U.S. patent application number 13/621968 was filed with the patent office on 2014-03-20 for gas turbine engine component cooling circuit.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is UNITED TECHNOLOGIES CORPORATION. Invention is credited to Steven Bruce Gautschi, Lane Thornton.
Application Number | 20140075947 13/621968 |
Document ID | / |
Family ID | 50273010 |
Filed Date | 2014-03-20 |
United States Patent
Application |
20140075947 |
Kind Code |
A1 |
Gautschi; Steven Bruce ; et
al. |
March 20, 2014 |
GAS TURBINE ENGINE COMPONENT COOLING CIRCUIT
Abstract
A component for a gas turbine engine, according to an exemplary
aspect of the present disclosure includes, among other things, a
body portion and a cooling circuit disposed inside of the body
portion. The cooling circuit includes a first baffle received
within a first core cavity that extends inside of the body portion,
a second baffle received within a second core cavity that extends
inside of the body portion, and a first rib disposed between the
first core cavity and the second core cavity. The first baffle is
in fluid communication with the second baffle through the first
rib.
Inventors: |
Gautschi; Steven Bruce;
(Naugatuck, CT) ; Thornton; Lane; (Meriden,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
UNITED TECHNOLOGIES CORPORATION |
Hartford |
CT |
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
50273010 |
Appl. No.: |
13/621968 |
Filed: |
September 18, 2012 |
Current U.S.
Class: |
60/726 ; 415/1;
415/177; 416/1; 416/96R |
Current CPC
Class: |
F05D 2240/126 20130101;
Y02T 50/60 20130101; F01D 5/189 20130101; Y02T 50/676 20130101 |
Class at
Publication: |
60/726 ; 415/177;
416/96.R; 415/1; 416/1 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F01D 25/12 20060101 F01D025/12; F01D 9/00 20060101
F01D009/00 |
Claims
1. A component for a gas turbine engine, comprising: a body
portion; and a cooling circuit disposed inside of said body
portion, wherein said cooling circuit includes: a first baffle
received within a first core cavity that extends inside of said
body portion; a second baffle received within a second core cavity
that extends inside of said body portion; and a first rib disposed
between said first core cavity and said second core cavity, wherein
said first baffle is in fluid communication with said second baffle
through said first rib.
2. The component as recited in claim 1, wherein the component is a
vane.
3. The component as recited in claim 1, wherein the component is a
blade.
4. The component as recited in claim 1, wherein said first rib
includes a plurality of openings that fluidly connect said first
core cavity and said second core cavity.
5. The component as recited in claim 4, wherein said plurality of
openings are positioned in a staggered relationship across a radial
span of said first rib.
6. The component as recited in claim 4, wherein said plurality of
openings each axially extend through said first rib in a direction
that extends from a leading edge toward a trailing edge of said
body portion.
7. The component as recited in claim 1, wherein said first baffle
and said second baffle each include a plurality of feed openings
that extend through said first baffle and said second baffle.
8. The component as recited in claim 7, wherein said plurality of
feed openings extend through each wall of said first baffle and
said second baffle.
9. The component as recited in claim 1, wherein a space extends
between an interior wall of said first core cavity and said first
baffle.
10. The component as recited in claim 1, wherein said cooling
circuit includes a third baffle received within a third core cavity
that extends inside of said body portion.
11. The component as recited in claim 10, wherein said third baffle
is in fluid communication with said second baffle through a second
rib.
12. The component as recited in claim 10, wherein said cooling
circuit includes a trailing edge cavity in fluid communication with
said third core cavity.
13. A gas turbine engine, comprising: a compressor section; a
combustor section in fluid communication with said compressor
section; a turbine section in fluid communication with said
combustor section; and wherein at least one of said compressor
section and said turbine section includes at least one component
having a body portion and a cooling circuit disposed inside of said
body portion, wherein said cooling circuit includes: a first baffle
received within a first core cavity that extends inside of said
body portion; a second baffle received within a second core cavity
that extends inside of said body portion; and a first rib disposed
between said first core cavity and said second core cavity, wherein
said first baffle is in fluid communication with said second baffle
through said first rib.
14. The gas turbine engine as recited in claim 13, wherein said at
least one component is a vane.
15. The gas turbine engine as recited in claim 13, wherein said
first rib includes a plurality of openings.
16. The gas turbine engine as recited in claim 13, wherein said
first baffle and said second baffle each include a plurality of
feed openings that extend through said first baffle and said second
baffle.
17. The gas turbine engine as recited in claim 13, wherein said
cooling circuit includes a third baffle received within a third
core cavity that extends inside of said body portion.
18. A method of cooling a component of a gas turbine engine,
comprising the steps of: feeding a cooling airflow into a first
core cavity of a body portion of the component; and expelling the
cooling airflow from the body portion through a second core cavity
that is in fluid communication with the first core cavity.
19. The method as recited in claim 18, wherein the step of feeding
includes: communicating the cooling airflow through a plurality
feed openings in a first baffle positioned within the first core
cavity; and impingement cooling at least one interior wall of the
body portion with the cooling airflow that is communicated through
the plurality of feed openings prior to the step of expelling.
20. The method as recited in claim 18, comprising the step of:
communicating the cooling airflow through a first rib that is
disposed between the first core cavity and the second core cavity
prior to the step of expelling.
Description
BACKGROUND
[0001] This disclosure relates to a gas turbine engine, and more
particularly to a cooling circuit for cooling a gas turbine engine
component.
[0002] Gas turbine engines typically include a compressor section,
a combustor section and a turbine section. In general, during
operation, air is pressurized in the compressor section and is
mixed with fuel and burned in the combustor section to generate hot
combustion gases. The hot combustion gases flow through the turbine
section, which extracts energy from the hot combustion gases to
power the compressor section and other gas turbine engine
loads.
[0003] The compressor and turbine sections of the gas turbine
engine typically include alternating rows of rotating blades and
stationary vanes. The rotating blades either create or extract
energy from the hot combustion gases that are communicated through
the gas turbine engine, and the vanes convert the velocity of the
airflow into pressure and prepare the airflow for the next set of
blades. The hot combustion gases are communicated over airfoils of
the blades and the vanes. The airfoils may include internal cooling
circuits that receive a cooling airflow to cool the various
internal and external surfaces of the airfoils.
SUMMARY
[0004] A component for a gas turbine engine, according to an
exemplary aspect of the present disclosure includes, among other
things, a body portion and a cooling circuit disposed inside of the
body portion. The cooling circuit includes a first baffle received
within a first core cavity that extends inside of the body portion,
a second baffle received within a second core cavity that extends
inside of the body portion, and a first rib disposed between the
first core cavity and the second core cavity. The first baffle is
in fluid communication with the second baffle through the first
rib.
[0005] In a further non-limiting embodiment of the foregoing
component, the component is a vane.
[0006] In a further non-limiting embodiment of either of the
foregoing components, the component is a blade.
[0007] In a further non-limiting embodiment of any of the foregoing
components, the first rib includes a plurality of openings that
fluidly connect the first core cavity and the second core
cavity.
[0008] In a further non-limiting embodiment of any of the foregoing
components, the plurality of openings are positioned in a staggered
relationship across a radial span of the first rib.
[0009] In a further non-limiting embodiment of any of the foregoing
components, the plurality of openings each axially extend through
the first rib in a direction that extends from a leading edge
toward a trailing edge.
[0010] In a further non-limiting embodiment of any of the foregoing
components, the first baffle and the second baffle each include a
plurality of feed openings that extend through the first baffle and
the second baffle.
[0011] In a further non-limiting embodiment of any of the foregoing
components, the plurality of feed openings extend through each wall
of the first baffle and the second baffle.
[0012] In a further non-limiting embodiment of any of the foregoing
components, a space extends between an interior wall of the first
core cavity and the first baffle.
[0013] In a further non-limiting embodiment of any of the foregoing
components, the cooling circuit includes a third baffle received
within a third core cavity that extends inside of the body
portion.
[0014] In a further non-limiting embodiment of any of the foregoing
components, the third baffle is in fluid communication with the
second baffle through a second rib.
[0015] In a further non-limiting embodiment of any of the foregoing
components, the cooling circuit includes a trailing edge cavity in
fluid communication with the third core cavity.
[0016] A gas turbine engine, according to an exemplary aspect of
the present disclosure includes, among other things, a compressor
section, a combustor section in fluid communication with the
compressor section and a turbine section in fluid communication
with the combustor section. At least one of the compressor section
and the turbine section includes at least one component having a
body portion and a cooling circuit disposed inside of the body
portion. The cooling circuit includes a first baffle received
within a first core cavity that extends inside of the body portion,
a second baffle received within a second core cavity that extends
inside of the body portion, and a first rib disposed between the
first core cavity and the second core cavity. The first baffle is
in fluid communication with the second baffle through the first
rib.
[0017] In a further non-limiting embodiment of the foregoing gas
turbine engine, the at least one component is a vane.
[0018] In a further non-limiting embodiment of either of the
foregoing gas turbine engines, the first rib includes a plurality
of openings.
[0019] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the first baffle and the second baffle each
include a plurality of feed openings that extend through the first
baffle and the second baffle.
[0020] In a further non-limiting embodiment of any of the foregoing
gas turbine engines, the cooling circuit includes a third baffle
received within a third core cavity that extends inside of the body
portion.
[0021] A method of cooling a component of a gas turbine engine,
according to another exemplary aspect of the present disclosure
includes, among other things, feeding a cooling airflow into a
first core cavity of a body portion of the component and expelling
the cooling airflow from the body portion through a second core
cavity that is in fluid communication with the first core
cavity.
[0022] In a further non-limiting embodiment of the foregoing method
of cooling a component of a gas turbine engine, the step of feeding
includes communicating the cooling airflow through a plurality feed
openings in a first baffle positioned within the first core cavity
and impingement cooling at least one interior wall of the body
portion with the cooling airflow that is communicated through the
plurality of feed openings prior to the step of expelling.
[0023] In a further non-limiting embodiment of either of the
foregoing methods of cooling a component of a gas turbine engine,
the method may comprise the step of communicating the cooling
airflow through a first rib that is disposed between the first core
cavity and the second core cavity prior to the step of
expelling.
[0024] The various features and advantages of this disclosure will
become apparent to those skilled in the art from the following
detailed description. The drawings that accompany the detailed
description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[0025] FIG. 1 illustrates a schematic, cross-sectional view of a
gas turbine engine.
[0026] FIG. 2 illustrates a component that can be incorporated into
a gas turbine engine.
[0027] FIG. 3 illustrates a cross-sectional view through section
A-A of the component of FIG. 2.
[0028] FIG. 4 illustrates a cooling circuit that can be
incorporated into an airfoil of a component.
[0029] FIG. 5 illustrates various features of a cooling circuit
that can be incorporated into an airfoil of a component.
[0030] FIGS. 6A, 6B and 6C schematically illustrate cooling an
airfoil using an exemplary cooling circuit.
DETAILED DESCRIPTION
[0031] FIG. 1 schematically illustrates a gas turbine engine 20.
The exemplary gas turbine engine 20 is a two-spool turbofan engine
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 augmenter section (not shown) among other
systems for features. The fan section 22 drives air along a bypass
flow path B, while the compressor section 24 drives air along a
core flow path C for compression and communication into the
combustor section 26. The hot combustion gases generated in the
combustor section 26 are expanded through the turbine section 28.
Although depicted as a turbofan gas turbine engine in the disclosed
non-limiting embodiment, it should be understood that the concepts
described herein are not limited to turbofan engines and these
teachings could extend to other types of engines, including but not
limited to, three-spool engine architectures.
[0032] The gas turbine engine 20 generally includes a low speed
spool 30 and a high speed spool 32 mounted for rotation about an
engine centerline longitudinal axis A. The low speed spool 30 and
the high speed spool 32 may be mounted relative to an engine static
structure 33 via several bearing systems 31. It should be
understood that other bearing systems 31 may alternatively or
additionally be provided.
[0033] The low speed spool 30 generally includes an inner shaft 34
that interconnects a fan 36, a low pressure compressor 38 and a low
pressure turbine 39. The inner shaft 34 can be connected to the fan
36 through a geared architecture 45 to drive the fan 36 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 35 that interconnects a high pressure compressor 37
and a high pressure turbine 40. In this embodiment, the inner shaft
34 and the outer shaft 35 are supported at various axial locations
by bearing systems 31 positioned within the engine static structure
33.
[0034] A combustor 42 is arranged between the high pressure
compressor 37 and the high pressure turbine 40. A mid-turbine frame
44 may be arranged generally between the high pressure turbine 40
and the low pressure turbine 39. The mid-turbine frame 44 can
support one or more bearing systems 31 of the turbine section 28.
The mid-turbine frame 44 may include one or more airfoils 46 that
extend within the core flow path C.
[0035] The inner shaft 34 and the outer shaft 35 are concentric and
rotate via the bearing systems 31 about the engine centerline
longitudinal axis A, which is co-linear with their longitudinal
axes. The core airflow is compressed by the low pressure compressor
38 and the high pressure compressor 37, is mixed with fuel and
burned in the combustor 42, and is then expanded over the high
pressure turbine 40 and the low pressure turbine 39. The high
pressure turbine 40 and the low pressure turbine 39 rotationally
drive the respective high speed spool 32 and the low speed spool 30
in response to the expansion.
[0036] In a non-limiting embodiment, the gas turbine engine 20 is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 bypass ratio is greater than about six (6:1). The
geared architecture 45 can include an epicyclic gear train, such as
a planetary gear system or other gear system. The example epicyclic
gear train has a gear reduction ratio of greater than about 2.3,
and in another example is greater than about 2.5:1. The geared
turbofan enables operation of the low speed spool 30 at higher
speeds, which can increase the operational efficiency of the low
pressure compressor 38 and low pressure turbine 39 and render
increased pressure in a fewer number of stages.
[0037] The pressure ratio of the low pressure turbine 39 can be
pressure measured prior to the inlet of the low pressure turbine 39
as related to the pressure at the outlet of the low pressure
turbine 39 and prior to an exhaust nozzle of the gas turbine engine
20. In one non-limiting embodiment, the bypass ratio of the gas
turbine engine 20 is greater than about ten (10:1), the fan
diameter is significantly larger than that of the low pressure
compressor 38, and the low pressure turbine 39 has a pressure ratio
that is greater than about five (5:1). 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] In this embodiment of the exemplary gas turbine engine 20, a
significant amount of thrust is provided by the bypass flow path B
due to the high bypass ratio. The fan section 22 of the gas turbine
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. This flight
condition, with the gas turbine engine 20 at its best fuel
consumption, is also known as bucket cruise Thrust Specific Fuel
Consumption (TSFC). TSFC is an industry standard parameter of fuel
consumption per unit of thrust.
[0039] Fan Pressure Ratio is the pressure ratio across a blade of
the fan section 22 without the use of a Fan Exit Guide Vane system.
The low Fan Pressure Ratio according to one non-limiting embodiment
of the example gas turbine engine 20 is less than 1.45. Low
Corrected Fan Tip Speed is the actual fan tip speed divided by an
industry standard temperature correction of ("T"/518.7).sup.0.5,
where T represents the ambient temperature in degrees Rankine. The
Low Corrected Fan Tip Speed according to one non-limiting
embodiment of the example gas turbine engine 20 is less than about
1150 fps (351 m/s).
[0040] Each of the compressor section 24 and the turbine section 28
may include alternating rows of rotor assemblies and vane
assemblies (shown schematically) that carry airfoils that extend
into the core flow path C. For example, the rotor assemblies can
carry a plurality of rotating blades 25, while each vane assembly
can carry a plurality of vanes 27 that extend into the core flow
path C. The blades 25 of the rotor assemblies create or extract
energy (in the form of pressure) from the core airflow that is
communicated through the gas turbine engine 20 along the core flow
path C. The vanes 27 of the vane assemblies direct the core air
flow to the blades 25 to either add or extract energy.
[0041] Various components of a gas turbine engine 20, such as the
airfoils of the blades 25 and the vanes 27 of the compressor
section 24 and the turbine section 28, may be subjected to
repetitive thermal cycling under widely ranging temperatures and
pressures. The hardware of the turbine section 28 is particularly
subjected to relatively extreme operating conditions. Therefore,
some components may require internal cooling circuits for cooling
the parts during engine operation. Example cooling circuits for
cooling an airfoil of a component are discussed below.
[0042] FIG. 2 illustrates a component 50 that can be incorporated
into a gas turbine engine, such as the gas turbine engine 20 of
FIG. 1. The component 50 includes a body portion 52 that axially
extends between a leading edge 54 and a trailing edge 56 and
circumferentially extends between a pressure side 58 and a suction
side 60. In this embodiment, the body portion 52 is an airfoil that
extends across a span S between an inner platform 61 and an outer
platform 63. In other words, the component 50 is illustrated as a
vane. However, the body portion 52 could also include an airfoil
that extends from a platform and a root portion connected to the
platform where the component is a blade. In yet another embodiment,
the body portion 52 could include a seal body of a blade outer air
seal (BOAS).
[0043] A gas path 62 is communicated axially downstream through the
gas turbine engine 20 along a core flow path C (FIG. 1) in a
direction that extends from the leading edge 54 toward the trailing
edge 56 of the body portion 52. The gas path 62 is schematically
represented by an arrow and represents the communication of core
airflow across the body portion 52.
[0044] The component 50 may include a cooling circuit 64 for
cooling the internal and/or external surfaces of the body portion
52. The cooling circuit 64 can include one or more core cavities 72
(that can be formed by using ceramic cores) that are radially,
axially and/or circumferentially disposed inside the body portion
52 to establish cooling passages for receiving a cooling airflow 68
to cool the body portion 52. In this particular embodiment, the
cooling circuit 64 includes two core cavities 72. However, any
number of core cavities 72 can be disposed inside of the body
portion 52.
[0045] The cooling circuit 64 can receive the cooling airflow 68
from one or more airflow sources 70 that are external to the body
portion 52. The cooling airflow 68 is generally a lower temperature
than the airflow of the gas path 62 that is communicated across the
body portion 52. In one embodiment, the cooling airflow 68 is a
bleed airflow that can be sourced from the compressor section 24 or
any other portion of the gas turbine engine 20 that is upstream
from the component 50. The cooling airflow 68 can be circulated
through the cooling circuit 64, including through one or more of
the core cavities 72, to transfer thermal energy from the component
50 to the cooling airflow 68 to cool the body portion 52. In one
embodiment, separate airflow sources 70A and 70B can be used to
communicate separate cooling airflows 68 to each of the core
cavities 72.
[0046] The cooling circuit 64 illustrated in this embodiment could
be incorporated into any component where dedicated cooling is
desired, including but not limited to any component that extends
into the core flow path C of the gas turbine engine 20 (see FIG.
1). It should be understood that the cooling circuit 64 depicted in
the illustrated embodiments of this disclosure could be
incorporated into vanes or blades of the compressor section 24
and/or the turbine section 28. Other components, such as the
airfoils of a mid-turbine frame or non-airfoil components such as
BOAS, could also benefit from the teachings of this disclosure.
[0047] FIG. 3, with continued reference to FIG. 2, illustrates one
exemplary cooling circuit 64 that can be incorporated into the
component 50. The cooling circuit 64 is generally defined inside of
the body portion 52 and may extend axially between the leading edge
54 and the trailing edge 56 and circumferentially between the
pressure side 58 and the suction side 60. In this exemplary
embodiment, the cooling circuit 64 includes a first core cavity 72A
and a second core cavity 72B.
[0048] In one embodiment, the first core cavity 72A is positioned
at the leading edge 54 of the body portion 52 and the second core
cavity 72B is positioned downstream from the first core cavity 72A
(i.e., at a mid-portion of the body portion 52 that is between the
leading edge 54 and the trailing edge 56). A first rib 74 separates
the first core cavity 72A from the second core cavity 72B. The
first rib 74 radially extends inside of the body portion 52 and
divides the core cavities 72A, 72B from one another.
[0049] A first baffle 76A may be received within the first core
cavity 72A, and a second baffle 76B may be received within the
second core cavity 72B. The exemplary first and second baffles 76A,
76B are inserts that can be bonded at one or both of the inner
platform 61 and the outer platform 63 within the first core cavity
72A and the second core cavity 72B. In one embodiment, the first
baffle 76A (and the first core cavity 72A) is in fluid
communication with the second baffle 76B (and the second core
cavity 72B) through the first rib 74. The first baffle 76A and the
second baffle 76B are hollow structures. Therefore, cooling airflow
68 can be communicated directly through the first baffle 76A and
the second baffle 76B.
[0050] The first baffle 76A and the second baffle 76B may include a
plurality of feed openings 80 that allow cooling airflow 68 to
escape from the first baffle 76A and the second baffle 76B and
impinge on interior walls 84 of the body portion 52. The feed
openings 80 may be arranged in a staggered relationship across a
radial span of the first baffle 76A and second baffle 76B (see FIG.
4). In addition, the first rib 74 may include a plurality of
openings 82 through which the first core cavity 72A fluidly
connects to the second core cavity 72B. The cooling airflow 68 can
be circulated throughout the core cavities 72A, 72B, the baffles
76A, 76B and the first rib 74 to cool the internal surfaces of the
body portion 52, as is discussed in greater detail below with
reference to FIGS. 6A, 6B and 6C.
[0051] The cooling circuit 64 may also include a trailing edge
cooling circuit 99 positioned to cool the trailing edge 56 of the
body portion 52. Together, in this embodiment, the first core
cavity 72A, the second core cavity 72B, the baffles 76A, 76B, the
first rib 74, and the trailing edge cooling circuit 99 establish
the cooling circuit 64. These features cooperate to cool the body
portion 52 with a minimum amount of dedicated cooling airflow.
[0052] FIG. 4 illustrates another exemplary cooling circuit 164
that can be incorporated into an airfoil 152 of a component 150. In
this embodiment, the platforms are removed from the component 150
to better illustrate the various features of the cooling circuit
164. The exemplary cooling circuit 164 includes a first core cavity
172A (near a leading edge 154), a second core cavity 172B
(downstream from the first core cavity 172A and between the leading
edge 154 and a trailing edge 156), and a third core cavity 172C
(downstream from the second core cavity 172B). Although illustrated
having three core cavities 172A, 172B and 172C, the cooling circuit
164 could include two or more core cavities.
[0053] A first baffle 176A is received within the first core cavity
172A, a second baffle 176B is received within the second core
cavity 172B, and a third baffle 176C is received within the third
core cavity 172C. The baffles 176A, 176B and 176C are shaped to
generally mirror the shape of the first core cavity 172A, the
second core cavity 172B, and the third core cavity 172C,
respectively, and are positioned in a spaced relationship relative
to the interior wall 84 of the airfoil 152.
[0054] A first rib 174A extends between the first core cavity 172A
and the second core cavity 172B and connects the pressure side 158
to the suction side 160 of the airfoil 152. A second rib 174B is
positioned between the second core cavity 172B and the third core
cavity 172C and also connects the pressure side 158 to the suction
side 160 of the airfoil 152. A third rib 174C may be positioned
between the third core cavity 172C and a trailing edge cavity
95.
[0055] Each of the baffles 176A, 176B and 176C can include a
plurality of feed openings 80. In one embodiment, a plurality of
feed openings 80 extend through each of the multiple walls 86 of
the first baffle 176A, the second baffle 176B and the third baffle
176C. Accordingly, cooling airflow 68 can be communicated through
the plurality of feed openings 80 to impinge upon the interior
walls 84 of the airfoil 152.
[0056] FIG. 5 illustrates the cooling circuit 164 of FIG. 4 with
the baffles 176A, 176B and 176C removed. In this embodiment, each
of the first rib 174A, the second rib 174B and the third rib 174C
includes a plurality of openings 82. The plurality of openings 82
of the first rib 174A fluidly connect the first core cavity 172A
with the second core cavity 172B, the plurality of openings 82 of
the second rib 174B fluidly connect the second core cavity 172B
with the third core cavity 172C, and the plurality of openings 82
of the third rib 174C fluidly connect the third core cavity 172C
with the trailing edge cavity 95. Therefore, adjacent baffles 176A,
176B and 176C may also be fluidly connected (see FIG. 4). In one
embodiment, the plurality of openings 82 are arranged in a
staggered relationship across a radial span of each of the first
rib 174A, the second rib 174B and the third rib 174C. The actual
number of openings 82 and the arrangement of these features can
vary depending on the cooling requirements of the airfoil 152,
among other criteria. The plurality of openings 82 extend axially
through the ribs 174A, 174B and 174C (i.e., in a direction that
extends from the leading edge 154 toward the trailing edge
156).
[0057] FIGS. 6A, 6B and 6C schematically illustrate cooling a
component 150 by using a cooling circuit, such as the cooling
circuit 164 described above. Cooling airflow 68 is communicated
into the cooling circuit 164 by feeding the cooling airflow 68 into
the first core cavity 172A. Although not necessary, a separate
cooling airflow 68 may also be simultaneously communicated into the
second core cavity 172B and/or the third core cavity 172C. The
cooling airflow 68 that is fed into the core cavities 172A, 172B
and/or 172C is radially communicated through the hollow portions of
the baffles 176A, 176B, and 176C. As it travels radially, the
cooling airflow 68 may also be communicated through the feed
openings 80 of each baffle 176A, 176B and 176C. The cooling airflow
68 that is communicated through the feed openings 80 may impinge
upon the interior walls 84 and the ribs 174A, 174B and 174C of the
airfoil 152 to cool the airfoil 152 at these locations (shown
schematically via arrows in FIG. 6A).
[0058] Next, as shown in FIG. 6B, a portion P1 of the cooling
airflow 68 within each core cavity 172A, 172B and 172C may be
expelled from the airfoil 152 into the gas path 62 through cooling
holes 88 that may be formed in the leading edge 154, the pressure
side 158 and/or the suction side 160 of the body portion 52. In
this embodiment, at least a portion of the cooling airflow 68 that
is communicated into the first core cavity 172A can be expelled
from the airfoil 152 through another cavity, such as the second
core cavity 172B. Likewise, a portion of the cooling airflow 68
that is communicated into the second core cavity 172B can be
expelled from the airfoil 152 through the third core cavity 172C
and so on. A second portion P2 of the cooling airflow 68 can flow
around the baffles 176A, 176B and 176C in a space 92 that extends
between the baffles 176A, 176B and 176C and the interior walls 84
of each core cavity 172A, 172B and 172C.
[0059] Subsequently, as shown in FIG. 6C, the cooling airflow 68
may be communicated through the plurality of openings 82 in the
ribs 174A, 174B and 174C before again impinging on the interior
walls 84 of any downstream cavity 172 (here, the second and third
core cavities 172B and 172C) of the body portion 52. The cooling
airflow 68 may then be communicated through the trailing edge
cavity 95 of the airfoil 152 and into the gas path 62.
[0060] Although the different non-limiting embodiments are
illustrated as having specific components, the embodiments of this
disclosure are not limited to those particular combinations. It is
possible to use some of the components or features from any of the
non-limiting embodiments in combination with features or components
from any of the other non-limiting embodiments.
[0061] It should be understood that like reference numerals
identify corresponding or similar elements throughout the several
drawings. It should also be understood that although a particular
component arrangement is disclosed and illustrated in these
exemplary embodiments, other arrangements could also benefit from
the teachings of this disclosure.
[0062] The foregoing description shall be interpreted as
illustrative and not in any limiting sense. A worker of ordinary
skill in the art would understand that certain modifications could
come within the scope of this disclosure. For these reasons, the
following claims should be studied to determine the true scope and
content of this disclosure.
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