U.S. patent application number 16/552450 was filed with the patent office on 2021-03-18 for airfoil having impingement leading edge.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Carey Clum, Timothy J. Jennings, Tracy A. Propheter-Hinckley.
Application Number | 20210079793 16/552450 |
Document ID | / |
Family ID | 1000005021736 |
Filed Date | 2021-03-18 |
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United States Patent
Application |
20210079793 |
Kind Code |
A1 |
Propheter-Hinckley; Tracy A. ;
et al. |
March 18, 2021 |
AIRFOIL HAVING IMPINGEMENT LEADING EDGE
Abstract
Airfoils for a gas turbine engines and gas turbine engines are
described. The airfoils include an airfoil body extending in a
radial direction from a first end to a second end, and axially from
a leading edge to a trailing edge. A radially extending leading
edge channel is formed in the leading edge of the airfoil body,
having first and second channel walls that join at a channel base.
A first leading edge impingement cavity is located within the
airfoil body proximate the leading edge and is defined, in part, by
the first channel wall. A leading edge feed cavity is arranged aft
of the first leading edge impingement cavity to supply air into the
first leading edge impingement cavity. A first leading edge
impingement hole is formed in the first channel wall and angled
toward a portion of the second channel wall.
Inventors: |
Propheter-Hinckley; Tracy A.;
(Manchester, CT) ; Clum; Carey; (East Hartford,
CT) ; Jennings; Timothy J.; (West Hartford,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Farmington |
CT |
US |
|
|
Family ID: |
1000005021736 |
Appl. No.: |
16/552450 |
Filed: |
August 27, 2019 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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15791464 |
Oct 24, 2017 |
10584593 |
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16552450 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2240/303 20130101;
F05D 2220/32 20130101; F05D 2260/201 20130101; F01D 5/186 20130101;
F01D 5/141 20130101; F01D 5/187 20130101; F05D 2260/202 20130101;
F05D 2240/121 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 5/14 20060101 F01D005/14 |
Claims
1. An airfoil for a gas turbine engine, the airfoil comprising: an
airfoil body extending in a radial direction from a first end to a
second end, and extending axially from a leading edge to a trailing
edge; a leading edge channel formed in the leading edge of the
airfoil body, the leading edge channel having a first channel wall
and a second channel wall that join at a channel base to define the
leading edge channel, the leading edge channel extending in a
radial direction along the leading edge of the airfoil body; a
first leading edge impingement cavity located within the airfoil
body proximate the leading edge, wherein the first channel wall
forms a portion of the airfoil body defining the first leading edge
impingement cavity; a leading edge feed cavity arranged aft of the
first leading edge impingement cavity and arranged to supply air
into the first leading edge impingement cavity; and a first leading
edge impingement hole formed in the first channel wall and angled
such that air flowing from the first leading edge impingement
cavity and through the first leading edge impingement hole impinges
upon a portion of the second channel wall.
2. The airfoil of claim 1, wherein the first end is a root of the
airfoil body and the second end is a tip of the airfoil, wherein
the leading edge channel extends less than a full length of a
distance between the root and the tip along the leading edge of the
airfoil body.
3. The airfoil of claim 1, further comprising a second leading edge
impingement cavity located within the airfoil body proximate the
leading edge, wherein the second channel wall forms a portion of
the airfoil body defining the second leading edge impingement
cavity.
4. The airfoil of claim 3, further comprising a second leading edge
impingement hole formed in the second channel wall and angled such
that air flowing from the second leading edge impingement cavity
and through the second leading edge impingement hole impinges upon
a portion of the first channel wall.
5. The airfoil of claim 4, wherein the leading edge channel
comprises a plurality of first leading edge impingement holes
formed in the first channel wall, wherein the plurality of first
leading edge impingement holes extend in an array radially along
the first channel wall, and wherein the leading edge channel
comprises a plurality of second leading edge impingement holes
formed in the second channel wall, wherein the plurality of second
leading edge impingement holes extend in an array radially along
the second channel wall.
6. The airfoil of claim 3, wherein the leading edge feed cavity is
arranged aft of the second leading edge impingement cavity and
arranged to supply air into the second leading edge impingement
cavity.
7. The airfoil of claim 1, wherein the leading edge channel had a
depth in an axial direction that is at least twice a diameter of
the first leading edge impingement hole.
8. The airfoil of claim 1, wherein the leading edge channel
comprises a plurality of first leading edge impingement holes
formed in the first channel wall, wherein the plurality of first
leading edge impingement holes extend in an array radially along
the first channel wall.
9. The airfoil of claim 1, further comprising at least one
additional channel formed on an exterior surface of the airfoil and
extending radially.
10. The airfoil of claim 1, wherein the airfoil body forms a blade
of a turbine section of the gas turbine engine.
11. A gas turbine engine comprising: an airfoil having an airfoil
body extending in a radial direction from a first end to a second
end, and extending axially from a leading edge to a trailing edge;
a leading edge channel formed in the leading edge of the airfoil
body, the leading edge channel having a first channel wall and a
second channel wall that join at a channel base to define the
leading edge channel, the leading edge channel extending in a
radial direction along the leading edge of the airfoil body; a
first leading edge impingement cavity located within the airfoil
body proximate the leading edge, wherein the first channel wall
forms a portion of the airfoil body defining the first leading edge
impingement cavity; a leading edge feed cavity arranged aft of the
first leading edge impingement cavity and arranged to supply air
into the first leading edge impingement cavity; and a first leading
edge impingement hole formed in the first channel wall and angled
such that air flowing from the first leading edge impingement
cavity and through the first leading edge impingement hole impinges
upon a portion of the second channel wall.
12. The gas turbine engine of claim 11, wherein the first end is a
root of the airfoil body and the second end is a tip of the
airfoil, wherein the leading edge channel extends less than a full
length of a distance between the root and the tip along the leading
edge of the airfoil body.
13. The gas turbine engine of claim 11, further comprising a second
leading edge impingement cavity located within the airfoil body
proximate the leading edge, wherein the second channel wall forms a
portion of the airfoil body defining the second leading edge
impingement cavity.
14. The gas turbine engine of claim 13, further comprising a second
leading edge impingement hole formed in the second channel wall and
angled such that air flowing from the second leading edge
impingement cavity and through the second leading edge impingement
hole impinges upon a portion of the first channel wall.
15. The gas turbine engine of claim 14, wherein the leading edge
channel comprises a plurality of first leading edge impingement
holes formed in the first channel wall, wherein the plurality of
first leading edge impingement holes extend in an array radially
along the first channel wall, and wherein the leading edge channel
comprises a plurality of second leading edge impingement holes
formed in the second channel wall, wherein the plurality of second
leading edge impingement holes extend in an array radially along
the second channel wall.
16. The gas turbine engine of claim 13, wherein the leading edge
feed cavity is arranged aft of the second leading edge impingement
cavity and arranged to supply air into the second leading edge
impingement cavity.
17. The gas turbine engine of claim 11, wherein the leading edge
channel had a depth in an axial direction that is at least twice a
diameter of the first leading edge impingement hole.
18. The gas turbine engine of claim 11, wherein the leading edge
channel comprises a plurality of first leading edge impingement
holes formed in the first channel wall, wherein the plurality of
first leading edge impingement holes extend in an array radially
along the first channel wall.
19. The gas turbine engine of claim 11, further comprising at least
one additional channel formed on an exterior surface of the airfoil
and extending radially.
20. The gas turbine engine of claim 11, wherein the airfoil is a
blade of a turbine section of the gas turbine engine.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation application of U.S.
application Ser. No. 15/791,464 filed Oct. 24, 2017, the contents
of which are incorporated by reference herein in their
entirety.
BACKGROUND
[0002] Illustrative embodiments pertain to the art of
turbomachinery, and specifically to turbine rotor components.
[0003] Gas turbine engines are rotary-type combustion turbine
engines built around a power core made up of a compressor,
combustor and turbine, arranged in flow series with an upstream
inlet and downstream exhaust. The compressor compresses air from
the inlet, which is mixed with fuel in the combustor and ignited to
generate hot combustion gas. The turbine extracts energy from the
expanding combustion gas, and drives the compressor via a common
shaft. Energy is delivered in the form of rotational energy in the
shaft, reactive thrust from the exhaust, or both.
[0004] The individual compressor and turbine sections in each spool
are subdivided into a number of stages, which are formed of
alternating rows of rotor blade and stator vane airfoils. The
airfoils are shaped to turn, accelerate and compress the working
fluid flow, or to generate lift for conversion to rotational energy
in the turbine.
[0005] Airfoils may incorporate various cooling cavities located
adjacent external side walls. Such cooling cavities are subject to
both hot material walls (exterior or external) and cold material
walls (interior or internal). Although such cavities are designed
for cooling portions of airfoil bodies, various cooling flow
characteristics can cause hot sections where cooling may not be
sufficient. Accordingly, improved means for providing cooling for
an airfoil may be desirable.
BRIEF DESCRIPTION
[0006] According to some embodiments, airfoils for gas turbine
engines are provided. The airfoils have an airfoil body extending
in a radial direction from a first end to a second end, and
extending axially from a leading edge to a trailing edge, a leading
edge channel formed in the leading edge of the airfoil body, the
leading edge channel having a first channel wall and a second
channel wall that join at a channel base to define the leading edge
channel, the leading edge channel extending in a radial direction
along the leading edge of the airfoil body, a first leading edge
impingement cavity located within the airfoil body proximate the
leading edge, wherein the first channel wall forms a portion of the
airfoil body defining the first leading edge impingement cavity, a
second leading edge impingement cavity located within the airfoil
body proximate the leading edge, wherein the second channel wall
forms a portion of the airfoil body defining the second leading
edge impingement cavity, and a first leading edge impingement hole
formed in the first channel wall and angled such that air flowing
from the first leading edge impingement cavity and through the
first leading edge impingement hole impinges upon a portion of the
second channel wall.
[0007] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the first end is a root of the airfoil body and the
second end is a tip of the airfoil, wherein the leading edge
channel extends less than a full length of a distance between the
root and the tip along the leading edge of the airfoil body.
[0008] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include a second leading edge impingement hole formed in the second
channel wall and angled such that air flowing from the second
leading edge impingement cavity and through the second leading edge
impingement hole impinges upon a portion of the first channel
wall.
[0009] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the leading edge channel comprises a plurality of
first leading edge impingement holes formed in the first channel
wall, wherein the plurality of first leading edge impingement holes
extend in an array radially along the first channel wall, and
wherein the leading edge channel comprises a plurality of second
leading edge impingement holes formed in the second channel wall,
wherein the plurality of second leading edge impingement holes
extend in an array radially along the second channel wall.
[0010] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include a leading edge feed cavity arranged aft of the first and
second leading edge impingement cavities and arranged to supply air
into at least one of the first leading edge impingement cavity and
the second leading edge impingement cavity.
[0011] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the leading edge feed cavity supplies air into both of
the first leading edge impingement cavity and the second leading
edge impingement cavity.
[0012] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the leading edge channel had a depth in an axial
direction that is at least twice a diameter of the first leading
edge impingement hole.
[0013] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the leading edge channel comprises a plurality of
first leading edge impingement holes formed in the first channel
wall, wherein the plurality of first leading edge impingement holes
extend in an array radially along the first channel wall.
[0014] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include at least one additional channel formed on an exterior
surface of the airfoil and extending radially.
[0015] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the airfoils may
include that the airfoil body forms a blade of a turbine section of
the gas turbine engine.
[0016] According to some embodiments, gas turbine engines are
provided. The gas turbine engines include an airfoil having an
airfoil body extending in a radial direction from a first end to a
second end, and extending axially from a leading edge to a trailing
edge, a leading edge channel formed in the leading edge of the
airfoil body, the leading edge channel having a first channel wall
and a second channel wall that join at a channel base to define the
leading edge channel, the leading edge channel extending in a
radial direction along the leading edge of the airfoil body, a
first leading edge impingement cavity located within the airfoil
body proximate the leading edge, wherein the first channel wall
forms a portion of the airfoil body defining the first leading edge
impingement cavity, a second leading edge impingement cavity
located within the airfoil body proximate the leading edge, wherein
the second channel wall forms a portion of the airfoil body
defining the second leading edge impingement cavity, and a first
leading edge impingement hole formed in the first channel wall and
angled such that air flowing from the first leading edge
impingement cavity and through the first leading edge impingement
hole impinges upon a portion of the second channel wall.
[0017] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the first end is a root of the airfoil
body and the second end is a tip of the airfoil, wherein the
leading edge channel extends less than a full length of a distance
between the root and the tip along the leading edge of the airfoil
body.
[0018] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include a second leading edge impingement hole formed
in the second channel wall and angled such that air flowing from
the second leading edge impingement cavity and through the second
leading edge impingement hole impinges upon a portion of the first
channel wall.
[0019] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the leading edge channel comprises a
plurality of first leading edge impingement holes formed in the
first channel wall, wherein the plurality of first leading edge
impingement holes extend in an array radially along the first
channel wall, and wherein the leading edge channel comprises a
plurality of second leading edge impingement holes formed in the
second channel wall, wherein the plurality of second leading edge
impingement holes extend in an array radially along the second
channel wall.
[0020] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include a leading edge feed cavity arranged aft of the
first and second leading edge impingement cavities and arranged to
supply air into at least one of the first leading edge impingement
cavity and the second leading edge impingement cavity.
[0021] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the leading edge feed cavity supplies air
into both of the first leading edge impingement cavity and the
second leading edge impingement cavity.
[0022] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the leading edge channel had a depth in an
axial direction that is at least twice a diameter of the first
leading edge impingement hole.
[0023] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the leading edge channel comprises a
plurality of first leading edge impingement holes formed in the
first channel wall, wherein the plurality of first leading edge
impingement holes extend in an array radially along the first
channel wall.
[0024] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include at least one additional channel formed on an
exterior surface of the airfoil and extending radially.
[0025] In addition to one or more of the features described herein,
or as an alternative, further embodiments of the gas turbine
engines may include that the airfoil is a blade of a turbine
section of the gas turbine engine.
[0026] The foregoing features and elements may be combined in
various combinations without exclusivity, unless expressly
indicated otherwise. These features and elements as well as the
operation thereof will become more apparent in light of the
following description and the accompanying drawings. It should be
understood, however, the following description and drawings are
intended to be illustrative and explanatory in nature and
non-limiting.
BRIEF DESCRIPTION OF THE DRAWINGS
[0027] The following descriptions should not be considered limiting
in any way. With reference to the accompanying drawings, like
elements are numbered alike: The subject matter is particularly
pointed out and distinctly claimed at the conclusion of the
specification. The foregoing and other features, and advantages of
the present disclosure are apparent from the following detailed
description taken in conjunction with the accompanying drawings in
which like elements may be numbered alike and:
[0028] FIG. 1 is a schematic cross-sectional illustration of a gas
turbine engine;
[0029] FIG. 2 is a schematic illustration of a portion of a turbine
section of the gas turbine engine of FIG. 1;
[0030] FIG. 3A is a schematic illustration of an airfoil in
accordance with an embodiment of the present disclosure; and
[0031] FIG. 3B is an enlarged illustration of a leading edge of the
airfoil shown in FIG. 3A.
DETAILED DESCRIPTION
[0032] Detailed descriptions of one or more embodiments of the
disclosed apparatus and/or methods are presented herein by way of
exemplification and not limitation with reference to the
Figures.
[0033] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 is disclosed herein as a two-spool
turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmentor section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B in a bypass duct, while the compressor
section 24 drives air along a core flow path C for compression and
communication into the combustor section 26 then expansion through
the turbine section 28. Although depicted as a two-spool turbofan
gas turbine engine in the disclosed non-limiting embodiment, it
should be understood that the concepts described herein are not
limited to use with two-spool turbofans as the teachings may be
applied to other types of turbine engines including three-spool
architectures.
[0034] The exemplary engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided, and the location of bearing systems 38
may be varied as appropriate to the application.
[0035] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 is connected to the fan 42
through a speed change mechanism, which in exemplary gas turbine
engine 20 is illustrated as a geared architecture 48 to drive the
fan 42 at a lower speed than the low speed spool 30. The high speed
spool 32 includes an outer shaft 50 that interconnects a high
pressure compressor 52 and high pressure turbine 54. A combustor 56
is arranged in exemplary gas turbine 20 between the high pressure
compressor 52 and the high pressure turbine 54. An engine static
structure 36 is arranged generally between the high pressure
turbine 54 and the low pressure turbine 46. The engine static
structure 36 further supports bearing systems 38 in the turbine
section 28. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via bearing systems 38 about the engine
central longitudinal axis A which is collinear with their
longitudinal axes.
[0036] The core airflow is compressed by the low pressure
compressor 44 then the high pressure compressor 52, mixed and
burned with fuel in the combustor 56, then expanded over the high
pressure turbine 54 and low pressure turbine 46. The turbines 46,
54 rotationally drive the respective low speed spool 30 and high
speed spool 32 in response to the expansion. It will be appreciated
that each of the positions of the fan section 22, compressor
section 24, combustor section 26, turbine section 28, and fan drive
gear system 48 may be varied. For example, gear system 48 may be
located aft of combustor section 26 or even aft of turbine section
28, and fan section 22 may be positioned forward or aft of the
location of gear system 48.
[0037] The engine 20 in one example is a high-bypass geared
aircraft engine. In a further example, the engine 20 bypass ratio
is greater than about six (6), with an example embodiment being
greater than about ten (10), the geared architecture 48 is an
epicyclic gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3 and
the low pressure turbine 46 has a pressure ratio that is greater
than about five. In one disclosed embodiment, the engine 20 bypass
ratio is greater than about ten (10:1), the fan diameter is
significantly larger than that of the low pressure compressor 44,
and the low pressure turbine 46 has a pressure ratio that is
greater than about five 5:1. Low pressure turbine 46 pressure ratio
is pressure measured prior to inlet of low pressure turbine 46 as
related to the pressure at the outlet of the low pressure turbine
46 prior to an exhaust nozzle. The geared architecture 48 may be an
epicycle gear train, such as a planetary gear system or other gear
system, with a gear reduction ratio of greater than about 2.3:1. It
should be understood, however, that the above parameters are only
exemplary of one embodiment of a geared architecture engine and
that the present disclosure is applicable to other gas turbine
engines including direct drive turbofans.
[0038] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The
flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with
the engine at its best fuel consumption--also known as "bucket
cruise Thrust Specific Fuel Consumption (`TSFC`)"--is the industry
standard parameter of lbm of fuel being burned divided by lbf of
thrust the engine produces at that minimum point. "Low fan pressure
ratio" is the pressure ratio across the fan blade alone, without a
Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as
disclosed herein according to one non-limiting embodiment is less
than about 1.45. "Low corrected fan tip speed" is the actual fan
tip speed in ft/sec divided by an industry standard temperature
correction of [(Tram .degree. R)/(514.7.degree. R)].sup.0.5. The
"Low corrected fan tip speed" as disclosed herein according to one
non-limiting embodiment is less than about 1150 ft/second (350.5
m/sec).
[0039] Although the gas turbine engine 20 is depicted as a
turbofan, it should be understood that the concepts described
herein are not limited to use with the described configuration, as
the teachings may be applied to other types of engines such as, but
not limited to, turbojets, turboshafts, and three-spool (plus fan)
turbofans wherein an intermediate spool includes an intermediate
pressure compressor ("IPC") between a low pressure compressor
("LPC") and a high pressure compressor ("HPC"), and an intermediate
pressure turbine ("IPT") between the high pressure turbine ("HPT")
and the low pressure turbine ("LPT").
[0040] FIG. 2 is a schematic view of a portion of the turbine
section 28 that may employ various embodiments disclosed herein.
Turbine section 28 includes a plurality of airfoils 60, 62
including, for example, one or more blades and vanes. The airfoils
60, 62 may be hollow bodies with internal cavities defining a
number of channels or cores, hereinafter airfoil cores, formed
therein and extending from an inner diameter 66 to an outer
diameter 68, or vice-versa. The airfoil cores may be separated by
partitions within the airfoils 60, 62 that may extend either from
the inner diameter 66 or the outer diameter 68 of the airfoil 60,
62. The partitions may extend for a portion of the length of the
airfoil 60, 62, but may stop or end prior to forming a complete
wall within the airfoil 60, 62. Thus, each of the airfoil cores may
be fluidly connected and form a fluid path within the respective
airfoil 60, 62. The airfoils 60, 62 may include platforms 70
located proximal to the inner diameter 66 thereof. Located below
the platforms 70 (e.g., radially inward with respect to the engine
axis) may be airflow ports and/or bleed orifices that enable air to
bleed from the internal cavities of the airfoils 60, 62. A root of
the airfoil may connect to or be part of the platform 70.
[0041] The turbine section 28 is housed within a case 80, which may
have multiple parts (e.g., turbine case, diffuser case, etc.). In
various locations, components, such as seals, may be positioned
between airfoils 60, 62 and the case 80. For example, as shown in
FIG. 2, blade outer air seals 82 (hereafter "BOAS") are located
radially outward from the blade 60. As will be appreciated by those
of skill in the art, the BOAS 82 may include BOAS supports that are
configured to fixedly connect or attach the BOAS 82 to the case 80
(e.g., the BOAS supports may be located between the BOAS 82 and the
case 80). As shown in FIG. 2, the case 80 includes a plurality of
case hooks 84 that engage with BOAS hooks 86 to secure the BOAS 82
between the case 80 and a tip of the airfoil 60.
[0042] As shown and labeled in FIG. 2, a radial direction is upward
on the page (e.g., radial with respect to an engine axis) and an
axial direction is to the right on the page (e.g., along an engine
axis). Thus, radial cooling flows will travel up or down on the
page and axial flows will travel left-to-right (or vice versa).
[0043] Turning to FIGS. 3A-3B, schematic illustrations of an
airfoil 300 in accordance with an embodiment of the present
disclosure are shown, with FIG. 3B being an enlarged illustration
of a leading edge of the airfoil 300. The airfoil 300 is defined by
an airfoil body 300a having a leading edge 302 and extends aftward
to a trailing edge 304. The airfoil 300, as shown, has a first
leading edge impingement cavity 306 and a second leading edge
impingement cavity 308. Aft of the leading edge impingement
cavities 306, 308 is a leading edge feed cavity 310. The airfoil
300 includes additional cooling cavities 312, as shown.
[0044] The leading edge impingement cavities 306, 308 are fed from
the leading edge feed cavity 310 by one or more respective
impingement holes that fluidly connect the leading edge feed cavity
310 leading edge impingement cavities 306, 308. For example, as
shown, a first impingement hole 314 enables air from the leading
edge feed cavity 310 to impinge into the first leading edge
impingement cavity 306 and a second impingement hole 316 enables
air from the leading edge feed cavity 310 to impinge into the
second leading edge impingement cavity 308. The impinging air from
the leading edge feed cavity 310 will provide cooling to the
leading edge 302 of the airfoil 300.
[0045] As shown, the airfoil 300 includes a stagnation divot,
hollow, trench, or channel (hereinafter "leading edge channel 318")
which is positioned at the tip or front of the leading edge 302 of
the airfoil 300. The leading edge channel 318, in some embodiments,
extends from a root to a tip (e.g., along a radial length) of the
airfoil 300, as will be appreciated by those of skill in the art.
In other embodiments, the leading edge channel can extend over a
partial extent of the radial length, e.g., less than a full length
of a distance between the first end (e.g., root) and the second
(e.g., tip) along the leading edge of the airfoil body. That is, in
some embodiments, the leading edge channel may not extend along the
entire radial length of the leading edge. For example, in one
non-limiting embodiment, the leading edge channel may extend along
about half of the radial length of the airfoil leading edge, with
no channel present on the other half of the leading edge. In one
such embodiment, the leading edge channel can extend from a tip of
the airfoil in a radially downward direction and stop approximately
at the halfway point. In other embodiments, the leading edge
channel can be less-than-full-radial-distance in length, but
positioned such that neither end of the channel is located at the
root or tip (e.g., positioned at a mid-point or partial mid-point
along the radial length of the leading edge of the airfoil).
Further, in some embodiments, multiple less-than-full length
leading edge channels can be formed on a leading edge of the
airfoil, without departing from the scope of the present
disclosure.
[0046] A portion of the air from the first and second leading edge
impingement cavities 306, 308 will exit the respective first and
second leading edge impingement cavities 306, 308 through leading
edge impingement holes 320, 322. A first leading edge impingement
hole 320 is formed in a first channel wall 324 that partially
defines the first leading edge impingement cavity 306. A second
leading edge impingement hole 322 is formed in a second channel
wall 326 that partially defines the second leading edge impingement
cavity 308. Although a single leading edge impingement hole is
shown in the illustration of FIGS. 3A-3B, those of skill in the art
will appreciate that an array of impingement holes can be formed
extending along a radial extent of the airfoil 300 along the
channel walls 324, 326 of the leading edge channel. The first
channel wall 324 and the second channel wall 326 form and define
the leading edge channel 318. A channel base 328 is formed where
the first and second channel walls 324, 326 meet at the base of the
leading edge channel 318.
[0047] Air from the first and second leading edge impingement
cavities 306, 308 will flow into and/or impinge into the leading
edge channel 318 along the leading edge 302. Once in the leading
edge channel 318, the cooling air diffuses into cooling air already
in the leading edge channel 318 and distributes spanwise along the
leading edge channel 318. One of the advantages of distributing
cooling air within the leading edge channel 318 is that the
pressure difference problems characteristic of conventional cooling
orifices are minimized. For example, the difference in pressure
across a cooling orifice is a function of a local internal cavity
pressure and a local gaspath gas pressure adjacent the orifice.
Both of these pressures vary as a function of time. If the gaspath
gas pressure is high and the internal cavity pressure is low
adjacent a particular cooling orifice in a conventional scheme,
undesirable hot core gas in-flow can occur (e.g., into one of the
leading edge impingement or other cooling cavities). Embodiments
provided herein can minimize the opportunity for the undesirable
in-flow due to impingement air being distributed within the leading
edge channel 318, thereby decreasing the opportunity for any low
pressure zones to occur. Likewise, the distribution of cooling air
within the leading edge channel 318 also avoids cooling air
pressure spikes which, in a conventional scheme, can jet cooling
air into the gaspath gas rather than add it to a film of cooling
air downstream along exterior surfaces of the airfoil 300.
[0048] Additionally, air impinging from the leading edge
impingement holes 320, 322 can provide impingement cooling to the
opposing channel walls 324, 326. For example, as shown in FIG. 3B,
air from the first leading edge impingement cavity 306 will flow
through the first leading edge impingement hole 320 and impinge
upon the second channel wall 326. Similarly, air from the second
leading edge impingement cavity 308 will flow through the second
leading edge impingement hole 322 and impinge upon the first
channel wall 324. That is, the first leading edge impingement hole
320 is angled such that air passing through the first leading edge
impingement hole 320 is directed such that air will impinge upon
the material of the second channel wall 326 and the second leading
edge impingement hole 322 is angled such that air passing through
the second leading edge impingement hole 322 is directed such that
air will impinge upon the material of the first channel wall
324.
[0049] In some embodiments, the leading edge channel 318 has a
depth 330 that is at least twice the diameter 332 of the leading
edge impingement holes 320, 322, as schematically shown. This
dimension enables formation of the leading edge impingement holes
320, 322 within the channel walls 324, 326. The depth 330 is
measured from the leading edge 302 to the channel base 328 of the
leading edge channel 318.
[0050] Although shown herein with two leading edge impingement
holes (one for each leading edge impingement cavity), those of
skill in the art will appreciate that other arrangements are
possible without departing from the scope of the present
disclosure. For example, in some embodiments, one of the leading
edge impingement holes can be omitted such that only one leading
edge impingement hole is provided to supply air into the leading
edge channel. Further, in some embodiment, multiple channels can be
formed along the leading edge or other surfaces of the airfoil. In
some such embodiments, each of the channels can be arranged with
one or more impingement and/or feed holes to supply air into the
channel, as will be appreciated by those of skill in the art. For
example, in one non-limiting embodiment, a second channel can be
formed on the exterior surface of the airfoil 300 adjacent the
first leading edge impingement cavity 306 and the leading edge feed
cavity 310, with the second channel being sourced from the first
leading edge impingement cavity 306 and the leading edge feed
cavity 310.
[0051] Embodiments provided herein are directed to airfoils having
a leading edge channel that is supplied with impingement air from
multiple different leading edge impingement cavities.
Advantageously, such arrangement can enable the leading edge
channel to be supplied with impingement air from any one (or more)
of the leading edge impingement cavities, thus ensuring constant
impinging air within the leading edge channel. Further,
advantageously, embodiment provided herein are directed to angled
impingement holes within the leading edge channel such that a
portion of the impinging air that flows into the leading edge
channel impinges upon and cools the material of the opposing
channel wall.
[0052] As used herein, the term "about" is intended to include the
degree of error associated with measurement of the particular
quantity based upon the equipment available at the time of filing
the application. For example, "about" may include a range of
.+-.8%, or 5%, or 2% of a given value or other percentage change as
will be appreciated by those of skill in the art for the particular
measurement and/or dimensions referred to herein.
[0053] The terminology used herein is for the purpose of describing
particular embodiments only and is not intended to be limiting of
the present disclosure. As used herein, the singular forms "a,"
"an," and "the" are intended to include the plural forms as well,
unless the context clearly indicates otherwise. It will be further
understood that the terms "comprises" and/or "comprising," when
used in this specification, specify the presence of stated
features, integers, steps, operations, elements, and/or components,
but do not preclude the presence or addition of one or more other
features, integers, steps, operations, element components, and/or
groups thereof. It should be appreciated that relative positional
terms such as "forward," "aft," "upper," "lower," "above," "below,"
"radial," "axial," "circumferential," and the like are with
reference to normal operational attitude and should not be
considered otherwise limiting.
[0054] While the present disclosure has been described with
reference to an illustrative embodiment or embodiments, it will be
understood by those skilled in the art that various changes may be
made and equivalents may be substituted for elements thereof
without departing from the scope of the present disclosure. In
addition, many modifications may be made to adapt a particular
situation or material to the teachings of the present disclosure
without departing from the essential scope thereof. Therefore, it
is intended that the present disclosure not be limited to the
particular embodiment disclosed as the best mode contemplated for
carrying out this present disclosure, but that the present
disclosure will include all embodiments falling within the scope of
the claims.
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