AIRFOIL FOR TURBINE SYSTEM

Abstract

An airfoil for a turbine system is disclosed. The airfoil includes a first body having exterior surfaces defining a first portion of an aerodynamic contour of the airfoil and formed from a first material. The airfoil further includes a second body having exterior surfaces defining a second portion of an aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

Description

FIELD OF THE INVENTION

The subject matter disclosed herein relates generally to turbine systems, and more particularly to airfoils for turbine systems.

BACKGROUND OF THE INVENTION

Turbine systems are widely utilized in fields such as power generation. For example, a conventional gas turbine system includes a compressor section, a combustor section, and at least one turbine section. The compressor section is configured to compress air as the air flows through the compressor section. The air is then flowed from the compressor section to the combustor section, where it is mixed with fuel and combusted, generating a hot gas flow. The hot gas flow is provided to the turbine section, which utilizes the hot gas flow by extracting energy from it to power the compressor, an electrical generator, and other various loads.

Various components, such as buckets and nozzles, are typically included in the various sections of the turbine system for interacting with flows through these sections. For example, various stages of buckets and nozzles may extend into the hot gas path of the turbine section of a turbine system. During operation of the turbine system, the portions of such components that are exposed to a hot gas path may be at risk of damage due to the high temperatures in the hot gas path. Thus, such components generally require cooling.

Previous attempts to cool components such as buckets and nozzles have involved the use of cooling passages internal to the buckets and nozzles. Cooling fluids are flowed through these internal cooling passages to cool portions of the external surfaces of the buckets and nozzles. However, these cooling attempts have a variety of disadvantages. For example, in many cases it may be difficult to manufacture cooling passages within the buckets and nozzles that are sufficiently close to some portions of the external surfaces to cool those surfaces. For example, the surfaces at and near to the trailing edges of the airfoils portions of buckets and nozzles may be particularly difficult to cool, due to a lack of internal space in which to manufacture a cooling passage.

Accordingly, improved hot gas path components would be desired in the art. For example, an improved airfoil for a hot gas path component would be advantageous.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

In one embodiment, an airfoil for a turbine system is disclosed. The airfoil includes a first body having exterior surfaces defining a first portion of an aerodynamic contour of the airfoil and formed from a first material. The airfoil further includes a second body having exterior surfaces defining a second portion of an aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

In another embodiment, a nozzle for a turbine section of a turbine system is disclosed. The nozzle includes an airfoil having exterior surfaces defining an aerodynamic contour, the aerodynamic contour comprising a pressure side and a suction side extending between a leading edge and a trailing edge. The airfoil includes a first body having exterior surfaces defining a first portion of the aerodynamic contour of the airfoil and formed from a first material. The airfoil further includes a second body having exterior surfaces defining a second portion of the aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a schematic view of a gas turbine system according to one embodiment of the present disclosure;

FIG. 2 is a cross-sectional view of a turbine section of a gas turbine system according to one embodiment of the present disclosure;

FIG. 3 is a close-up cross-sectional view of various components of a turbine section of a gas turbine system according to one embodiment of the present disclosure;

FIG. 4 is a top cross-sectional view of an airfoil according to one embodiment of the present disclosure;

FIG. 5 is a partial perspective view of a plurality of nozzles according to one embodiment of the present disclosure; and

FIG. 6 is a cross-sectional view of a nozzle according to one embodiment of the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.

FIG. 1 is a schematic diagram of a gas turbine system 10. It should be understood that the turbine system 10 of the present disclosure need not be a gas turbine system 10, but rather may be any suitable turbine system 10, such as a steam turbine system or other suitable system. The gas turbine system 10 may include a compressor section 12, a combustor section 14, and a turbine section 16. The compressor section 12 and turbine section 16 may be coupled by a shaft 18. The shaft 18 may be a single shaft or a plurality of shaft segments coupled together to form shaft 18.

As is generally known in the art, air or another suitable working fluid is flowed through and compressed in the compressor section 12. The compressed working fluid is then supplied to the combustor section 14, wherein it is combined with fuel and combusted, creating hot gases of combustion. After the hot gases of combustion are flowed through the combustor section 14, they may be flowed into the turbine section 18.

FIG. 2 illustrates one embodiment of portions of a turbine section 18 according to the present disclosure. A hot gas path 20 may be defined within the turbine section 18. Various hot gas path components, such as shrouds 22, nozzles 24, and buckets 26, may be at least partially disposed in the hot gas path 20.

For example, as shown, the turbine section 18 may include a plurality of buckets 26 and a plurality of nozzles 24. Each of the plurality of buckets 26 and nozzles 24 may be at least partially disposed in the hot gas path 20. Further, the plurality of buckets 26 and the plurality of nozzles 24 may be disposed in one or more annular arrays, each of which may define a portion of the hot gas path 20.

The turbine section 16 may include a plurality of turbine stages. Each stage may include a plurality of buckets 26 disposed in an annular array and a plurality of nozzles 24 disposed in an annular array. For example, in one embodiment, the turbine section 16 may have three stages, as shown in FIG. 2. For example, a first stage of the turbine section 16 may include a first stage nozzle assembly 31 and a first stage buckets assembly 32. The nozzles assembly 31 may include a plurality of nozzles 24 disposed and fixed circumferentially about the shaft 18. The bucket assembly 32 may include a plurality of buckets 26 disposed circumferentially about the shaft 18 and coupled to the shaft 18. A second stage of the turbine section 16 may include a second stage nozzle assembly 33 and a second stage buckets assembly 34. The nozzles 24 included in the nozzle assembly 33 may be disposed and fixed circumferentially about the shaft 18. The buckets 26 included in the bucket assembly 34 may be disposed circumferentially about the shaft 18 and coupled to the shaft 18. The second stage nozzle assembly 33 is thus positioned between the first stage bucket assembly 32 and second stage bucket assembly 34 along the hot gas path 20. A third stage of the turbine section 16 may include a third stage nozzle assembly 35 and a third stage bucket assembly 36. The nozzles 24 included in the nozzle assembly 35 may be disposed and fixed circumferentially about the shaft 18. The buckets 26 included in the bucket assembly 36 may be disposed circumferentially about the shaft 18 and coupled to the shaft 18. The third stage nozzle assembly 35 is thus positioned between the second stage bucket assembly 34 and third stage bucket assembly 36 along the hot gas path 20.

It should be understood that the turbine section 16 is not limited to three stages, but rather that any number of stages are within the scope and spirit of the present disclosure.

It should be understood that hot gas path components according to the present disclosure are not limited to components in turbine sections 16. Rather, hot gas path components may be components at least partially disposed in flow paths for compressor sections 12 or any other suitable sections of a system 10.

As further shown in FIGS. 2 and 3, each hot gas path component, such as each nozzle 24 and each bucket 26, may include an airfoil 40. For example, airfoils 40 of buckets 26 may extend outward from shanks 42, while airfoils 40 of nozzles 24 may extend outward from end caps 44. An airfoil 40 according to the present disclosure has exterior surfaces defining a generally aerodynamic contour. For example, the exterior surfaces may define a pressure side 52 and a suction side 54 each extending between a leading edge 56 and a trailing edge 58.

As shown in FIGS. 3 through 6, an airfoil 40 according to the present disclosure may further include a first body 62 and a second body 64. Each body 62, 64 may define a portion of the exterior surfaces of the airfoil 40. Thus, for example, the first body 62 may have exterior surfaces defining a first portion of the aerodynamic contour of the airfoil 40, while the second body 64 has exterior surfaces defining a second portion of the aerodynamic contour of the airfoil 40.

In exemplary embodiments, the exterior surfaces of the first body 62 may define at least a portion of the pressure side 52 and suction side 54 as well as the leading edge 56, while the exterior surfaces of the second body 64 may define the trailing edge 58. The exterior surfaces of the second body 64 may further define the remaining portions of the pressure side 52 and suction side 54, or these entire sides may be defined by the first body 62. In other embodiments, however, the first body 62 may include exterior surfaces defining any one or more of the pressure side 52, suction side 54, leading edge 56, and trailing edge 58, or any portion thereof, while the second body 64 may include exterior surfaces defining the remaining such sides and/or edges, or portions thereof.

The first body 62 according to the present disclosure is formed from a first material. The first material may be any suitable material, such as a suitable metal or metal alloy. For example, in some embodiments, the first material may include aluminum, nickel, iron, carbon, chromium, and/or any other suitable metal. It should further be understood that the first material is not limited to metals, but rather that the first material may be formed from any suitable material.

Further, in exemplary embodiments as shown in FIG. 4, the first body 62 may define one or more cooling passages 65 therein. The cooling passages 65 may be generally linear or curvilinear, and may extend through any suitable portions of the first body 62. For example, in some embodiments, the cooling passages 65 may be generally serpentine cooling passages. Such cooling passages 65 may provide convective cooling, impingement cooling, and/or any other suitable form of cooling.

The second body 64 of an airfoil 40 according to the present disclosure is formed from a second material. The second material according to the present disclosure has a different temperature capability than the first material. Temperature capability means ability to withstand a certain temperature before failing, such that failure of the material occurs at such temperature. Thus, a second material fails at a temperature that is different from the temperature at which the first material fails. In exemplary embodiments, the second material has a temperature capability that is higher than that of the first material.

The second material may be any suitable metal or metal alloy, as discussed above, or a suitable non-metal material, alloy, or composite. For example, in exemplary embodiments, the second material is a ceramic matrix composite (“CMC”) material. CMC materials are designed to withstand relatively increased temperatures, such as those temperatures in a hot gas path 20 during operation of a system 10. CMC materials are typically formed from ceramic fibers embedded in a ceramic matrix. The fibers and/or matrix may be formed from carbon, silicon carbide, alumina, mullite, or any other suitable materials. The use of CMC materials or other suitable second materials to form the second body 64 may, if the second materials have higher temperature capabilities than the first materials, advantageously allow the second body 64 to withstand the temperatures in the hot gas path 20 during operation. Thus, no cooling of the second body 64 may be required. This is particularly advantageous in embodiments wherein the second body 64 includes the trailing edge 58 of the airfoil 40, which may otherwise not be adequately cooled during operation due to the lack of internal space in the airfoil 40 in which to manufacture cooling passages or other suitable cooling apparatus.

It should be understood that a body formed from a first material or second material may include other materials that are covered with a layer of the first or second material, or may be formed solely from a first or second material.

In some exemplary embodiments, as shown in FIG. 4, the second body 64 is continuous throughout a cross-sectional profile. Thus, the second body 64 has a generally solid interior with no passages or other bore holes or apertures therein. In other embodiments, however, suitable passages, bore holes, or apertures may be defined through the second body 64 or portions thereof.

Further, in some embodiments, the second body 64 may be a single, unitary component. In other embodiments as shown in FIG. 3, however, the second body 64 may comprise a plurality of second body sections 66. Each section 66 may be formed from a second material. It should be understood that the second materials utilized for each section 66 may vary or be identical. The sections 66 may be stacked on one another in the generally radial direction, as shown, or may otherwise abut one another to form the second body 64. The use of a plurality of second body sections 66 to form the second body 64 may advantageously allow for thermal growth of the second body 64 before, during, and after operation of the system 10 due to changes in temperatures in the hot gas path 20.

As shown in FIGS. 4 through 6, the second body 64 may be coupled to the first body 62. When coupled together, the second body 64 and first body 62 may form the generally aerodynamic contour of the airfoil 40.

In some embodiments, for example, one of the first body 62 and the second body 64 may define a recess 72, while the other of the first body 62 and the second body 64 defines a mating protrusion 74. As shown in FIG. 4, for example, the recess 72 may be a dovetail recess, while the protrusion 74 is a dovetail protrusion. It should be understood, however, that the present disclosure is not limited to dovetails, and rather that any suitable female and male portions may comprise the recess 72 and protrusion 74. Engagement of the protrusion 74 in the recess 72 may mate them together to couple the first body 62 to the second body 64, as shown.

In other embodiments, end caps 44 or other suitable coupling apparatus may couple the second body 64 to the first body 62. For example, as shown in FIGS. 5 and 6, an end cap 44 may include openings 76 defined therein for both the first body 62 and the second body 64. The first body 62 and second body 64 may be positioned relative to each other such that they are coupled together by the end cap 44, as shown. Portions of the first body 62 and second body 64 that are disposed in the openings 76 may be secured therein. For example, as shown, welds 78 may cover the openings 76 or secure cover plates 79 to the openings 76 to secure the first body 62 and second body 64 to the end cap 44.

It should be understood that the present disclosure is not limited to recesses and protrusions or end caps for coupling the second body 64 to the first body 62 of an airfoil 40. Rather, any suitable design of the bodies or suitable coupling apparatus for coupling the bodies together is within the scope and spirit of the present disclosure.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

1. An airfoil for a turbine system, the airfoil comprising:

a first body having exterior surfaces defining a first portion of an aerodynamic contour of the airfoil and formed from a first material; and

a second body having exterior surfaces defining a second portion of an aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

2. The airfoil of claim 1, wherein the second material is a ceramic matrix composite.

3. The airfoil of claim 1, wherein the exterior surfaces of the first body at least partially define a pressure side and a suction side and further define a leading edge, and wherein the exterior surfaces of the second body define a trailing edge.

4. The airfoil of claim 1, wherein the second body is continuous throughout a cross-sectional profile.

5. The airfoil of claim 1, wherein the second body comprises a plurality of second body sections, each of the second body sections formed from a second material.

6. The airfoil of claim 1, wherein one of the first body or the second body defines a recess and the other of the first body or the second body comprises a protrusion, and wherein engagement of the protrusion in the recess couples the second body to the first body.

7. The airfoil of claim 1, further comprising an end cap coupling the second body to the first body.

8. The airfoil of claim 1, wherein the first body defines a cooling passage.

9. A nozzle for a turbine section of a turbine system, the nozzle comprising:

an airfoil having exterior surfaces defining an aerodynamic contour, the aerodynamic contour comprising a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil comprising: a first body having exterior surfaces defining a first portion of the aerodynamic contour of the airfoil and formed from a first material; and a second body having exterior surfaces defining a second portion of the aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

10. The nozzle of claim 9, wherein the second material is a ceramic matrix composite.

11. The nozzle of claim 9, wherein the exterior surfaces of the first body at least partially define a pressure side and a suction side and further define a leading edge, and wherein the exterior surfaces of the second body define a trailing edge.

12. The nozzle of claim 9, wherein the second body is continuous throughout a cross-sectional profile.

13. The nozzle of claim 9, wherein one of the first body or the second body defines a recess and the other of the first body or the second body comprises a protrusion, and wherein engagement of the protrusion in the recess couples the second body to the first body.

14. The nozzle of claim 9, further comprising an end cap coupling the second body to the first body.

15. A turbine system, comprising:

a turbine section, the turbine section comprising a plurality of hot gas path components, at least one of the plurality of hot gas path components comprising an airfoil having exterior surfaces defining an aerodynamic contour, the aerodynamic contour comprising a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil comprising: a first body having exterior surfaces defining a first portion of the aerodynamic contour of the airfoil and formed from a first material; and a second body having exterior surfaces defining a second portion of the aerodynamic contour of the airfoil, the second body coupled to the first body and formed from a second material having a different temperature capability than the first material.

16. The turbine system of claim 15, wherein the second material is a ceramic matrix composite.

17. The turbine system of claim 15, wherein the exterior surfaces of the first body at least partially define a pressure side and a suction side and further define a leading edge, and wherein the exterior surfaces of the second body define a trailing edge.

18. The turbine system of claim 15, wherein the second body is continuous throughout a cross-sectional profile.

19. The turbine system of claim 15, wherein one of the first body or the second body defines a recess and the other of the first body or the second body comprises a protrusion, and wherein engagement of the protrusion in the recess couples the second body to the first body.

20. The turbine system of claim 15, further comprising an end cap coupling the second body to the first body.