US20140075947A1

US20140075947A1 - Gas turbine engine component cooling circuit

Info

Publication number: US20140075947A1
Application number: US13/621,968
Authority: US
Inventors: Steven Bruce Gautschi; Lane Thornton
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 2012-09-18
Filing date: 2012-09-18
Publication date: 2014-03-20
Also published as: EP2898203A1; WO2014047022A1; EP2898203A4

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.

Description

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularly to a cooling circuit for cooling a gas turbine engine component.
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.
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

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.
In a further non-limiting embodiment of the foregoing component, the component is a vane.
In a further non-limiting embodiment of either of the foregoing components, the component is a blade.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
In a further non-limiting embodiment of the foregoing gas turbine engine, the at least one component is a vane.
In a further non-limiting embodiment of either of the foregoing gas turbine engines, the first rib includes a plurality of openings.
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.
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.
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.
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.
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.
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

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbine engine.

FIG. 2 illustrates a component that can be incorporated into a gas turbine engine.

FIG. 3 illustrates a cross-sectional view through section A-A of the component of FIG. 2.

FIG. 4 illustrates a cooling circuit that can be incorporated into an airfoil of a component.

FIG. 5 illustrates various features of a cooling circuit that can be incorporated into an airfoil of a component.

FIGS. 6A, 6B and 6C schematically illustrate cooling an airfoil using an exemplary cooling circuit.

DETAILED DESCRIPTION

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.
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.
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.
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.
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.
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.
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.
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.
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)^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).
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.
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.
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).
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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.
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).
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).
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.
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.
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.
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.
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.

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:

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.