Core assemblies and gas turbine engine components formed therefrom

Core assemblies and methods for manufacturing components of gas turbine engines include a first core body having a first trunk configured to attach to a first location of a cavity core structure, a first branch extending from the first trunk and configured to form a first portion of a first cooling circuit, the first branch having a first joining surface, and a second core body having a second trunk configured to attach to a second location of a cavity core structure, a first branch of the second core body extending from the second trunk and configured to form a first portion of a second cooling circuit in the component. The first branches of the core bodies joined to form a junction. The junction defines a merger of the first cooling circuit and the second cooling circuit proximate to an exit of the first and second cooling circuits.

BACKGROUND

The subject matter disclosed herein generally relates to gas turbine engine components and, more particularly, to core assemblies, core subassemblies, and core bodies for manufacturing components of gas turbine engines.

Turbine engine components, such as turbine blades and vanes, are operated in high temperature environments. To avoid deterioration in the components resulting from their exposure to high temperatures, it is necessary to provide cooling circuits within the components. Turbine blades and vanes are subjected to high thermal loads on both the suction and pressure sides of their airfoil portions and at both the leading and trailing edges. The regions of the airfoils having the highest thermal load can differ depending on engine design and specific operating conditions.

Refractory metal core technology offers the potential to provide higher specific cooling passages for turbine components such as blade and vane airfoils and seals. Refractory metal core technology allows cooling circuits to be placed just under the surface of the airfoil through which cooling air flows and is expelled into the gaspath. Improved cooling circuits within turbine components may be advantageous.

SUMMARY

In accordance with an embodiment, core assemblies for manufacturing components of gas turbine engines are provided. The core assemblies include a first core body having a first trunk configured to attach to a first location of a cavity core structure, a first branch of the first core body extending from the first trunk and configured to form a first portion of a first cooling circuit in the component, the first branch having a first joining surface and a second core body having a second trunk configured to attach to a second location of a cavity core structure, a first branch of the second core body extending from the second trunk and configured to form a first portion of a second cooling circuit in the component, the first branch of the second core body having a second joining surface joined to the first joining surface to form a junction. The junction defines a merger of the first cooling circuit and the second cooling circuit proximate to an exit of the first and second cooling circuits from the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the first core body includes a second branch extending from the first trunk to define a second exit of the first cooling circuit.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the second exit is formed in one of a pressure side surface or a suction side surface of the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the second core body includes a second branch extending from the second trunk to define a second exit of the second cooling circuit.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the second exit of the first cooling circuit and the second exit of the second cooling circuit are on opposite side surfaces of the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that at least one of the first core body and the second core body is a refractory metal core.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the first core body and the second core body are attached at the junction by at least one of welding, gluing, forging, pressing, laser operations, or mechanical attachment.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that at least one of the first core body and the second core body includes a plurality of openings configured to form a plurality of air disturbance features in the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the first location is on a first internal cavity core structure and the second location is on a second internal cavity core structure that is different from the first internal cavity core structure.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the first location and the second location are different locations on a single internal cavity core structure.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the core assemblies may include that the first location and the second location are different from each other.

According to another embodiment, components for gas turbine engines are provided. The components include a cavity formed inside the component and defining a cooling flow path within the component, a first cooling circuit fluidly connecting the cavity to an exterior of the component, wherein the first cooling circuit comprises a first portion and a second portion wherein the first portion of the cooling circuit and the second portion of the cooling circuit are configured to define a first exit and a second exit at two different locations on the exterior of the component, and wherein the first portion and the second portion extend from a trunk portion of the first cooling circuit, and a second cooling circuit formed within the component and merging with the first cooling circuit proximate the first exit of the first cooling circuit.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that at least one of the trunk portion, the first portion of the cooling circuit, or the second portion of the cooling circuit includes a plurality of air disturbance features in the cooling circuit.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the second cooling circuit is fluidly connected to the cavity at a location different from a location where the first cooling circuit fluidly connects to the cavity.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the second cooling circuit is fluidly connected to a second cavity different from the cavity the first cooling circuit is fluidly connected to.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the first exit is on an end of the component and the second exit of the first cooling circuit is on a pressure side surface or suction side surface of the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the second cooling circuit has a second exit separate from the location of the merging with the first cooling circuit.

According to another embodiment, methods of manufacturing components for gas turbine engines are provided. The methods include forming a core assembly having a first core body with a trunk that attaches to a cavity core structure, a first branch extending from the trunk and configured to form a first portion of a first cooling circuit in the component, and a second branch extending from the trunk and configured to form a second portion of the first cooling circuit in the component, the first branch of the first core body having a first joining surface, attaching a second core body to the first core body at the first joining surface to form a junction, the second core body configured to define a second cooling circuit within the component, and attaching the first core body to a cavity core structure. The junction of the first core body and the second core body define a merging of the first and second cooling circuits proximate an end of the component.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the methods may include forming the component having an interior cavity based on the cavity core structure and cooling circuits defined by the first and second core bodies.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the methods may include attaching the second core body to the same cavity core structure as the first core body.

In addition to one or more of the features described herein, or as an alternative, further embodiments of the methods may include attaching the second core body to a different cavity core structure than the cavity core structure the first core body is attached to.

Technical effects of embodiments of the present disclosure include core assemblies and core bodies, such as refractory metal cores, for manufacturing components of gas turbine engines having a trunk and multiple branches extending therefrom. Further technical effects include components for gas turbine engines having a cavity and a branch portion of a cooling circuit extending therefrom with multiple branch portions of the cooling circuit extending from the trunk to define multiple, different exits on an exterior of the component. Further technical effects include cooling circuits of gas turbine engine components that can start at different locations within the component and merge proximate to an exit from the component and associated cores and core bodies to form such configurations.

DETAILED DESCRIPTION

As shown and described herein, various features of the disclosure will be presented. Various embodiments may have the same or similar features and thus the same or similar features may be labeled with the same reference numeral, but preceded by a different first number indicating the figure to which the feature is shown. Thus, for example, element “a” that is shown in FIG. X may be labeled “Xa” and a similar feature in FIG. Z may be labeled “Za.” Although similar reference numbers may be used in a generic sense, various embodiments will be described and various features may include changes, alterations, modifications, etc. as will be appreciated by those of skill in the art, whether explicitly described or otherwise would be appreciated by those of skill in the art.

FIG. 1Aschematically illustrates a gas turbine engine20. The exemplary gas turbine engine20is a two-spool turbofan engine that generally incorporates a fan section22, a compressor section24, a combustor section26, and a turbine section28. Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section22drives air along a bypass flow path B, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26. Hot combustion gases generated in the combustor section26are expanded through the turbine section28. 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 engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine centerline longitudinal axis A. The low speed spool30and the high speed spool32may be mounted relative to an engine static structure33via several bearing systems31. It should be understood that other bearing systems31may alternatively or additionally be provided.

The low speed spool30generally includes an inner shaft34that interconnects a fan36, a low pressure compressor38and a low pressure turbine39. The inner shaft34can be connected to the fan36through a geared architecture45to drive the fan36at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft35that interconnects a high pressure compressor37and a high pressure turbine40. In this embodiment, the inner shaft34and the outer shaft35are supported at various axial locations by bearing systems31positioned within the engine static structure33.

A combustor42is arranged between the high pressure compressor37and the high pressure turbine40. A mid-turbine frame44may be arranged generally between the high pressure turbine40and the low pressure turbine39. The mid-turbine frame44can support one or more bearing systems31of the turbine section28. The mid-turbine frame44may include one or more airfoils46that extend within the core flow path C.

The inner shaft34and the outer shaft35are concentric and rotate via the bearing systems31about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor38and the high pressure compressor37, is mixed with fuel and burned in the combustor42, and is then expanded over the high pressure turbine40and the low pressure turbine39. The high pressure turbine40and the low pressure turbine39rotationally drive the respective high speed spool32and the low speed spool30in response to the expansion.

The pressure ratio of the low pressure turbine39can be pressure measured prior to the inlet of the low pressure turbine39as related to the pressure at the outlet of the low pressure turbine39and prior to an exhaust nozzle of the gas turbine engine20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine20is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor38, and the low pressure turbine39has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only examples 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 example gas turbine engine20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section22of the gas turbine engine20is 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 engine20at 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.

Each of the compressor section24and the turbine section28may 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 blades25, while each vane assembly can carry a plurality of vanes27that extend into the core flow path C. The blades25of the rotor assemblies extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine20along the core flow path C. The vanes27of the vane assemblies direct the core airflow to the blades25to either add or extract energy.

Various components of a gas turbine engine20, including but not limited to the airfoils of the blades25and the vanes27of the compressor section24and the turbine section28, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section28is 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 that include features such as airflow bleed ports are discussed below.

FIG. 1Bis a schematic view of a turbine section that may employ various embodiments disclosed herein. Turbine100includes a plurality of airfoils, including, for example, one or more blades101and vanes102. The airfoils101,102may be hollow bodies with internal cavities defining a number of channels or cavities, hereinafter airfoil cavities, formed therein and extending from an inner diameter106to an outer diameter108, or vice-versa. The airfoil cavities may be separated by partitions within the airfoils101,102that may extend either from the inner diameter106or the outer diameter108of the airfoil101,102. The partitions may extend for a portion of the length of the airfoil101,102, but may stop or end prior to forming a complete wall within the airfoil101,102. Thus, each of the airfoil cavities may be fluidly connected and form a fluid path within the respective airfoil101,102. The blades101and the vanes may include platforms110located proximal to the inner diameter or outer diameter thereof. Located below the platforms110may be airflow ports and/or bleed orifices that enable air to bleed from the internal cavities of the airfoils101,102. A root of the airfoil may connected to or be part of the platform110.

Although an aero or aircraft engine application is shown and described above, those of skill in the art will appreciate that airfoil configurations as described herein may be applied to industrial applications and/or industrial gas turbine engines, land based or otherwise.

As noted, turbine airfoils can operate in high temperature environments that, in some circumstances, may exceed the melting point of the material of the airfoil. In order to cool the airfoil, cooling air is passed through it from the compressor. The coolant travels through the airfoil cavities which are designed to promote convective heat transfer. The cooling air can then be discharged out the airfoil through cavity slots (e.g., exits). The air exiting the slots can form a film of cooler air along surfaces of the airfoil and thus shield the airfoil from incoming hot fluids (e.g., combustion gases).

In accordance with various embodiments of the present disclosure, discharge cooling air flows can be provided to or on multiple surfaces of an airfoil by using stacked core subassemblies or bodies to form a core assembly during formation and/or manufacture of the airfoils. In some embodiments the core bodies, core subassemblies, and/or core assemblies can include refractory metal cores (RMCs). RMCs are a tool that makes a negative in the final airfoil body and the airfoil is produced from positive material. As provided herein, the cores and/or core bodies can be bent in singular concepts, shapes, geometries, etc. and in multi-core configurations, the cores and/or core bodies can be glued, pancaked, welded, brazed, mechanically joined, or otherwise joined to create a desired stack and/or airfoil cavity configuration.

Stacks of core bodies as provided herein can be optimized for heat transfer and desired flow characteristics through and along an airfoil. The component (e.g., an airfoil), in some embodiments, can be additive manufactured with the desired internal cavity and/or flow path geometries to allow discharge on more than one surface and/or at multiple locations on a single surface of the airfoil. Accordingly, advantageously, discharge of cooling air can be provided onto multiple surfaces of an airfoil thus improving convective and conductive heat transfer by utilizing a single cavity and reducing the number and complexity of multiple cores, resulting in efficient film cooling.

Turning now toFIGS. 2A and 2B, schematic illustrations of an airfoil (FIG. 2A) and a core assembly (FIG. 2B) are shown.FIG. 2Ais a cross-sectional illustration of an airfoil201cast using the core assembly212shown inFIG. 2B.FIG. 2Bis a perspective illustration of the core assembly212used to form the airfoil201ofFIG. 2A;

FIG. 2Aillustrates a cross-sectional illustration of the airfoil201cast using the core assembly212illustrated inFIG. 2B. Airfoil201includes leading edge surface214, trailing edge216, pressure side surface218, suction side surface220, leading edge cavity222, midchord cavity224, trailing edge cavity226, and cooling circuits228a,228b, and228c. As a result of manufacturing process, leading edge cavity222is formed by a leading edge ceramic core, midchord cavity224is formed by a midchord ceramic core, and trailing edge cavity226is formed by a trailing edge ceramic core (see, e.g.,FIG. 2B). Each cavity222,224,226is bounded by a respective cavity wall222a,224a,226a. One or more ribs 230 separate the cavities222,224,226. Each cooling circuit228a,228b,228cis formed by one or more core bodies, such as RMCs (seeFIG. 2B). As shown inFIG. 2A, cooling circuits228a,228bare positioned between a downstream cavity (e.g.,224,226) and a side (e.g., pressure side surface218) of airfoil201.

As shown, cooling circuit228afluidly connects leading edge cavity222along a downstream portion of cavity wall222aand extends between midchord cavity224and pressure side surface218. As such, cooling fluid flowing through leading edge cavity222exits the cavity222and flows through cooling circuit228ato cool the pressure side surface218of airfoil201. Similarly, cooling circuit228bjoins with midchord cavity224along a downstream portion of cavity wall224aand extends between trailing edge cavity226and pressure side surface218. Cooling fluid exits midchord cavity224and flows through cooling circuit228bto cool the pressure side surface218of airfoil201farther downstream of cooling circuit228a. WhileFIG. 2Aillustrates cooling circuits near the pressure side surface218of airfoil201, cooling circuits can also be located between the cavities222,224,226and the suction side surface220of airfoil201.

FIG. 2Billustrates a perspective view of one embodiment of a core assembly212for forming an airfoil201. Core assembly212includes leading edge ceramic core232, midchord ceramic core234, and trailing edge ceramic core236. A first core body238ais configured with the leading edge ceramic core232, a second core body238bis configured with the midchord ceramic core234, and a third core body238cis configured with the trailing edge ceramic core236. The ceramic cores232,234,236are used to form inner passages (e.g., cavities222,224,226) for cooling fluid within the airfoil201. The core bodies238a,238b,238care used to form cooling circuits (e.g., a network of cooling passages including cooling circuits228a,228b,228c) within the airfoil201. The cooling circuits228a,228b,228cin the cast airfoil201will receive cooling fluid from the inner passage(s)222,224,226with which they are fluidly connected. In order for the cooling circuits228a,228b,228cin the cast airfoil201to receive cooling fluid from the inner passages222,224,226, the ceramic cores232,234,236and the core bodies238a,238b,238care in contact with one another. The core bodies238a,238b,238care secured to the appropriate ceramic core232,234,236to maintain contact during a casting process. In some embodiments, core assembly212can contain more than one midchord ceramic core234and associated downstream core body238b.

In some non-limiting embodiments, the core bodies of the present disclosure, described above and below, can be refractory metal cores. However, those of skill in the art will appreciate that other materials can be used to form the core bodies without departing from the scope of the present disclosure.

Each of the core bodies238a,238b,238ccan include a plurality of openings240, as shown. Once cast, openings240form a plurality of air disturbance features, include pedestals or other features, which direct cooling fluid through a respective cooling circuit228a,228b,228c. Openings240can be circular, oblong, racetrack-shaped, teardrop-shaped, or any other shape depending on the flow control needs of the specific cooling circuit228a,228b,228c. Although described above with respect to casting, those of skill in the art will appreciate that other manufacturing processes can be used without departing from the scope of the present disclosure. For example, additive manufacturing techniques can be used to form the structures and configurations of airfoils as provided herein.

Turning now toFIGS. 3A-4B, schematic illustrations of core bodies and/or core subassemblies in accordance with non-limiting embodiments of the present disclosure are shown.FIGS. 3A-3Eare top-down illustrations of core bodies and core subassemblies in various configurations.FIG. 4Ais a top-down illustration of a core body in accordance with a non-limiting embodiment andFIG. 4Bis an elevational illustration of the core body ofFIG. 4A. The core bodies and subassemblies ofFIGS. 3A-4Bcan be used with ceramic cores or other structures to form core assemblies and define a positive structure of the internal cavities (i.e., negative space) within an airfoil, such as described above.

FIG. 3Ais a first illustration of a core subassembly338athat is constructed of two core bodies that are attached together. For example, as shown, a first core body342aand a second core body344aare bonded together to form the core subassembly338a, thus forming a core body stack. The attachment between the first core body342aand the second core body344acan be by any known mechanism including, but not limited to, gluing, brazing, pancaking, welding (e.g., friction, heat, etc.), laser operations, forging, pressing, mechanical fixing, and/or other joining processes or mechanisms. Further, in some embodiments, the core bodies and/or the core assembly can be additively manufactured.

The core subassembly338aincludes a first portion346aand a second portion348a. In the embodiment ofFIG. 3A, the first portion346aforms a trunk350athat is defined as the portion or section of the core subassembly338awhere the first core body342aand the second core body344aare attached. The trunk350ahas a first end352aand a second end354a. The trunk350ais a structure or portion of the core subassembly338athat is configured to join with or attach to a ceramic core or other core structure used to form a cavity within an airfoil. The first end352aof the trunk350ais thus free to be engaged with or otherwise interact with a cavity core structure and the second end354ais opposite therefrom.

The second portion348ais defined by one or more branches356athat extend from the trunk350awhere the first core body342aand the second core body344aare not joined or attached (i.e., are separated from each other). In this embodiment, the first core body342adefines one branch356aand the second core body344adefines another branch356a, each extending from the second end354aof the trunk350a.

During manufacture of an airfoil based on, in part, the core subassembly338a, the trunk350awill form a relatively wide cooling circuit that can extend from a cavity of the airfoil (e.g., as shown and described above). However, due to the configuration and structure of the core subassembly338a, the cooling circuit can include multiple passages that extend to different locations and/or surfaces of the airfoil and thus provide cooling at multiple locations on the exterior of the airfoil. The multiple passages are based on the configuration of the branches356aof the second portion348aof the core subassembly338a.

FIG. 3Bshows another configuration of a core subassembly338bin accordance with an embodiment of the present disclosure. The core subassembly338bis substantially similar to the core subassembly338aofFIG. 3Aand forms a similar cooling circuit in a manufactured airfoil. However, as shown inFIG. 3B, the core subassembly338bis a unitary or single core body342b. In this case, a trunk350bis a portion of the single core body342band the branches356bextend therefrom, with each branch356bpart of the single core body342b. The trunk350bis the portion of the core subassembly338bthat is configured to connect to or join with a ceramic core or other core structure that is used to form cavities of an airfoil.

FIG. 3Cshows another configuration of a core subassembly338cin accordance with an embodiment of the present disclosure. In this embodiment, a first core body342cand multiple second core bodies344care connected on one side of the first core body342c. In this configuration, a trunk350chas multiple branches356cextending therefrom. As shown, one of the branches356cextends from a point between a first end352cand a second end354cof the trunk350c. In an alternative configuration, the two second core bodies344cshown inFIG. 3Ccan be formed as a single, second core body without departing from the scope of the present disclosure.

FIG. 3Dshows another configuration of a core subassembly338din accordance with an embodiment of the present disclosure. The core subassembly338dofFIG. 3Dis a unitary, single core body342dwith a trunk350dand multiple branches356dextending from the trunk350d. As shown in the embodiment ofFIG. 3D, the branches356dextend at different angles and form a “Y” configuration with the trunk350d.

FIG. 3Eshows another configuration of a core subassembly338ein accordance with an embodiment of the present disclosure.FIG. 3Eillustrates that multiple second core bodies344ecan be attached to a first core body342e. As shown, one second core body344eis attached on a first side of the first core body342eand multiple second core bodies344eare attached on a second (and opposite) side of the first core body342e. The trunk350eis defined as any section of the core subassembly338ewhere different core bodies342e,344eare attached. As shown, the branches356ecan extend from the trunk350efrom multiple locations along the length of the trunk350e.

Turning now toFIGS. 4A-4B, another configuration of a core subassembly438in accordance with an embodiment of the present disclosure is shown.FIG. 4Ashows a top-down illustration of the core subassembly438andFIG. 4Bshows an elevational illustration of the core subassembly438. As shown, the core subassembly438includes a single core body442defining a trunk450and a plurality of branches456a,456bextending therefrom. A first set of branches456aare configured to extend in a first direction relative to the trunk450and a second set of branches456bare configured to extend in a second direction relative to the trunk450. As shown, each branch456a,456bis angled from a second end454of the trunk450(as illustrated by the dashed line inFIGS. 4A-4B). In the embodiment ofFIGS. 4A-4B, the two sets of branches456a,456bare shown alternating in configuration. However, those of skill in the art will appreciate that any pattern or configuration of branches can be employed without departing from the scope of the present disclosure.

Those of skill in the art will appreciate that various alternative configurations and/or geometries are enabled by the present disclosure. Further, several of the above described embodiments can be combined and/or altered to form a desired cooling circuit within an airfoil. In some embodiments, a portion of the trunk (e.g., where at least two core bodies are joined) can extend the full length of the core assembly, such that a wider cooling circuit passageway can be formed for the length of the trunk. Additionally, in some configurations, the trunk can bend, turn, or otherwise have a different geometry than a relatively straight line/body, as shown above.

Further, in some embodiments, one or more of the core bodies used to form a core subassembly or core assembly of the present disclosure can include openings (e.g., openings240ofFIG. 2B) to form a plurality of pedestals or other features that direct cooling fluid through a respective cooling circuit. The openings can be circular, oblong, racetrack-shaped, teardrop-shaped, or any other shape. Further, in some embodiments, two joined core bodies can be configured to match or align openings of the two core bodies. Moreover, in some embodiments, one core body can include openings while the other core body does not include openings, thus forming a unique interior structure to the cooling circuits when the airfoil is formed from the core assembly including the two different core bodies. Still further, multiple of the core bodies can be configured with openings that do not align, or some openings that align and other that do not, thus enabling unique pedestal structures and/or configurations.

Turning now toFIGS. 5-7, schematic illustrations of airfoils formed from core assemblies and subassemblies as provided herein are shown.FIGS. 5-6each show a trailing edge of respective airfoils andFIG. 7shows an airfoil extending from a leading edge to a trailing edge. Each ofFIGS. 5-7is a cross-sectional, top-down view of the interior structure of the respective airfoils.

With reference toFIG. 5, airfoil501includes a trailing edge cavity526that is fluidly connected to exterior surfaces of the airfoil501by cooling circuit528. As shown, cooling circuit528has two exits with one exit configured near the trailing edge516but being open on a pressure side surface518. The other of the exits of the cooling circuit528is formed in the trailing edge516of the airfoil501. The airfoil501is manufactured using at least one core subassembly similar to that shown and described above. For example, the configuration and geometry of the cooling circuit528ofFIG. 5could be formed using a core subassembly similar to that shown in eitherFIG. 3AorFIG. 3B. As shown, the trunk of the core subassembly would be connected to a ceramic core that forms the trailing edge cavity526, and the branches of the core subassembly would extend toward the trailing edge.

FIG. 6shows an alternative configuration of a trailing edge616of an airfoil601in accordance with an embodiment of the present disclosure. The airfoil601includes a trailing edge cavity626with a cooling circuit628extending from the trailing edge cavity626toward the trailing edge616. However, in this embodiment, as shown, the cooling circuit628has two exits that open onto each of the pressure surface side618and the suction surface side620. The cooling circuit628can be formed, for example, by a core subassembly similar to that shown inFIG. 3D.

Turning now toFIG. 7, an airfoil701is shown. The airfoil701extends from a leading edge714to a trailing edge716. As shown, the airfoil701includes a leading edge cavity722, two midchord cavities724, and a trailing edge cavity726. The trailing edge cavity726includes two separate cooling circuits728c, with one exiting onto a pressure side surface718and one exiting at the trailing edge716. Further, as shown in the embodiment ofFIG. 7, the airfoil701does not include any cooling circuits connected to the midchord cavities724, although those of skill in the art will appreciate that cooling circuits could be formed therewith (e.g., fluidly exiting from the midchord cavities724to a suction side surface720.

The leading edge cavity722includes multiple cooling circuits728a. As shown, a first cooling circuit728a′ can connect the leading edge cavity722to the pressure side surface718by a single passaged cooling circuit. Additionally, the leading edge cavity722is fluidly connected to the pressure side surface718by a second cooling circuit728a″ that is formed by a core subassembly in accordance with the present disclosure. As shown, the second cooling circuit728a″ has a larger section near the leading edge cavity722(e.g., formed by the trunk of the RMC) and two separate exits exiting onto the pressure surface side718(e.g., each formed by a branch of the core subassembly).

Turning now toFIG. 8, a flow process for forming or manufacturing an airfoil having cooling circuits as shown and described above is shown. The flow process800involves a casting process for the airfoil based on ceramic cores and core assemblies using RMCs. However, alternative manufacturing techniques can be used to form an airfoil having internal structures and/or configurations as described herein. For example, although discussed with respect to RMC, this is merely for explanatory purposes and other materials or types of subassemblies can be employed without departing from the scope of the present disclosure. Thus, the flow process800is not intended to be limiting, but rather is provided for illustrative purposes.

At block802, an RMC having a trunk and branches formed thereon is formed. In some configurations, the formation of the RMC can be by additive manufacturing, with the trunk and branches integrally formed in a single piece or component. In other embodiments, the formation of the RMC can involve attaching or joining multiple RMC bodies to form the RMC having a trunk at the portions where at least two RMC bodies are joined or attached and branches where an RMC body is not attached to another RMC body. The attachment or joining of the RMC bodies can be by any known means and can include welding, gluing, laser operations, mechanically fixing, etc. In some configurations, a branch (as described above) can be attached to another RMC body and the trunk can be a portion of the RMC body that is configured to interact with a cavity core structure, as shown and described above. In some embodiments, the branches can be formed by bending a portion or portions of the RMC body (e.g., as shown inFIGS. 4A-4B).

The RMC can then be attached to a cavity core structure, as shown at block804. In some embodiments, the cavity core structure may be a ceramic core. The attachment between the RMC and the cavity core structure may be by any means, as will be appreciated by those of skill in the art.

At block806, an airfoil can be formed from the RMC and cavity core structure. The formed airfoil includes cooling circuits fluidly connecting internal cavities to exterior surfaces of the airfoil at multiple locations based on the branches of the RMC. That is, the cavity core structure can form the internal cavities and the RMC (trunk and branches) can form the cooling circuits, as shown and described herein.

Turning now toFIG. 9, a schematic illustration of an airfoil901having features in accordance with a non-limiting embodiment of the present disclosure is shown. The airfoil901extends from a leading edge (not shown) to a trailing edge916. As shown, the airfoil901includes a trailing edge cavity926. The trailing edge cavity926includes two separate cooling circuits960a,960bthat fluidly connect to the trailing edge cavity926at two separate locations962a,962b(e.g., a pressure side opening962aand a suction side opening962b). As shown, the two separate cooling circuits960a,960bfluidly connect at a junction964and exit at the trailing edge916at a trailing edge exit966.

Further, as shown in the embodiment ofFIG. 9, each of the cooling circuits960a,960bincludes a side exit (e.g., pressure side exit968aor suction side exit968b). For example, in this embodiment, as shown, a pressure side cooling circuit960aon a pressure side of the airfoil901has a pressure side exit968athat opens on or through a pressure side surface918of the airfoil901. Similarly, a suction side cooling circuit960bon a suction side of the airfoil901has a suction side exit968bthat opens on or through a suction side surface920of the airfoil901.

The separate and then joined cooling circuits960a,960bcan be formed within the airfoil901through use of core assemblies, core subassemblies, and/or core bodies similar to that shown and described above. For example, turning now toFIGS. 10A-10B, schematic illustrations of a core subassembly1070used to form the cooling circuits shown inFIG. 9are shown.FIG. 10Ais a cross-sectional view illustration of a first core body1070aand a second core body1070bthat are separated, but can be joined to form the cooling circuit described above with respect toFIG. 9.FIG. 10Bis an isometric illustration of the first core body1070a, with the second core body1070bbeing a mirror of the first core body1070a. The first and second core bodies1070a,1070bcan be joined to form the core subassembly1070.

As shown inFIGS. 10A-10B, the core bodies1070a,1070beach have a respective trunk1072a,1072bthat is configured to join with or attach to a ceramic core or other core structure used to form a cavity within an airfoil. One or more first branches1074a,1074band one or more second branches1076a,1076bcan extend from the respective trunks1072a,1072b(e.g., laterally or vertically on the refractory metal core body). As will be appreciated by those of skill in the art, the first branches1074a,1074bcan form the portion of the cooling circuit within the airfoil that exits at a trailing edge of the airfoil (e.g., as shown inFIG. 9). The second branches1076a,1076bcan form portions of a cooling circuit that exits on a pressure or suction side of the airfoil. For example, the second branch1076aof the first core body1070acan be used to form a cooling circuit that exits an airfoil on the pressure side surface of the airfoil. Similarly, for example, the second branch1076bof the second core body1070bcan be used to form a cooling circuit that exits the airfoil on the suction side surface of the airfoil.

To form the core subassembly1070, the two core bodies1070a,1070bare joined or attached. For example, a first joining surface1078acan be defined on the first branch1074aof the first core body1070aand a second joining surface1078bcan be defined on the first branch1074bof the second core body1070b. The joining surfaces1078a,1078bcan be configured to enable joining of the first core body1070ato the second core body1070bby means of glue, pancaking (e.g., press-fusion), welding, brazing, mechanical joining, or other joining mechanism or procedure to create a desired stack and thus form the core subassembly1070.

As shown inFIG. 10B, the first core body1070ais illustrated in an isometric view, illustrating the nature of the first and second branches1074a,1076a. As shown, the second branches1076aare bent or skewed relative to the first branches1074a, and thus enable different exits for cooling cavities on one or more surfaces of an airfoil formed based on the core subassembly1070.

Turning now toFIG. 11, an airfoil1101is shown. The airfoil1101extends from a leading edge (not shown) to a trailing edge1116. As shown, the airfoil1101includes a midchord cavity1124and a trailing edge cavity1126. In this embodiment, each of the cavities1124,1126, is fluidly connected to a respective cooling circuit1180,1182. As shown, the two cooling circuits1180,1182fluidly connect at a junction1164and exit at a trailing edge1116at a trailing edge exit1166of the airfoil1101. Each of the cooling circuits1180,1182can be formed, in part, by separate core bodies or core subassemblies, as described above. However, as shown and similar to that described with respect toFIGS. 9-10B, the two core bodies can be joined toward a trailing edge to form the trailing edge exit1166where the junction1164is formed and the two cooling circuits1180,1182merge.

Further, as shown in the embodiment ofFIG. 11, each of the cooling circuits1180,1182includes one or more side exits (e.g., pressure side exit1184a,1184bor suction side exit1186). For example, in this embodiment, as shown, a pressure side cooling circuit1180on a pressure side of the airfoil1101has a first pressure side exit1184aand a second pressure side exit1184bthat each open on or through a pressure side surface1118of the airfoil1101. A suction side cooling circuit1182on a suction side of the airfoil1101has a suction side exit1186that opens on or through a suction side surface1120of the airfoil1101.

As shown inFIGS. 9-11, two or more core bodies or core subassemblies can be configured to fluidly connect to one or more internal cavities of an airfoil, and then join proximate to an exit from the airfoil, thus forming cooling circuits as shown and described herein. Although certain geometries are shown and described herein, those of skill in the art will appreciate that various other geometries of core assemblies, core subassemblies, core bodies, and/or cooling circuits can be formed using embodiments of the present disclosure.

Advantageously, embodiments described herein can provide improved high temperature applications for airfoil. For example, using core subassemblies as provided herein can be employed to optimize pressure side film cooling, while allowing for a more conventional serpentine cavity (e.g., midchord cavities) to be dedicated on the suction side of the airfoil.

The double-stack core subassembly (or stack of core bodies) of some embodiments provided herein can be configured to provide superposition of slot film effectiveness on the pressure side of the airfoil and greatly benefit the trailing edge temperatures. Further, multiple core body and/or core subassembly insertions into the ceramic cores can be minimized. That is, core subassemblies and core bodies as provided herein can be joined to a ceramic core at a single location (e.g., single trunk) and still provide multiple exits (e.g., branches) at various locations on the exteriors surfaces of the airfoil.

Further, advantageously, embodiments provided herein can provide cooling discharge on pressure side, suction side, trailing edge, top and/or bottom platform, and/or combinations thereof. That is, advantageously, cooling flow is enabled on multiple sides of a component from a single internal cavity of the component. Advantageously, such cooling can enable product life improvement which can decrease product life cycle costs.

For example, although shown and described with respect to airfoils (e.g., vanes and blades) embodiments provided herein can be used in the manufacture of blade outer air seals, combustor panels, or other components that employ fluid cooling. Moreover, although primarily described with respect to conventional casting, additive manufacturing and machining methods can be used without departing from the scope of the present disclosure.