U.S. patent application number 13/228516 was filed with the patent office on 2013-03-14 for turbine endwall with grooved recess cavity.
The applicant listed for this patent is Zhihong Gao, Joseph B. Gilliam, GEORGE LIANG, Brian J. Wessell. Invention is credited to Zhihong Gao, Joseph B. Gilliam, GEORGE LIANG, Brian J. Wessell.
Application Number | 20130064680 13/228516 |
Document ID | / |
Family ID | 47682743 |
Filed Date | 2013-03-14 |
United States Patent
Application |
20130064680 |
Kind Code |
A1 |
LIANG; GEORGE ; et
al. |
March 14, 2013 |
TURBINE ENDWALL WITH GROOVED RECESS CAVITY
Abstract
A vane assembly for a gas turbine engine including an endwall
and an airfoil extending from the endwall. An inner rail extends
radially inwardly of the endwall, and an overhang portion extends
axially from a location of the inner rail to a downstream edge. A
recess cavity is defined in the overhang portion between the inner
rail and the downstream edge. The recess cavity extends radially
into the overhang portion and defines a cavity surface. A plurality
of grooves extend radially into the cavity surface and have an
elongated dimension extending in a direction from the inner rail
toward the downstream edge. A plurality of cooling passages extend
axially through the overhang portion, and are located between the
grooves.
Inventors: |
LIANG; GEORGE; (Palm City,
FL) ; Gao; Zhihong; (Orlando, FL) ; Wessell;
Brian J.; (Orlando, FL) ; Gilliam; Joseph B.;
(Deltona, FL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
LIANG; GEORGE
Gao; Zhihong
Wessell; Brian J.
Gilliam; Joseph B. |
Palm City
Orlando
Orlando
Deltona |
FL
FL
FL
FL |
US
US
US
US |
|
|
Family ID: |
47682743 |
Appl. No.: |
13/228516 |
Filed: |
September 9, 2011 |
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F05D 2260/2214 20130101;
F01D 9/04 20130101; F05D 2240/81 20130101; F01D 25/12 20130101 |
Class at
Publication: |
416/97.R |
International
Class: |
F01D 5/18 20060101
F01D005/18 |
Claims
1. A component in a gas turbine engine comprising: an airfoil
extending radially outwardly from an endwall associated with said
airfoil, said endwall extending between an upstream edge and a
downstream edge and defining a cool side and a gas side; a recess
cavity defined in an overhang portion extending from a location
adjacent to said downstream edge toward said upstream edge, said
recess cavity extending radially into said overhang portion from
said cool side toward said gas side and defining a cavity surface
located a first distance into said endwall from a peripheral
radially inner surface of said endwall; a plurality of grooves
extending radially into said cavity surface and having an elongated
dimension extending in a direction from said downstream edge toward
said upstream edge, said grooves including a groove bottom surface
located a second distance radially into said endwall greater than
said first distance; and cooling passages extending through said
overhang portion and located: between pairs of said grooves; and
radially between said groove bottom surface and said cavity
surface.
2. The component of claim 1, wherein said endwall includes opposing
lateral sides extending in an axial direction between said upstream
and downstream edges and said recess cavity extends
circumferentially between said lateral sides of said endwall, and
said plurality of grooves are spaced circumferentially across said
recess cavity.
3. The component of claim 1, wherein said endwall comprises a
radially inner endwall and includes an inner diameter endwall
post-impingement cooling chamber located adjacent to said recess
cavity, said cooling passages extending from said inner diameter
endwall post-impingement cooling chamber to said downstream
edge.
4. (canceled)
5. The component of claim 1, wherein a radially extending raised
portion of said recess cavity, between each pair of grooves,
includes one of said cooling passages.
6. The component of claim 3, including an inner rail extending
generally circumferentially between said inner diameter endwall
post-impingement cooling chamber and said recess cavity, and said
cooling passages extend through said inner rail.
7. (canceled)
8. (canceled)
9. The component of claim 1, wherein said airfoil comprises a
leading edge and a trailing edge, and said trailing edge of said
airfoil is joined to said gas side of said endwall at an axial
location aligned with a portion of said recess cavity.
10. A vane assembly for a gas turbine engine comprising: an inner
endwall extending between an upstream edge and a downstream edge,
and defining a cool side and a gas side; an outer endwall spaced
radially outward of said inner endwall; an airfoil extending from
said inner endwall to said outer endwall, said airfoil including a
leading edge and a trailing edge; an inner rail extending generally
circumferentially along said inner endwall and radially inwardly of
said cool side of said inner endwall; said inner endwall including
an overhang portion extending axially from a location of said inner
rail; a recess cavity defined between said inner rail and said
downstream edge, and extending radially into said overhang portion
from said cool side toward said gas side and defining a cavity
surface located a first distance into said inner endwall from a
peripheral radially inner surface of said inner endwall; a
plurality of grooves extending radially into said cavity surface
and having an elongated dimension extending in a direction from
said inner rail toward said downstream edge, said grooves including
a groove bottom surface located a second distance radially into
said inner endwall greater than said first distance; and cooling
passages extending through said overhang portion and located:
between pairs of said grooves; radially between said groove bottom
surface and said cavity surface.
11. The vane assembly of claim 10, wherein said endwall includes
opposing lateral sides extending in an axial direction between said
upstream and downstream edges and said recess cavity extends
circumferentially between said lateral sides of said inner endwall,
and said plurality of grooves are spaced circumferentially across
said recess cavity.
12. The vane assembly of claim 10, wherein said inner endwall
includes an inner diameter endwall post-impingement cooling chamber
located adjacent to said recess cavity, said cooling passages
extending from said inner diameter endwall post-impingement cooling
chamber to said downstream edge.
13. (canceled)
14. The vane assembly of claim 10, wherein a radially extending
raised portion of said recess cavity, between each pair of grooves,
includes one of said cooling passages.
15. The vane assembly of claim 12, wherein said cooling passages
extend through said inner rail.
16. (canceled)
17. (canceled)
18. The vane assembly of claim 10, wherein said trailing edge of
said airfoil is joined to said gas side of said inner endwall at an
axial location aligned with a portion of said recess cavity.
19. A component in a gas turbine engine comprising: an airfoil
extending radially outwardly from an endwall associated with said
airfoil, said endwall extending between an upstream edge and a
downstream edge and defining a cool side and a gas side; a recess
cavity defined in an overhang portion extending from a location
adjacent to said downstream edge toward said upstream edge, said
recess cavity extending radially into said overhang portion from
said cool side toward said gas side and defining a cavity surface;
a plurality of grooves extending radially into said cavity surface
and having an elongated dimension extending in a direction from
said downstream edge toward said upstream edge; and cooling
passages extending through: said overhang portion; and radially
extending raised portions of said recess cavity located between
respective pairs of grooves.
20. The component of claim 19, wherein said endwall includes
opposing lateral sides extending in an axial direction between said
upstream and downstream edges and said recess cavity extends
circumferentially between said lateral sides of said endwall, and
said plurality of grooves are spaced circumferentially across said
recess cavity.
21. The component of claim 19, wherein said endwall comprises a
radially inner endwall and includes an inner diameter endwall
post-impingement cooling chamber located adjacent to said recess
cavity, said cooling passages extending from said inner diameter
endwall post-impingement cooling chamber to said downstream
edge.
22. The component of claim 19, including an inner rail extending
generally circumferentially between said inner diameter endwall
post-impingement cooling chamber and said recess cavity, and said
cooling passages extend through said inner rail.
23. The component of claim 19, wherein said cavity surface is
located a first distance into said endwall from a peripheral
radially inner surface of said endwall, and said grooves include a
groove bottom surface located a second distance radially into said
endwall greater than said first distance.
24. The component of claim 23, wherein said cooling passages are
located: between pairs of said grooves; and radially between said
groove bottom surface and said cavity surface.
Description
FIELD OF THE INVENTION
[0001] The present invention relates generally to gas turbine
engines and, more particularly, to a cooling configuration for
cooling an endwall of a component, such as a vane assembly, in a
gas turbine engine.
BACKGROUND OF THE INVENTION
[0002] In gas turbine engines, compressed air discharged from a
compressor section and fuel introduced from a source of fuel are
mixed together and burned in a combustion section, creating
combustion products defining a high temperature working gas. The
working gas is directed through a hot gas path in a turbine
section, where the working gas expands to provide rotation of a
turbine rotor. The turbine rotor may be linked to an electric
generator, wherein the rotation of the turbine rotor can be used to
produce electricity in the generator.
[0003] In view of high pressure ratios and high engine firing
temperatures implemented in modern engines, certain components,
such as airfoil assemblies, e.g., stationary vane assemblies and
rotating blade assemblies within the turbine section, must be
cooled with cooling fluid, such as compressor discharge air, to
prevent overheating of the components and to reduce thermal stress
in the components.
SUMMARY OF THE INVENTION
[0004] In accordance with an aspect of the invention, a component
is provided in a gas turbine engine. The component comprises an
airfoil extending radially outwardly from a endwall associated with
the airfoil. The endwall extends between an upstream edge and a
downstream edge and defines a cool side and a gas side. A recess
cavity is defined in an overhang portion extending from a location
adjacent to the downstream edge toward the upstream edge. The
recess cavity extends radially into the overhang portion from the
cool side toward the gas side and defines a cavity surface. A
plurality of grooves extend radially into the cavity surface and
have an elongated dimension extending in a direction from the
downstream edge toward the upstream edge.
[0005] In accordance with yet further aspects of the invention, the
endwall may include opposing lateral sides extending in an axial
direction between the upstream and downstream edges, and the recess
cavity may extend circumferentially between the lateral sides of
the endwall. Additionally, the plurality of grooves may be spaced
circumferentially across the recess cavity.
[0006] The endwall may comprise a radially inner endwall and may
include an inner diameter endwall post-impingement cooling chamber
located adjacent to the recess cavity. A plurality of cooling
passages may be provided extending from the inner diameter endwall
post-impingement cooling chamber to the downstream edge. Each of
the cooling passages may extend through the overhang portion and
may be located between a pair of the grooves. A radially extending
raised portion of the recess cavity, between each pair of grooves,
may include one of the cooling passages. An inner rail may be
provided extending generally circumferentially between the inner
diameter endwall post-impingement cooling chamber and the recess
cavity, and the cooling passages may extend through the inner
rail.
[0007] The cavity surface may be located a first distance into the
endwall from a peripheral radially inner surface of the endwall,
and the grooves may include a groove bottom surface located a
second distance radially into the endwall greater than the first
distance. The cooling passages may be located radially between the
groove bottom surface and the cavity surface.
[0008] The airfoil may comprise a leading edge and a trailing edge,
and the trailing edge of the airfoil may be joined to the gas side
of the endwall at an axial location aligned with a portion of the
recess cavity.
[0009] In accordance with another aspect of the invention, a vane
assembly is provided for a gas turbine engine. The vane assembly
comprises an inner endwall extending between an upstream edge and a
downstream edge, and defining a cool side and a gas side. An outer
endwall is spaced radially outward of the inner endwall, and an
airfoil extends from the inner endwall to the outer endwall and
includes a leading edge and a trailing edge. An inner rail extends
generally circumferentially along the inner endwall and radially
inwardly of the cool side of the inner endwall. The inner endwall
includes an overhang portion extending axially from a location of
the inner rail. A recess cavity is defined between the inner rail
and the downstream edge. The recess cavity extends radially into
the overhang portion from the cool side toward the gas side and
defines a cavity surface. A plurality of grooves extend radially
into the cavity surface and have an elongated dimension extending
in a direction from the inner rail toward the downstream edge.
[0010] Additionally, the inner endwall may include an inner
diameter endwall post-impingement cooling chamber located adjacent
to the recess cavity. A plurality of cooling passages may be
provided extending from the inner diameter endwall post-impingement
cooling chamber to the downstream edge, and the cooling passages
may extend through the inner rail.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] While the specification concludes with claims particularly
pointing out and distinctly claiming the present invention, it is
believed that the present invention will be better understood from
the following description in conjunction with the accompanying
Drawing Figures, in which like reference numerals identify like
elements, and wherein:
[0012] FIG. 1 is a partial cross-sectional view of a gas turbine
engine incorporating a vane assembly formed in accordance with
aspects of the present invention;
[0013] FIG. 2 is a perspective view of a vane assembly for a gas
turbine incorporating aspects of the present invention;
[0014] FIG. 3 is a perspective view of the vane assembly viewed
from an inner diameter side of the vane assembly;
[0015] FIG. 4 is an enlarged detail perspective view of a recess
portion of the vane assembly shown in FIG. 3; and
[0016] FIG. 5 is a cross-sectional view taken along line 5-5 in
FIG. 4, and including an impingement plate.
DETAILED DESCRIPTION OF THE INVENTION
[0017] In the following detailed description of the preferred
embodiment, reference is made to the accompanying drawings that
form a part hereof, and in which is shown by way of illustration,
and not by way of limitation, a specific preferred embodiment in
which the invention may be practiced. It is to be understood that
other embodiments may be utilized and that changes may be made
without departing from the spirit and scope of the present
invention.
[0018] Referring to FIG. 1, a gas turbine engine 10 is illustrated
including a compressor section 12, a combustor 14, and a turbine
section 16. The compressor section 12 compresses ambient air 18
that enters an inlet 20. The combustor 14 combines the compressed
air with a fuel and ignites the mixture creating combustion
products comprising a hot working gas defining a working fluid. The
working fluid travels to the turbine section 16. Within the turbine
section 16 are rows of stationary vanes 22 and rows of rotating
blades 24 coupled to a rotor 26, each pair of rows of vanes 22 and
blades 24 forming a stage in the turbine section 16. The rows of
vanes 22 and rows of blades 24 extend radially into an axial flow
path 28 extending through the turbine section 16. The working fluid
expands through the turbine section 16 and causes the blades 24,
and therefore the rotor 26, to rotate. The rotor 26 extends into
and through the compressor 12 and may provide power to the
compressor 12 and output power to a generator (not shown).
[0019] Referring to FIG. 2, an airfoil structure 30 comprising two
of the vanes of the row of vanes 22 is illustrated for the purpose
of describing aspects of the present invention. However, it should
be understood that the following description is not limited to
implementation on an airfoil structure comprising a vane, and the
described aspects of the invention may be implemented on other
airfoil structures, such as may be implemented on a blade of the
row of blades 24.
[0020] Further, it should be understood that the terms "inner",
"outer", "radial", "axial", "circumferential", and the like, as
used herein, are not intended to be limiting with regard to an
orientation or particular use of the elements recited for aspects
of the present invention.
[0021] The airfoil structure 30 may comprise a vane assembly
including first and second airfoils or vanes 32 adapted to be
supported to extend radially across the flow path 28. The vanes 32
each include a generally concave sidewall 34 defining a pressure
side of the vane 32, and include an opposing generally convex
sidewall 36 defining a suction side of the vane 32. The sidewalls
34, 36 extend radially between an outer diameter endwall 38 and an
inner diameter endwall 40, and extend generally axially in a
chordal direction between a leading edge (not seen in FIGS. 2-5)
and a trailing edge 42 of each of the vanes 32. The endwalls 38, 40
are located at opposing ends of the vanes 32 and are positioned at
locations where they form a boundary, i.e., outer and inner
boundaries, defining a portion of the flow path 28 for the working
fluid.
[0022] The inner endwall 40 includes a gas or hot side 44 facing
radially outwardly toward the flow path 28, and a cool side 46
facing radially inwardly toward the center of the turbine engine
10. The hot side 44 and cool side 46 of the inner endwall 40 extend
circumferentially between opposing lateral sides 48, 50 of the
endwall 40, and the lateral sides 48, 50 extend axially between an
upstream edge 52 and a downstream edge 54 of the inner endwall 40.
As seen in FIGS. 3 and 4, the airfoil structure 30 may further
include an inner rail 56 extending generally circumferentially
along the inner endwall 40 and extending radially inwardly of the
cool side 46 of the inner endwall 40 for retaining the airfoil
structure 30 in position at a radially inner location.
[0023] As seen in FIG. 2, the vanes 32 may be provided with
radially extending cooling channels 60 for providing cooling to the
side walls 34, 36 and the leading and trailing edges (only trailing
edges 42 shown). The trailing edges 42 may be provided with a
plurality of trailing edge cooling slots 62 spaced radially along
the trailing edge 42 to provide convective cooling to the trailing
edge 42.
[0024] It may be noted that, due to migration of hot gases along
the vanes 32 radially inwardly from the radially outer portions
toward the radially inner portions of the vanes 32, joints between
the vanes 32 and the inner endwall 40 defined at fillet portions 64
adjacent to the aft portions of the vanes 32, i.e., adjacent to the
trailing edges 42, experience elevated temperatures. That is, due
to a trailing edge wake effect of the hot gases flowing past the
vanes 32, the temperature of the aft fillet portions 64 may be
substantially greater than temperatures radially outwardly from the
inner endwall 40 and axially forward of the trailing edges 42. It
is normally anticipated that an increased thermal stress will exist
in the region where the trailing edge 42 meets the endwall 40, and
it has generally been the practice to not provide the trailing edge
cooling slots 62 in the areas of the trailing edges 42 closely
adjacent to the endwall 40 in order to provide sufficient material
to withstand the thermal stress. As a result, it has been difficult
to provide effective convective cooling to the region of the
junction between the trailing edge 42 and the endwall 40, i.e., at
a trailing edge corner 66, which may be a further contributing
factor in the formation of thermal stress at this location. In
addition, different convective cooling mechanisms are provided to
the endwall 40 and to the vanes 32, resulting in a differential
cooling of these components which, in combination with a difference
in the mass distribution of the metal forming the trailing edge
corner 66 relative to the thicker or more massive endwall, may
result in a substantial thermally induced strain at the trailing
edge corner 66 during transient thermal cycles.
[0025] In accordance with an aspect of the invention, an overhang
portion 68 of the inner endwall 40 may be configured to reduce
thermal stress at the trailing edge corner 66, such as during
transient thermal cycles. Referring to FIGS. 3 and 4, the overhang
portion 68 generally comprises a portion of the inner endwall 40
that extends axially upstream from the downstream edge 54 toward
the upstream edge 52. In particular, the overhang portion 68
extends between the downstream edge 54 and a location at or
adjacent to the inner rail 56, generally corresponding to the
location of the trailing edge corner 66.
[0026] The overhang portion 68 may be provided with a recess cavity
70 extending from a downstream boundary 72, adjacent to and axially
spaced from the downstream edge 54, toward the upstream edge 52.
The recess cavity 70 may have an upstream boundary 74 extending
circumferentially and located adjacent to the inner rail 56. The
recess cavity 70 may additionally extend circumferentially between
lateral boundaries 76, 78 located adjacent to and circumferentially
spaced from the lateral edges 48, 50. A remaining portion of the
cool side 46 at the overhang portion 68 defines a peripheral
radially inner surface 80 surrounding the recess cavity 70.
[0027] The recess cavity 70 is generally defined by a hollowed out
area formed in the cool side 46 of the inner endwall 40, and
includes a cavity surface 82 spaced radially outwardly from a plane
p.sub.1 (FIG. 5) defined by the peripheral radially inner surface
80. As seen in FIG. 5, the recess cavity 70 extends radially into
the overhang portion 68 a first distance d.sub.1, measured radially
from a peripheral radially inner surface 80 to the cavity surface
82. A plurality of grooves 84 extend radially into the cavity
surface 82, and extend in the axial direction from the downstream
edge 54 toward the upstream edge 52. Each of the grooves 84 include
a groove bottom surface 86 located a second distance d.sub.2
radially into the endwall, as measured from the peripheral radially
inner surface 80, that is greater than the first distance d.sub.1.
It should be noted that peripheral radially inner surface 80, or
the plane p.sub.1 defined by the peripheral radially inner surface
80, may extend at an angle extending radially inwardly from the
downstream edge 54 toward the upstream edge 52, as may be seen in
FIG. 5. Accordingly, the comparative first and second distances
d.sub.1, d.sub.2 referred to herein are distances measured at
substantially the same axial location, and the values of distances
d.sub.1, d.sub.2 may comprise varying distances, i.e., varying with
axial location of the measurement, due to the angle of the cool
side 46 at the overhang portion 68.
[0028] As seen in FIG. 4, the grooves 84 are defined by opposing
radially extending surfaces 88, 90 extending between the cavity
surface 82 and the groove bottom surface 86. The cavity surface 82
and associated pairs of the radially extending surfaces 88, 90
define radially extending raised portions, or ridges 92, within the
recess cavity 70 wherein the ridges 92 and grooves 84 alternate
through the recess cavity 70. The width of the ridges 92 may be
substantially the same as or greater than the width of the grooves
84.
[0029] Referring to FIGS. 2 and 5, cooling passages 94 may extend
through the overhang portion 68. The cooling passages 94 are
illustrated as being located between pairs of the grooves 84, and
extend within the ridges 92. Further, the cooling passages 94 are
located radially between the groove bottom surface 86 and the
cavity surface 82, such that all or a portion of each of the
passages 94 is located radially inwardly beyond the groove bottom
surface 86. The grooves 84 extending into the recess cavity 70 the
distance d.sub.2 provide a substantially thinned mass of material
between the groove bottom surface 86 and the surface of the endwall
40 defined by the hot side 44, as depicted by thickness t.sub.1 in
FIG. 5. The thickness t.sub.1 is preferably substantially constant
along the axial length of the recess cavity 70, as well as at
circumferential locations across the recess cavity 70.
[0030] It should be noted that the ridges 92 provide sufficient
material for defining the cooling passages 94, while the grooves 84
minimize or reduce an amount of material in the recess cavity 70
extending on either side of and surrounding the cooling passages
94. Removal of material of the endwall 40 to form the recess cavity
70 reduces the structural rigidity of the endwall 40, and
particularly reduces the rigidity or structural stiffness of the
overhang portion 68 adjacent to the trailing edge 42 of the vanes
32. Additionally, removal of the material between the cooling
passages 94 to form the grooves 84 further reduces the structural
rigidity of the overhang portion 68. Hence, the mass of material of
the endwall 40 adjacent to the trailing edges 42 of the vanes 32 at
the trailing edge corners 66 is reduced permitting a greater degree
of flexure in the endwall 40, effecting a reduced material strain
at this location. It should also be noted that the reduced mass of
material associated with the cooling passages, i.e., in the area of
the grooves 84 and ridges 92, permits greater cooling effectiveness
from the cooling passages 94 which may reduce the cooling
differential between the convective cooling provided in the
airfoils 32 at the trailing edges 42 and the convective cooling,
provided to the overhang portion 68, additionally effecting the
reduced strain at and/or near the location of the trailing edge
corners 66, such as may occur during thermal cycles during
operation of the gas engine 10.
[0031] An inner diameter endwall post-impingement cooling chamber
96 is located adjacent to the recess cavity 70, and the inner rail
56 is located axially between the recess cavity 70 and the
post-impingement cooling chamber 96, as may be seen in FIGS. 3 and
5. The post-impingement cooling chamber 96 receives cooling air for
cooling a portion of the endwall 40 corresponding to the location
of the post-impingement cooling chamber 96, and for cooling the
overhang portion 68. In the embodiment shown, cooling air enters
the chamber 96 from an inner diameter seal housing 98 (FIG. 5)
located radially inwardly from the endwall 40. At least a portion
of the cooling air provided to the inner diameter seal housing 98
may be provided from the cooling air passing through the radially
extending cooling channels 60 in the vanes 32. The cooling air from
the inner diameter seal housing 98 enters the chamber 96 through
impingement holes 100 formed in one or more impingement plates 102
associated with the post-impingement cooling chamber 96, which
impingement plates 102 define a radially inner boundary for the
post-impingement cooling chamber 96. A portion of the cooling air
passing through the holes 100 impinges on an inner surface 104 of
the endwall 40.
[0032] The cooling passages 94 extend axially through or past the
inner rail 56 to the post-impingement cooling chamber 96. The
cooling air metering through the impingement cooling holes 100,
comprising post-impingement air, may pass into entry openings 106
of the cooling passages 94 and flow through the cooling passages 94
to exit openings 108 (FIGS. 4 and 5) at the downstream edge 54 of
the endwall 40 to provide convective cooling to the overhang
portion 68 of the endwall 40. The cooling air passing out the exit
openings 108 may further provide convective cooling to the surfaces
of the downstream edge 54. In addition, a portion of the
post-impingement cooling air may be provided from the
post-impingement cooling chamber 96 to cooling passages 110 (FIG.
3) at the upstream edge 52 of the endwall and to cooling passages
(not shown) providing cooling to the mate faces at the lateral
edges 48, 50 of the endwall 40.
[0033] It should be understood that, although the recess cavity 70
illustrating aspects of the present invention is shown as a
rectangular cavity extending substantially the axial and
circumferential extent of the overhang portion 68, other
configurations of the recess cavity 70 may be provided. For
example, the overhang portion 68 may be provided with one or more
recess cavities configured or shaped to address particular
structural rigidity and cooling requirements associated with a
specific airfoil structure 30.
[0034] While particular embodiments of the present invention have
been illustrated and described, it would be obvious to those
skilled in the art that various other changes and modifications can
be made without departing from the spirit and scope of the
invention. It is therefore intended to cover in the appended claims
all such changes and modifications that are within the scope of
this invention.
* * * * *