U.S. patent application number 10/855049 was filed with the patent office on 2005-12-01 for cooled rotor blade.
This patent application is currently assigned to United Technologies Corporation. Invention is credited to Gregg, Shawn J., Mongillo, Dominic J. JR..
Application Number | 20050265839 10/855049 |
Document ID | / |
Family ID | 34941471 |
Filed Date | 2005-12-01 |
United States Patent
Application |
20050265839 |
Kind Code |
A1 |
Mongillo, Dominic J. JR. ;
et al. |
December 1, 2005 |
Cooled rotor blade
Abstract
A rotor blade is provided that includes a root and a hollow
airfoil. The hollow airfoil has a cavity, a leading edge, and a
tip. An internal passage configuration is disposed within the
cavity that includes a first radial passage, a second radial
passage, and a rib disposed between the passages. The passages and
the rib are contiguous with a tip endwall. The first radial passage
is disposed contiguous with the leading edge. A plurality of
crossover apertures are disposed in the rib. One of the crossover
apertures is disposed flush with the tip endwall. A conduit is
disposed within the root that is operable to permit airflow through
the root and into the passages. In some embodiments, an aperture is
disposed within the tip endwall aligned with the first radial
passage.
Inventors: |
Mongillo, Dominic J. JR.;
(West Hartford, CT) ; Gregg, Shawn J.;
(Wethersfield, CT) |
Correspondence
Address: |
MCCORMICK, PAULDING & HUBER LLP
CITY PLACE II
185 ASYLUM STREET
HARTFORD
CT
06103
US
|
Assignee: |
United Technologies
Corporation
Hartford
CT
|
Family ID: |
34941471 |
Appl. No.: |
10/855049 |
Filed: |
May 27, 2004 |
Current U.S.
Class: |
416/97R ;
416/96R |
Current CPC
Class: |
F01D 5/20 20130101; F01D
5/187 20130101; F05D 2260/202 20130101 |
Class at
Publication: |
416/097.00R ;
416/096.00R |
International
Class: |
B63H 001/14 |
Claims
What is claimed is:
1. A rotor blade, comprising: a root; a hollow airfoil having a
cavity, a leading edge, and a tip; an internal passage
configuration disposed within the cavity, which configuration
includes a first radial passage, a second radial passage, and a rib
disposed between the passages, wherein the passages and the rib are
contiguous with a tip endwall, and the first radial passage is
disposed contiguous with the leading edge, and a plurality of
crossover apertures are disposed in the rib, and one of the
crossover apertures is disposed flush with the tip endwall; and a
conduit disposed within the root that is operable to permit airflow
through the root and into the passages.
2. The rotor blade of claim 1, wherein the first radial passage is
a cavity.
3. The rotor blade of claim 1, wherein the first radial passage
extends down and opens to the conduit.
4. The rotor blade of claim 1, further comprising at least one
aperture disposed in the endwall, aligned with the first radial
passage.
5. The rotor blade of claim 4, further comprising a tip pocket
disposed in the tip of the rotor blade.
6. The rotor blade of claim 5, wherein the aperture disposed in the
endwall aligned with the first radial passage extends between the
first radial passage and the tip pocket.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Technical Field
[0002] This invention applies to gas turbine rotor blades in
general, and to cooled gas turbine rotor blades in particular.
[0003] 2. Background Information
[0004] Turbine sections within an axial flow turbine engine include
rotor assemblies that each include a rotating disc and a number of
rotor blades circumferentially disposed around the disk. Rotor
blades include an airfoil portion for positioning within the gas
path through the engine. Because the temperatures within the gas
path very often negatively affect the durability of the airfoil, it
is known to cool an airfoil by passing cooling air through the
airfoil. The cooled air helps decrease the temperature of the
airfoil material and thereby increase its durability.
[0005] Prior art cooled rotor blades very often utilize internal
passage configurations that include a leading edge passage 10a that
either dead-ends adjacent the tip (see FIG. 7), or is connected to
an axially extending passage that dead-ends prior to the trailing
edge. All of these internal passage configurations suffer from
airflow stagnation regions, or regions of relatively low velocity
flow that inhibit internal convective cooling. The airfoil wall
regions adjacent these regions of low cooling effectiveness are
typically at a higher temperature than other regions of the
airfoil, and are therefore more prone to undesirable oxidation,
thermal mechanical fatigue (TMF), creep, and erosion.
[0006] What is needed, therefore, is an airfoil having an internal
passage configuration that promotes desirable cooling of the
airfoil and thereby increases the durability of the blade.
DISCLOSURE OF THE INVENTION
[0007] According to the present invention, a rotor blade is
provided that includes a root and a hollow airfoil. The hollow
airfoil has a cavity, a leading edge, and a tip. An internal
passage configuration is disposed within the cavity that includes a
first radial passage, a second radial passage, and a rib disposed
between the passages. The passages and the rib are contiguous with
a tip endwall. The first radial passage is disposed contiguous with
the leading edge. A plurality of crossover apertures are disposed
in the rib. One of the crossover apertures is disposed flush with
the tip endwall. A conduit is disposed within the root that is
operable to permit airflow through the root and into the
passages.
[0008] In some embodiments, an aperture is disposed within the tip
endwall aligned with the first radial passage.
[0009] One of the advantages of the present rotor blade and method
is that airflow stagnation regions within the radial passages are
decreased or eliminated adjacent the tip. Another advantage is that
the convective cooling of the tip endwall is improved relative to
many prior art internal passage configurations. The crossover
aperture disposed flush with the tip endwall permits cooling air to
travel into the radial end of the leading edge radial passage. As a
result, undesirable stagnation regions 12a (see FIG. 7) typically
present in prior art dead-end radial passages 14a are eliminated.
For the case where the cavity supplying the coolant air to the
flush crossover does not dead-end, but feeds another cavity (e.g.,
axial extending cavity or aftward flowing serpentine), the
stagnation/recirculation region 16a (see FIG. 7) typical in prior
art is also eliminated. In addition, the cooling airflow through
flush crossover aperture improves convective cooling of the tip
endwall. The airfoil tip is consequently able to accommodate high
temperature environments with greater resistance to oxidation, TMF,
creep, and erosion.
[0010] Additional advantages are provided in the embodiment of the
present invention where an aperture is disposed within the tip
endwall aligned with the first radial passage, contiguous with the
leading edge. The aperture provides a cooling air path out of the
first radial passage, which facilitates the elimination of
stagnation regions within the first radial passage. The position of
the aperture relative to the first radial passage, leading edge,
and tip also enables it to cool a region of the airfoil where
cooling has historically been problematic. The position of the
aperture at the radial end of the first radial passage also enables
it to act as a debris purge. Debris that is carried within the
cooling air or dislodged from a surface within the airfoil will be
forced outward by centrifugal forces as the blade rotates. The
aperture at the radial end of the first radial passage is
positioned to receive and pass debris outside the airfoil.
[0011] These and other objects, features and advantages of the
present invention will become apparent in light of the detailed
description of the best mode embodiment thereof, as illustrated in
the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] FIG. 1 is a diagrammatic perspective view of the rotor
assembly section.
[0013] FIG. 2 is a diagrammatic sectional view of a rotor blade
having an embodiment of the internal passage configuration.
[0014] FIG. 3 is a diagrammatic sectional view of a rotor blade
having an embodiment of the internal passage configuration.
[0015] FIG. 4 is a diagrammatic sectional view of a rotor blade
having an embodiment of the internal passage configuration.
[0016] FIG. 5 is a diagrammatic sectional view of a rotor blade
having an embodiment of the internal passage configuration.
[0017] FIG. 6 is an enlarged diagrammatic sectional view of part of
a rotor blade having an embodiment of the internal passage
configuration.
[0018] FIG. 7 is an enlarged diagrammatic sectional view of part of
a prior art rotor blade.
DETAILED DESCRIPTION OF THE INVENTION
[0019] Referring to FIG. 1, a rotor blade assembly 10 for a gas
turbine engine is provided having a disk 12 and a plurality of
rotor blades 14. The disk 12 includes a plurality of recesses 16
circumferentially disposed around the disk 12 and a rotational
centerline 18 about which the disk 12 may rotate. Each blade 14
includes a root 20, an airfoil 22, a platform 24, and a radial
centerline 25. The root 20 includes a geometry (e.g., a fir tree
configuration) that mates with that of one of the recesses 16
within the disk 12. As can be seen in FIGS. 2-5, the root 20
further includes conduits 26 through which cooling air may enter
the root 20 and pass through into the airfoil 22.
[0020] Referring to FIGS. 1-5, the airfoil 22 includes a base 28, a
tip 30, a leading edge 32, a trailing edge 34, a pressure side wall
36 (see FIG. 1), and a suction side wall 38 (see FIG. 1), and an
internal passage configuration 40. FIGS. 2-5 diagrammatically
illustrate an airfoil 22 sectioned between the leading edge 32 and
the trailing edge 34. The pressure side wall 36 and the suction
side wall 38 extend between the base 28 and the tip 30 and meet at
the leading edge 32 and the trailing edge 34.
[0021] The internal passage configuration 40 includes a first
conduit 42, a second conduit 44, and a third conduit 46 extending
through the root 20 into the airfoil 22. The first conduit 42 is in
fluid communication with one or more leading edge passages 48 ("LE
passages") disposed adjacent the leading edge 32. The first conduit
42 provides the primary path into these LE passages 48 for cooling
air, and therefore the leading edge 32 is primarily cooled by the
cooling air that enters the airfoil 22 through the first conduit
42.
[0022] Referring to FIG. 2, in a first embodiment of the one or
more LE passages 48, the first conduit 42 is in fluid communication
with a single LE passage 50, and that passage 50 is contiguous with
the leading edge 32. At the outer radial end of the LE passage 50
(i.e., the end of the LE passage 50 opposite the first conduit 42),
the LE passage 50 is connected to an axially extending passage 52
("AE passage") that extends between the LE passage 50 and the
trailing edge 34 of the airfoil 22, adjacent the tip 30 of the
airfoil 22. As can be seen from FIG. 2, the cross-sectional area
within the transition between the passages 50,52 is approximately
the same as or greater than the adjacent regions of the passages
50,52. Hence, there is no flow impediment within the transition
that is attributable to a decrease in cross-sectional area. The LE
passage 50 is connected to the exterior of the airfoil 22 by a
plurality of cooling apertures 54 disposed along the leading edge
32.
[0023] Referring to FIG. 3, in a second embodiment of the one or
more LE passages 48, the first conduit 42 is in fluid communication
with a first LE passage 56 and a second LE passage 58. The first LE
passage 56 is contiguous with the leading edge 32, and the second
LE passage 58 is immediately aft and adjacent the first LE passage
56. The first LE passage 56 is connected to the exterior of the
airfoil 22 by a plurality of cooling apertures 54 disposed along
the leading edge 32. In some embodiments, the first LE passage 56
is also connected to the tip 30 or a tip pocket 60 by one or more
apertures 62. At the outer radial end of the second LE passage 58
(i.e., the end of the second LE passage 58 opposite the first
conduit 42), the second LE passage 58 is connected to an AE passage
52 that extends to the trailing edge 34 of the airfoil 22, adjacent
the tip 30 of the airfoil 22. As can be seen from FIG. 3, the
cross-sectional area within the transition between the passages
58,52 is approximately the same as or greater than the adjacent
regions of the passages 58,52. Hence, there is no flow impediment
within the transition that is attributable to a decrease in
cross-sectional area.
[0024] Referring to FIG. 4, in a third embodiment of the one or
more LE passages 48, the first conduit 42 is in fluid communication
with a first LE passage 64 and a second LE passage 66. The first LE
passage 64 is contiguous with the leading edge 32, and the second
LE passage 66 is immediately aft and adjacent the first LE passage
64. The first LE passage 64 is connected to the exterior of the
airfoil 22 by a plurality of cooling apertures 54 disposed along
the leading edge 32. At the outer radial end of the first LE
passage 64 (i.e., the end of the first LE passage 64 opposite the
first conduit 42), the first LE passage 64 is connected to an AE
passage 52 that extends to the trailing edge 34 of the airfoil 22,
adjacent the tip 30 of the airfoil 22. As can be seen from FIG. 4,
the cross-sectional area within the transition between the passages
64,52 is approximately the same as or greater than the adjacent
regions of the passages 64,52. Hence, there is no flow impediment
within the transition that is attributable to a decrease in
cross-sectional area. The second LE passage 66 ends radially below
the AE passage 52. One or more apertures 68 disposed in the rib
between the AE passage 52 and the second LE passage 66 permits
airflow therebetween.
[0025] Referring to FIG. 5, in a fourth embodiment of the one or
more LE passages 48, the first conduit 42 is in fluid communication
with a single LE passage 70. One or more cavities 72 are disposed
forward of the LE passage 70, connected to the LE passage 70 by a
plurality of crossover apertures 74. The one or more cavities 72
are contiguous with the leading edge 32. The one or more cavities
72 are connected to the exterior of the airfoil 22 by a plurality
of cooling apertures 54 disposed along the leading edge 32. In some
embodiments, the cavity 72 (or the outer most radial cavity if more
than one cavity) is also connected to the tip 30 or a tip pocket 60
by one or more apertures 76. At the outer radial end of the LE
passage 70 (i.e., the end of the LE passage 70 opposite the first
conduit 42), the LE passage 70 is connected to an AE passage 52
that extends to the trailing edge 34 of the airfoil 22, adjacent
the tip 30 of the airfoil 22. As can be seen from FIG. 5, the
cross-sectional area within the transition between the passages
70,52 is approximately the same as or greater than the adjacent
regions of the passages 70,52. Hence, there is no flow impediment
within the transition that is attributable to a decrease in
cross-sectional area.
[0026] Referring to FIG. 6, in a preferred embodiment of an
internal passage configuration like those shown in FIGS. 3 and 5,
the internal passage configuration 40 includes a first radial
passage 92 (e.g., first LE passage 56--FIG. 3; cavity 72--FIG. 5),
a second radial passage 94 (e.g., second LE passage 58--FIG. 3; LE
passage 70--FIG. 5), and a rib 96 disposed therebetween. The first
radial passage 92, second radial passage 94, and rib 96 are
contiguous with a tip endwall 98. A plurality of crossover
apertures 74 are disposed in the rib 96, including a crossover
aperture 100 that is disposed flush with the tip endwall 98. As
stated above, in some applications an aperture 62,76 is disposed in
the radial end of the first radial passage. This preferred
embodiment of an internal passage configuration is not limited to
the internal passage configurations shown in FIGS. 3 and 5.
[0027] Referring to FIGS. 2-5, the second conduit 44 is in fluid
communication with a serpentine passage 78 disposed immediately aft
of the LE passages, in the mid-body region of the airfoil 22. The
second conduit 44 provides the primary path into the serpentine
passage 78 for cooling air, and therefore the mid-body region is
primarily cooled by the cooling air that enters the airfoil 22
through the second conduit 44. The serpentine passage 78 has an odd
number of radial segments 80, which number is greater than one;
e.g., 3, 5, etc. The odd number of radial segments 80 ensures that
the last radial segment 82 in the serpentine 78 ends adjacent the
AE passage 52. The "last radial segment" is defined as the last
possible segment within the serpentine passage that can receive
cooling air along the serpentine. The radial segments 80 are
connected to one another by turns of approximately 180.degree.;
e.g., the first radial segment is connected to the second radial
segment by a 180.degree. turn, the second radial segment is
connected to the third radial segment by a 180.degree. turn, etc.
The serpentine passage 78 shown in FIGS. 2-5 is oriented so that
the path through the serpentine 78 directs the cooling air forward;
i.e., toward the leading edge 32 of the airfoil 22. In alternative
embodiments, the serpentine 78 can also be oriented so that cooling
air is directed aft, toward the trailing edge 34 of the airfoil 22.
In some embodiments, a cooling air sink 84, typically in the form
of one or more cooling apertures, is disposed within the exterior
wall (e.g., the suction side wall) of the last segment 82, sized to
permit cooling airflow out of the airfoil 22. In a preferred
embodiment, the one or more cooling apertures are film holes. One
or more apertures 85 extend through the rib separating the last
radial segment 82 and the AE passage, thereby permitting fluid
communication therebetween.
[0028] The third conduit 46 is in fluid communication with one or
more passages 86 disposed between the serpentine passage 78 and the
trailing edge 34 of the airfoil 22. With the exception of portion
of the trailing edge 34 adjacent the tip 30 of the airfoil 22, the
third conduit 46 provides the primary path for cooling air into the
trailing edge 34, and therefore the trailing edge 34 is primarily
cooled by the cooling air that enters the airfoil 22 through the
third conduit 46. As stated above, the portion of the trailing edge
34 adjacent the tip 30 of the airfoil 22 is cooled by cooling air
passing through the AE passage 52.
[0029] The AE passage 52 includes a tapered segment 88 adjacent the
trailing edge 34 that decrease in cross-sectional area. The rate of
decrease in cross-sectional area is chosen to cause the cooling
airflow exiting the AE passage 52 to choke. The specific rate of
decrease in cross-sectional area is chosen to suit the application
at hand.
[0030] In the embodiments shown in FIGS. 2-5, the transition
between the LE passage(s) and the AE passage 52 is approximately a
ninety degree (90.degree.) turn that has been optimized to minimize
pressure loss as cooling air travels between the LE passage(s) and
the AE passage 52. For example, the LE passage 50,58,64,70
increases in width as it approaches the turn. As a result, the
interior boundary 90 of the turn forms an angle that is greater
than 90.degree.. The obtuse angle facilitates the cooling airflow
therethrough, and consequently causes a pressure loss which is less
than would be in a similar channel having a 90.degree. turn.
[0031] All of the foresaid passages may include one or more cooling
apertures and/or cooling features (e.g., trip strips, pedestals,
pin fins, etc.) to facilitate heat transfer within the particular
passage. The exact type(s) of cooling aperture and/or cooling
feature can vary depending on the application, and more than one
type can be used. The present invention can be used with a variety
of different cooling aperture and cooling feature types and is not,
therefore, limited to any particular type.
[0032] Some embodiments further include a tip pocket 60 disposed
radially outside of the AE passage 52. The tip pocket 60 is open to
the exterior of the airfoil 22. One or more apertures extend
through a wall portion of the airfoil 22 disposed between the tip
pocket 60 and the LE passage and/or the AE passage 52.
[0033] The above-described rotor blade 14 can be manufactured using
a casting process that utilizes a ceramic core to form the cooling
passages within the airfoil 22. The ceramic core is advantageous in
that it is possible to create very small details within the
passages; e.g., cooling apertures, trip strips, etc. A person of
skill in the art will recognize, however, that the brittleness of a
ceramic core makes it is difficult to use. The above-described
rotor blade internal passage configurations 40 facilitate the
casting process by including features that increase the durability
of the ceramic core. For example, the first and second LE passage
embodiments permit the use of a rod extending from the tip pocket
60, through the AE passage 52, and into the serpentine passage 78.
The rod supports: 1) the core portion that forms the tip pocket 60;
2) the core portion that forms the AE passage 52; and 3) the core
portion that forms the serpentine passage 78. The rod is removed at
the same time the ceramic core is removed, leaving apertures
between the tip pocket 60 and the AE passage 52, and between the AE
passage 52 and the serpentine passage 78. Core-ties can also be
used between core portions.
[0034] Another feature of the present internal passage
configurations that increases the durability of the ceramic core is
the AE passage 52 adjacent the tip 30 of the airfoil 22. The
extension of the passage 52 to the trailing edge 34 enables the
passage 52 and the trailing edge 34 core portion to be tied
together by a stringer that is disposed outside the exterior of the
airfoil 22. The core portions representing internal cooling
passages (e.g., one of more segments of the serpentine passage 78)
may also be supported by the AE passage 52 via rods or
core-ties.
[0035] In the operation of the invention, the airfoil 22 portion of
the rotor blade 14 is disposed within the core gas path of the
turbine engine. The airfoil 22 is subject to high temperature core
gas passing by the airfoil 22. Cooling air, that is substantially
lower in temperature than the core gas, is fed into the airfoil 22
through the conduits 42,44,46 disposed in the root 20.
[0036] Cooling air traveling through the first conduit 42 passes
directly into the one or more LE passages 48 disposed adjacent the
leading edge 32, and subsequently into the AE passage 52 adjacent
the tip 30 of the airfoil 22. The relatively large and unobstructed
LE passages 48 permit a volume rate of flow that provides a
desirable amount of cooling to the leading edge 32, and yet still
has sufficient heat transfer capacity to adequately cool other
regions of the airfoil 22; e.g., the tip 30 and a portion of the
serpentine passage 78. The first conduit 42 provides the primary
path into these LE passages 48 for cooling air, although the exact
path depends upon the particular LE passage 48 embodiment.
[0037] Cooling air traveling through the first conduit 42 into the
first embodiment of the one or more LE passages 48 incurs
relatively low pressure losses, and will enter the AE passage 52 at
a relatively high pressure and velocity. Because the first
embodiment of the one or more LE passages 48 is a single passage 50
contiguous with the leading edge 32, the cooling air is subject to
heat transfer from the leading edge 32, the pressure side wall 36,
and the suction side wall 38. In this embodiment, the AE passage 52
extends across the entire chord of the airfoil 22.
[0038] Cooling air traveling through the first conduit 42 into the
second embodiment of the one or more LE passages 48 is divided
between the first LE passage 56 and the second LE passage 58. The
cooling air entering the first LE passage 56 travels contiguous
with the leading edge 32, and is subject to heat transfer from the
leading edge 32, the pressure side wall 36, and the suction side
wall 38. The cooling air traveling within the first LE passage 56
exits via cooling apertures 54 disposed along the radial length of
the leading edge 32, and through one or more cooling apertures 62
disposed between the radial end of the passage 56 and the tip 30
(or tip pocket 60). The apertures 62 disposed at the radial end
prevent cooling airflow stagnation within the first LE passage 56.
Cooling air traveling within the second LE passage 58 incurs
relatively low pressure losses, and will enter the AE passage 52 at
a relatively high pressure and velocity. Because the second LE
passage 58 is aft of the first LE passage 56 (and therefore the
leading edge 32), the cooling air traveling through the second LE
passage 58 is subject to less heat transfer from the leading edge
32. As a result, the cooling air reaches the AE passage 52
typically at a lower temperature than it would be if it were in
contact with the leading edge 32. In this embodiment, the AE
passage 52 extends across nearly the entire chord of the airfoil
22.
[0039] Cooling air traveling through the first conduit 42 into the
third embodiment of the one or more LE passages 48 is divided
between the first LE passage 64 and the second LE passage 66. The
cooling air entering the first LE passage 64 incurs relatively low
pressure losses, and will enter the AE passage 52 at a relatively
high pressure and velocity. The cooling air entering the second LE
passage 66 will likewise flow substantially unobstructed until the
radial end is reached. Cooling air can exit the second LE passage
66 through one or more cooling apertures 68 disposed in the rib
separating the second LE passage 66 and the AE passage 52, or
through cooling apertures disposed within the walls of the airfoil
22. The apertures 68 disposed at the radial end prevent cooling
airflow stagnation within the second LE passage 66. In this
embodiment, the AE passage 52 extends across the entire chord of
the airfoil 22.
[0040] Cooling air traveling through the first conduit 42 into the
fourth embodiment of the one or more LE passages 48 incurs
relatively low pressure losses, and will enter the AE passage 52 at
a relatively high pressure and velocity. A portion of the cooling
air traveling within the LE passage 48 enters the cavity(ies) 72
disposed between the LE passage 70 and the leading edge 32. The
cooling air traveling within the cavity 72 exits via cooling
apertures 54 disposed along the radial length of the leading edge
32, and through one or more cooling apertures 76 disposed between
the radial end of the cavity 72 and the tip 30 (or tip pocket 60).
The apertures 76 disposed at the radial end prevent cooling airflow
stagnation within the cavity 72. Because the LE passage 70 is aft
of cavity(ies) 72 (and therefore the leading edge 32), the cooling
air traveling through the LE passage 70 is subject to less heat
transfer from the leading edge 32. As a result, the cooling air
reaches the AE passage 52 typically at a lower temperature than it
would be if it were in contact with the leading edge 32.
[0041] In the preferred embodiment of the internal passage
configuration shown in FIG. 6, a portion of the cooling air
traveling within the second radial passage 94 (e.g., second LE
passage 58--FIG. 3; LE passage 70--FIG. 5) exits the second radial
passage 94 and enters the first radial passage 92 (e.g., first LE
passage 56--FIG. 3; cavity 72--FIG. 5) via the crossover apertures
74 disposed in the rib 96. Cooling air traveling through the
flush-mounted crossover aperture 100 passes along the surface of
the tip endwall 98, providing desirable convective cooling. The
cooling air entering the first radial passage 92 through the
flush-mounted crossover aperture 100 helps to eliminate a
stagnation/recirculation zone within the first radial passage 92
adjacent the tip endwall 98 (see flow 12a, FIG. 7) and within the
second radial passage 94 adjacent the tip endwall 98 (see flow 16a,
FIG. 7). In those applications wherein an aperture 62,76 is
disposed in the radial end of the first radial passage 92, cooling
air exiting via the aperture 62, 76 also facilitates the
elimination of undesirable stagnation/recirculation zones within
the first radial passage 92; and the aperture 62, 76 reduces the
risk of plugging the region of the passage adjacent to the tip
endwall 98 by providing foreign particles (e.g., dirt) a path
through which to exit the airfoil. Additionally, these applications
increase local heat transfer adjacent the leading edge portion 102
adjacent the tip 30 typically prone to distress (e.g., oxidation)
in prior art.
[0042] In all of the above embodiments, a portion of the cooling
air passing through the AE passage 52 typically exits the AE
passage 52 via cooling apertures; e.g., the cooling apertures
extending between the tip 30 and/or tip cavity and the AE passages
52. An advantage provided by the present internal passage
configuration, and in particular by the AE passage 52 extending the
length or nearly the length of the chord, is that manufacturability
of the airfoil 22 is increased since cooling apertures can be
drilled through the tip 30 without interference from ribs
separating radial segments.
[0043] Cooling air traveling through the second conduit 44 enters
the serpentine passage 78 at P.sub.1. The cooling air passes
through each radial segment 80 and 180.degree. turn. A portion of
the cooling air that enters the passage 78, exits the passage 78
via cooling apertures disposed in the walls of the airfoil 22. The
remainder of the cooling air that enters the serpentine passage 78
will enter the last radial segment 82 of the passage 78. With the
present internal passage configurations, the cooling air that
reaches the last radial segment 82 will typically be at a pressure
P.sub.3 that is lower than the pressure P.sub.2 of the cooling air
in the adjacent region of the AE passage 52 (e.g., because of head
losses incurred within the serpentine passage 78), wherein
P.sub.1>P.sub.2>P.sub.3. In those instances, cooling air will
enter the last radial segment 82 from the AE passage 52 via the one
or more apertures 85 extending between the last radial segment 82
and the AE passage 52 (P.sub.2>P.sub.3). To accommodate the
inflow from the AE passage 52, a cooling air sink 84 (e.g., film
holes) is disposed within the exterior wall of the last segment
(e.g., the suction side wall 38), sized to permit cooling airflow
out of the airfoil 22. The cooling air sink 84 prevents undesirable
flow stagnation within the last radial segment 82 of the serpentine
passage 78. The two opposing flows of cooling air within the
serpentine passage 78 will come to rest at a location where the
static pressure of each flow equals that of the other. Preferably,
the cooling air sink 84 is positioned adjacent that rest location.
The pressure P.sub.1 of the cooling air entering the serpentine
passage 78 prevents the AE passage 52 inflow from traveling
completely through the serpentine passage 78
(P.sub.1>P.sub.2).
[0044] Cooling air traveling through the third conduit 46 enters
one or more passage(s) 86 disposed between the serpentine passage
78 and the trailing edge 34. All of the cooling air that enters
these passages exits via cooling apertures disposed in the walls of
the airfoil 22 or along the trailing edge 34.
[0045] Although this invention has been shown and described with
respect to the detailed embodiments thereof, it will be understood
by those skilled in the art that various changes in form and detail
thereof may be made without departing from the spirit and the scope
of the invention.
* * * * *