Patent Description:
Turbomachines are utilized in a variety of industries and applications for energy transfer purposes. For example, a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section progressively increases the pressure of a working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and a fuel (e.g., natural gas) mix within the combustion section and burn in a combustion chamber to generate high pressure and high temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where they expand to produce work. For example, expansion of the combustion gases in the turbine section may rotate a rotor shaft connected, e.g., to a generator to produce electricity. The spent combustion gases then exit the gas turbine via the exhaust section.

During operation of the turbomachine, various hot gas path components in the system are subjected to high temperature flows, which can cause the hot gas path components to fail. Since higher temperature flows generally result in increased performance, efficiency, and power output of the turbomachine, the hot gas path components that are subjected to high temperature flows must be cooled to allow the gas turbine system to operate with flows at increased temperatures.

As the maximum local temperature of the hot gas path components approaches the melting temperature of the hot gas path components, forced air cooling becomes necessary. For this reason, airfoils of turbine rotor blades and stationary nozzles often require complex cooling schemes in which air, typically bleed air from the compressor section, is forced through internal cooling passages within the airfoil and then discharged through cooling holes at the airfoil surface to transfer heat from the hot gas path component.

Many complex cooling schemes use small cooling passages, or microchannels, to deliver cooling fluid through the airfoil. Such cooling schemes present a considerable fabrication challenge for cores and castings, which can significantly increase the manufacturing cost of the hot gas path components using such known near wall cooling systems. To address the fabrication challenges with complex and/or small cooling channels near the component surface, many hot gas path components with such features may be additively manufactured. Additive manufacturing is capable of producing components with intricate and varied cooling features. However, additively manufacturing a hot gas path component, such as a rotor blade or stator vane, as a single component may be costly and time-consuming. Additionally, manufacturing errors in a single portion of the hot gas path component may result in the scrapping of the entire component.

As such, manufacturing a hot gas path component as multiple sub-components may be advantageous. However, due to the complex geometries of the sub-components, joining the sub-components to form the hot gas path component may be difficult. Additionally, the joints formed between the sub-components may be particularly weak and/or fail when exposed to the hot combustion gases produced during operation of the turbomachine and transmitted through the hot gas path through the turbine section.

Accordingly, an improved hot gas path component, having one or more subcomponents joined together and capable of being subject to hot combustion gases without risk of joint failure, is desired and would be appreciated in the art. <CIT> discloses a stator assembly comprising an annular platform defining a boundary of an annulus and a plurality of guide vanes for arranging in a circumferential array on the annular platform. <CIT> discloses a modular blade or vane for a gas turbine. <CIT> discloses a turbine vane for a gas turbine engine incorporating a ceramic matrix composite airfoil.

Aspects and advantages of the stator vanes and turbomachines in accordance with the present invention will be set forth in part in the following description, or may be obvious from the description.

In accordance with the invention, a stator vane according to claim <NUM> is provided.

In accordance with another embodiment, a turbomachine is provided. The turbomachine includes a compressor section, a combustion section, and a turbine section. At least one stator vane is disposed in the turbine section. The at least one stator vane includes a platform that defines an opening. The stator vane further includes an airfoil that has a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base and a tip. At least one of the base or the tip includes a protrusion. The protrusion extends into the opening of the platform such that the platform surrounds the protrusion of the airfoil. The stator vane further includes a braze joint disposed between and fixedly coupling the platform and the protrusion of the airfoil. The stator vane further includes a cooling circuit defined in one of the protrusion or the platform to cool the braze joint.

These and other features, aspects, and advantages of the present stator vanes and turbomachines will become better understood with reference to the following description and appended claims.

A full and enabling disclosure of the present stator vanes and turbomachines, including the best mode of making and using the present systems and methods, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:.

Reference now will be made in detail to embodiments of the present stator vanes and turbomachines, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation, rather than limitation of, the technology. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present technology without departing from the scope or spirit of the claimed technology. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.

The term "fluid" may be a gas or a liquid. The term "fluid communication" means that a fluid is capable of making the connection between the areas specified.

As used herein, the terms "upstream" (or "forward") and "downstream" (or "aft") refer to the relative direction with respect to fluid flow in a fluid pathway. The term "radially" refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component; the term "axially" refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component; and the term "circumferentially" refers to the relative direction that extends around the axial centerline of a particular component.

Terms of approximation, such as "about," "approximately," "generally," and "substantially," are not to be limited to the precise value specified. For example, the approximating language may refer to being within a <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, or <NUM> percent margin in either individual values, range(s) of values and/or endpoints defining range(s) of values. When used in the context of an angle or direction, such terms include within ten degrees greater or less than the stated angle or direction. For example, "generally vertical" includes directions within ten degrees of vertical in any direction, e.g., clockwise or counter-clockwise.

For example, a process, method, article, or apparatus that comprises a list of features is not necessarily limited only to those features but may include other features that are not expressly listed or that are inherent to such process, method, article, or apparatus. Further, unless expressly stated to the contrary, "or" refers to an inclusive- or and not to an exclusive- or. For example, a condition "A or B" is satisfied by any one of the following: A is true (or present) and B is false (or not present), A is false (or not present) and B is true (or present), and both A and B are true (or present).

Referring now to the drawings, <FIG> illustrates a schematic diagram of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine engine <NUM>. Although an industrial or land-based gas turbine is shown and described herein, the present disclosure is not limited to an industrial or land-based gas turbine engine, unless otherwise specified in the claims. For example, the invention as described herein may be used in any type of turbomachine including, but not limited to, a steam turbine, an aircraft gas turbine, or a marine gas turbine.

As shown in <FIG>, the gas turbine <NUM> generally includes a compressor section <NUM>. The compressor section <NUM> includes a compressor <NUM>. The compressor includes an inlet <NUM> that is disposed at an upstream end of the gas turbine engine <NUM>. The gas turbine engine <NUM> further includes a combustion section <NUM> having one or more combustors <NUM> disposed downstream from the compressor section <NUM>. The gas turbine engine <NUM> further includes a turbine section <NUM> (i.e., an expansion turbine) that is downstream from the combustion section <NUM>. A shaft <NUM> extends generally axially through the gas turbine engine <NUM>.

The compressor section <NUM> may generally include a plurality of rotor disks <NUM> and a plurality of rotor blades <NUM> extending radially outwardly from and connected to each rotor disk <NUM>. Each rotor disk <NUM> in turn may be coupled to or form a portion of the shaft <NUM> that extends through the compressor section <NUM>. The rotor blades <NUM> of the compressor section <NUM> may include turbomachine airfoils that define an airfoil shape (e.g., having a leading edge, a trailing edge, and side walls extending between the leading edge and the trailing edge).

The turbine section <NUM> may generally include a plurality of rotor disks <NUM> and a plurality of rotor blades <NUM> extending radially outwardly from and being interconnected to each rotor disk <NUM>. Each rotor disk <NUM> in turn may be coupled to or form a portion of the shaft <NUM> that extends through the turbine section <NUM>. The turbine section <NUM> further includes an outer casing <NUM> that circumferentially surrounds the portion of the shaft <NUM> and the rotor blades <NUM>. The turbine section <NUM> may include stator vanes or stationary nozzles <NUM> extending radially inward from the outer casing <NUM>. The rotor blades <NUM> and stator vanes <NUM> may be arranged in alternating stages along an axial centerline <NUM> of gas turbine <NUM>. Both the rotor blades <NUM> and the stator vanes <NUM> may include turbomachine airfoils that define an airfoil shape (e.g., having a leading edge, a trailing edge, and side walls extending between the leading edge and the trailing edge).

In operation, ambient air <NUM> or other working fluid is drawn into the inlet <NUM> of the compressor <NUM> and is progressively compressed to provide a compressed air <NUM> to the combustion section <NUM>. The compressed air <NUM> flows into the combustion section <NUM> and is mixed with fuel to form a combustible mixture. The combustible mixture is burned within a combustion chamber <NUM> of the combustor <NUM>, thereby generating combustion gases <NUM> that flow from the combustion chamber <NUM> into the turbine section <NUM>. Energy (kinetic and/or thermal) is transferred from the combustion gases <NUM> to the rotor blades <NUM>, causing the shaft <NUM> to rotate and produce mechanical work. The spent combustion gases <NUM> (also called "exhaust gases") exit the turbine section <NUM> and flow through the exhaust diffuser <NUM> across a plurality of struts or main airfoils <NUM> that are disposed within the exhaust diffuser <NUM>.

The gas turbine <NUM> may define a cylindrical coordinate system having an axial direction A extending along the axial centerline <NUM>, a radial direction R perpendicular to the axial centerline <NUM>, and a circumferential direction C extending around the axial centerline <NUM>.

<FIG> is a partial cross-sectional side view of the turbine section <NUM> of the gas turbine engine <NUM>, in accordance with embodiments of the present disclosure. The turbine section <NUM> may include one or more stages <NUM> that each include a set of rotor blades <NUM> coupled to a rotor disk <NUM> that may be rotatably attached to the shaft <NUM>. The one or more stages <NUM> may further include a set of stator vanes <NUM>. The stator vane <NUM> described herein may be employed in a first stage, a second stage, a third stage, or combinations thereof, in which "first" refers to the stage immediately downstream of the combustion section <NUM>.

Each stator vane <NUM> may include at least one airfoil <NUM> that extends in the radial direction R between an inner platform or endwall <NUM> and an outer platform or endwall <NUM>. The circumferentially adjacent outer platforms <NUM> of each stator vane <NUM> may be coupled together to form an outer annular ring extending around an inner annular ring of the circumferentially adjacent inner platforms <NUM> of each stator vane <NUM>. The at least one airfoil <NUM> may extend between the two annular rings formed by the platforms <NUM>, <NUM>. The turbine section <NUM> may also include shroud segments <NUM>, which may be disposed downstream of the outer platform <NUM> to direct combustion gases <NUM> flowing past the stator vanes <NUM> to the rotor blades <NUM>.

Structures or components disposed along the flow path of the combustion gases <NUM> may be referred to as hot gas path components. In one example, the hot gas path component may be the stator vane <NUM> and/or the rotor blade <NUM>. In some embodiments, to cool the hot gas path components, cooling features, such as impingement sleeves, cooling channels, cooling holes, etc. may be disposed within the hot gas path components, as indicated by the dashed line <NUM>. For example, cooling air as indicated by an arrow <NUM> may be routed from the compressor section <NUM> or elsewhere and directed through the cooling features as indicated by arrows <NUM>.

<FIG> illustrates an exploded view of a stator vane <NUM>, in which the inner platform <NUM>, the outer platform <NUM>, and the airfoil are separated from one another along an axial centerline <NUM> of the stator vane <NUM>, in accordance with embodiments of the present disclosure. The axial centerline <NUM> may be generally parallel to the radial direction of the gas turbine <NUM> when the stator vane <NUM> is installed in the turbine section <NUM>. As shown, the stator vane <NUM> may include an inner platform <NUM>, an outer platform <NUM>, and an airfoil <NUM> extending between the inner platform <NUM> and the outer platform <NUM>. The airfoil <NUM> may define a generally aerodynamic shape or contour. For example, the airfoil <NUM> may include a leading edge <NUM> that engages a flow of combustion gases and side walls that guide the combustion gases along the airfoil <NUM> to a trailing edge <NUM>. For example, the airfoil <NUM> may include a pressure side wall <NUM> and a suction side wall <NUM> that each extend from the leading edge <NUM> to the trailing edge <NUM>. Additionally, the airfoil <NUM> may extend radially between a base <NUM> and a tip <NUM>. The base <NUM> is configured to be coupled to the inner platform <NUM>, and the tip <NUM> is configured to be coupled to the outer platform <NUM>.

In exemplary embodiments, the inner platform <NUM>, the airfoil <NUM>, and the outer platform <NUM> may be separate components (e.g., manufactured as separate components) that are brazed or welded to one another via one or more braze joints. For example, in particular embodiments, the inner platform <NUM>, the airfoil <NUM>, and the outer platform <NUM> may each be separately additively manufactured (e.g., 3D printed) and subsequently joined to one another via braze joints. Forming the stator vane <NUM> as three separate components advantageously increases the repairability of the stator vane <NUM>. For example, if a portion of the airfoil <NUM> is damaged, then the entire stator vane <NUM> would not need to be replaced. Rather, the braze connection could be undone (e.g., via reheating and melting the braze joint) to decouple the airfoil <NUM> from the platforms <NUM>, <NUM>, and a new airfoil <NUM> could be employed or the old airfoil could be repaired.

Additionally, in instances where additive manufacturing is used to produce some portion of or all of the stator vane <NUM>, the build (i.e., print) time of the airfoil <NUM> as a separate component from the inner platform <NUM> and the outer platform <NUM> is significantly shorter as compared to the build time of an integral nozzle in which the inner and outer platforms <NUM>, <NUM> are printed with the airfoil <NUM>. Moreover, forming the stator vane <NUM> from three separate components permits different manufacturing techniques and/or different materials to be used for the various components.

As shown in <FIG>, the airfoil <NUM> may include a camber axis <NUM>, which may extend from the leading edge <NUM> to the trailing edge <NUM> and which may be defined halfway between the pressure side wall <NUM> and the suction side wall <NUM>. The camber axis <NUM> may be curved and/or contoured to correspond with the curve of the pressure side wall <NUM> and the suction side wall <NUM>. A transverse direction T may be defined orthogonally with respect to the camber axis <NUM>. As shown in <FIG>, the stator vane <NUM> may further include one or more protrusions <NUM> that extend outwardly from the airfoil <NUM> in both the radial direction R and in the transverse direction T. For example, the stator vane <NUM> may include a first protrusion <NUM> that extends from the airfoil <NUM> at the base <NUM> and a second protrusion <NUM> that extends from the airfoil <NUM> at the tip <NUM>. Each protrusion <NUM> may extend generally perpendicularly from the pressure side wall <NUM> and the suction side wall <NUM> in the transverse direction T.

Both the inner platform <NUM> and the outer platform <NUM> may define an opening <NUM>. The opening <NUM> may be sized and shaped to correspond with the protrusion <NUM>, such that the protrusion <NUM> may be inserted into each opening <NUM> (and subsequently brazed to the respective platform), thereby coupling the airfoil <NUM> to the inner platform <NUM> and the outer platform <NUM>. Particularly, the protrusion <NUM> extending from the base <NUM> of the airfoil <NUM> may be inserted into the opening <NUM> of the inner platform <NUM> (and subsequently brazed to the inner platform <NUM>), and the protrusion <NUM> of extending from the tip <NUM> of the airfoil <NUM> may be inserted into the opening <NUM> of the outer platform <NUM> (and subsequently brazed to the outer platform <NUM>).

For example, the protrusions <NUM> of the airfoil <NUM> may be brazedly coupled to the inner and outer platforms <NUM> and <NUM> via a braze joint <NUM> (<FIG> and <FIG>). That is, as shown in <FIG>, a braze material <NUM> may be disposed between the protrusion <NUM> and the boundary of the opening <NUM> (at both the inner and outer platforms <NUM> and <NUM>). The entire stator vane <NUM> may then be placed in a braze oven to melt the braze material <NUM>, and subsequently, the braze material <NUM> may solidify thereby joining the inner and outer platforms <NUM> and <NUM> to a respective protrusion <NUM> of the airfoil <NUM>. The braze material may contain copper, nickel, silver, gold, aluminum, or other suitable braze metals.

In many embodiments, as shown in <FIG>, the inner platform <NUM> and/or the outer platform <NUM> includes a main body <NUM>. The main body <NUM> may extend generally perpendicularly to the airfoil <NUM>. The main body <NUM> of the inner platform <NUM> may define a radially inward flow boundary for combustion gases in the turbine section <NUM>. Similarly, the main body <NUM> of the outer platform <NUM> may define a radially outer flow boundary for combustion gases in the turbine section <NUM>. Additionally, the main body <NUM> of the inner platform <NUM> and the outer platform <NUM> may at least partially define the respective openings <NUM>. In exemplary embodiments, one or both of the inner platform <NUM> and/or the outer platform <NUM> may include a raised wall <NUM> extending radially from the main body <NUM>. The raised wall <NUM> may at least partially define the opening <NUM>. In many embodiments, the protrusion <NUM> may be brazedly coupled to the raised wall <NUM>. In some embodiments, the inner and outer platforms <NUM> and <NUM> may include one or more rails <NUM> (which may extend across the main body <NUM> generally in the circumferential direction C).

In certain embodiments, the protrusion <NUM> that extends from the base <NUM> of the airfoil <NUM> may extend into the opening <NUM> of the inner platform <NUM> such that the inner platform <NUM> surrounds the protrusion <NUM> of the airfoil <NUM>. Particularly, the main body <NUM> and/or the raised wall <NUM> of the inner platform <NUM> may surround the protrusion <NUM> extending from the base <NUM> of the airfoil <NUM> (i.e., the raised wall <NUM> extends around a perimeter of the protrusion <NUM>). Similarly, the protrusion <NUM> that extends from the tip <NUM> of the airfoil <NUM> may extend into the opening <NUM> of the outer platform <NUM> such that the outer platform <NUM> surrounds the protrusion <NUM> of the airfoil <NUM>. Particularly, the main body <NUM> and/or the raised wall <NUM> of the outer platform <NUM> may surround the protrusion <NUM> extending from the tip <NUM> of the airfoil <NUM>.

In many embodiments, the protrusion <NUM> of the airfoil <NUM> may include a leading edge portion <NUM>, a trailing edge portion <NUM>, a pressure side portion <NUM>, and a suction side portion <NUM>. The leading edge portion <NUM>, the pressure side portion <NUM>, and the suction side portion <NUM> may at least partially define a cavity <NUM> that extends into and is further defined in the airfoil <NUM>. The cavity <NUM> may be exposed by the opening <NUM> of the platform <NUM>, such that air (e.g., bleed air from the compressor <NUM>) may enter the cavity <NUM>. A rib <NUM> (<FIG>, <FIG>) may extend across the cavity <NUM>, e.g., between the pressure side portion <NUM> and the suction side portion <NUM> of the protrusion <NUM>, and between the pressure side wall <NUM> and the suction side wall <NUM> of the airfoil <NUM>, to partition the cavity <NUM> into multiple portions and to provide structural support for the airfoil <NUM>. The protrusion <NUM> may further include a solid tail <NUM>, which may provide a flat, and/or smooth, surface for additional braze material to bond the protrusion <NUM> to the platform <NUM> (as shown by the grid in <FIG>).

The leading edge portion <NUM> of the protrusion <NUM> may extend from the leading edge <NUM> of the airfoil <NUM>. The trailing edge portion <NUM> of the protrusion <NUM> may extend from the trailing edge <NUM> of the airfoil <NUM>. The pressure side portion <NUM> of the protrusion <NUM> may extend from the pressure side wall <NUM> of the airfoil <NUM>. The suction side portion <NUM> of the protrusion <NUM> may extend from the suction side wall <NUM> of the airfoil <NUM>.

<FIG> illustrates an enlarged perspective view of a stator vane <NUM>, in accordance with embodiments of the present disclosure. As shown, the protrusion <NUM> of the airfoil <NUM> may extend into the opening <NUM> of the platform <NUM> such that the platform <NUM> surrounds the protrusion <NUM> of the airfoil <NUM>. A braze joint <NUM> (as shown by the dashed lines arranged in grids) is disposed between and fixedly couples the platform <NUM> and the protrusion <NUM> of the airfoil <NUM>.

<FIG> each illustrate a cross-sectional view of a stator vane <NUM> in accordance with one or more embodiments of the present disclosure. Particularly, <FIG> may be a cross-sectional view of a stator vane <NUM> from along the line <NUM>-<NUM> shown in <FIG>, and <FIG> may be an enlarged planar view of the stator vane <NUM> in <FIG>. As shown, the stator vane <NUM> may include a platform <NUM> (such as the inner platform <NUM> or the outer platform <NUM> shown in <FIG>). The platform <NUM> may define an opening <NUM>, and an airfoil <NUM> may extend into the opening <NUM> and couple to the platform <NUM>. A braze joint <NUM> may be disposed between the airfoil <NUM> and the platform <NUM>, thereby coupling the airfoil <NUM> to the platform <NUM>. The braze joint <NUM> may include a braze material <NUM> that bonds the platform <NUM> to the airfoil <NUM>. The braze material <NUM> may have a melting temperature that is lower than a melting temperature of a material from which the airfoil <NUM> and/or the platform <NUM> are formed. For example, the braze material <NUM> may be copper, nickel, silver, gold, aluminum, or combinations thereof. The airfoil <NUM> and the platform <NUM> may be formed from a metal material having a high melting temperature, such as nickel and cobalt alloys or other materials.

In exemplary embodiments, the stator vane <NUM> may include a cooling circuit <NUM> defined in one of the airfoil <NUM> and/or the platform <NUM> to cool the braze joint <NUM> (e.g., during operation of the gas turbine engine <NUM>). In some embodiments, the cooling circuit <NUM> may be defined in one or both of the protrusion <NUM> and/or the airfoil <NUM>. In other embodiments, the cooling circuit <NUM> may be defined in the platform <NUM>, such as in one or more of the main body <NUM>, the raised wall <NUM>, and/or the rail <NUM> described above with reference to <FIG>. The cooling circuit <NUM> may be in fluid communication with the compressor section <NUM>, such that the cooling circuit <NUM> may utilize a flow of compressed air from the compressor to cool the braze joint <NUM>.

In many embodiments, the platform <NUM> may include an inner wall <NUM> and an outer wall <NUM>, and the raised wall <NUM> may extend between the inner wall <NUM> and the outer wall <NUM>. In exemplary embodiments, as shown, the braze joint <NUM> may be disposed between, and fixedly couple, the raised wall <NUM> of the platform <NUM> to the protrusion <NUM> of the airfoil <NUM>. In many embodiments, an inlet plenum <NUM> may be defined between the inner wall <NUM>, the outer wall <NUM>, and the raised wall <NUM>. One or more inlet apertures <NUM> may be defined in the outer wall <NUM> to allow compressed air to enter the inlet plenum <NUM> (e.g., from a compressed air source such as bleed air from the compressor section <NUM>). The inner wall <NUM> may define an outer surface <NUM> partially defining the inlet plenum <NUM> and an inner surface <NUM> partially defining a hot gas path boundary of the combustion gases. In other words, the inner surface <NUM> may be exposed to the combustion gases during operation of the gas turbine <NUM>.

In many embodiments, as shown in <FIG> and <FIG>, the cooling circuit <NUM> may include one or more inlet channels <NUM>, a plenum <NUM>, and one or more outlet channels <NUM> each defined in the raised wall <NUM> and/or the inner wall <NUM> of the platform <NUM>. The one or more inlet channels <NUM> may extend from the inlet plenum <NUM> to a plenum <NUM>. In some embodiments, the plenum <NUM> may be annular, such that it surrounds the protrusion <NUM> (thereby surrounding the axial centerline <NUM> shown in <FIG>). Each inlet channel <NUM> of the one or more inlet channels <NUM> may be sized and oriented to direct coolant or air <NUM> to impinge upon a joint boundary wall <NUM>. The joint boundary wall <NUM> may form a portion of the raised wall <NUM> and may be disposed between the plenum <NUM> and the braze joint <NUM>. For example, the joint boundary wall <NUM> of the raised wall <NUM> may be oriented generally radially, and the inlet channel <NUM> may include an axially oriented portion immediately upstream of an outlet of the inlet channel <NUM>, such that coolant exits the inlet channel <NUM> traveling generally axially (or perpendicularly to the joint boundary wall <NUM>). The air <NUM> may then impinge upon, or strike, the joint boundary wall <NUM>, thereby transferring heat from the joint boundary wall <NUM> to the air <NUM>. Because the joint boundary wall <NUM> may be a portion of the raised wall <NUM>, the joint boundary wall <NUM> may both contact the braze joint <NUM> and partially define the plenum <NUM>.

Particularly, inlet channels <NUM> may be sized and oriented to direct the air <NUM> in discrete jets to impinge upon the joint boundary wall <NUM>. The discrete jets of air <NUM> may impinge (or strike) an impingement surface <NUM> of the joint boundary wall <NUM> and cool the impingement surface <NUM>, which allows for heat transfer from the joint boundary wall <NUM> and the braze joint <NUM> to the air. The impingement surface <NUM> may partially define a boundary of the plenum <NUM>. For example, the inlet channels <NUM> may include an exit portion that extends generally perpendicularly to the impingement surface <NUM> of the joint boundary wall <NUM>, such that the discrete jets of air <NUM> exiting the inlet channels <NUM> are perpendicular to the surface upon which they strike, e.g., the impingement surface <NUM> of the joint boundary wall <NUM>. Once the discrete jets of air have impinged upon the impingement surface <NUM>, they may be referred to as "post-impingement air" and/or "spent cooling air" because the air has undergone an energy transfer and therefore has different characteristics (e.g., higher temperature and lower pressure than prior to impingement).

In various embodiments, the plenum <NUM> may be generally rectangularly shaped in cross-section. For example, the plenum <NUM> may be bounded by an upstream surface <NUM> (with respect to the flow of coolant thought the cooling circuit <NUM>), the impingement surface <NUM> opposite the upstream surface <NUM>, a radially outer surface <NUM>, and a radially inner surface <NUM>. The radially outer surface <NUM> and the radially inner surface <NUM> may be spaced apart (e.g., radially spaced apart) and may each extend generally parallel to one another. The impingement surface <NUM> and the upstream surface <NUM> may be spaced apart from one another, generally parallel to one another and may each extend generally radially.

In exemplary embodiments, the cooling circuit <NUM> may include one or more outlet channels <NUM> defined in the joint boundary wall <NUM> (e.g., defined partially in the raised wall <NUM> and defined partially in the inner wall <NUM>) and in fluid communication with the plenum <NUM>. The outlet channel <NUM> may extend from an inlet <NUM> defined in the radially outer surface <NUM> and in fluid communication with the plenum <NUM> to an outlet <NUM> defined in the inner surface <NUM> of the inner wall <NUM>. The one or more outlet channels <NUM> may include at least a portion that extends between the impingement surface <NUM> and the braze joint <NUM> to cool the braze joint <NUM> (and the braze material <NUM>). For example, the one or more outlet channels <NUM> may include a U-shaped portion <NUM> and a straight portion <NUM> (which may extend radially). The U-shaped portion <NUM> may extend from the inlet <NUM> defined in the radially outer surface <NUM> to the straight portion <NUM>. The straight portion <NUM> may be disposed axially between the impingement surface <NUM> and the braze joint <NUM>, and the straight portion <NUM> of the outlet channel <NUM> may extend from the U-shaped portion <NUM> to an outlet <NUM> defined in the inner surface <NUM> of the inner wall <NUM>. In exemplary implementations, the air <NUM> exiting the outlet <NUM> of the outlet channel <NUM> may film cool the inner surface <NUM> (and/or the airfoil <NUM>).

Referring now specifically to the embodiment shown in <FIG>, the one or more outlet channels <NUM> and the one or more inlet channels <NUM> may be alternately arranged. For example, the one or more outlet channels <NUM> and the one or more inlet channels <NUM> may not lie in the same plane defined by the radial direction R and the direction of the cooling air <NUM> exiting the inlet channel <NUM>. Particularly, each outlet channel <NUM> of the one or more outlet channels <NUM> may be disposed between two inlet channels <NUM> of the one or more inlet channels <NUM>. Similarly, each inlet channel <NUM> of the one or more inlet channels <NUM> may be disposed between two outlet channels <NUM> of the one or more outlet channels <NUM>. Stated otherwise, the inlet channels <NUM> and the outlet channels <NUM> may be spaced apart from one another with respect to a direction <NUM> that is perpendicular to the cooling air <NUM> exiting the inlet channel <NUM>. In some embodiments, the direction <NUM> may be generally parallel to the upstream surface <NUM> partially defining plenum <NUM>. Particularly, the direction <NUM> may extend along a perimeter of the protrusion <NUM>. For example, in many embodiments, each outlet channel <NUM> of the one or more outlet channels <NUM> may be offset from each inlet channel <NUM> (e.g., each neighboring inlet channel <NUM>) of the one or more inlet channels <NUM> both radially (as shown in <FIG>) and in the direction <NUM> that extends along a perimeter of the protrusion <NUM> (as shown in <FIG>).

As shown in <FIG>, the joint boundary wall <NUM> may include one or more projections <NUM> extending from the impingement surface <NUM>. For example, the projections <NUM> may extend into the plenum <NUM> towards the upstream surface <NUM>. In exemplary embodiments, the outlet channels <NUM> may be at least partially defined in the projections <NUM> of the joint boundary wall <NUM>. Particularly, the straight portion <NUM> of the outlet channels <NUM> may be defined in the projections <NUM> of the joint boundary wall <NUM>. Each of the projections <NUM> may have a semi-circular cross-sectional shape (or other suitable cross-sectional shape). An advantage of the embodiment of <FIG> is that the impingement surface <NUM> and the straight portion <NUM> of the outlet channel <NUM> may be disposed in closer proximity to the braze joint <NUM>, thereby improving the cooling of the braze joint <NUM>.

As shown in <FIG>, in some embodiments, the cooling circuit <NUM> may include a direct channel <NUM> extending between an inlet <NUM> defined in the raised wall <NUM> (or in the inner wall <NUM> in other embodiments) to an outlet <NUM>. In such embodiments, the direct channel <NUM> may be disposed proximate to the braze joint <NUM> to provide convective cooling to the braze joint <NUM>. Additionally, as shown in <FIG>, the platform <NUM> may further include a raised outlet portion <NUM> extending radially inward from the inner surface <NUM> of the inner wall <NUM> of the platform <NUM> in an area radially inward of the raised wall <NUM> and proximate to the braze joint <NUM>. The direct channel <NUM> may be defined at least partially in the raised outlet portion <NUM>. For example, the outlet <NUM> of the direct channel <NUM> may be defined in the raised outlet portion <NUM> of the platform <NUM>. The raised outlet portion <NUM> may advantageously prevent any melted braze material <NUM> from wicking or sliding into the direct channel <NUM> during braze application and brazing. As should be appreciated, the raised outlet portion <NUM> may be incorporated into one or more of the outlet channels <NUM> described above with reference to <FIG>, in order to prevent braze material <NUM> from entering the cooling circuit <NUM> during braze application and brazing.

Referring now to <FIG>, the cooling circuit <NUM> may be defined in the protrusion <NUM> of the airfoil <NUM> rather than (or in addition to) the cooling circuit <NUM> being defined in the platform(s) <NUM>, <NUM>, in accordance with embodiments of the present disclosure. In such embodiments, the cooling circuit <NUM> may be disposed in the protrusion <NUM> proximate the braze joint <NUM>, to provide cooling (e.g., convective cooling) thereto during operation of the gas turbine <NUM>. <FIG> illustrates a perspective view of an airfoil <NUM> decoupled from the inner and outer platforms <NUM>, <NUM>, in order to show details of the cooling circuit <NUM>. <FIG> illustrates a cross-sectional perspective view of the stator vane <NUM> having a cooling circuit <NUM> defined in the protrusion <NUM>. <FIG> illustrates an enlarged cross-sectional view of the stator vane <NUM> shown in <FIG>.

As shown, the cooling circuit <NUM> may include a cooling passage <NUM> defined in the protrusion <NUM> and extending between an inlet <NUM> and an outlet <NUM>. Particularly, the inlet <NUM> may be defined in a leading edge portion <NUM> of the protrusion <NUM>, and the outlet may be defined in a trailing edge portion <NUM> of the protrusion <NUM>. In this way, the passage may extend through the pressure side portion <NUM> (and/or the suction side portion <NUM>, as shown in <FIG>) of the protrusion from the inlet <NUM> to the outlet <NUM>. The cooling passage <NUM> along the pressure side portion <NUM> may have its own inlet <NUM> and its own outlet <NUM> relative to the cooling passage <NUM> along the suction side portion <NUM>, or the cooling passages <NUM> along the pressure side portion <NUM> and the suction side portion <NUM> may share a common inlet <NUM> and/or a common outlet <NUM>. Additionally, while a single cooling passage <NUM> is shown along each of the pressure side portion <NUM> and the suction side portion <NUM>, it should be understood that more than one cooling passage <NUM> may be disposed in one or both of the pressure side portion <NUM> and the suction side portion <NUM> (e.g., as radially stacked passages).

While <FIG> illustrate a cooling passage <NUM> defined in the protrusion <NUM> at the tip <NUM> of the airfoil <NUM>, it should be appreciated that the cooling passage <NUM> may alternatively (or additionally) be employed in the protrusion <NUM> at the base <NUM> of the airfoil <NUM>. Because the braze joint <NUM> bonds the protrusion <NUM> to the outer platform <NUM>, and the cooling passage <NUM> extends through the protrusion <NUM> proximate the braze joint <NUM>, the cooling passage <NUM> may advantageously provide convective cooling to the braze joint <NUM> during operation of the gas turbine <NUM>. Any aspects of the cooling passage <NUM> described above with respect to tip <NUM> of the airfoil <NUM> are equally applicable to cooling passage(s) <NUM> in the base <NUM> of the airfoil <NUM>.

Referring now to <FIG>, cooling circuit <NUM> may be defined in the protrusion <NUM> rather than (or in addition to) the cooling circuit <NUM> being defined in the platform(s) <NUM>, <NUM>, in accordance with embodiments of the present disclosure. In such embodiments, the cooling circuit <NUM> may be disposed in the protrusion <NUM> proximate the braze joint <NUM>, to provide convective cooling thereto during operation of the gas turbine <NUM>. <FIG> illustrates a perspective view of an airfoil <NUM> decoupled from the inner platforms <NUM>, <NUM>, in order to show details of the cooling circuit <NUM>. , and <FIG> illustrates an enlarged cross-sectional view of the stator vane shown in <FIG>.

As shown, the cooling circuit <NUM> may include a cooling passage <NUM> defined in the protrusion <NUM> and extending between an inlet <NUM> and a plurality of outlets <NUM>. Particularly, the inlet <NUM> may be defined in a leading edge portion <NUM> of the protrusion <NUM>, and the plurality of outlets <NUM> may be defined in one of the pressure side wall <NUM> or the suction side wall <NUM>. In this way, the cooling passage <NUM> may include a first portion <NUM> that extends through the pressure side portion <NUM> (or the suction side portion <NUM>) of the protrusion <NUM> and a second portion <NUM> that extends through the pressure side wall <NUM> (or the suction side wall <NUM>) of the airfoil <NUM>. The second portion <NUM> of the cooling passage <NUM> may be oriented generally radially and may be disposed closer to the trailing edge <NUM> than the leading edge <NUM> of the airfoil <NUM>.

As shown in <FIG>, the plurality of outlets <NUM> may be radially spaced apart from one another and may extend through the pressure side wall <NUM> (or the suction side wall <NUM>). In this way, the first portion <NUM> of the cooling passage <NUM> may provide convective cooling to the braze joint <NUM>, and the second portion <NUM> of the cooling passage <NUM> may provide film cooling to one of the pressure side wall <NUM> and/or the suction side wall <NUM> via air <NUM> exiting the plurality of outlets <NUM>.

Referring now to <FIG>, the stator vane may further include a cap <NUM> and an insert <NUM>. The cap <NUM> may be coupled to the protrusion <NUM> and may extend across the cavity <NUM> defined by the protrusion <NUM> and the airfoil <NUM>. For example, the cap <NUM> may provide a barrier and/or boundary for air outside of the cavity <NUM>. The cap may be a generally flat plate that is shaped as an airfoil (e.g., an airfoil without a trailing edge). For example, the cap <NUM> may include a leading edge that contacts the leading edge portion <NUM> of the protrusion <NUM>, a pressure side that contacts the pressure side portion <NUM> of the protrusion <NUM>, and a suction side that contacts the suction side portion <NUM> of the protrusion <NUM>. The insert <NUM> may be coupled to the protrusion <NUM> and may extend into the cavity <NUM>. The insert may include a first portion <NUM> that extends generally parallel to the cap <NUM> and a second portion <NUM> that extends generally perpendicularly to the cap <NUM> (e.g., the second portion <NUM> may extend radially into the cavity <NUM>).

In exemplary embodiments, the cap <NUM> and the insert <NUM> may define an outlet plenum <NUM> within the cavity <NUM> of the stator vane <NUM>. The cooling circuit <NUM> may be in fluid communication with the outlet plenum <NUM>. For example, as shown in <FIG>, the cooling circuit <NUM> may be defined in the protrusion <NUM> rather than (or in addition to) the cooling circuit <NUM> being defined in the platform(s) <NUM>, <NUM>, in accordance with embodiments of the present disclosure. In such embodiments, the cooling circuit <NUM> may be disposed in the protrusion <NUM> proximate the braze joint <NUM>, to provide convective cooling thereto during operation of the gas turbine engine <NUM>. As shown in <FIG>, the cooling circuit <NUM> may include an inlet channel <NUM> and an outlet channel <NUM>. The inlet channel <NUM> may extend from an inlet <NUM> defined in a radially outer surface <NUM> of the protrusion <NUM>, towards the braze joint <NUM>, to the outlet channel <NUM>. The outlet channel <NUM> may extend from the inlet channel <NUM> to an outlet <NUM> defined in an interior surface <NUM> of the protrusion <NUM>. The outlet <NUM> may exhaust air from the cooling circuit <NUM> into the outlet plenum <NUM>.

In various embodiments, as shown collectively in <FIG> and <FIG>, the protrusion <NUM> may include a plurality of cooling circuits <NUM> each in fluid communication with the outlet plenum <NUM>. For example, each cooling circuit <NUM> may extend from an inlet <NUM> defined in the radially outer surface <NUM> of the protrusion <NUM> to the outlet <NUM> defined in the interior surface <NUM> of the protrusion <NUM>. For example, as shown in <FIG>, the plurality of cooling circuits <NUM> may be spaced apart from one another (e.g., equally or unequally spaced apart) around a perimeter of the protrusion <NUM>. Particularly, the inlets <NUM> of each cooling circuit <NUM> of the plurality of cooling circuits <NUM> may be spaced apart from one another around the perimeter of the protrusion <NUM>.

In many embodiments, an annular plenum <NUM> may be defined between the insert <NUM> and a wall <NUM> (e.g., the pressure side wall <NUM> and/or the suction side wall <NUM>) of the airfoil <NUM>. For example, the insert <NUM> may be annular, such that the annular plenum <NUM> is defined around the entire perimeter of the airfoil <NUM> between the wall <NUM> and the insert <NUM>. In exemplary embodiments, the annular plenum <NUM> may be fluidly coupled to the outlet plenum <NUM> via an insert aperture <NUM>. For example, the insert aperture <NUM> may be defined in the second portion <NUM> of the insert <NUM> to fluidly couple the outlet plenum <NUM> to the annular plenum <NUM>, thereby providing air to the annular plenum <NUM> for cooling the wall <NUM> during operation of the turbomachine.

Alternatively, or additionally, as shown in <FIG>, the cooling circuit <NUM> may include a plurality of inlet channels <NUM>, a plenum <NUM>, and an outlet channel <NUM>. A joint boundary wall <NUM> may be disposed between the plenum <NUM> and the braze joint <NUM>, and an impingement surface <NUM> of the joint boundary wall <NUM> may partially define the plenum <NUM>. Each inlet channel <NUM> may extend from an inlet <NUM> defined in the interior surface <NUM> of the protrusion <NUM> to the plenum <NUM>, and the outlet channel <NUM> may extend from the plenum <NUM> to an outlet defined in the interior surface <NUM>. Particularly, inlet channels <NUM> may be sized and oriented to direct the air in discrete jets to impinge upon an impingement surface <NUM>. The discrete jets of air may impinge (or strike) an impingement surface <NUM> which allows for heat transfer between the air and the braze joint <NUM>.

Claim 1:
A stator vane (<NUM>) for a turbomachine comprising:
a platform (<NUM>, <NUM>, <NUM>) defining an opening (<NUM>);
an airfoil (<NUM>) having a leading edge (<NUM>), a trailing edge (<NUM>), a suction side wall (<NUM>), and a pressure side wall (<NUM>), the airfoil (<NUM>) extending radially between a base (<NUM>) and a tip (<NUM>), wherein at least one of the base (<NUM>) or the tip (<NUM>) includes a protrusion (<NUM>), the protrusion (<NUM>) extending into the opening (<NUM>) of the platform (<NUM>, <NUM>) such that the platform (<NUM>, <NUM>) surrounds the protrusion (<NUM>) of the airfoil (<NUM>);
a braze joint (<NUM>) disposed between and fixedly coupling the platform (<NUM>, <NUM>, <NUM>) and the protrusion (<NUM>) of the airfoil (<NUM>); and
a cooling circuit (<NUM>) defined in the protrusion (<NUM>) or the platform (<NUM>, <NUM>, <NUM>) to cool the braze joint (<NUM>), characterized in having:
the cooling circuit (<NUM>) comprising one or more inlet channels (<NUM>), a plenum (<NUM>), and one or more outlet channels (<NUM>),
wherein the inlet channels (<NUM>), the outlet channels (<NUM>) and the plenum (<NUM>) are each defined in the platform (<NUM>, <NUM>, <NUM>), or wherein the inlet channels (<NUM>), the outlet channels (<NUM>) and the plenum (<NUM>) are each defined in the protrusion (<NUM>).