Stator vane assembly having mechanical retention device

Abstract

To provide a stator vane assembly having a mechanical retention device.SOLUTION: A stator vane assembly 200 includes a platform 201 that defines an opening 102. The stator vane assembly 200 further includes an airfoil segment 206 that includes an airfoil 56 and a boss 216. The airfoil 56 includes a leading edge 60, a trailing edge 62, a negative pressure side wall 66, and a positive pressure side wall 64. The airfoil 56 extends radially between a base end 68 and a tip 70. The boss 216 extends from at least one of the base end 68 or the tip 70 of the airfoil 56. The boss 216 is disposed in the opening 102 of the platform 201. The stator vane assembly further includes a mechanical retention device 210 that couples the platform 201 to the boss 216 of the airfoil segment 206.SELECTED DRAWING: Figure 3

Description

[Technical field]

[0001] SUMMARY The present disclosure relates generally to a stator vane having multiple components held together by mechanical retention devices. [Background technology]

[0002] Turbomachinery is used in various industries and applications for energy transfer. For example, a gas turbine engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section gradually increases the pressure of the working fluid entering the gas turbine engine and supplies the compressed working fluid to the combustion section. The compressed working fluid and fuel (e.g., natural gas) are mixed in the combustion section and combusted in the combustion chamber to generate high-pressure, high-temperature combustion gases. The combustion gases enter the turbine section from the combustion section and expand to generate work. For example, the expansion of the combustion gases in the turbine section rotates a rotor shaft connected to, for example, a generator, generating electricity. The spent combustion gases are exhausted from the turbine section through the exhaust section.

[0003] During operation of the turbomachine, various hot gas path components of the system are exposed to high temperature streams that can cause the hot gas path components to fail. Since high temperature streams generally improve the performance, efficiency, and power output of the turbomachine, the hot gas path components that are exposed to the high temperature streams must be cooled so that the gas turbine system can operate with the elevated temperature streams.

[0004] Forced air cooling is required when the maximum localized temperatures of hot gas path components approach the melting temperature of the hot gas path components. This often requires complex cooling schemes for turbine rotor blades and stationary nozzle airfoils. In this cooling configuration, air (typically bleed air from the compressor section) flows through internal cooling passages in the airfoil and is exhausted through cooling holes in the airfoil surface to transfer heat to the hot gas path components.

[0005] Many complex cooling schemes use small cooling passages (or microchannels) to deliver cooling fluid to the airfoil. Such cooling schemes create significant fabrication challenges for cores and castings and can significantly increase the manufacturing cost of hot gas path components that use such known near wall cooling systems. To address the manufacturing challenges due to complex and / or small cooling channels near the surface of the component, many hot gas path components with such features may be additively manufactured. Additive manufacturing can produce components with complex and diverse cooling sections. However, additively manufacturing a hot gas path component (such as a rotor blade or stator vane) as a single part can be costly and time consuming. Furthermore, a manufacturing error in one portion of the hot gas path component can result in the entire part being scrapped.

[0006] As such, it may be advantageous to manufacture the hot gas path component as multiple subcomponents. However, joining the subcomponents to form the hot gas path component may be difficult due to the complex geometries of the subcomponents. Furthermore, the joints formed between the subcomponents may be particularly sensitive and / or susceptible to failure when exposed to the hot combustion gases generated during operation of the turbomachine and flowing through the hot gas path in the turbine section.

[0007] Accordingly, it is believed that an improved hot gas path component having one or more subcomponents joined together and capable of withstanding the high temperatures of combustion gases without risk of joint failure would be desirable and appreciated in the art. Summary of the Invention

[0008] Aspects and advantages of the stator vane assembly and turbomachine according to the present disclosure may be set forth in part in the description which follows, may be obvious from the description, or may be learned by practice of the art.

[0009] According to one embodiment, a stator vane assembly is provided. The stator vane assembly includes a platform defining an opening. The stator vane assembly further includes an airfoil segment including an airfoil and a boss. The airfoil includes a leading edge, a trailing edge, a suction sidewall, and a pressure sidewall. The airfoil extends radially between a base end and a tip end. The boss extends from at least one of the base end and the tip end of the airfoil. The boss is disposed in an opening in the platform. The stator vane assembly further includes a mechanical retention device coupling the platform to the boss of the airfoil segment.

[0010] According to another embodiment, a turbomachine is provided. The turbomachine defines an axial direction, a radial direction, and a circumferential direction. The turbomachine includes a compressor section, a combustion section, and a turbine section. At least one of the compressor section or the turbine section includes a stator vane assembly. The stator vane assembly includes a platform defining an opening. The stator vane assembly further includes an airfoil segment including an airfoil and a boss. The airfoil includes a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base end and a tip end. The boss extends from at least one of the base end and the tip end of the airfoil. The boss is disposed in an opening in the platform. The stator vane assembly further includes a mechanical retention device coupling the platform to the boss of the airfoil segment.

[0011] These and other features, aspects and advantages of the stator vane assembly and method of the present invention will become better understood with reference to the following description and claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present technology and, together with the description, serve to explain the principles of the technology. [Brief description of the drawings]

[0012] A complete and enabling disclosure of the stator vane assembly and turbomachine of the present invention, including the best mode, directed to one of ordinary skill in the art, of making and using the system and method, is set forth in the specification, which makes reference to the following figures:

[0013] [Figure 1] 1 is a schematic diagram of a turbomachine (e.g., a gas turbine engine) according to an embodiment of the present disclosure. [Diagram 2] 2 is a partial cross-sectional side view of a turbine section of the gas turbine engine shown in FIG. 1 according to an embodiment of the present disclosure. [Diagram 3] FIG. 2 is an exploded view of a stator vane according to an embodiment of the present disclosure. [Figure 4] FIG. 2 illustrates a perspective view of an airfoil segment according to an exemplary embodiment of the present disclosure. [Diagram 5] FIG. 2 is a perspective view of an outer portion including an airfoil segment and an outer platform of a stator vane assembly according to an exemplary embodiment of the present disclosure; [Figure 6] 6 illustrates a cross-sectional perspective view of a stator vane assembly taken along line 6-6 shown in FIG. 5 according to an exemplary embodiment of the present disclosure. [Figure 7] FIG. 2 is a plan view of a stator vane assembly according to an exemplary embodiment of the present disclosure. [Figure 8] FIG. 8 is a cross-sectional perspective view of a stator vane assembly taken along line 8-8 shown in FIG. 7 according to an exemplary embodiment of the present disclosure. [Figure 9] FIG. 2 is a perspective view of an inner portion including an airfoil segment and an inner platform of a stator vane assembly according to an exemplary embodiment of the present disclosure; [Figure 10] 10 illustrates a cross-sectional perspective view of a stator vane assembly taken along line 10-10 shown in FIG. 9 according to an exemplary embodiment of the present disclosure. [Figure 11] FIG. 2 is a cross-sectional perspective view of a stator vane assembly according to an embodiment of the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0014] Reference will now be made in detail to embodiments of the stator vane assembly and turbomachine of the present invention, one or more examples of which are illustrated in the drawings. Each example is provided for purposes of explanation, not limitation of the technology. Indeed, it will be apparent to those skilled in the art that modifications and variations of the technology can be made without departing from the scope or spirit of the claimed technology. For example, features illustrated or described as part of one embodiment can be used with another embodiment to yield still a further embodiment. Thus, it is intended that the present disclosure cover such modifications and variations as come within the scope of the claims and their equivalents.

[0015] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Moreover, unless otherwise specified, all embodiments described herein should be considered exemplary.

[0016] In the detailed description, numerical and letter designations are used to refer to features in the drawings. Like or similar designations in the drawings and description are used to represent like or similar parts of the invention. In this specification, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, and are not intended to imply a position or importance of the individual components.

[0017] The term "fluid" may be a gas or a liquid. The term "fluid communication" means that two or more regions defining a flow path are joined to one another to allow fluid communication between the defined regions.

[0018] As used herein, the terms "upstream" (or "forward") and "downstream" (or "aft") refer to relative directions with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which fluid flows and "downstream" refers to the direction to which fluid flows. The term "radially" refers to relative directions that are substantially perpendicular to an axial centerline of a particular component, the term "axially" refers to relative directions that are substantially parallel and / or coaxial to an axial centerline of a particular component, and the term "circumferentially" refers to relative directions that extend around an axial centerline of a particular component.

[0019] Approximate terms (such as "approximately," "about," "generally," "substantially," etc.) are not intended to be limited to the exact value specified. In at least some examples, approximation language may correspond to the precision of an instrument for measuring a value, or the precision of a method or machine for constructing or manufacturing a component and / or system. In at least some examples, approximation language may correspond to the precision of an instrument for measuring a value, or the precision of a method or machine for constructing or manufacturing a component and / or system. For example, approximation language may express within a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% for either an individual value, a range of values, and / or an end value defining a range of values. Such terms, when used in the context of an angle or direction, include a range of 10 degrees greater or less than the stated angle or direction. For example, "approximately vertical" includes a direction within 10 degrees in any direction (e.g., clockwise or counterclockwise) from vertical.

[0020] Terms such as "coupled," "fixed," "attached," and the like, unless otherwise specified herein, refer to both direct coupling, fixing, or attachment, and indirect coupling, fixing, or attachment through one or more intermediate components or parts. Terms such as "directly coupled," "directly fixed," "directly attached," and the like, mean that two components are attached in contact with each other and there are no intermediate components or parts present.

[0021] As used herein, the terms "comprises, comprising, including, including, having, having" or other variations thereof are intended to cover a non-exclusive inclusion. For example, a process, method, article, or apparatus that includes multiple features is not necessarily limited to only those features, but may include other features not expressly recited or inherent to such process, method, article, or apparatus. Additionally, unless expressly stated to the contrary, "and / or" refers to a condition that is satisfied by any one of the following: A is true (or present), B is false (or absent), A is false (or absent), B is true (or present), and both A and B are true (or present).

[0022] Throughout this specification and claims, range limitations are combinable and interchangeable, and such ranges are specified and include all subranges contained therein, unless otherwise indicated by context or language. For example, all ranges disclosed herein are inclusive of the end values, and the end values ​​are independently combinable with each other.

[0023] The phrase "near" refers to one end being closer than the opposite end. For example, when used with first and second ends, high and low pressure sides, leading and trailing edges, etc., the phrases "near the first end" or "near the high pressure side" refer to a location closer to the first end than the second end, or closer to the high pressure side than the low pressure side, respectively.

[0024] As used herein, a "solid" is a component that has no voids, cavities, pores, or other openings, and may refer to the component as being impermeable, such that air or other fluids do not pass through the component.

[0025] Referring now to the drawings, Figure 1 shows a schematic diagram of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine engine 10. Although an industrial or land-based gas turbine engine is illustrated and described herein, the present disclosure is not limited to industrial or land-based gas turbine engines unless otherwise specified in the claims. For example, the invention described herein may be used with any type of turbomachine, such as, but not limited to, a steam turbine, an aircraft gas turbine, or a marine gas turbine.

[0026] As shown in Figure 1, a gas turbine engine 10 generally includes a compressor section 12. The compressor section 12 includes a compressor 14. The compressor section 12 includes an inlet 16 disposed at an upstream end of the gas turbine engine 10. The gas turbine engine 10 further includes a combustion section 18 having one or more combustors 20 disposed downstream from the compressor section 12. The gas turbine engine 10 further includes a turbine section 22 (i.e., expansion turbine) downstream from the combustion section 18. A shaft 24 extends generally axially through the gas turbine engine 10.

[0027] The compressor section 12 may generally include a plurality of rotor disks 21 and a plurality of rotor blades 23 extending radially outward from and connected to each rotor disk 21. Each rotor disk 21 may in turn be coupled to or form a portion of a shaft 24 that passes through the compressor section 12. The rotor blades 23 of the compressor section 12 may include turbomachine airfoils that define an airfoil shape (e.g., a shape having a leading edge, a trailing edge, and sidewalls extending between the leading edge and the trailing edge). Additionally, in many embodiments, the compressor section 12 may include stator vanes 19 disposed between the rotor blades 23. The stator vanes 19 may extend from and be coupled to the compressor casing 11.

[0028] The turbine section 22 may generally include a plurality of rotor disks 27 and a plurality of rotor blades 28 extending radially outward from and interconnected to each rotor disk 27. Each rotor disk 27 may in turn be coupled to or may form a portion of the shaft 24 that passes through the turbine section 22. The turbine section 22 may further include an outer casing 32 that circumferentially surrounds the portion of the shaft 24 and the rotor blades 28. The turbine section 22 may include stator vanes or static nozzles 26 that extend radially inward from the outer casing 32. The rotor blades 28 and stator vanes 26 may be arranged in alternating stages along an axial centerline 30 of the gas turbine 10. Both the rotor blades 28 and the stator vanes 26 may include turbomachine airfoils that define an airfoil shape (e.g., a shape having a leading edge, a trailing edge, and sidewalls extending between the leading edge and the trailing edge).

[0029] In operation, ambient air 36 or other working fluid is drawn into the inlet 16 of the compressor 14 and gradually compressed to provide compressed air 38 to the combustion section 18. The compressed air 38 enters the combustion section 18 and is mixed with fuel to form a combustible mixture. The combustible mixture is combusted in a combustion chamber 40 of the combustor 20, which generates combustion gases 42 that pass from the combustion chamber 40 to the turbine section 22. Energy (kinetic and / or thermal energy) is transferred from the combustion gases 42 to the rotor blades 28, which rotate the shaft 24 to produce mechanical work. The spent combustion gases 42 (also referred to as "exhaust gases") exit the turbine section 22 and flow through the exhaust diffuser 34, traversing a number of struts or main airfoils 44 disposed within the exhaust diffuser 34.

[0030] The gas turbine 10 may define a cylindrical coordinate system having an axial direction A extending along the axial centerline 30, a radial direction R perpendicular to the axial centerline 30, and a circumferential direction C extending around the axial centerline 30.

[0031] 2 is a partial cross-sectional side view of turbine section 22 of gas turbine engine 10 according to an embodiment of the present disclosure. Turbine section 22 may have one or more stages 50. Each stage 50 may include a set of rotor blades 28 coupled to a rotor disk 27, which may be rotatably mounted on shaft 24. Each of the one or more stages 50 further includes a set of stator vanes 26 axially adjacent to the set of rotor blades 28. The stator vanes 26 described herein may be employed in a first stage, a second stage, a third stage, a fourth stage, or multiple stages among the first through fourth stages, where the "first stage" refers to the stage immediately downstream of combustion section 18.

[0032] Each stator vane 26 may include at least one airfoil 56 extending in a radial direction R between an inner platform or endwall 52 and an outer platform or endwall 54. The circumferentially adjacent outer platforms 54 of each stator vane 26 may be coupled together to form an outer annular ring extending around an inner annular ring formed by the circumferentially adjacent inner platforms 52 of each stator vane 26. The at least one airfoil 56 may extend between the two annular rings formed by the platforms 52, 54. The turbine section 22 may also include a shroud segment 58 disposed downstream of the outer platforms 54 to direct the combustion gases 42 flowing through the stator vanes 26 to the rotor blades 28.

[0033] Structures or components disposed along the flow path of the combustion gases 42 may be referred to as hot gas path components. In one example, the hot gas path components may be the stator vanes 26 and / or rotor blades 28. In some embodiments, cooling sections (impingement sleeves, cooling channels, cooling holes, etc.) may be disposed within the hot gas path components as indicated by dashed lines 78 to cool the hot gas path components. For example, cooling air, indicated by arrows 79, may be routed from the compressor section 12 or elsewhere and flow through the cooling sections as indicated by arrows 81. Although such cooling flows are illustrated in FIG. 2 with respect to the stator vanes 2, it should be understood that similar cooling flows may be used for the rotor blades 28.

[0034] FIG. 3 illustrates an exploded view of a stator vane assembly 200. The stator vane assembly 200 may be representative of a stator vane 19 of the compressor section 12 and / or a stator vane 26 of the turbine section 22. However, in an exemplary embodiment, the stator vane assembly 200 is a stator vane 26 of the turbine section 22. As shown in FIG. 3, the stator vane assembly 200 includes a platform 201 and an airfoil segment 206. For example, the stator vane assembly 200 may include an inner platform 202, an outer platform 204, and an airfoil segment 206. The inner platform 202, the outer platform 204, and the airfoil segment 206 are separated from one another along the axial centerline 55 of the stator vane assembly 200 to illustrate various features of the stator vane assembly 200 that are not visible once the stator vane assembly is constructed. The axial centerline 55 may be generally parallel to a radial direction of the gas turbine 10 when the stator vane assembly 200 is installed in the turbine section 22. The airfoil segment 206 may include an airfoil 56 that generally defines an aerodynamic shape or contour. For example, the airfoil 56 may include a leading edge 60 that contacts the flow of combustion gases and sidewalls that guide the combustion gases along the airfoil 56 to a trailing edge 62. For example, the airfoil 56 may include a pressure sidewall 64 and a suction sidewall 66 that extend from the leading edge 60 to the trailing edge 62. Further, the airfoil 56 may extend radially between a base end 68 and a tip end 70.

[0035] In an exemplary embodiment, the inner platform 202, the airfoil 56, and the outer platform 204 may be separate components (e.g., manufactured as separate components) that are coupled together by a mechanical retention fixture 210. In an exemplary embodiment, the mechanical retention fixture 210 may include one or more retention pins 212 that pass through bosses 216 of the platform 201 and the airfoil segment 206. In some embodiments, the one or more retention pins 212 may include a leading edge pin 213 and a trailing edge pin 214.

[0036] While this exemplary embodiment of stator vane assembly 200 includes a mechanical retention device 210 in the form of one or more retention pins 212, the advantage of which is ease of assembly, it should be understood that other mechanical retention devices may be used without departing from the scope of the present disclosure. For example, in other embodiments, mechanical retention device 210 may include one or more of a seam profile, a snap ring, a machined tongue and groove, or other means. The present disclosure should not be limited to embodiments having one or more retention pins as the mechanical retention device, unless expressly recited in the claims.

[0037] Each of the one or more retaining pins 212 may be a solid (e.g., not hollow, not including voids) cylindrically shaped pin. In this configuration, each of the retaining pins 212 may extend linearly along a centerline 215, and each of the retaining pins 212 may have a circular cross-sectional shape. However, in other embodiments, the retaining pins 212 may have other cross-sectional shapes, and the present disclosure should not be limited to a particular shape unless specifically recited in the claims. In various embodiments, the one or more retaining pins 212 may be formed from a metal (such as a nickel superalloy or other metal or metal alloy). However, in other embodiments, the retaining pins 212 may be formed from another material (such as a ceramic), and the pins of the present disclosure should not be limited to a particular material unless specifically recited in the claims.

[0038] For example, in certain embodiments, each of the inner platform 202, the airfoil 56, and the outer platform 204 may be additively manufactured (e.g., 3D printed) separately and then joined together by a mechanical holding fixture 210 and / or by one or more metallurgical bonds. In various embodiments, the metallurgical bonds may include weld bonds, braze bonds, sinter bonds, etc. In an exemplary embodiment, the metallurgical bonds may be braze bonds. Forming the stator vane assembly 200 as three separate pieces advantageously increases the repairability of the stator vane assembly 200. For example, if a portion of the airfoil segment 206 is damaged, the entire stator vane assembly 200 does not need to be replaced. The mechanical connections can be released (e.g., by removing one or more pins 212) and the metallurgical connections can be released (e.g., by reheating the components to melt them apart) to separate the airfoil segment 206 from the platforms 202, 204, and a new airfoil segment 206 can be employed or the old airfoil can be repaired.

[0039] Furthermore, in examples where additive manufacturing is used to manufacture some or all of the stator vane assembly 200, the build time (i.e., printing) of the airfoil segments 206, which are separate components from the inner and outer platforms 202, 204, is significantly shorter as compared to the build time of a unitary nozzle in which the inner and outer platforms 202, 204 are printed along with the airfoil segments 206. Additionally, forming the stator vane assembly 200 from three separate components simplifies the manufacturing process and allows for the use of different manufacturing techniques and / or different materials for the various components.

[0040] 3, the airfoil segment 206 may include a camber axis 57 that may extend from the leading edge 60 to the trailing edge 62 and may be defined midway between the pressure sidewall 64 and the suction sidewall 66 of the airfoil 56. The camber axis 57 may be curved and / or contoured to correspond to the curvatures of the pressure sidewall 64 and the suction sidewall 66. A lateral direction T may be defined as perpendicular to the camber axis 57.

[0041] 3, the airfoil segment 206 may further include a boss 216 extending from at least one of the base end 68 and the tip end 70 of the airfoil 56. The boss 216 may extend outwardly from the airfoil 56 in both the radial direction R and the lateral direction T. For example, the stator vane assembly 200 may include an inner boss 73 extending radially inwardly from the airfoil 56 at the base end 68 and an outer boss 75 extending radially outwardly from the airfoil 56 at the tip end 70. Each boss 73, 75 may extend generally perpendicularly in the lateral direction T from the pressure sidewall 64 and the suction sidewall 66.

[0042] Both the inner platform 202 and the outer platform 204 may define an opening 102. The boss 216 is disposed within the opening 102 of the platforms 202, 204. The opening 102 is sized and shaped to correspond to the respective boss 73, 75, which may be inserted into the opening 102 (and then connected to the respective platform 202, 204 by mechanical retention devices 208 and / or brazed connections) to couple the airfoil segment 206 to the inner platform 202 and the outer platform 204. In particular, the inner boss 73 may be inserted into the opening 102 of the inner platform 202 (and then connected to the inner platform 202), and the outer boss 75 may be inserted into the opening 102 of the outer platform 204 (and then connected to the outer platform 204).

[0043] In many embodiments, in addition to or as an alternative to the mechanical retention 208, the bosses 216 of the airfoil segments 206 may be brazed to the inner and outer platforms 52, 54 by a metallurgical bond 211 (FIGS. 6 and 10). In an exemplary embodiment, the metallurgical bond 211 may be a brazed bond. In such an embodiment, the braze material 107 may be disposed between the bosses 216 and the boundaries of the openings 102 (on both the inner and outer platforms 202, 204). The entire stator vane assembly 200 may then be placed in a brazing oven to melt the braze material 107, which may then be solidified, thereby joining the inner and outer platforms 202, 204 to their respective bosses 216 of the airfoil segments 206. The braze material 107 may include copper, nickel, silver, gold, aluminum, or other suitable braze metal.

[0044] In many embodiments, as shown in FIG. 3 , the inner platform 202 and / or the outer platform 204 include a floor 53. The floor 53 may extend generally perpendicular to the airfoil 56. The floor 53 of the inner platform 202 may define a radially inner flow boundary of the combustion gases of the turbine section 22. Similarly, the floor 53 of the outer platform 204 may define a radially outer flow boundary of the combustion gases of the turbine section 22. Additionally, the floors 53 of the inner platform 202 and the outer platform 204 may at least partially define the respective openings 102. In an exemplary embodiment, one or both of the inner platform 202 and / or the outer platform 204 may include one or more protrusions 80 extending radially from the floor 53. The one or more protrusions 80 may at least partially define the openings 102. In many embodiments, the boss 216 may be brazed to the one or more protrusions 80.

[0045] 3, the one or more protrusions 80 may be a single protrusion 80 that annularly surrounds the opening 102. However, in other embodiments, the one or more protrusions 80 may include multiple protrusions 80 that are spaced apart from one another and that are non-contiguous about the opening 102.

[0046] In certain embodiments, the boss 73, 216 extending from the base end 68 of the airfoil 56 may span the opening 102 of the inner platform 202, such that the inner platform 202 surrounds the boss 73, 216 of the airfoil 56. In particular, the floor 53 and / or one or more protrusions 80 of the inner platform 202 may surround the boss 73, 216 extending from the base end 68 of the airfoil 56. Similarly, the boss 75, 216 extending from the tip end 70 of the airfoil 56 may span the opening 102 of the outer platform 204, such that the outer platform 204 surrounds the boss 75, 216 of the airfoil 56. In particular, the floor 53 and / or one or more protrusions 80 of the outer platform 204 may surround the boss 75, 216 extending from the tip end 70 of the airfoil 56.

[0047] In some embodiments, as shown, the boss 216 may at least partially define the cavity 92. The cavity 92 extends into and is defined within the airfoil 56. The cavity 92 is exposed by an opening 102 in the platform 54 to allow air (e.g., bleed air from the compressor 14) to enter the cavity 92. In other embodiments (not shown), the boss 216 may be solid (i.e., the boss 216 does not have a void, cavity, or opening defined therein).

[0048] 4, a perspective view of an airfoil segment 206 is shown in accordance with an embodiment of the present disclosure. As shown, the airfoil segment 206 may include an airfoil 56 and at least one boss 216. The airfoil 56 may include a leading edge 60 that contacts the working fluid and a trailing edge 62 opposite the leading edge 60. Additionally, the airfoil 56 may include a pressure sidewall 64 (FIG. 3) and a suction sidewall 66, each of which extends from the leading edge 60 to the trailing edge 62. The airfoil 56 may extend radially between a base end 68 and a tip end 70. In many embodiments, the boss 216 may extend from at least one of the base end 68 and the tip end 70 of the airfoil 56. In particular, in an exemplary embodiment, the boss 216 may extend from both the base end 68 and the tip end 70. In such an embodiment, the airfoil segment 206 may include an inner boss 73 extending radially inward from the base end 68 and an outer boss 75 extending radially outward from the tip end 70. The inner boss 73 and the outer boss 75 may have different radial heights and may include different features as described below.

[0049] In many embodiments, the boss 216 may include a shoulder 218 and a shank portion 220. The shoulder 218 may extend (e.g., radially) from and surround the tip 70 of the airfoil 56. In particular, the outer boss 75 may include both the shoulder 218 and the shank portion 220, and the inner boss 73 may include only the shoulder 218 (without a shank portion). The shank portion 220 may extend radially (e.g., radially outward) from the shoulder 218. As shown in FIG. 4, the shank portion 220 may include a shank leading edge 222, a shank trailing edge 224, a shank suction side 226, and a shank pressure side 228. The shank leading edge 222 is proximate the leading edge 60 of the airfoil 56, while the shank trailing edge 224 is offset from the trailing edge 62 of the airfoil 56. Additionally, the shank portion 220 may include a solid shank aft portion 230 that extends from the shank suction side 226 to the shank pressure side 228 and proximate the shank trailing edge 224. The shank portion 220 of the boss 216 may at least partially define the cavity 92, which extends into and is defined in the shoulder 218 and the airfoil 56.

[0050] In various embodiments, the shoulder 218 may include a shoulder leading edge 232, a shoulder trailing edge 234, a shoulder suction side 236, and a shoulder pressure side 238. The shoulder leading edge 232 is adjacent to the leading edge 60 of the airfoil 56 and aligned with the shank leading edge 222, and the shoulder trailing edge 234 is adjacent to the trailing edge 62 of the airfoil and offset from the shank trailing edge 224. Additionally, the shoulder 218 may include a solid shoulder aft portion 240. The solid shoulder aft portion 240 may extend downstream of both the shank trailing edge 224 and the trailing edge 62 of the airfoil 56 such that the solid shoulder aft portion 240 defines a support surface 242. The support surface 242 may contact (and be brazed coupled to) a radially inner surface of the outer platform 204 when the stator vane assembly 200 is constructed.

[0051] The boss 216 may have a generally similar profile as the airfoil 56, although the boss 216 may have a larger size than the airfoil 56. In many embodiments, the boss 216 may define one or more generally planar (e.g., flat) undercut surfaces 244. The undercut surfaces 244 may extend along the shank portion 220 and / or shoulder portion 218 of the boss 216. At least one of the retaining pins 212 may pass through one or more of the undercut surfaces 244. An advantage of the undercut surfaces 244 is that they may provide a secure connection between the platforms 202, 202 and the boss 216.

[0052] 4, the one or more retaining pins 212 may include a leading edge pin 213, a trailing edge pin 214, and an inner retaining pin 217. The leading edge pin 213 may pass through the outer boss 75 near the leading edge 60 of the airfoil 56. The leading edge pin 213 may be located in an undercut surface 244 (e.g., along the shank suction side 226) near the shank leading edge 222. The trailing edge pin 214 may pass through the solid shank aft portion 230 of the shank portion 220 of the outer boss 75 near the shank trailing edge 224. The inner retaining pin 217 (or inner leading edge pin) may pass through the inner boss 73 near the leading edge 60 of the airfoil 56. The inner retaining pin 217 may be located in the undercut surface 244.

[0053] 5-10, various views of a stator vane assembly 200 are illustrated in accordance with an exemplary embodiment of the present disclosure. In particular, FIG. 5 illustrates a perspective view of an outer portion of the stator vane assembly 200, including the airfoil segments 206 and the outer platform 204. FIG. 6 illustrates a cross-sectional perspective view of the stator vane assembly 200 taken along line 6-6 as shown in FIG. 5. FIG. 7 illustrates a top view of the stator vane assembly 200. FIG. 8 illustrates a cross-sectional perspective view of the stator vane assembly 200 taken along line 8-8 as shown in FIG. 7. FIG. 9 illustrates a perspective view of an inner portion of the stator vane assembly 200, including the airfoil segments 206 and the inner platform 202. FIG. 10 illustrates a cross-sectional perspective view of the stator vane assembly 200 taken along line 10-10 as shown in FIG. 9.

[0054] 5-10, each platform 202, 204 may include a floor 53 and one or more protrusions 80 extending radially from the floor 53. The one or more protrusions 80 may be generally perpendicular to the floor 53. In some embodiments, as shown in FIGS. 5-8, the one or more protrusions 80 may include a plurality of protrusions 246. The plurality of protrusions 246 are spaced apart from one another and disposed about the boss 216 of the airfoil segment 206. In particular, the outer platform 204 may include a plurality of protrusions 246. The plurality of protrusions 246 are spaced apart from one another and disposed about the shank portion 220 of the boss 216, with each protrusion 246 in contact with the shank portion 220 of the boss 216. Each protrusion 246 of the plurality of protrusions 246 may define a hole through which a respective retaining pin 212 passes. 9 and 10 , the one or more protrusions 80 may be a continuous protrusion 248 that surrounds the boss 216. In particular, the inner platform 202 includes a continuous protrusion 248 that extends radially inward from the floor 53, where the continuous protrusion 248 may contact the inner boss 75.

[0055] Each platform 202, 204 may include a pressure side slashface 250, a suction side slashface 252 opposite the pressure side slashface 250, a leading edge face 254, and a trailing edge face 256 opposite the leading edge face 254. Each of the leading edge face 254 and the trailing edge face 256 may extend between the pressure side slashface 250 and the suction side slashface 252. The suction side slashface 252 and the pressure side slashface 250 may contact an adjacent stator vane assembly 200 when positioned within a gas turbine. The suction side slashface 252 and the pressure side slashface 250 may be generally parallel to one another, and the leading edge face 254 and the trailing edge face 256 may be generally parallel to one another.

[0056] In many embodiments, each platform 202, 204 may include a leading edge hook rail 258 and a trailing edge hook rail 260 extending (e.g., radially) from the floor 53. The leading edge hook rail 258 may extend radially from the floor 53 at or near the leading surface 254. The trailing edge hook rail 260 may extend radially from the floor 53 at or near the trailing surface 256. The hook rails 258, 260 may extend generally parallel to the leading surface 254 and the trailing surface 256. The hook rails 258, 260 may be located on either side (e.g., forward and aft) of the shank portion 220 of the boss 216. The hook rails 258, 260 may be attached to an adjacent shroud assembly, a turbine or compressor casing, or a stationary retaining ring.

[0057] As shown in FIGS. 5-10 , the one or more retaining pins 212 may include a leading edge pin 213, a trailing edge pin 214, and an inner retaining pin 217. The leading edge pin 213 may pass through the outer boss 75 near the leading edge 60 of the airfoil 56. The leading edge pin 213 may be located at an undercut surface 244, which may be located at the shank suction side 226 of the shank portion 220. The trailing edge pin 214 may pass through the solid shank aft portion 230 of the shank portion 220 of the outer boss 75 near the trailing edge 62 of the airfoil 56. The inner retaining pin 217 (or inner leading edge pin) may pass through the inner boss 73 near the leading edge 60 of the airfoil 56. The inner retaining pin 217 may be located at the undercut surface 244.

[0058] As best shown in FIG. 6 , the leading edge pin 213 may extend through a first protrusion 262 of the plurality of protrusions 246 and through the boss 216 into a second protrusion 264 of the plurality of protrusions 246. The first protrusion 262 may extend from the floor 53 and may be located near the leading edge hook rail 258 and / or may be located near the leading edge 60 of the airfoil 56. The second protrusion 264 may extend from the floor 53 and may be located at (or near) the suction side slashface 252. In particular, the leading edge pin 213 may extend through the first protrusion 262, through the shank pressure side 228, across the cavity 92, through the shank suction side 226 (at the undercut surface 244), and into the second protrusion 264.

[0059] Further, with respect to the leading edge pin 213, as best shown in FIG. 7, the leading edge pin 213 may be oriented generally perpendicular to the axial direction A of the gas turbine engine 10. Stated another way, the leading edge pin 213 may be oriented generally perpendicular to the flow direction 268 ( FIG. 5 ) of the fluid contacting the leading edge 60 of the airfoil 56. The aerodynamic loads on the airfoil 56 are generally along the flow direction 268 (and / or axial direction A) and create a building moment on the stator vane assembly 200. Thus, orienting the leading edge pin 213 perpendicular to the axial direction A and positioning the leading edge pin 213 near the leading edge 60 may advantageously allow the leading edge pin 213 to cushion the aerodynamic loads, thereby improving the robustness (and overall hardware life) of the stator vane assembly 200.

[0060] 7 and 8 , the trailing edge pin 214 may extend through a third protrusion 266 of the plurality of protrusions 246, through the solid shank aft portion 230, and into the trailing edge hook rail 260. The trailing edge pin 214 may be oriented to counteract a moment resulting from the leading edge pin 213. In particular, the trailing edge pin 213 may be oriented obliquely relative to the axial direction A and / or the flow direction 268.

[0061] 9 and 10, an inner retaining pin 217 (or inner leading edge pin) may extend through the inner boss 73 near the leading edge 60 of the airfoil 56. However, unlike the leading edge pins 213 and the trailing edge pins 214 on the outer platform 204, the inner retaining pin 217 may not extend completely through the continuous projection 248 such that at least one end of the retaining pin 217 is received within the projection 248. For example, the inner retaining pin 217 may be mounted through an opening on one side of the inner boss 73 and extend through the inner boss 73 into (but not through) the continuous projection 248 on the opposite side of the inner boss 73. The inner boss 73 may include mounting holes 280 by which the inner retaining pin 217 may be mounted as shown. The mounting holes 280 may be generally radially oriented and defined in the inner boss 73. The inner retaining pin 217 may be disposed in the undercut surface 244. The inner retaining pin 217 may be oriented generally parallel to the leading edge pin 213 (eg, generally perpendicular to the axial direction A).

[0062] In various embodiments, a metallurgical bond 211 may be disposed between the platforms 202, 204 and the respective bosses 73, 75 of the airfoil segment 206, and the metallurgical bond 211 may fixedly couple the platforms 202, 204 to the respective bosses 73, 75 of the airfoil segment 206. In many embodiments, the metallurgical bond 211 may extend annularly around (or surround) the shoulders 218 of the bosses 73, 75 to couple the shoulders 218 to the respective platforms 202, 204. In an exemplary embodiment, the metallurgical bond 211 may be a braze bond. In such an embodiment, the braze material 107 may be disposed between the bosses 216 and the boundary of the opening 102 (on both the inner platform 202 and the outer platform 204). The entire stator vane assembly 200 may then be placed in a brazing oven to melt the braze material 107, which may then be solidified, thereby joining the inner and outer platforms 202 and 204 to their respective bosses 216 of the airfoil segments 206. The braze material 107 may include copper, nickel, silver, gold, aluminum, or other suitable braze metal.

[0063] 6, the height 270 (e.g., radial height) of the metallurgical bond 211 may be equal to (e.g., about 100%) or greater than (e.g., greater than 200% or greater than 300%) the thickness 272 of the walls 64, 66 of the airfoil 56. In particular, for the connection between the outer boss 75 and the outer platform 204, the height 270 of the metallurgical bond 211 may be equal to (e.g., about 100%) or greater than (e.g., greater than 200% or greater than 300%) the thickness 272 of the walls 64, 66 of the airfoil 56 at about 90% span of the airfoil 56. Similarly, for the connection between the inner boss 73 and the inner platform 202, the height 270 of the metal bond 211 may be equal to the thickness 272 of the walls 64, 66 of the airfoil 56 (e.g., about 100%) at about 10% span of the airfoil 56, or may be greater than the thickness 272 (e.g., greater than 200%, or greater than 300%).

[0064] As used herein, "% span" may represent a spanwise position of the airfoil 56, which extends radially from the base end 68 to the tip end 70 of the airfoil 56. For example, 0% span may represent the base end 68 of the airfoil 56, 100% span may represent the tip end 70 of the airfoil 56, and 50% span may represent a position halfway between the base end 68 and the tip end 70 of the airfoil 56. Having a height 270 of the metallurgical bond 211 that is greater than 300% of the thickness 272 of the walls 64, 66 of the airfoil 56 may advantageously ensure that the metallurgical bond 211 is stronger (e.g., in shear strength) than the walls 64, 66 of the airfoil 56, such that the walls 64, 66 will fail before the metallurgical bond 211 (if failure occurs).

[0065] 11, a cutaway perspective view of a stator vane assembly 200 according to an alternative embodiment of the present disclosure is shown illustrating the connection between the inner boss 73 and the inner platform 202. As shown in FIG. 11, in some embodiments, the inner retaining pin 217 can penetrate completely through the continuous protrusion 248 on both sides such that both ends of the inner retaining pin 217 are exposed. For example, the inner retaining pin 217 can penetrate the single protrusion 248 at a first location, penetrate the inner boss 73, and penetrate the single protrusion 248 at a second location different from the first location.

[0066] An advantage of forming the stator vane assembly 200 as three separate pieces is that it improves the repairability of the stator vane assembly 200. For example, if a portion of the airfoil segment 206 is damaged, the entire stator vane assembly 200 does not need to be replaced. The airfoil segment 206 can be separated from the platforms 202, 204 by removing the mechanical connections (e.g., by removing one or more pins 212) and a new airfoil segment 206 can be employed or the old airfoil can be repaired. Similarly, if an improved airfoil segment cooling scheme is developed, the improved airfoil segment can be attached to the existing platform, thereby improving turbine efficiency and minimizing manufacturing costs.

[0067] This specification uses examples to disclose the invention, including the best mode, and also enables any person skilled in the art to practice the invention (e.g., make and use any devices or systems, and perform any methods incorporating the same). 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 contain structural elements that do not differ from the literal language of the claims, or if they contain equivalent structural elements that are not substantially different from the literal language of the claims.

[0068] Further aspects of the invention are provided by the following embodiments. [Embodiment 1] 1. A stator vane assembly comprising: an airfoil segment having a platform, an airfoil and a boss defining an opening, the airfoil including a leading edge, a trailing edge, a suction sidewall and a pressure sidewall, the airfoil extending radially between a base end and a tip end, the boss extending from at least one of the base end and the tip end of the airfoil, the boss being disposed in an opening in the platform; and a mechanical retention device coupling the platform to the boss of the airfoil segment. [Embodiment 2] 2. The stator vane assembly of embodiment 1, wherein a metallurgical bond is disposed between the platform and a boss of the airfoil segment, the metallurgical bond fixedly coupling the platform to the boss of the airfoil segment. [Embodiment 3] 3. A stator vane assembly as described in embodiment 1 or 2, wherein a height of the metallurgical bond is greater than about 300% of a thickness of one of the pressure side wall and the suction side wall. [Embodiment 4] A stator vane assembly as described in any one of embodiments 1 to 3, wherein the mechanical retention device includes one or more retention pins that pass through the platform and bosses of the airfoil segments. [Embodiment 5] The stator vane assembly of any one of embodiments 1 to 4, wherein the one or more retaining pins include a leading edge pin and a trailing edge pin. [Embodiment 6] A stator vane assembly as described in any one of embodiments 1 to 5, wherein the boss includes a shoulder portion and a shank portion, the shoulder portion extending from at least one of a base end and a tip end of the airfoil and surrounding one of the base end and the tip end, and the shank portion extending radially from the shoulder portion. [Embodiment 7] A stator vane assembly as described in any one of embodiments 1 to 6, wherein the shank portion includes a shank leading edge, a shank trailing edge, a shank suction side, and a shank pressure side, and the shank portion includes a solid shank aft portion extending from the shank suction side to the shank pressure side near the shank trailing edge. [Embodiment 8] A stator vane assembly as described in any one of embodiments 1 to 7, wherein a leading edge pin penetrates the platform, the shank pressure side, and the shank suction side near the shank leading edge, and a trailing edge pin penetrates the platform and the solid shank aft portion. [Embodiment 9] A stator vane assembly as described in any one of embodiments 1 to 8, wherein the platform includes a floor and one or more protrusions extending radially from the floor, and the mechanical retention device couples one of the one or more protrusions of the platform to a boss of the airfoil segment. [Embodiment 10] A stator vane assembly as described in any one of embodiments 1 to 9, wherein the platform includes a pressure side slashface, a suction side slashface, a leading edge surface, and a trailing edge surface, each of the leading edge surface and the trailing edge surface extending between the pressure side slashface and the suction side slashface, and the platform includes leading edge hook rails and trailing edge hook rails extending from the floor. [Embodiment 11] 11. The stator vane assembly of any one of embodiments 1-10, wherein the mechanical retention device is a leading edge pin, the leading edge pin extending through a first protrusion of the one or more protrusions and the boss into a second protrusion of the one or more protrusions, the second protrusion of the one or more protrusions being disposed on the pressure side slash face. [Embodiment 12] 12. The stator vane assembly of any one of embodiments 1 to 11, wherein the mechanical retention device is a trailing edge pin, the trailing edge pin extending through a first protrusion of the one or more protrusions and the solid shank aft portion, and into the trailing edge hook rail. [Embodiment 13] A stator vane assembly as described in any one of embodiments 1 to 12, wherein the platform defining an opening is an outer platform, and the boss extending from one of the base end or the tip end of the airfoil is a first boss extending from the tip end of the airfoil, and the stator vane assembly further includes an inner platform defining an inner platform opening, and a second boss extending from the base end of the airfoil and coupled to the inner platform by an inner retaining pin. [Embodiment 14] 1. A turbomachine defining an axial, radial, and circumferential direction, the turbomachine including a compressor section, a combustion section downstream of the compressor section, and a turbine section downstream of the combustion section, at least one of the compressor section or the turbine section including a stator vane assembly, the stator vane assembly including an airfoil segment having a platform, an airfoil, and a boss defining an opening, the airfoil including a leading edge, a trailing edge, a suction side wall, and a pressure side wall, the airfoil extending radially between a base end and a tip end, the boss extending from at least one of the base end and the tip end of the airfoil, the boss being disposed in an opening in the platform, and a mechanical retention device coupling the platform to the boss of the airfoil segment. [Embodiment 15] 15. The turbomachine of embodiment 14, wherein a metallurgical bond is disposed between the platform and a boss of the airfoil segment, the metallurgical bond fixedly coupling the platform to the boss of the airfoil segment. [Embodiment 16] 16. The turbomachine of embodiment 14 or 15, wherein the mechanical retention device comprises one or more retention pins extending through the platform and bosses of the airfoil segments. [Embodiment 17] 17. The turbomachine of any one of embodiments 14 to 16, wherein the one or more retaining pins include leading edge pins and trailing edge pins, the leading edge pins being oriented perpendicular to the axial direction. [Embodiment 18] 18. The turbomachine of any of embodiments 14-17, wherein the boss includes a shoulder and a shank portion, the shoulder extending from at least one of a base end and a tip end of the airfoil and surrounding one of the base end and the tip end, the shank portion extending radially from the shoulder, the shank portion including a shank leading edge, a shank trailing edge, a shank suction side, and a shank pressure side, the shank portion including a solid shank aft portion extending from the shank suction side to the shank pressure side near the shank trailing edge. [Embodiment 19] A turbomachine according to any one of embodiments 14 to 18, wherein the mechanical retention device includes a leading edge pin and a trailing edge pin, the leading edge pin penetrating the platform, the shank pressure side, and the shank suction side near the shank leading edge, and the trailing edge pin penetrating the platform and the solid shank aft portion. [Embodiment 20] A turbomachine according to any one of embodiments 14 to 19, wherein the platform defining an opening is an outer platform, and the boss extending from one of the base end or the tip end of the airfoil is a first boss extending from the tip end of the airfoil, and the stator vane assembly further includes an inner platform defining an inner platform opening, and a second boss extending from the base end of the airfoil and coupled to the inner platform by an inner retaining pin. [Explanation of symbols]

[0069] 2 stator vanes 6 wires 8 lines 11 Compressor casing 12 Compressor section 14 Compressor 18 Combustion section 19 Stator vane 20 Combustor 21 Rotor disk 22 Turbine section 23 Rotor Blades 24 Shaft 27 Rotor disc 28 Rotor Blades 30 Axial centerline 32 Outer casing 34 Exhaust diffuser 36 Ambient Air 38 Compressed Air 40 Combustion chamber 42 Combustion Gas 44 Airfoil 50 steps 53 Floor 55 Axial center line 56 Airfoil 57 Camber axis 58 Shroud Segment 60 Leading Edge 62 Trailing edge 64 Pressure side wall 66 Pressure side wall 68 Proximal end 70 Tip 80 Protrusion 81 Arrow 92 Cavity 102 Opening 107 Brazing material 200 Stator Vane Assembly 201 Platform 202 Inner Platform 204 Outer Platform 206 Airfoil Segment 208 Mechanical holding devices 210 Mechanical retention devices 211 Metallurgical Bond 212 Retaining pin 213 Leading edge pin 215 Center line 216 Boss 217 Inner retaining pin 218 Shoulder 220 Shank 222 Shank leading edge 224 Shank rear edge 226 Pressure side 228 Pressure side 230 Shank rear 232 Shoulder front edge 234 Posterior edge of shoulder 236 Pressure side 238 Compression side 240 Back of shoulder 242 Support surface 244 Undercut Surface 246 Protrusion 250 Compression side slash surface 252 Compression side slash surface 254 Leading edge surface 256 Trailing edge surface 258 Leading edge hook rail 260 Edge hook rail 262 First protrusion 264 Second protrusion 266 Third protrusion 268 Flow Direction 270 Height 280 Mounting hole

Claims

1. A stator vane assembly (200) comprising: A platform (201) defining an opening (102); an airfoil segment (206) having an airfoil (56) and a boss (216), the airfoil (56) including a leading edge (60), a trailing edge (62), a suction sidewall (66), and a pressure sidewall (64), the airfoil (56) extending radially between a base end (68) and a tip end (70), the boss (216) extending from one of the base end (68) and the tip end (70) of the airfoil (56), the boss (216) being disposed in an opening (102) in the platform (201); and A mechanical retention device (210) that connects the platform (201) to a boss (216) of the airfoil segment (206). A stator vane assembly (200) comprising:

2. 2. The stator vane assembly of claim 1, wherein a metallurgical bond is disposed between the platform and a boss of the airfoil segment, the metallurgical bond fixedly coupling the platform to the boss of the airfoil segment.

3. The stator vane assembly (200) of claim 2, wherein a height of the metallurgical bond (211) is greater than about 300% of a thickness of one of the pressure sidewall (64) and the suction sidewall (66).

4. The stator vane assembly (200) of claim 1, wherein the mechanical retention device (210) comprises one or more retention pins (212) that extend through the platform and bosses of the airfoil segments.

5. 5. The stator vane assembly (200) of claim 4, wherein the one or more retaining pins (212) include leading edge pins (213) and trailing edge pins (214), the leading edge pins (213) being oriented perpendicular to an axial direction of a turbomachine (10) in which the stator vane assembly (200) is installed.

6. 2. The stator vane assembly of claim 1, wherein the boss includes a shoulder and a shank, the shoulder extending from one of the base and tip ends of the airfoil and surrounding the one of the base and tip ends, and the shank extending radially from the shoulder.

7. 7. The stator vane assembly of claim 6, wherein the shank portion includes a shank leading edge, a shank trailing edge, a shank suction side, and a shank pressure side, the shank portion including a solid shank aft portion proximate the shank trailing edge and extending from the shank suction side to the shank pressure side.

8. 8. The stator vane assembly (200) of claim 7, wherein a leading edge pin (213) penetrates the platform (201), the shank pressure side (228), and the shank suction side (226) near the shank leading edge (222), and a trailing edge pin (214) penetrates the platform (201) and the solid shank aft portion (230).

9. 2. The stator vane assembly of claim 1, wherein the platform includes a floor and one or more protrusions extending radially from the floor, and the mechanical retention device couples one of the one or more protrusions of the platform to a boss of the airfoil segment.

10. 10. The stator vane assembly of claim 9, wherein the platform includes a pressure side slashface, a suction side slashface, a leading edge surface, and a trailing edge surface, each of the leading edge surface and the trailing edge surface extending between the pressure side slashface and the suction side slashface, and the platform includes leading and trailing edge hook rails extending from the floor.

11. 11. The stator vane assembly of claim 10, wherein the mechanical retention device is a leading edge pin that extends through a first one of the one or more protrusions and the boss, and into a second one of the one or more protrusions, the second one of the one or more protrusions being disposed on the pressure side slash face.

12. 11. The stator vane assembly (200) of claim 10, wherein the mechanical retention device (210) is a trailing edge pin (214) that extends through a first protrusion (266) of the one or more protrusions (80, 246) and the solid shank aft portion (230) and into the trailing edge hook rail (260).

13. 2. The stator vane assembly of claim 1, wherein the platform (201) defining an opening (102) is an outer platform (204), and a boss (216) extending from one of the base end (68) or the tip end (70) of the airfoil (56) is a first boss extending from the tip end (70) of the airfoil (56), the stator vane assembly (200) further includes an inner platform (202) defining an inner platform opening (102), a second boss (216) extending from the base end (68) of the airfoil (56) and coupled to the inner platform (202) by an inner retention pin (217).

14. A turbomachine (10) defining an axial direction (A), a radial direction (R), and a circumferential direction (C) relative to an axial centerline (30) of the turbomachine (10), comprising: A compressor section (12), a combustion section (18) provided downstream of the compressor section (12); and A turbine section (22) provided downstream of the combustion section (18).

14. A turbomachine (10) comprising: a compressor section (12) or a turbine section (22) comprising a stator vane assembly (200) according to any one of claims 1 to 13.

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