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Patent Searching and Data


Title:
TURBINE VANE INSERT
Document Type and Number:
WIPO Patent Application WO/2017/039607
Kind Code:
A1
Abstract:
The turbine component includes an endwall (46, 50) and an elongated airfoil (32) extending radially from the endwall (46, 50), the airfoil (32) having a leading edge (54), a trailing edge (56), and an intermediate portion (58) between the leading edge (54) and the trailing edge (58), and a pressure side (40) and a suction side (44) defined between the leading (54) and the trailing (58) edges. The intermediate portion (58) may be formed from a metal material (61) and the airfoil (32) further comprises a non-metallic insert (60) forming a portion of the airfoil (32) at the leading edge (54), the trailing edge (56), or both.

Inventors:
WIEBE DAVID J (US)
Application Number:
PCT/US2015/047718
Publication Date:
March 09, 2017
Filing Date:
August 31, 2015
Export Citation:
Click for automatic bibliography generation   Help
Assignee:
SIEMENS ENERGY INC (US)
International Classes:
F01D5/14; B23P6/00; F01D5/00; F01D5/28
Foreign References:
US5358379A1994-10-25
US20060228211A12006-10-12
US20140023483A12014-01-23
US20050175444A12005-08-11
US3315941A1967-04-25
DE3821005A11989-12-28
US8058191B22011-11-15
US7745022B22010-06-29
US7153096B22006-12-26
US7093359B22006-08-22
US6733907B22004-05-11
US8528339B22013-09-10
US7255535B22007-08-14
US7402347B22008-07-22
US7247002B22007-07-24
US7247003B22007-07-24
US7198458B22007-04-03
US20140023483A12014-01-23
Attorney, Agent or Firm:
SCOTT, Mark W. (3501 Quadrangle Blvd. Ste. 230Orlando, Florida, US)
Download PDF:
Claims:
CLAIMS

What we claim is:

1 . A turbine component (30) comprising:

an endwall (46, 50); and

an elongated airfoil (32) extending radially from the endwall (46, 50), the airfoil (32) having a leading edge (54), a trailing edge (56), and an intermediate portion (58) between the leading edge (54) and the trailing edge (56), and a pressure side (40) and a suction side (44) defined between the leading (54) and the trailing (56) edges;

wherein the intermediate portion (58) is formed from a metal material (61 ); and wherein the airfoil (32) further comprises a non-metallic insert (60) forming a portion of the airfoil (32) at the leading edge (54), the trailing edge (56), or both.

2. The turbine component (30) of claim 1 , further comprising a protuberance (68) extending radially from the endwall (46, 50) and disposed at a junction (70) between the endwall (46, 50) and the non-metallic insert, the protuberance (68) effective to provide a stop (72) for securement of the insert (60) and to reduce formation of vortices (2) at the junction (70).

3. The turbine component (30) of claim 2, wherein the protuberance (58) is disposed on a platform (74), and wherein the platform (74) is fixed to a surface (76) of the endwall (46, 50) and abuts the insert (60).

4. The turbine component (30) of claim 3, wherein the platform (74) is secured to the endwall surface (76) by welding, brazing, or mechanical attachment.

5. The turbine component (30) of claim 3, wherein the platform (74) further comprises at least one fluid injection passage (86) extending therethrough and a fluid source (88) in communication with the fluid injection passage (86), the fluid source (88) configured to introduce a fluid (92) through the passage (86) toward the junction (70) to at least reduce formation of vortices (2) at the junction (70).

6. The turbine component (30) of claim 2, wherein the protuberance (58) extends from a region comprising a portion of the suction side (44) and a portion of the pressure side (40) of the airfoil (32) at the junction (70).

7. The turbine component (30) of claim 2, wherein the protuberance (58) extends around an entire perimeter (84) of the airfoil (32) at the junction (70) between the airfoil (32) and the endwall (46, 50).

8. The turbine component (30) of claim 8, wherein the insert (60) comprises a radial support rod (94) extending radially from the endwall (46, 50) and through the insert (60) for structural support of the insert (60).

9. The turbine component (30) of claim 8, wherein the radial support rod (94) comprises a channel (96) extending therethrough for travel of a cooling fluid

therethrough.

10. The turbine component (30) in any of one of claims 1 to 9, wherein the insert (60) comprises a lateral support rod (98) fixed between the insert (60) and the intermediate portion (58) of the airfoil (32) to secure the insert (60) to the intermediate portion (58).

1 1 . The turbine component (30) of claim 10, wherein the lateral support rod (98) comprises a flanged end (100A), wherein the airfoil (32) comprises a cavity (59), and wherein the flanged end (100A) extends into the cavity (59) and is secured directly or indirectly against the intermediate body portion (58).

12. The turbine component (30) of claim 1 1 , further comprising a spring member (104) disposed between the flanged end (100A) and the intermediate portion (58) to allow for thermal expansion between the non-metallic insert (60) and the intermediate portion (58).

13. The turbine component (30) in any one of claims 1 1 to 12, wherein the component (30) comprises a plurality of lateral support rods (98) having at least one flanged end (100A), and wherein the component (30) further comprises a retaining structure (106) extending from the endwall (46, 50) and through the cavity (59) of the airfoil (32), the retaining structure (106) comprising openings (108) arranged to secure flanged ends (100A) of the lateral support rods (98) therein.

14. The turbine component (30) of claim 1 1 , wherein the retaining structure (106) further comprises a cap (1 10) comprising a first arm (1 12) configured to snap over the flanged end (100A) to secure the lateral support rod (98) in a fixed position.

15. The turbine component (30) of claim 14, wherein the cap (1 10) further comprises a second arm (1 14) configured to retain a base (1 16) of the flanged end (100A) to further secure the flanged end (100A) of the lateral support rod (98).

16. The turbine component (30) of claim 15, wherein the turbine component (30) comprises a first endwall (46) and a second endwall (50), and wherein the retaining structure (106) extends radially between the two endwalls (46, 50).

17. The turbine component (30) in any one of claims 1 to 17, wherein the turbine component (30) comprises a turbine vane (22), wherein the vane (22) comprises a first endwall (46) and a second endwall (50), and wherein the elongated airfoil (32) and the non-metallic insert (60) extend radially between the first and second endwalls (46, 50).

18. The turbine component (30) in any of claims 1 to 17, wherein the intermediate body portion (58) comprises a superalloy material (61 ).

19. The turbine component (30) in any one of claims 1 to 18, wherein the turbine component (30) comprises a stationary turbine blade.

20. The turbine component (30) in any of claims 1 to 19, wherein the airfoil (32) comprises the non-metallic insert (60) at both the leading edge (54) and the trailing edge (56) of the airfoil (32).

21 . The turbine component (30) in any one of claims 1 to 19, wherein the airfoil (32) comprises a non-metallic insert (60) at the leading edge (54) of the airfoil.

22. The turbine component (30) in any one of claims 1 to 21 , wherein the non- metallic insert (60) comprises a ceramic matrix composite material.

23. A process for modifying a turbine component (30) comprising an elongated airfoil (32) extending radially from an endwall (46, 50), the airfoil (32) having a leading edge (54), a trailing edge (56), and an intermediate portion (58) between the leading edge (54) and the trailing edge (56), and a pressure side (40) and a suction side (44) defined between the leading (54) and the trailing (56) edges, the method

comprising:

removing a first portion (1 18) of the airfoil (32) at a leading edge (54) or a trailing edge (56) of the airfoil (32) to define a first void portion (120) in the airfoil (32); and

securing a first non-metallic insert (60) within the first void portion (120).

24. The process of claim 23, further comprising providing a first protuberance (58) extending radially from the endwall (46, 50) at a junction (70) between the endwall (46, 50) and the first insert (60), the first protuberance (58) effective to provide a stop (72) for securement of the insert (60) to the airfoil (32) and to reduce formation of vortices (2) at the junction (70).

25. The process of claim 24, further comprising:

providing the first protuberance (58) on a first platform (74); and securing the first platform (74) to a surface (76) of the endwall (46, 50) against the first insert (60).

26. The process of claim 25, wherein the platform (24) further comprises at least one fluid injection passage (86) extending therethrough and a fluid source (88) in communication with the fluid injection passage (86), and wherein the process further comprises introducing a fluid (92) through the passage (92) toward the junction (70) to at least reduce formation of vortices (2) at the junction (70).

27. The process in any of claims 23 to 26, further comprising:

removing a second portion of the airfoil (32) at the other of the leading edge (54) and the trailing edge (56) of the airfoil (32) to define a second void portion (124) in the airfoil (32); and

securing a second non-metallic insert (60) within the second void portion (124).

28. The process of claim 27, wherein the component (30) further comprises an endwall (46, 50) at an end of the airfoil (32) and the second non-metallic insert (60), and wherein the method further comprises:

providing a second protuberance (58) extending radially from the endwall (46, 50) at a junction (70) between the endwall (46, 50) and the second non-metallic insert (60), the second protuberance (58) effective to provide a stop (72) for securement of the second non-metallic insert (60) to the airfoil (32) and to reduce formation of vortices (2) at the junction (70) .

29. The process of claim 28, further comprising:

providing the second protuberance (58) on a second platform (74); and securing the second platform (74) to a surface of the endwall (46, 50) against the second non-metallic insert (60).

30. The process of claim 29, wherein the second platform (74) further comprises at least one fluid injection passage (86) extending therethrough and a fluid source (88) in communication with the fluid injection passage (86), and wherein the process further comprises introducing a fluid (92) through the passage (86) toward the junction (70) to at least reduce formation of vortices (2) at the junction (70).

31 . The process of claims 28 to 30, wherein at least one the first and second protuberances (58) extend from a portion of the suction side (44) to a portion of the pressure side (40) of the airfoil.

32. The process of claims 28 to 31 , wherein the first and second

protuberances (58) collectively extend around an entire perimeter of the airfoil (32) at the junction (70) between the airfoil (32) and the endwall (46, 50).

33. The process of claims 23 to 32, further comprising securing the first or second platform (74) by welding, brazing, or mechanical attachment.

34. The process of claims 27 to 33, wherein at least one of the first or second non-metallic inserts (60) comprises a radial support rod (94) extending radially from the endwall (46, 50) and through the respective insert (60) for structural support of the insert (60).

35. The process of claim 34, wherein the radial support rod (94) comprises a channel (96) extending therethrough for travel of a cooling fluid therethrough.

36. The process in any of claims 27 to 35, wherein at least one of the first or second non-metallic inserts (60) comprises a lateral support rod (98) fixed between the respective insert (60) and the intermediate portion (58) of the airfoil (32) to secure the first or second non-metallic inserts (60) to the intermediate portion (58).

37. The process in any of claims 23 to 37, wherein the intermediate portion (58) comprises a superalloy material.

38. The process of claims 27 to 37, further comprising fixing a lateral support rod (98) between the first or second non-metallic insert (60) and the intermediate portion (58) of the airfoil (32) to secure the first or second non-metallic insert (60) to the intermediate portion (58).

39. The process of claim 38, wherein the lateral support rod (98) comprises a flanged end (100A), wherein the airfoil (32) comprises a cavity (59), and wherein the flanged end (100A) extends into the cavity (59), and wherein the method further comprises securing the flanged end (100A) directly or indirectly against the intermediate portion (58).

40. The process of claim 39, wherein the securing is done by fixing a cap (1 10) over the flanged end (100A) to secure the lateral support rod (98) in a fixed position.

41 . The process of claims 38 to 40, further comprising disposing a spring member (104) between the flanged end (100A) and the intermediate body portion (58) to allow for thermal expansion between the first or second non-metallic insert (60) and the intermediate body portion (58).

42. The process of claims 27 to 41 , wherein the turbine component (30) comprises a turbine vane (22), wherein the vane (22) comprises a first endwall (46) and a second endwall (50), and wherein the elongated airfoil (32) and the first or second non-metallic insert (60) extends radially between the first and second endwalls (46, 50).

43. The process of claims 23 to 42, wherein the turbine component (30) comprises a stationary turbine blade.

44. The process in any one of claims 27 to 43, wherein the first or the second non-metallic insert (60) comprises a ceramic matrix composite material.

45. The process in any one of claims 23 to 44, wherein the first portion (1 18) of the airfoil (32) is removed at the leading edge (54).

Description:
TURBINE VANE INSERT

STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT

Development for this invention was supported in part by Contract No. DE- FE0023955, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention.

FIELD

The present invention relates generally to turbine engine components, and more particularly to turbine components partially comprising a non-metallic insert at a leading edge and/or a trailing edge thereof, and to methods of producing the same.

BACKGROUND

A gas turbine engine typically includes a compressor section, a combustor, and a turbine section. The compressor section compresses ambient air that enters an inlet. The combustor combines the compressed air with a fuel and ignites the mixture, thereby creating combustion products defining a working fluid. The working fluid travels to the turbine section where it is expanded to produce a work output. Within the turbine section are rows of stationary vanes directing the working fluid to rows of rotating blades coupled to a rotor. Each pair of a row of vanes and a row of blades form a stage in the turbine section.

Given the costs of the gas turbine and the components therein, it is desirable to maximize the efficiency of the system and minimize efficiency losses. It is well known that in order to maximize efficiency, the gas turbine is typically operated at as high of a temperature as is practical. In so doing, the leading and trailing edges of the turbine vane or blades may be subjected to the greatest amount of heat which can pose a substantial risk of structural damage to the components at those locations.

Further, it is known that efficiency of a gas turbine can be improved by reducing aerodynamic losses in the turbine. For example, as shown in FIG. 1 , one source of aerodynamic losses may be the formation of vortex flows 2 that may occur as a result of a boundary layer that is formed between a hot working gas flow and an endwall 3 located at the end of an airfoil 4. The boundary layer tends to adhere to the endwall 3 with a resulting lower velocity than the main body of the gas flow. For, example, vortex flows 2 known as horseshoe vortices may form at upstream leading edge locations 5 where the airfoil 4 attaches to the endwall 3, and may extend a substantial distance downstream between adjacent airfoils. These vortex flows 2 may decrease the efficiency of the turbine engine. Solutions which can improve the gas turbine's operation in extreme temperatures and reduce efficiency or aerodynamic losses are thus needed.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 is a plan view of an airfoil of a turbine stage illustrating a typical vortex flow around a leading edge of the airfoil.

FIG. 2 is a partial cross-sectional view of a gas turbine engine incorporating turbine components in accordance with an aspect of the present invention.

FIG. 3 is a perspective view of a vane having an airfoil comprising a non-metallic insert as described herein in accordance with an aspect of the present invention.

FIG. 4 is a cross-sectional view of a turbine airfoil having an inert at a leading edge thereof taken along line 2-2 of FIG. 3 in accordance with an aspect of the present invention.

FIG. 5 illustrates an embodiment of an airfoil having inserts at both a leading edge and a trailing edge in accordance with an aspect of the present invention.

FIG. 6 illustrates an insert comprising a continuously solid body in accordance with an aspect of the present invention.

FIG. 7 illustrates an insert comprising a plurality of stacked laminate plates in accordance with an aspect of the present invention.

FIG. 8 illustrates a side view of a turbine vane comprising protuberances in accordance with an aspect of the present invention. FIG. 9 illustrates a side view of a turbine vane comprising a platform having protuberances mechanically attached thereto in accordance with an aspect of the present invention.

FIG. 10 illustrates a cross-sectional view of an airfoil having protuberances spaced apart at locations about the leading edge and the trailing edge of the airfoil in accordance with an aspect of the present invention.

FIG. 1 1 illustrates a cross-sectional view of an airfoil having a protuberance about an entire perimeter of the airfoil in accordance with an aspect of the present invention.

FIG. 12 illustrates a platform having a protuberance and a channel in

communication with a fluid source for introducing a fluid in front of the protuberance in accordance with an aspect of the present invention.

FIG. 13 is a side view of a turbine component comprising a non-metallic insert with a radial support rod having a cooling channel therein in accordance with an aspect of the present invention.

FIG. 14 is a cross-sectional view of an airfoil comprising a non-metallic insert with a lateral support rod in accordance with an aspect of the present invention.

FIG. 15 is a cross-sectional view showing mechanical attachment of the lateral support rod to an insert in accordance with an aspect of the present invention.

FIG. 16 is a side view of a retaining structure which engages flanged ends of multiple lateral support rods in accordance with another aspect of the present invention.

FIG. 17 is a side view of a retaining structure 106 may comprise a structure configured to clamp respective ends of the lateral support rods in accordance with another aspect of the present invention.

FIGS. 18A-18C illustrate a process for retrofitting an existing airfoil with an insert as described herein in accordance with another aspect of the present invention.

DETAILED DESCRIPTION

The present inventor has developed improved turbine components for use in turbine engines that improve efficiency of the turbine, minimize cooling needs, and/or reduce the fornnation of horseshoe vortices at a junction between an airfoil portion of the component and an endwall of the turbine component, such as a blade or vane. In accordance with one aspect, there is provided a turbine component comprising an insert comprising a non-metallic material, such as a ceramic matrix composite material, at least at a leading edge of an airfoil of the component. The non-metallic insert may reduce cooling needs for the component, and thus operating and component costs, as well as allow operation of the associated gas turbine at higher temperatures. In particular embodiments, the non-metallic insert may be fixed to an intermediate portion of a respective airfoil with the aid of a protuberance disposed at a leading or trailing edge of the airfoil. The protuberance may define a stop for securement of the insert during installation and may reduce horseshoe vortices at a junction between the airfoil and an endwall to which the airfoil is secured.

Now referring to the figures, FIG. 2 shows a gas turbine engine 10 including a compressor section 12, a combustor 14, and a turbine section 16. The compressor section 12 compresses ambient air 18 that enters an inlet 20. The combustor 14 combines the compressed air with a fuel and ignites the mixture creating combustion products comprising a hot working gas defining a working fluid. The working fluid travels to the turbine section 16. Within the turbine section 16 are rows of stationary vanes 22 and rows of rotating blades 24 coupled to a rotor 26. Each pair of rows of vanes 22 and blades 24 forms a stage in the turbine section 16. The rows of vanes 22 and rows of blades 24 extend radially into an axial flowpath 28 extending through the turbine section 16. The working fluid expands through the turbine section 16 and causes the blades 24, and therefore the rotor 26, to rotate. The rotor 26 extends into and through the compressor 12, and may provide power to the compressor 12 and output power to a generator. In addition to providing air to the combustor 14, bleed air from the compressor 12 may be provided to components of the turbine section 16, such as for providing a cooling fluid to the turbine components.

Referring to FIG. 3, there is shown a turbine component 30 for the engine 10 in the form of a stationary turbine vane 22. The vane 22 includes an elongated airfoil 32 having a body 34 with an outer wall 36. The vane 22 may be configured for use, for example, in turbine engine 10. The outer wall 36 may have a generally concave- shaped portion 38 forming a pressure side 40 and a generally convex shaped portion (opposite side) 42 forming the suction side 44. The turbine vane 22 may also include an outer endwall 46 at a first end 48 of the component 30 configured to be coupled to a hook attachment and may include an inner endwall 50 at a second end 52 of the component 10. The airfoil 32 also includes a leading edge 54 and a trailing edge 56. Although a vane 22 is shown, it is appreciated that the turbine component 30 is not limited to the vane 22, but may include any other static turbine component in a turbine engine, such as a stationary turbine blade, and may have further structural features as is necessary for the component.

In accordance with an aspect of the present invention, the turbine component 30 may be constructed with at least one portion formed from a non-metallic material, such as a ceramic matrix composite (CMC) material, instead of being composed entirely of a single material (e.g., superalloy material) as in prior art airfoil structures. Referring to FIGS 4-5, there is shown a vane 22 that includes a non-metallic insert in accordance with an aspect of the present invention. FIG. 4 is a cross-sectional view of the airfoil 32 of vane 22 taken at line 2-2 in FIG. 3. As shown, the body 34 of the airfoil 32 includes an intermediate portion 58 formed from a metallic material 61 , such as a superalloy, and at least one insert 60 comprised of the non-metallic material 62, e.g. a CMC material. Each of the intermediate portion 58 and the insert 60 are disposed between the leading edge 54 and the trailing edge 56. In an embodiment, the airfoil 32 may include an insert 60 at only one of the leading edge 54 and the trailing edge 56 of the airfoil 32 as is shown in FIG. 4.

In another embodiment, as shown in FIG. 5, the airfoil 32 may include an insert 60 at each of the trailing edge 54 and the leading 56 of the airfoil 32. By forming at least a portion of the airfoil 32 from a non-metallic material via the inserts 60, the cooling needs for the component 30 may be significantly reduced yet a metallic portion (e.g., intermediate portion 58) of the airfoil 32 remains to provide the component 30 with a desired mechanical strength. For ease of description, it is appreciated that although the singular term "insert" is used, the term "insert" may refer to one or more inserts at the leading edge 54 and/or the trailing edge 56.

The insert 60 may be of any suitable width, size, shape, and dimension to render the component 30, e.g., vane 22, suitable for its intended purpose. In an embodiment, the insert 60 at the leading edge 54 and/or trailing edge 56 may be configured to extend radially from between the endwalls 46, 50 of the vane 22 along with the intermediate portion 58 such that the insert 60 and intermediate portion 58 extend into the axial flowpath 28 (FIG. 2). Typically, the area of the vane 22 about the leading edge 54 and the trailing edge 56 may be exposed to the highest temperatures in the gas turbine engine 10. Forming the insert 60 from a non-metallic material 62 with a high degree of thermal conductivity may significantly reduce cooling needs for the vane 22 as the non- metallic material 62 may be one with a high degree of thermal conductivity, such as a CMC material. Thus, in certain embodiments, the use of an insert 60 as described herein may eliminate the need for cooling in or adjacent to the area occupied by a respective insert 60.

The intermediate portion 58 of the body 34 may define one or more cavities, e.g., cavity 59, to allow for a cooling fluid to flow therethrough for cooling of the intermediate portion 58, or to allow for other elements such as support beams, impingement tubes, or the like to be housed. To provide adequate structural support for the turbine component 30, the intermediate portion 58 may comprise an alloy material such as an Fe-based alloy, a Ni-based alloy, a Co-based alloy as are well known in the art. In certain embodiments, the alloy may comprise a superalloy. The term "superalloy" may be understood to refer to a highly corrosion-resistant and oxidation-resistant alloy that exhibits excellent mechanical strength and resistance to creep even at high

temperatures. Exemplary superalloy materials are commercially available and are sold under the trademarks and brand names Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g. Rene N5, Rene 41 , Rene 80, Rene 108, Rene 142, Rene 220), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-750, ECY 768, 262, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys, GTD 1 1 1 , GTD 222, MGA 1400, MGA 2400, PSM 1 16, CMSX-8, CMSX-10, PWA 1484, IN 713C, Mar-M- 200, PWA 1480, IN 100, IN 700, Udimet 600, Udimet 500 and titanium aluminide, for example.

The non-metallic material 62 of the insert 60 may comprise any suitable material that has a higher degree of thermal conductivity than the metal of the intermediate portion 58 of the body 34 of the airfoil 32. In an embodiment, the non-metallic material 62 may comprise a suitable ceramic matrix material that hosts a plurality of reinforcing fibers as is known in the art. In certain embodiments, the CMC material may be anisotropic, at least in the sense that it can have different strength characteristics in different directions. It is appreciated that various factors, including material selection and fiber orientation, can affect the strength characteristics of a CMC material. In addition, the CMC material may comprise oxide as well as non-oxide CMC materials.

In a particular embodiment, the CMC material may comprise alumina, and the fibers may comprise an aluminosilicate composition consisting of approximately 70% alumina; 28% silica; and 2% boron (sold under the name NEXTELâ„¢ 312). The fibers may be provided in various forms, such as a woven fabric, blankets, unidirectional tapes, and mats. A variety of techniques are known in the art for making a CMC material and such techniques can be used in forming the CMC material 62 for use herein. Exemplary CMC materials 62 are described in US Patent Nos. 8,058,191 , 7,745,022, 7,153,096; 7,093,359; and 6,733,907, the entirety of each of which is hereby incorporated by reference. As mentioned, the selection of materials may not be the only factor which governs the properties of the CMC material as the fiber direction may also influence the mechanical strength of the material, for example. As such, the fibers for the CMC material may have any suitable orientation, such as those described in U.S. Patent No. 7,153,096.

In one embodiment, the insert 60 may comprise a uniform, continuous body as shown in FIG. 6. In another embodiment, as shown in FIG. 7, the insert 60 may comprise a plurality of stacked laminate plates 66 formed from the non-metallic material, e.g., a CMC material. In this embodiment, each of the stacked laminate plates 66 may be cut to a desired shape, such as an airfoil shape, via a laser cutting process and stacked to provide a respective insert 60. In certain embodiments, the stacked laminate plates 66 may be provided with a support structure, such as a tie rod, extending through the stacked laminates as will be described in detail further below. The plates 66 may further include suitable structures, such as retainers, for radial compression of the plates. Exemplary processes for forming a body of stacked laminates from a non-metal material, such as a CMC material, and associated structures are set forth in U.S. Patent Nos. 8,528,339, 7,255,535, 7,402,347, 7,247,002, 7,247,003, 7,198,458 and 7,153,096, for example, the entirety of each of which is hereby incorporated by reference. Some advantages a stacked laminate structure include enabling the non-metallic insert 60 to itself bear some structural loading via the individual plates, as well as increasing a number of possible dimensions and

configurations for an associated component by controlling the structure of the

component on a level -by-level basis.

Now referring to FIG. 8 (which is a cross-sectional view of a component 30, e.g., vane 22), the component 30 may further comprise one or more protuberances 68 (protuberance 68) extending radially from a respective endwall (e.g., 46, 50) and disposed at a junction 70 between the endwalls 46, 50 and a respective non-metallic insert 60 at a leading edge 54 and a trailing edge 56 of the airfoil 32. In the

embodiment shown in FIG. 8, there is a protuberance 68 disposed at each junction 70 of an endwall 46, 50 and an insert 60 at the leading edge 54 and the trailing edge 56 of the airfoil 32, although it is understood that the present invention is not so limited. In certain embodiments, a protuberance 68 may be only disposed at the leading edge 54 or the trailing edge 56 of the airfoil 32 at one or more locations thereof. For example, a protuberance 68 may be disposed only at a leading edge 54 of the airfoil 32. In addition, a protuberance 68 may be disposed only at one of endwalls 46, 50, but not the other, or at both.

Each protuberance 68 may provide two beneficial characteristics. First, each protuberance 68 may be effective to define a stop 72 for securement of a respective insert 60 as will be explained further below. Second, each protuberance 68 may be effective to reduce formation of vortices at the junction 70, in particular horseshoe vortices at the junction 70 as were explained with reference to FIG. 1 . In this way, the protuberance 68 may function as an anti-horseshoe vortex structure for the component 30, as well as a structure to help secure the non-metallic insert 60 in place in the component 30. In an aspect, any of the protuberances 68 may be formed integrally with a portion of the component 30, such as integrally with either of endwalls 46, 50. For example, in the embodiment shown in FIG. 8, the protuberances 68 may be formed integrally with the endwalls 46, 50.

In another aspect, referring to FIG. 9, any of the protuberances 68 located at the junctions 70 of the leading edge 54 and the trailing edge 56 of the endwalls 46, 50 may be disposed on a platform 74 which is directly or indirectly secured to the a surface 76 of a respective endwall, such as endwall 50 as shown. In this way, the protuberance 68 may be fixed so as to directly abut a respective insert 60 to provide the stop 72 for the respective insert 60. Further, the platform 74 may be configured to further provide a compressive force on a respective insert 60 to help lock the insert 60 in place as will be described in further detail below.

Each protuberance 68 may be formed from any suitable material, such as the same material of an endwall 46, 50, insert 60, or intermediate portion 58 of the turbine component 30. In addition, each protuberance 68 may be of any suitable size, shape, and dimension for the particular application. In an embodiment, each protuberance 68 may each be of a size, shape, and dimension effective to provide a stop 72 for the installation of a respective insert 60 as described herein and/or effective to reduce the formation of horseshoe vortices at one or more junctions of an insert 60 and a respective endwall 46, 50.

It is contemplated that each platform 74 may be secured to a respective endwall surface 76 by any suitable structure or method known in the art, such as by welding, brazing, or mechanical attachment, for example. As shown in FIG. 9, for example, the platform 74 may be secured to a respective endwall, e.g., endwall 50, by mechanical attachment, such as via a bolt 78 extending through the platform 74 and the endwall 50. In certain embodiments, the bolt 78 may be configured to be threaded into and through a corresponding threaded portion integrally formed within the endwall and into the platform 74 as shown. In some embodiments, an embedded nut having threads to receive the bolt may be embedded with the platform. Such bolt-in retainers may be utilized for securing an insert 60 at one or both of the leading edge 54 and the trailing edge 56 of the component 30. In addition, in certain embodiments, a depth of either endwall 46, 50 may be removed to provide a channel 82 in which the insert 60 may be disposed as shown in FIG. 9.

Each protuberance 68 may be disposed at any suitable location about a perimeter of the airfoil 32 at a respective junction 70. In an embodiment, as shown in FIG. 10 (taken at line 3-3 of FIG. 9), the protuberance 68 may be disposed about a perimeter of the airfoil 32. For example, each protuberance 68 may extend from a location on the pressure side 40 to a location on the suction side 44 of the airfoil 32 about at least a portion of the junction 70. In other embodiments, as shown in FIG. 1 1 , a single protuberance 68 may be provided which extends about an entire perimeter 84 of the airfoil 32, including about any (one or more) inserts 60 present.

In accordance with another aspect, as shown in FIG. 12, the platform 74 may further comprise at least one fluid injection passage 86 extending therethrough and a fluid source 88 in communication with the fluid injection passage 86. The fluid injection passage 86 may be oriented at an angle less than ninety degrees relative to a top surface 90 of the platform 74. In this embodiment, the fluid source 88 may be configured to introduce a fluid 92, such as air, through the passage 86 toward the junction 70 between a respective endwall, e.g., endwall 50 as shown, to at least reduce formation of vortex flows 2 (FIG. 1 ) at the junction 70 (FIG. 9). An exemplary configuration is set forth in U.S. Published Patent Application No. 2014/0023483, the entirety of which is hereby incorporated by reference.

In accordance with another aspect, any insert 60 may comprise a radial support rod 94 extending therethrough to provide additional structural support to the insert 60. For example, as shown in FIG. 13, there is provided a radial support rod 94 extending radially between the outer endwall 46 and the inner endwall 50, and through the insert 60 for structural support of the insert 60. In certain embodiments, as is also shown in FIG. 13, the radial support rod 94 may also comprise a channel 96 extending through the radial support rod 94 and one or both of endwalls 46, 50 for travel of a cooling fluid, such as air, therethrough from a suitable source when cooling of the insert 60 is desired. It is understood, however, that the presence of the cooling channel 96 is not necessary and that, in some embodiments, the insert 60 may require no cooling structures therein.

In accordance with another aspect, as shown in FIG. 14, the insert 60 may further or instead comprise a lateral support rod 98 fixed between the insert 60 and the intermediate portion 50 of the airfoil 32 to secure the insert 60 to the intermediate portion 58. FIG. 14 is a cross-sectional view of an airfoil 32 showing a positioning of one lateral support rod 98, though it is understood that a plurality of lateral support rods 98 may be provided spaced apart and stacked in a radial direction of the body 34 of the insert 60. The lateral support rod 98 may further provide additional structural support to the insert 60 in a direction different from the radial support rod 94, such as an angle parallel to one of endwalls 46, 50.

FIG. 14 illustrates an embodiment comprising both a radial support rod 94 and a lateral support rod 98; however, it is understood that the present invention is not so limited and that the insert 60 may comprise either one or the other of the radial support rod 94 and the lateral support rod 98, or both. The radial support rod 94 and the lateral support rod 98 may be formed from any suitable material that provides a degree of structural support to the insert, such as a metallic material, e.g., an alloy or superalloy. In certain embodiments, the intermediate portion 58 and the insert 60 may comprise corresponding beveled edges 99 to facilitate attachment of the components.

The lateral support rod 98 may be fixed in its desired position by any suitable method, such as by a joining process (welding, soldering, or the like) or via mechanical attachment to secure the lateral support rod 98 directly or indirectly against the intermediate body portion 58. Referring to FIG. 15, there is shown an exemplary mechanical arrangement 102 for securing a lateral support rod 98 extending through the insert 60 and the intermediate portion 58. In the embodiment shown, the lateral support rod 98 comprises a flanged end 100A which extends into the cavity 59 of the intermediate portion 58. In some embodiments, the cavity 59 may also include one or more impingement tubes 103 therein as are know in the art. The flanged end 100A may be secured directly or indirectly against the intermediate body portion 58. In addition, in certain embodiments, the lateral support rod 98 may further comprise flanged end 100B, which is shaped to abut an endwall 102 of the insert 60.

In an aspect, as shown in FIG. 15, the mechanical arrangement 102 may also comprise a spring member 104 between the flanged end 100A and an inner edge 101 of the insert 60. In an embodiment, the spring member 104 allows for thermal expansion between the non-metallic insert 60 and the intermediate portion 58. In certain embodiments, the lateral support rod 98 may be formed within the insert 60 as the insert 60 is manufactured. Alternatively, at least a portion of a body of the insert 60 may be machined away to accommodate each lateral support rod 98 provided.

In another aspect, as shown at the trailing edge 56 of the airfoil 32 of FIG. 15, any lateral support rod 98 may instead or further comprise a threaded body 128 which is configured to mate with a corresponding nut 128. The nut 128 may be embedded within the insert 60 during manufacture of the insert 60. In this way, the lateral support rod 98 may be secured in place.

In accordance with another aspect, as shown in FIG. 16, the component 30 may comprise a plurality of lateral support rods 98, each having at least one flanged end 100A extending into the cavity 59 (FIG. 15) of the intermediate portion 58 of the body 32. The lateral support rods 98 may be stacked in a radial direction through the component 30. In an embodiment, the component 30 further comprises a lateral insert retaining structure 106 formed from any suitable and relatively rigid material. The retaining structure 106 may be in the form of a substrate, e.g., flat plate, which extends between endwalls 46, 50 in between the flanged end 100A as describe and the intermediate portion 58. In an embodiment, the retaining structure 106 further includes a plurality of openings 108 sized and arranged to secure accept a portion of the lateral rods 98 therein. In a particular embodiment, the larger diameter portions of the openings 108 are sized to fit over the flanged ends 100A. Thereafter, the retaining structure 106 may slide down into a fixed position, wherein the smaller diameter portion of the openings 108 engage the flanged ends 100A. In this way, via the retaining structure 106, the lateral support rods 98 may be more securely held in place between the endwalls 46, 50. In an embodiment, once in its desired position, the retaining structure 106 may be fixed to the endwalls 46, 50 by a joining process such as soldering, welding, or brazing. If necessary, either or both of the endwalls 46, 50 may include channels within which the retaining structure 106 is received.

In accordance with another aspect, as shown in FIG. 17, the retaining structure 106 may comprise a structure configured to clamp respective ends (flanged ends 100A) of the lateral support rods 98 to aid in their retention. In an embodiment, the retaining structure 106 may include a cap 1 10 configured to engage a respective flanged end 100A of a lateral support rod 98. As shown in FIG. 17, the cap 106 may include a first arm 1 12 configured to move in place and snap over the flanged end 100A to secure the lateral support rod 98 in a fixed position. In certain embodiments, the cap 1 10 may also comprise comprises a second arm 1 14 that abuts a base 1 16 of the flange 100A to retain the flange 100A from an underside thereof to further secure the flange 100A of the lateral support rod 98.

It is appreciated that the above components 30 may be formed as a new manufacture. Alternatively, in other embodiments, the structural features described herein may be incorporated into existing turbine components as a retrofit, such as in repair of the component. For example, due to the high temperature extreme a portion of a leading edge and/or trailing edge of a blade or vane may be damaged. Aspects of the present allow for repair of the leading edge and/or the trailing edge with any of the structural features described herein.

In accordance with one aspect, there is provided a process for modifying a turbine component 30 which comprises an elongated airfoil 32 extending radially from the endwall (e.g., endwall 46 or 50) as shown in FIGS. 18A-18C. The airfoil 32 includes a leading edge 54, a trailing edge 56, and an intermediate portion 58 between the leading edge 54 and the trailing edge 56. The airfoil 32 further includes a pressure side 40 and a suction side 44 defined between the leading edge 54 and the trailing edge 56. As is illustrated in FIG. 18A, the process may comprise removing a first portion 1 18 of the airfoil 32 at a leading edge 54 or the trailing edge 56 of the airfoil 32 to define a first void portion (shown bound by dotted lines 120 in the airfoil 32 as shown in FIG. 18B). In the embodiment shown, the first void portion 120 is removed from the leading edge 54, although it is understood that the invention is not so limited. The removing may be done by any suitable process as by conventional machining techniques for removal of material, including but not limited to Electrical Discharge Machining (EDM) methods, or the like. When the airfoil 32 is disposed on an endwall, such as in the case of a blade or vane, the removal step my further including removing a corresponding portion of the endwall.

After the removal process, an intermediate component 125 is produced as shown in FIG. 18B. Thereafter, as shown in FIG. 18C, an insert 60 can be inserted within the void space 120 and secured to the intermediate body portion 58. The insert 60 may be of a structure as described previously herein and may further include any other additional components as described herein, such as protuberances 68, radial support rod 94, lateral support rods 98, beveled edges 99, any structures for attachment or securement of any of the components, and the like. In an embodiment, the removing may further comprise removing one or more existing impingement tubes in the area to be occupied by the insert 60.

In an embodiment, only an area adjacent the leading edge 54 of the airfoil 32 is replaced. However, it is understood that the present invention is not limited. In another embodiment, the process may also comprise removing a second portion 122 of the airfoil 32 at the other (relative to the first portion 1 18 removed) of the leading edge 54 and the trailing edge 56 of the airfoil 32 to define a second void portion (shown bound by dotted lines 124 in FIG. 18B). The removing of the second portion 122 may also be done by any suitable process as by conventional machining techniques for removal of material, including but not limited to Electrical Discharge Machining (EDM) methods, or the like. When the airfoil 32 is disposed on an endwall, such as in the case of a blade or vane, the removing of the second portion 122 may also further include removing a corresponding portion of the endwall. Thereafter as shown in FIG. 18C, another insert 60 may be inserted within the void space 124 and secured to the intermediate body portion 58 to form a component 30 having two non-metallic inserts 60. While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.