling apertures in the surface of the hot
wall. In some examples, the cooling air may help protect the
component in such high temperature gaseous environments by, for
example, reducing the relative temperature of the component,
creating an insulating film of cooling air passing over the surface
of the component exposed to the high temperature environment, or
reducing the temperature of the gas within the high temperature
environment. Dual wall structures may also reduce cooling airflow
needs compared to a single wall structure, so that a greater volume
of the airflow is available for operation of the turbine, for
example, for combustion.
[0016] Support structures may include features such as pins, fins,
pedestals, or the like between the hot wall and the cold wall in
the dual wall structure. In some examples, the support structures
also function as cooling features, the dual wall structure may
include additional cooling features (such as cooling channel)
between the hot wall and the cold wall, or both. Such cooling
features may improve the effectiveness of cooling, for example, by
providing additional surface area for convective cooling, by
increasing conduction area to draw heat away from the hot wall, by
routing cooling air through the space between the hot wall and cold
wall in selected flow patterns, or the like. In some examples, the
effectiveness of the cooling features may increase as the cooling
features are made finer due to an increase in exposed surface area
to volume of the cooling features.
[0017] While techniques such as integral casting, diffusion
bonding, and machining may be used to fabricate dual wall
structures, these techniques have drawbacks. For example, integral
casting with ceramic cores may utilize ceramic cores with very fine
features, which are difficult to reliably and repeatably form, may
have low manufacturing yields, may have limitations on feature
size, and may present difficulties in inspecting support structures
and cooling features between the hot wall and the cold wall to
check for defects, blockages, or other failures. Using refractory
metal cores may present similar difficulties in inspecting the
support structures and cooling features to check for defects,
blockages, or other failures.
[0018] Diffusion bonding of separate spars and coversheets may
present higher costs and increased complexity because of additional
machining of castings prior to diffusion bonding and the use of
multiple castings. Additionally, in some examples, bonding cycles
may lead to some loss in material capability. While DMLS (direct
metal laser sintering) may be used to fabricate separate cover
sheets on spars having cooling features, the resulting components
may have reduced material properties compared to single crystal
alloys used in hot section components, for example, because of
geometric discontinuities or compositional differences between
separately fabricated cover sheets and spars, and may exhibit
potential blockage of cooling circuits if filler material used to
define voids or other channels is not fully removed after DMLS is
complete. Aligning cooling holes in the hot wall with the
underlying cooling pattern also may be difficult when fabricating
the hot wall separately from the cold wall then joining the hot
wall and cold wall.
[0019] In some examples according to the disclosure, dual wall
components may be formed by additively depositing a unitary
structure on a substrate. The unitary structure may include the hot
wall, the support structures, and, optionally, one or more separate
cooling features. The disclosed examples and techniques described
herein may be used to manufacture dual wall structures with, in
various examples, intricate or fine cooling features, having higher
yields compared to integral casting, having lower costs than
diffusion bonded constructions, and/or providing better alignment
between cooling holes in the hot wall and support structures or
cooling features. Example dual wall components according to the
disclosure may have a bond line between the cold wall and the hot
wall that is farther from the hot wall compared to dual wall
components manufactured using other techniques, which may lead to
an increased working life and better mechanical stability.
[0020] FIG. 1 is a conceptual cross-sectional view of an example
component 10 for a gas turbine engine. Component 10 includes a
unitary structure 12, and is configured to separate a cooling air
plenum 14 from a heated gas environment 16 such that component 10
acts as a physical separation between the two environments.
[0021] In some examples, component 10 may include a hot section
component for a gas turbine engine that receives or transfers
cooling air as part of cooling system for a gas turbine engine.
Component 10 may include, for example, components of a combustor
such as a flame tube, combustion ring, the inner or outer casing,
liner, guide vanes, or the like; components of a turbine section
such as a nozzle guide vane, a turbine disc, a turbine blade, or
the like; or another component associated with the air-cooling
system of a gas turbine engine. In some examples, component 10 may
be constructed with a ceramic matrix composite, a superalloy
substrate, or other materials used, for example, in the aviation or
aerospace industry. However, component 10 may be formed of suitable
materials other than those mentioned above.
[0022] Cooling air plenum 14 and heated gas environment 16 may
represent different flow paths, chambers, or regions within the gas
turbine engine in which component 10 is installed. For example, in
some examples in which component 10 is a flame tube of a combustor
of a gas turbine engine, heated gas environment 16 may comprise the
combustion chamber within the flame tube and cooling air plenum 14
may be the by-pass/cooling air flow path that surrounds the
exterior of the flame tube. In some examples in which component 10
is a turbine blade or vane, heated gas environment 16 may represent
the environment exterior to and flowing past the turbine blade or
vane while cooling air plenum 14 may include one or more interior
chambers within the turbine blade or vane representing part of the
integral cooling system of the gas turbine engine.
[0023] Unitary structure 12 includes a cover layer 18 and a cooling
region 22. Unitary structure 12 is constructed of a composition 28
that may include any suitable material discussed above with
reference to the construction of component 10 that may be deposited
using additive manufacturing. In some examples, composition 28 may
include a superalloy, for example a nickel-based superalloy.
[0024] Unitary structure 12 may be disposed on and attached to a
major surface 31 of a substrate 30 of component 10. Cooling region
22 of unitary structure 12 may be disposed between cover layer 18
and substrate 30 such that cover layer 18 faces heated gas
environment 16 and substrate 30 faces cooling air plenum 14. As
such, in some examples, substrate 30 may be referred to as a cold
wall and cover layer 18 may be referred to as a hot wall. In some
examples, one or both of cover layer 18 and substrate 30 may define
a thickness between about 0.014 inches and about 0.300 inches
(e.g., about 0.36 mm to about 7.62 mm). In some examples, cooling
region 22 may have a thickness between about 0.25 mm and about 7
mm.
[0025] Cover layer 18 defines a hot wall surface 20 configured to
face heated gas environment 16. Substrate 30 defines a cold wall
surface 38 configured to face cooling air plenum 14. The terms
"cold wall surface" and "hot wall surface" are used merely to
orient which wall is adjacent to cooling air plenum 14 and which
wall is adjacent to heated gas environment 16, respectively, and
are not intended to limit the relative temperatures of the
different environments or wall. For example, while cold wall
surface 38 and cooling air plenum 14 may be described as "cold"
sections compared to hot wall surface 20 and heated gas environment
16, the respective temperatures of cold wall surface 38 or cooling
air plenum 14 may reach temperatures between about 390.degree. F.
to about 1830.degree. F. (e.g., about 200.degree. C. to about
1000.degree. C.) during routine operation.
[0026] Cooling region 22 may include a plurality of support
structures 24. The plurality of support structures 24 may define a
network of the plurality of cooling channels 26. For example, the
plurality of support structures 24 may include one or more of
pedestals, columns, spires, raised features, or channel walls. The
plurality of support structures 24 also may function as cooling
features, e.g., for conducting heat from cover layer 18 toward
substrate 30. In some examples, cooling region 22 may include one
or more additional cooling features, such as the plurality of
cooling channels 26. The plurality of support structures 24 and,
optionally, other cooling features, may take on any useful
configuration, size, shape, or pattern. In some such examples, the
height of plurality of support structures 24 may be between about
0.25 mm and about 7 mm to define the thickness of cooling region
22.
[0027] In some examples, the plurality of support structures 24 may
include a corrugated structure that defines the plurality of
cooling channels 26 between the respective walls of the corrugated
structure. In some examples, the plurality of support structures 24
may also include one or more dams that act as zone dividers within
the cooling region 22 thereby separating one cooling channel of the
plurality of cooling channels 26 from another cooling channel of
the plurality of cooling channels 26. The introduction of dams
within cooling region 22 may assist with maintaining a more uniform
temperature along hot wall surface 20 by controlling flow of
cooling air within the plurality of cooling channels 26. Thus, in
some examples, the plurality of support structures 24 provides a
conduit for heat transfer across hot wall surface 20 of cover layer
18 and cooling region 22 between cooling air plenum 14 and heated
gas environment 16, as part of the air-cooling system for a gas
turbine engine.
[0028] Unitary structure 12 is bonded to substrate 30, for example,
at bond surface 32 defined by cooling region 22, e.g., at
respective bases of the plurality of support structures 24 opposite
of cover layer 18. As seen in FIG. 1, bond surface 32 is between
cooling region 22 and substrate 30, such that bond surface 32 is on
the opposite side of cooling region 22 from cover layer 18.
[0029] The efficiency of heat transferred from heated gas
environment 16 to cooling air plenum 14 across cooling region 22
may depend on a variety of factors including, but not limited to,
the thermal conductivity of composition 28 of unitary structure 12,
the total area of hot wall surface 20 of cover layer 18, the
surface area defined by plurality of support structures 24 and
plurality of cooling channels 26, the thermal conductivity of
substrate 30, the total area of cold wall surface 38, the thermal
conductivity at bond surface 32, and the size of cooling channels
of the plurality of cooling channels 26. For example, the thermal
conductivity at bond surface 32 may be different (e.g., less) from
the respective thermal conductivity of one or both of unitary
structure 12 and substrate 30. In some examples, bond surface 32
may act as a thermal resistor that inhibits the transfer of heat
across bond surface 32, because of a lower thermal conductivity at
bond surface 32, or because of local variations in composition or
geometry. In examples where bond surface 32 has a lower thermal
conductivity or acts as a thermal resistor, positioning bond
surface 32 closer to hot wall surface 20, for example between cover
layer 18 and cooling region 22, may impede the transfer of heat
from hot section part 14 to pedestals 16 compared to when bond
surface 32 is positioned further from hot wall surface 20, for
example between cooling region 22 and substrate 30. The net
consequence of such a configuration may result in less heat being
transferred to cooling air 32. In contrast, positioning bond
surface 32 farther away from hot wall surface 20, for example,
between cooling region 22 and substrate 30 may improve the
efficiency of heat transfer. Therefore, positioning bond surface 32
between cooling region 22 and substrate 30 may result in more heat
being transferred from hot wall surface 20 across cooling region
22.
[0030] In some examples, cover layer 18 may define a plurality of
cooling apertures 34, as shown in FIG. 2. FIG. 2 is conceptual
cross-sectional view of an example component 10b for a gas turbine
engine that includes a unitary structure 12b adjacent a substrate
30b. Unitary structure 12b may be substantially similar to unitary
structure 12 discussed with reference to FIG. 1 above, while
including a cover layer 18b substantially similar to cover layer 18
of FIG. 1.
[0031] Unlike cover layer 18 of FIG. 1, cover layer 18b defines the
plurality of cooling apertures 34. Cooling apertures 34 may extend
between cooling region 22 and hot wall surface 20. In some
examples, substrate 30b may be substantially similar to substrate
30 discussed with reference to FIG. 1 above. However, unlike
substrate 30 of FIG. 1, substrate 30b may define a plurality of
impingement apertures 36 extending between cooling region 22 and
cold wall surface 38. In some examples, the diameter of one or both
of plurality of cooling apertures 34 and impingement apertures 36
may be between about 0.01 inches and about 0.12 inches (e.g., about
0.25 mm to about 3 mm). Thus, in some examples, article 10b may be
substantially similar to article 10 discussed above with reference
to FIG. 1, while further including one or both of plurality of
cooling apertures 34 or plurality of impingement apertures 36.
[0032] During operation of either component 10 or component 10b,
the temperature of the air within cooling air plenum 14 may be less
than that of the hot gas environment 16. During operation of
component 10b, cooling air may pass from cooling air plenum 14 to
heated gas environment 16 through one or both of the plurality of
cooling apertures 34 or the plurality of impingement apertures 36.
The cooling air may assist in maintaining the temperature of
component 10b at a level lower than that of heated gas environment
16. For example, the cooling air may enter heated gas environment
16 creating an insulating film of relatively cool gas along hot
wall surface 20 of component 10b that allows hot wall surface 20 of
component 10b to remain at a temperature less than that of the bulk
temperature of heated gas environment 16. In some examples, the
cooling air may also at least partially mix with the gas of heated
gas environment 16, thereby reducing the relative temperature of
heated gas environment 16. In some examples, the cooling region 22
may create a zoned temperature gradient between the respective
regions of cooling air plenum 14 and heated gas environment 16.
Additionally, or alternatively, the cooling gas may act as a
cooling reservoir that absorbs heat from component 10b as the gas
passes through cooling apertures 24 or along one or more of the
surfaces of component 10b, thereby dissipating the heat of
component 10b and allowing the relative temperature of component to
be maintained at a temperature less than that of heated gas
environment 16.
[0033] In some examples, the cooling air may be supplied to
component 10b (e.g., via cooling air plenum 14) at a pressure
greater than the gas path pressure within heated gas environment
16. The pressure differential between cooling air plenum 12 and
heated gas environment 16 may force cooling air 18 through the
plurality of cooling apertures 34. In some examples, the plurality
of cooling apertures 34 may include film cooling holes that are
shaped to reduce the use of cooling air. The plurality of cooling
apertures 34 may be positioned in any suitable configuration and
position about the surface of component 10b. For example, the
plurality of cooling apertures 34 may be positioned along the
leading edge of a gas turbine blade or vane. In some examples, the
plurality of cooling apertures 34 may define incidence angle less
than 90 degrees, i.e., non-perpendicular, to the hot wall surface
20 of component 10a. In some examples the angle of incidence may be
between about 10 degrees and about 75 degrees to hot wall surface
34 of component 10. In some such examples, adjusting the angle of
incidence of hot wall surface 34 may assist with creating a cooling
film of the cooling air along hot wall surface 20. Additionally, or
alternatively, one or more of the plurality of cooling apertures 34
may include a fanned Coanda ramp path at the point of exit from hot
wall surface 20 to help assist in the distribution or film
characteristics of the cooling air as it exits a respective cooling
aperture of the plurality of cooling apertures 34.
[0034] In some examples, component 10 or component 10b may be a
dual wall component (e.g., as illustrated in FIGS. 1 and 2). For
example, substrate 30 may include a spar, and cover layer 18 may
include a coversheet for the spar.
[0035] Unitary structure 12 of component 10, or unitary structure
12b of component 10b may be fabricated using example techniques and
example additive manufacturing systems, as described with reference
to FIGS. 6 and 7. For example, an additive manufacturing system
including a controller, a material stream and an energy beam may be
used to fabricate unitary structure 12. The controller may direct
the material stream and the energy beam along a tool path based on
a digital representation of unitary structure 12.
[0036] The energy beam may interact with portions of the material
stream, for example, by fusing, solidifying, or sintering material
from the material stream at a series of focal regions along the
tool path, to deposit volumes of material along the tool path. In
some examples, the controller may deposit successive layers of
material that ultimately form unitary structure 12. For example,
the controller may direct the material stream and the energy beam
to deposit material in layers, forming the plurality of support
structures 24 defining the plurality of cooling channels 26 in
cooling region 22.
[0037] In some examples, the controller may direct deposition of
material on surface 31 of substrate 30 in a predetermined build
direction, for example, a vertical direction pointing away from
surface 31. Thus, the controller may direct material along the
build direction, beginning with layers of material on or adjacent
surface 31 and then continuing to deposit layers that are
successively farther away from surface 31. For example, the
controller may first direct the material stream and the energy beam
to deposit layers of material forming the plurality of support
structures 24, and then continue depositing material along the tool
path forming layers of material supported by the plurality of
supporting structures 24, that eventually form cover layer 18.
Thus, additive manufacturing may be used to form cover layer 18
that is integrated, continuous, or otherwise unitary with cooling
region 22.
[0038] In some examples, before additively depositing material on
surface 31 of substrate 30, a filler composition may be deposited
in a predetermined pattern. The filler composition may be deposited
using extrusion, coating, stamping, masking, templating, or other
suitable technique for depositing the filler composition in the
predetermined pattern. The filler composition may be in the form of
a paste, a liquid, a gel, a solid, or any other form that may
support additively deposited material without collapsing. In some
examples, a layer of the filler composition of a predetermined
thickness may be deposited on surface 31, followed by machining or
removal of filler composition from predetermined portions of the
layer. The predetermined pattern of filler composition may define
the plurality of cooling channels 26, for example, by preventing
the deposition of additively deposited material within a volume
occupied by the filler composition. For example, the controller may
direct the deposition of additive material over the predetermined
pattern defined by the filler composition, such that the
predetermined pattern of filler composition defines the plurality
of cooling channels as layers of additively deposited material are
deposited over the predetermined pattern.
[0039] The filler composition may include a filler or sacrificial
material that may persist through the additive manufacturing, and
be removed after the additive manufacturing is complete. For
example, the filler composition may be susceptible to at least one
of heat, leaching, or oxidation. The filler composition may be
removable from the plurality of cooling channels 26, for example,
by subjecting the filler composition to at least one of a thermal
treatment, leaching composition, or an oxidizing environment. In
some examples, the filler composition includes one or more of
ceramic, metal, alloys, or other suitable refractory material.
Thus, in examples in which filler composition is used, at the end
of additive manufacturing, the plurality of cooling channels 26 may
be occupied with the filler composition. After the additive
manufacturing, part or whole of unitary structure 12 or component
10 may be subjected to one or more of heating, leaching, oxidation,
or other suitable treatment for substantially removing the filler
composition from unitary structure 12.
[0040] Thus, additive manufacturing may be used to fabricate
component 10, component 10b, or other example components including
unitary structures. For example, in addition to the components
described with reference to FIGS. 1 and 2, additive manufacturing
may be used to fabricate turbine components described with
reference to FIGS. 3-5.
[0041] FIG. 3 is a conceptual diagram of an example turbine airfoil
component (e.g., turbine blade or vane) for use in a gas turbine
engine. FIG. 3 illustrates an example turbine airfoil 50 that
includes a plurality of cooling apertures 52 arranged on a hot
section wall surface 54 of the airfoil. Turbine airfoil 50 may be a
dual or multi-walled structure as described above with respect to
FIGS. 1 and 2. For example, FIG. 4 illustrates a cross-sectional
view of an example dual wall turbine airfoil 70 that includes a
plurality of cooling apertures 72 along a hot section wall 84 and a
plurality of impingement apertures 80 along a cold section wall 86.
In some examples, dual wall turbine airfoil 70 may have
substantially the same structural configuration as component 12,
for example, including a cooling region including a plurality of
support structures extending between hot section wall 84 and cold
section wall 86, with the cooling region and hot section wall 84
forming a unitary structure. As shown, cooling air 78 may flow from
cooling air plenum 74 through impingement apertures 80 into cooling
channels 88 before exiting through cooling apertures 72 into heated
gas environment 76.
[0042] FIG. 5 illustrates a cross-sectional view of an example
combustor 90 that includes a flame tube 92 (e.g., combustion
chamber) with a sidewall defining a plurality of cooling apertures
94. In some examples, the gases within the combustor post
combustion, (e.g., heated gas environment 96) may exceed about
1,800.degree. C., which may be too hot for introduction against the
vanes and blade of the turbine (e.g., FIGS. 3 and 4). In some
examples, the combusted gases may be initially cooled prior to
being introduced against the vanes and blade of the turbine by
progressively introducing portions of the by-pass air (e.g.,
cooling air 98) into heated gas environment 96 of flame tube 92 via
ingress through plurality of cooling aperture 94 strategically
position around flame tube 94, fluidly connecting cooling air 98
within cooling air plenum 100 with heated gas environment 96.
[0043] In some examples, combustor 90 includes a dual wall
structure having substantially the same structural configuration as
component 12, for example, including a cooling region including a
plurality of support structures extending between a surface
adjacent heated gas environment 96 and a surface adjacent cooling
air 98, with the cooling region and the layer adjacent heated gas
environment 96 forming a unitary structure. In some example,
cooling air 98 may intimately mix with the combusted gases to
decease the resultant temperature of the volume of heated gas
environment 96. Additionally, or alternatively, cooling air 98 may
form an insulating cooling air film along the interior surface
(e.g., hot section surface) of flame tube 92. In some examples, the
wall of flame tube may include a dual wall (e.g., component 10 or
10b) structure.
[0044] Example gas turbine engine components including a unitary
structure and a substrate have been described above. As described
above, additive manufacturing may be used to fabricate a unitary
structure including an integrated or continuous unitary cover layer
and cooling region. For example, example components may be
fabricated using additive manufacturing, for example, using the
example system of FIG. 6, and the example technique of FIG. 7, as
discussed below. However, example components described with
reference to FIGS. 1-5 above may be fabricated using other suitable
example systems or other suitable example techniques.
[0045] FIG. 6 is a schematic and conceptual block diagram of an
example system 110 for fabricating a gas turbine engine component
that includes a unitary structure 112 disposed on a substrate 130.
Example system 110 includes a computing device 116 that may control
an additive manufacturing tool 114 for fabricating unitary
structure 112 that includes a material 120. In some examples,
computing device 116 may generate or store a digital representation
112a of unitary structure 112, or of the gas turbine engine
component that includes unitary structure 112. In some examples,
computing device 116 may control additive manufacturing tool 114 to
fabricate the component including unitary structure 112 based on
digital representation 112a. In some examples, additive
manufacturing tool 114 may include a controller 134 for controlling
one or more of a material source 124, and energy source 128, and an
imaging device 132.
[0046] Computing device 116 may send control signals to controller
134 for controlling additive manufacturing tool 114. For example,
computing device 116 may send operational signals to and receive
status signals from controller 134 to control and monitor the
operation of additive manufacturing tool 114. In some examples,
computing device 116 may not control additive manufacturing tool
114, and controller 134 may be configured to receive signals
indicative of digital representation 112a from computing device 116
and to control additive manufacturing tool 114 based on digital
representation 112a to fabricate the component including unitary
structure 112.
[0047] In some examples, controller 134 may control material source
124 of additive manufacturing tool 114 to direct a material stream
126 including material 120 at a build location on
partially-fabricated component 112, which is carried on a build
platform 122, or at an initial build location on a region of
substrate 130 on build platform 22. In some examples, material 120
may include metal or alloy, or any suitable material composition
discussed above with reference to component 10 or unitary structure
12 of FIG. 1.
[0048] Controller 134 also may control energy source 128 to direct
an energy beam 127 at the build location. Energy beam 127 may
interact with material 120 from material stream 126 at the build
location, for example, by fusing, melting, sintering, curing,
solidifying or otherwise modifying material 120 at the build
location to cause material 120 to be joined to other material of
component 112 at the build location, or to material of substrate
130. Energy beam 127 may include any energy, for example,
ultraviolet light, electron beam, plasma, or laser, that may
interact with material 120 to change a state of material 120. For
example, energy beam 127 may be focusable or directable towards
material 120 in material stream 126. In some examples, the build
location at which energy beam 127 interacts with material stream
126 is adjacent an existing surface of unitary structure 112 such
that material 120 is added to unitary structure 112. In some
examples, controller 134 may control energy source 128 to emit a
diffuse energy beam, or a patterned array of beams, for example, a
light pattern. The build location may change as unitary structure
112 is fabricated, for example, along regions or surfaces of partly
fabricated unitary structure 112. In some examples, controller 134
may cause additive manufacturing tool 114 to fabricate unitary
structure 112 by depositing material 120 at different build
locations along a tool path, so that material 120 is ultimately
deposited along a predetermined build direction, for example a
vertical build direction upwards (for example, against a
gravitational force) or downwards (for example, toward a
gravitational force).
[0049] In some examples, build platform 122 may remain stationary
as unitary structure 112 is fabricated. In other examples, build
platform 122 may be movable or rotatable, for example, along
multiple axis, and controller 134 may control the position of build
platform 122. In some examples, controller 134 may successively
move build platform 122 against the build direction, or to change
the build location by changing the orientation of build platform
122, and that of unitary structure 112, relative to material stream
126 and energy beam 127.
[0050] In some examples, controller 134 may separately control
material source 124 and energy source 128, for example, by
separately controlling material source 124 to direct material
stream 126 to deposit a layer or volume of material 120, and then
controlling energy source 128 to direct energy beam 127 along a
series of build locations within the deposited layer or volume of
material 120 to energize material 120 at the build locations to
fabricate unitary structure 112. Therefore, controller 134 may
direct build location along a two-dimensional or three-dimensional
tool path to fabricate unitary structure 112 based on digital
representation 112a.
[0051] In some examples, controller 134 may control imaging device
132 to image surfaces or regions or volumes of one or more of
unitary structure 112, the build location, or platform 122 to
generate respective build images periodically or continuously.
Controller 134 may periodically or continuously compare the build
images with the digital representation 112a to verify that unitary
structure 112 substantially conforms (e.g., conforms or nearly
conforms) to digital representation 112a. In some examples,
controller 134 may control one or more of material source 124,
energy source 128, and build platform 122 based on the build images
and the digital representation 112a. For example, controller 134
may be configured to control build platform 122 and material source
124, energy source 128, and/or imaging device 132 to translate
and/or rotate along at least one axis to position unitary structure
112 relative to material stream 126, energy beam 127, and/or
imaging device 132. Positioning unitary structure 112 relative to
material stream 126, energy beam 127, and/or imaging device 132 may
include positioning a predetermined surface (e.g., a surface to
which material is to be added) of unitary structure 112 in a
predetermined orientation relative to material source 124, energy
source 128, and/or imaging device 132, so that material is added in
regions or volumes defined by digital representation 112a.
[0052] In some examples, additive manufacturing tool 114 may not
include controller 134, and computing device 112 may control one or
more of material source 124, energy source 128, imaging device 132,
and build platform 122, instead of controller 134.
[0053] Example system 110 discussed above may be used to fabricate
example components described above with reference to FIGS. 1-5.
However, example system 110 may be used to fabricate other example
components according to the disclosure.
[0054] FIG. 7 is a flow diagram illustrating an example technique
for forming a component of a gas turbine engine that includes a
unitary structure including a cover layer and a cooling region.
While the below technique of FIG. 7 is described with respect to
components 10 and 10b of FIGS. 1 and 2, and system 110 of FIG. 6,
it will be understood from the context of the specification that
the technique of FIG. 7 may be implemented using other systems, or
applied to other components of a gas turbine engine including, for
example, components 50, 70, and 90, flame tubes, combustor rings,
combustion chambers, casings of combustion chambers, turbine
blades, turbine vanes, or the like; all of which are envisioned
within the scope of the technique of FIG. 7.
[0055] The example technique of FIG. 7 includes fabricating
component 10 by additively depositing a unitary structure 12 on a
surface of a substrate 30 (210). As discussed with reference to
FIG. 1 above, unitary structure 12 includes cooling region 22
opposing cover layer 18. Cooling region 22 is disposed between
cover layer 18 and substrate 30. Component 10 is configured to
separate cooling air plenum 14 from heated gas environment 16.
Cover layer 18 defines a hot wall surface 20 configured to face
heated gas environment 16. Cooling region 22 defines the plurality
of support structures 24 extending between cover layer 18 and
surface 31 of substrate 30. Cooling region 22 defines bond surface
32 bonded to substrate 30.
[0056] In some examples, additively depositing the unitary
structure includes directing material stream 126 and energy stream
127 at a focal region on a surface of substrate 122, or on a
surface of partially fabricated unitary structure 112. In some
examples, additively depositing the unitary structure includes
moving the focal region along a predetermined path. For example,
controller 134 may control energy beam 127 and material stream 126
such that energy beam 127 interacts with portions, volumes, or
packets of material from material stream 126, for example, by
fusing, solidifying, or sintering material from material stream 126
at a series of focal regions along the tool path, to deposit
volumes of material along the tool path. In some examples,
controller 134 may deposit successive layers of material that
ultimately form unitary structure 12. For example, controller 134
may direct material stream 126 and energy beam 127 to deposit
material in layers, forming the plurality of support structures 24
defining the plurality of cooling channels 26 in cooling region
22.
[0057] In some examples, controller 134 may direct deposition of
material on surface 31 of substrate 30 in a predetermined build
direction, for example, a vertical direction pointing away from
surface 31. Thus, controller 134 may direct material along the
build direction, beginning with layers of material on or adjacent
surface 31 and then continuing to deposit layers that are
successively farther away from surface 31. For example, controller
134 may first direct material stream 126 and energy beam 127 to
deposit layers of material forming the plurality of support
structures 24, and then continue depositing material along the tool
path forming layers of material supported by the plurality of
supporting structures 24, that eventually form cover layer 18.
[0058] In some examples, digital representation 112a may include a
representation of plurality of cooling apertures 34. Controller 134
may direct material stream 126 and laser beam 128 around regions
defining cooling apertures 34, for example, by depositing layers
that define successive cross-sections of predetermined channels
that eventually define cooling apertures 34. Thus, in some
examples, system 110 may fabricate cover layer 18 that defines
plurality of cooling apertures 34. In some examples, cover layer 18
may be fabricated without cooling apertures 34, and cooling
apertures may be formed after the additive manufacturing, for
example, by machining or drilling apertures through cover layer
18.
[0059] Thus, example techniques may include additive manufacturing
to form cover layer 18 that is integrated, continuous, or otherwise
unitary with cooling region 22.
[0060] Once formed, component 10 may be installed in a gas turbine
engine (212) and connected to the air cooling system of the engine.
In some examples, installing the unitary structure includes bonding
the component to a gas turbine engine component surface. In some
examples, bonding the component includes diffusion bonding. In some
examples, installing the unitary structure includes connecting the
component to an air-cooling system of the gas turbine engine.
[0061] While the example technique of FIG. 7 may be used to
fabricate example components described above with reference to
FIGS. 1-5 and with reference to example system 110 of FIG. 6, the
example technique of FIG. 7 may be used to fabricate other example
components.
[0062] Various examples have been described. These and other
examples are within the scope of the following claims.
* * * * *