Nozzle assembly with alternating inserted vanes for a turbine engine

Abstract

The present invention relates to nozzle assemblies with alternatingly inserted vanes for turbine engines. In particular, a nozzle assembly for a gas turbine engine and a method for assembling a nozzle assembly are provided. In one exemplary aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced apart from the outer wall. The outer wall defines a plurality of mounting openings circumferentially spaced apart from one another. The inner wall defines a plurality of mounting openings circumferentially spaced apart from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The nozzle assembly includes vanes inserted through the mounting openings of the outer wall in a radially inward direction and vanes inserted through the mounting openings of the inner wall in a radially outward direction in an alternating fashion.

Description

[0001] The present application is a divisional application of the patent application “Nozzle assembly with alternatingly inserted vanes for a turbine engine” (Application No: 202110270493.0, Applicant: General Electric Company) filed on March 12, 2021. TECHNICAL FIELD

[0002] The subject matter of the present disclosure relates generally to turbine engines. More particularly, the subject matter of the present disclosure relates to nozzle assemblies of turbine engines. BACKGROUND

[0003] Gas turbine engines generally include a fan and a core arranged in flow communication with one another. Further, the core of a gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air within the combustion section and burned to provide combustion gases. The combustion gases pass from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then passed through the exhaust section, e.g., to the atmosphere.

[0004] More particularly, the combustion section includes a combustor having a combustion chamber defined by a combustor liner. Downstream of the combustion chamber, the turbine section includes one or more stages, e.g., each stage can include a plurality of stationary nozzle airfoils and a plurality of blade airfoils attached to a rotor that is driven by the flow of combustion gases relative to the blade airfoils. The turbine section can also have other configurations. In any case, a flowpath is defined by an inner boundary and an outer boundary, both of which extend from the combustion chamber through the stages of the turbine section.

[0005] Generally, the inner and outer boundaries defining the flowpath are formed from separate components. For example, an outer liner of a combustor, a separate outer band of a nozzle portion of a turbine stage, and a separate shroud of a blade portion of a turbine stage generally define at least a portion of the outer boundary of the flowpath. However, utilizing separate components to form each of the outer and inner boundaries requires a large number of parts, such as one or more seals at each joint between the separate components to minimize fluid leakage to or from the flowpath, which increases the complexity and weight of the gas turbine engine. Accordingly, it would be desirable for a flowpath assembly to have an integral outer boundary structure in which two or more components of the outer boundary are integrated as a single piece and / or an integral inner boundary structure in which two or more components of the inner boundary are integrated as a single piece. While forming an integral boundary structure has many benefits, forming a nozzle assembly along such an integral boundary structure has proven to be challenging. For example, nozzle vanes can be integrally formed with the integral structure to form the nozzle assembly. However, such an integral structure can be complex and expensive to manufacture. Nozzle vanes can also be inserted through the boundary structure. However, the space between such vanes is limited, making it difficult to secure and seal the vanes with the boundary structure.

[0006] Accordingly, it would be desirable to have improved turbine engines and nozzle assemblies that address one or more of the above challenges. Moreover, it would be beneficial to have methods for assembling nozzle assemblies of turbine engines that address one or more of the above challenges. SUMMARY

[0007] Aspects and advantages of the application will be set forth in part in the following description, or can be obvious from the description, or can be learned through practice of the application.

[0008] In one aspect, a nozzle assembly for a turbine engine is provided, the turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending along the axial direction. The nozzle assembly includes an outer wall defining mounting openings spaced apart from one another along the circumferential direction. The nozzle assembly also includes an inner wall defining mounting openings spaced apart from one another along the circumferential direction, the inner wall and the outer wall defining a flowpath. Further, the nozzle assembly includes a plurality of first vanes, each first vane having an airfoil and a mounting flange, wherein the airfoil of each first vane extends through one of the mounting openings of the outer wall and is at least partially positioned within the flowpath, and wherein the mounting flange of each first vane is positioned radially outward of the outer wall relative to the axial centerline. Additionally, the nozzle assembly includes a plurality of second vanes, each second vane having an airfoil and a mounting flange, wherein the airfoil of each second vane extends through one of the mounting openings of the inner wall and is at least partially positioned within the flowpath, and wherein the mounting flange of each second vane is positioned radially inward of the inner wall relative to the axial centerline. The plurality of first vanes and the plurality of second vanes are arranged in an alternating manner along the circumferential direction.

[0009] In another aspect, a turbine engine is provided that defines an axial direction, a radial direction, and a circumferential direction. The turbine engine includes a combustion section and a turbine section positioned downstream of the combustion section. Further, the turbine engine includes an inner wall that defines mounting openings spaced apart from one another along the circumferential direction. Further, the turbine engine includes an outer wall that includes a combustor portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section, the combustor portion and the turbine portion being integrally formed as a single unitary structure, wherein the turbine portion of the outer wall defines the mounting openings spaced apart from one another along the circumferential direction, wherein each mounting opening of the inner wall is positioned between adjacent mounting openings of the outer wall along the circumferential direction. Further, the turbine engine includes a plurality of vanes extending between the outer wall and the inner wall along the radial direction, wherein the plurality of vanes are inserted inwardly through the mounting openings of the outer wall along the radial direction and outwardly through the mounting openings of the inner wall along the radial direction in an alternating fashion.

[0010] In another aspect, a turbine engine is provided that defines an axial direction, a radial direction, a circumferential direction, and an axial centerline extending along the axial direction. The turbine engine includes a combustion section and a turbine section positioned downstream of the combustion section. The turbine engine further includes an inner wall defining mounting openings. The turbine engine further has an outer wall that includes a combustor portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section, the combustor portion and the turbine portion being integrally formed as a single unitary structure, wherein the turbine portion of the outer wall defines the mounting openings. The turbine engine further has a first vane having an airfoil and a mounting flange, wherein the airfoil of the first vane extends through the mounting openings of the outer wall, and wherein the mounting flange of the first vane is positioned radially outward of the outer wall relative to the axial centerline. The turbine engine further includes a second vane positioned adjacent to the first vane along the circumferential direction and having an airfoil and a mounting flange, wherein the airfoil of the second vane extends through the mounting openings of the inner wall, and wherein the mounting flange of the second vane is positioned radially inward of the inner wall relative to the axial centerline.

[0011] In yet another aspect, a method for assembling a nozzle assembly for a turbine engine is provided, the turbine engine defining an axial direction, a radial direction, and a circumferential direction. The method includes inserting a plurality of vanes inwardly through an outer wall of the nozzle assembly along the radial direction, and inserting the plurality of vanes outwardly through an inner wall of the nozzle assembly along the radial direction, the inner wall being spaced apart from the outer wall along the radial direction, and wherein the plurality of vanes inserted inwardly through the outer wall along the radial direction alternate with the plurality of vanes inserted outwardly through the inner wall along the radial direction along the circumferential direction.

[0012] In further aspects, a method of assembling a nozzle assembly for a turbine engine is provided, the turbine engine defining an axial direction, a radial direction, and a circumferential direction. The method includes inserting a first vane inwardly through an outer wall of the nozzle assembly in the radial direction. The method also includes, adjacent the first vane in the circumferential direction, inserting a second vane outwardly through an inner wall of the nozzle assembly in the radial direction, the inner wall being spaced apart from the outer wall in the radial direction.

[0013] In some embodiments, the method further includes, adjacent the first vane in the circumferential direction and opposite the second vane, inserting a third vane inwardly through the outer wall of the nozzle assembly in the radial direction. In this manner, the first vane is positioned between the second vane and the third vane in the circumferential direction. Both the first vane and the third vane are inserted inwardly through the outer wall of the nozzle assembly in the radial direction, while the second vane is inserted outwardly through the inner wall of the nozzle assembly in the radial direction.

[0014] In yet other embodiments, the method further includes, adjacent the second vane in the circumferential direction and opposite the first vane, inserting a third vane outwardly through the inner wall of the nozzle assembly in the radial direction. In this manner, the second vane is positioned between the first vane and the third vane in the circumferential direction. Both the second vane and the third vane are inserted outwardly through the inner wall of the nozzle assembly in the radial direction, while the first vane is inserted inwardly through the outer wall of the nozzle assembly in the radial direction.

[0015] In yet another aspect, a nozzle assembly for a turbine engine is provided, the turbine engine defining an axial direction, a radial direction, and a circumferential direction. The flowpath assembly includes an outer wall defining mounting openings spaced apart from one another in the circumferential direction. The nozzle assembly also includes an inner wall defining mounting openings spaced apart from one another in the circumferential direction, wherein each mounting opening of the inner wall is positioned between adjacent mounting openings of the outer wall in the circumferential direction, the inner wall and the outer wall defining a flowpath. Further, the nozzle assembly includes a plurality of vanes inserted inwardly through the mounting openings of the outer wall in the radial direction and into the flowpath in an alternating manner and outwardly through the mounting openings of the inner wall in the radial direction and into the flowpath.

[0016] These and other features, aspects, and advantages of the present application will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the application and, together with the description, serve to explain the principles of the application.

[0017] Technical Solution 1. A nozzle assembly for a turbine engine, the turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending in the axial direction, the nozzle assembly comprising:

[0018] an outer wall defining mounting openings;

[0019] an inner wall defining the mounting opening, the inner wall and the outer wall defining a flowpath;

[0020] a first vane having an airfoil and a mounting flange, wherein the airfoil of the first vane extends through the mounting opening of the outer wall and is at least partially positioned within the flowpath, and wherein the mounting flange of the first vane is positioned exterior to the outer wall in the radial direction relative to the axial centerline;

[0021] a second vane positioned adjacent to the first vane in the circumferential direction and having an airfoil and a mounting flange, wherein the airfoil of the second vane extends through the mounting opening of the inner wall and is at least partially positioned within the flowpath, and wherein the mounting flange of the second vane is positioned interior to the inner wall in the radial direction relative to the axial centerline.

[0022] Technical Solution 2. The nozzle assembly of any preceding technical solution, wherein the mounting opening of the outer wall is one of a plurality of mounting openings and the mounting opening of the inner wall is one of a plurality of mounting openings, and wherein the first vane is one of a plurality of first vanes and the second vane is one of a plurality of second vanes, and wherein the airfoil of each first vane of the plurality of first vanes extends through one of the plurality of mounting openings of the outer wall and is at least partially positioned within the flowpath and wherein the mounting flange of each first vane of the plurality of first vanes is positioned exterior to the outer wall in the radial direction relative to the axial centerline, and wherein the airfoil of each second vane of the plurality of second vanes extends through one of the plurality of mounting openings of the inner wall and is at least partially positioned within the flowpath and wherein the mounting flange of each second vane of the plurality of second vanes is positioned interior to the inner wall in the radial direction relative to the axial centerline.

[0023] Technical Solution 3. The nozzle assembly of any preceding technical solution, wherein at least two pairs of adjacent vanes of the nozzle assembly include one first vane of the plurality of first vanes and one second vane of the plurality of second vanes.

[0024] Technical Solution 4. The nozzle assembly of any preceding technical solution, wherein the plurality of first vanes and the plurality of second vanes are arranged in an alternating manner in the circumferential direction.

[0025] TECHNICAL SOLUTION 5. The nozzle assembly of any preceding technical solution, wherein the inner wall defines positioning grooves spaced apart from one another along the circumferential direction and positioned opposite the mounting openings of the outer wall along the radial direction, and the outer wall defines positioning grooves spaced apart from one another along the circumferential direction and positioned opposite the mounting openings of the inner wall along the radial direction, and wherein each vane of the first and second plurality of vanes has a positioning end opposite the mounting flange of the vane, and wherein the positioning end of each vane of the first plurality of vanes is received within one of the positioning grooves defined by the inner wall, and wherein the positioning end of each vane of the second plurality of vanes is received within one of the positioning grooves defined by the outer wall.

[0026] TECHNICAL SOLUTION 6. The nozzle assembly of any preceding technical solution, wherein the mounting openings of the inner wall are positioned intermediate the mounting openings of the outer wall along the circumferential direction.

[0027] TECHNICAL SOLUTION 7. The nozzle assembly of any preceding technical solution, wherein the mounting openings of the outer wall are sized to receive and shaped to complement the radial cross-section of the airfoil of the first vanes, and the mounting openings of the inner wall are sized to receive and shaped to complement the radial cross-section of the second vanes.

[0028] TECHNICAL SOLUTION 8. The nozzle assembly of any preceding technical solution, wherein the outer wall has an inner surface and an outer surface spaced apart from the inner surface along the radial direction, and the mounting flange of the first vanes engages the outer surface of the outer wall, and wherein the inner wall has an inner surface and an outer surface spaced apart from the inner surface along the radial direction, and the mounting flange of the second vanes engages the inner surface of the inner wall.

[0029] TECHNICAL SOLUTION 9. The nozzle assembly of any preceding technical solution, wherein the first vanes have a sealing portion connecting the airfoil and the mounting flange, and the second vanes have a sealing portion connecting the airfoil and the mounting flange, and wherein the sealing portion of the first vanes wedges the first vanes into engagement with the outer wall, and the sealing portion of the second vanes wedges the second vanes into engagement with the inner wall.

[0030] TECHNICAL SOLUTION 10. The nozzle assembly of any preceding technical solution, wherein the nozzle assembly defines an outer diameter and an inner diameter, and wherein cooling fluid is delivered to the first vanes from the outer diameter of the nozzle assembly, and wherein cooling fluid is delivered to the second vanes from the inner diameter of the nozzle assembly.

[0031] Technical Solution 11. A turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending along the axial direction, the turbine engine comprising:

[0032] Combustion zone;

[0033] The turbine section is located downstream of the combustion section;

[0034] The inner wall that defines the installation opening;

[0035] The outer wall includes a burner portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section, the burner portion and the turbine portion being integrally formed as a single integral structure, wherein the turbine portion of the outer wall defines a mounting opening;

[0036] A first guide vane having an airfoil and a mounting flange, wherein the airfoil of the first guide vane extends through the mounting opening in the outer wall, and wherein the mounting flange of the first guide vane is positioned outside the outer wall in the radial direction relative to the axial centerline; and

[0037] A second guide vane is positioned adjacent to the first guide vane along the circumferential direction and has an airfoil and a mounting flange, wherein the airfoil of the second guide vane extends through the mounting opening of the inner wall, and wherein the mounting flange of the second guide vane is positioned inside the inner wall in the radial direction relative to the axial centerline.

[0038] Technical Solution 12. The turbine engine according to any of the foregoing technical solutions, characterized in that the outer wall, the inner wall, the first guide vane and the second guide vane are formed of ceramic matrix composite (CMC) material.

[0039] Technical Solution 13. The turbine engine according to any of the foregoing technical solutions, characterized in that the inner wall defines a positioning groove positioned relative to the mounting opening of the outer wall in the radial direction, and the outer wall defines a positioning groove positioned relative to the mounting opening of the inner wall in the radial direction.

[0040] Technical Solution 14. The turbine engine according to any of the foregoing technical solutions, characterized in that the first guide vane and the second guide vane each extend along the radial direction between the mounting end and the positioning end, and wherein the positioning end of the first guide vane is positioned in the positioning groove of the inner wall, and the positioning end of the second guide vane is positioned in the positioning groove of the outer wall.

[0041] Technical Solution 15. The turbine engine of any preceding technical solution, wherein the mounting opening of the inner wall and the mounting opening of the outer wall are aligned along the axial direction.

[0042] Technical Solution 16. A method for assembling a nozzle assembly for a turbine engine, the turbine engine defining an axial direction, a radial direction, and a circumferential direction, the method comprising:

[0043] inserting a first vane inwardly through an outer wall of the nozzle assembly along the radial direction; and

[0044] inserting a second vane outwardly through an inner wall of the nozzle assembly along the radial direction adjacent the first vane along the circumferential direction, the inner wall being spaced apart from the outer wall along the radial direction.

[0045] Technical Solution 17. The method of any preceding technical solution, further comprising:

[0046] inserting a third vane inwardly through an outer wall of the nozzle assembly along the radial direction adjacent the first vane along the circumferential direction and opposite the second vane.

[0047] Technical Solution 18. The method of any preceding technical solution, further comprising:

[0048] inserting a third vane outwardly through an inner wall of the nozzle assembly along the radial direction adjacent the second vane along the circumferential direction and opposite the first vane.

[0049] Technical Solution 19. The method of any preceding technical solution, wherein the first vane is one of a plurality of first vanes and the second vane is one of a plurality of second vanes, and wherein the method further comprises:

[0050] inserting the plurality of first vanes inwardly through an outer wall of the nozzle assembly along the radial direction; and

[0051] inserting the plurality of second vanes outwardly through an inner wall of the nozzle assembly along the radial direction,

[0052] wherein the plurality of first vanes inserted inwardly through the outer wall along the radial direction alternate with the plurality of second vanes inserted outwardly through the inner wall along the radial direction along the circumferential direction.Technical Solution 20. The method of any preceding technical solution, wherein the plurality of vanes inserted inwardly through the outer wall in the radial direction extend between the outer wall and the inner wall in the radial direction, and wherein the plurality of vanes inserted outwardly through the inner wall in the radial direction extend between the outer wall and the inner wall in the radial direction.

[0053] Technical Solution 21. The method of any preceding technical solution, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted inwardly through the outer wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the inner wall, and wherein each of the plurality of vanes inserted outwardly through the inner wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted outwardly through the inner wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the outer wall.

[0054] Technical Solution 22. The method of any preceding technical solution, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction has a sealing portion that is wedged into a mounting opening defined by the outer wall, and wherein each of the mounting openings has an insertion end and a flowpath end spaced apart from the insertion end in the radial direction, the insertion end being positioned outside of the flowpath end in the radial direction, and wherein the insertion end of each of the mounting openings has a greater radial cross-sectional area than the flowpath end.

[0055] Implementation 1. A nozzle assembly for a turbine engine, comprising:

[0056] a first wall defining a mounting opening;

[0057] a second wall spaced apart from the first wall, the first wall and the second wall defining a flowpath; and

[0058] a vane positioned at least partially within the flowpath and at least partially within the mounting opening, the vane having a mounting flange that is larger than the mounting opening to secure the vane to the first wall, the vane further having a retention feature that secures the vane to the second wall.

[0059] Implementation 2. The nozzle assembly of Implementation 1, wherein the first wall is spaced apart from the second wall along a radial direction defined by a turbine engine, and wherein the first wall is positioned inside of the second wall in the radial direction relative to an axial centerline of the turbine engine.

[0060] Embodiment 3. The nozzle assembly of embodiment 1, wherein the first wall is spaced apart from the second wall along a radial direction defined by the turbine engine, and wherein the first wall is positioned outside of the second wall along the radial direction relative to an axial centerline of the turbine engine.

[0061] Embodiment 4. The nozzle assembly of embodiment 1, wherein the second wall defines a through opening through which the vane extends, the vane being secured by the retention feature outside of the flowpath.

[0062] Embodiment 5. The nozzle assembly of embodiment 1, wherein the second wall has a first surface facing the flowpath and a second surface facing away from the flowpath, and wherein a radial space is defined between the retention feature and the second surface of the second wall.

[0063] Embodiment 6. The nozzle assembly of embodiment 1, wherein the retention feature is one of a ring clip, a retaining pin, and a nut.

[0064] Embodiment 7. The nozzle assembly of embodiment 1, wherein the retention feature is one or more plies.

[0065] Embodiment 8. The nozzle assembly of embodiment 7, wherein the one or more plies are laminated and secured to the vane outside of the flowpath.

[0066] Embodiment 9. The nozzle assembly of embodiment 1, wherein the retention feature is a bonding material.

[0067] Embodiment 10. The nozzle assembly of embodiment 9, wherein the bonding material is a braze-type bond.

[0068] Embodiment 11. The nozzle assembly of embodiment 1, wherein the vane defines a locating end recess, and wherein the retention feature is a locating barb ring received within the locating end recess and compressed between the vane and the second wall.

[0069] Embodiment 12. The nozzle assembly of embodiment 1, wherein the vane defines a mounting end recess, and wherein a mounting barb ring is received within the mounting end recess and compressed between the vane and the first wall.

[0070] Embodiment 13. The nozzle assembly of embodiment 1, wherein the second wall defines a second mounting opening, and wherein the nozzle assembly further comprises:

[0071] a second vane positioned at least partially within the flowpath and at least partially within the second mounting opening, the second vane having a second mounting flange that is larger than the second mounting opening to secure the second vane to the second wall, the second vane further having a second retention feature that secures the vane to the first wall.

[0072] Embodiment 14. The nozzle assembly of embodiment 13, wherein the first wall defines a second through opening through which the second vane extends, the second vane being secured by the second retention feature outside of the flowpath.

[0073] Embodiment 15. A nozzle assembly for a turbine engine, comprising:

[0074] a first wall defining a mounting opening;

[0075] a second wall spaced apart from the first wall, the first wall and the second wall defining a flowpath; and

[0076] a vane positioned at least partially within the flowpath and extending between a first end and a second end, the first end of the vane having a mounting flange positioned within the mounting opening and wedgingly engaging the first wall, the second end of the vane being secured to the second wall.

[0077] Embodiment 16. The nozzle assembly of embodiment 15, wherein the mounting flange of the vane is positioned entirely within the mounting opening of the first wall.

[0078] Embodiment 17. The nozzle assembly of embodiment 15, wherein the first wall has a first surface and a second surface, and wherein the mounting flange extends between the first surface and the second surface.

[0079] Embodiment 18. The nozzle assembly of embodiment 15, wherein the first wall has a radial thickness and the mounting flange has a radial thickness, and wherein the radial thickness of the mounting flange is less than the radial thickness of the first wall.

[0080] Embodiment 19. The nozzle assembly of embodiment 15, wherein the first wall has a mounting seat, and wherein the mounting flange is seated on the mounting seat.

[0081] Embodiment 20. A method for assembling a nozzle assembly for a turbine engine, the turbine engine defining a radial direction, the method comprising:

[0082] A first guide vane is inserted inward along the radial direction such that the locating end of the first guide vane is inserted through a locating opening defined by the inner wall of the nozzle assembly and such that the mounting flange of the first guide vane engages the outer wall of the nozzle assembly, the inner wall being spaced apart from the outer wall along the radial direction; and

[0083] The second guide vane is inserted outward along the radial direction such that the positioning end of the second guide vane is inserted through the positioning opening defined by the outer wall and such that the mounting flange of the second guide vane engages the inner wall of the nozzle assembly. Attached Figure Description

[0084] The complete and full disclosure of the invention, including its best mode, for those skilled in the art, is set forth in the description with reference to the accompanying drawings, in which:

[0085] Figure 1 Schematic cross-sectional views of exemplary gas turbine engines according to various embodiments of this subject are provided;

[0086] Figure 2 Provided Figure 1 A schematic cross-sectional view of the combustion section and high-pressure turbine section of a gas turbine engine;

[0087] Figure 3 Provided Figure 2 A schematic axial cross-sectional view of a portion of the first-stage nozzle assembly of the flow path component;

[0088] Figure 4 Provided passage by Figure 3 A partial magnified view of the guide vane inserted through the mounting opening defined by the outer wall of the first-stage nozzle assembly;

[0089] Figure 5 Provided passage by Figure 3 A partial magnified view of the guide vane inserted through the mounting opening defined by the inner wall of the first-stage nozzle assembly;

[0090] Figure 6 Provided Figure 2 A perspective cross-sectional view of a portion of the integral outer boundary structure and inner boundary structure of the flow path component;

[0091] Figure 7 A schematic axial cross-sectional view of a portion of the first-stage nozzle assembly of the flow path assembly is provided, illustrating various cooling features of the assembly according to exemplary embodiments of the subject matter;

[0092] Figure 8 A schematic axial cross-sectional view of a portion of a first-stage nozzle assembly for a flow path assembly of a turbine engine, according to an exemplary embodiment of this subject matter, is provided.

[0093] Figure 9 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine is provided in accordance with yet another exemplary embodiment of the present subject matter;

[0094] Figure 10 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine is provided in accordance with yet another exemplary embodiment of the present subject matter;

[0095] Figure 11 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine is provided in accordance with yet another exemplary embodiment of the present subject matter;

[0096] Figure 12 A flowchart of an exemplary method in accordance with an exemplary embodiment of the present subject matter is provided;

[0097] Figure 13 A schematic axial cross-sectional view of a portion of a nozzle assembly in accordance with an exemplary embodiment of the present subject matter is provided;

[0098] Figure 14 A schematic axial cross-sectional view of a portion of a nozzle assembly for a turbine engine in accordance with an exemplary embodiment of the present subject matter is provided; and

[0099] Figure 15 Another schematic axial cross-sectional view of a portion of a nozzle assembly for a turbine engine in accordance with an exemplary embodiment of the present subject matter is provided. DETAILED DESCRIPTION

[0100] Reference will now be made in detail to embodiments of the application, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to indicate like or similar portions of the application. As used herein, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms "upstream" and "downstream" refer to the relative direction with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which the fluid flow is coming, and "downstream" refers to the direction to which the fluid flow is going. Further, as used herein, the terms "axial" or "axially" refer to a dimension along an axial centerline of the engine. The term "forward" used in connection with "axial" or "axially" refers to a direction toward an engine inlet, or a component is relatively closer to the engine inlet than another component. The term "aft" used in connection with "axial" or "axially" refers to a direction toward an engine nozzle, or a component is relatively closer to the engine nozzle than another component. The terms "radial" or "radially" refer to a dimension extending between an axial centerline of the engine and an outer engine circumference or outer annulus. Radially inward is toward the axial centerline, and radially outward is away from the axial centerline.

[0101] Exemplary aspects of the present disclosure relate to a nozzle assembly for a gas turbine engine and a method for assembling the nozzle assembly. In one exemplary aspect, the nozzle assembly includes an outer wall and an inner wall radially spaced apart from the outer wall. The outer wall defines a plurality of mounting openings circumferentially spaced apart from one another. Likewise, the inner wall defines a plurality of mounting openings circumferentially spaced apart from one another. The mounting openings defined by the inner wall are positioned circumferentially between adjacent mounting openings defined by the outer wall. The mounting openings defined by the outer wall and the inner wall are axially aligned. The nozzle assembly includes vanes inserted through the mounting openings of the outer wall in a radially inward direction and vanes inserted through the mounting openings of the inner wall in a radially outward direction in an alternating fashion. For example, in some exemplary aspects, every other vane around an annulus of the nozzle assembly is inserted radially inward through the outer wall, and every other vane around the annulus of the nozzle assembly is inserted radially outward through the inner wall. The inserted vanes are positioned such that they extend across a flowpath defined between the outer wall and the inner wall. Further, the vanes inserted radially inward through the mounting openings of the outer wall can be positioned and constrained within a positioning groove positioned radially opposite the mounting opening through which the vane is inserted. Similarly, the vanes inserted radially outward through the mounting openings of the inner wall can be positioned and constrained within a positioning groove positioned radially opposite the mounting opening through which the vane is inserted.

[0102] Figure 1 A schematic cross-sectional view of a gas turbine engine in accordance with exemplary embodiments of the present disclosure is provided. More specifically, for Figure 1In an example embodiment, the gas turbine engine is a high-bypass turbofan jet engine, referred to herein as "turbofan engine 10." As shown in Figure 1 turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal or axial centerline 12 provided for reference) and a radial direction R. A circumferential direction C extends three-hundred and sixty degrees (360°) about the axial centerline 12.

[0103] The turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14. The exemplary core turbine engine 16 shown includes, generally, a substantially tubular outer casing 18 that defines an annular core inlet 20. The casing 18 encloses, in serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and a jet exhaust nozzle section 32. A high- pressure (HP) shaft or spool 34 drivably connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft or spool 36 drivably connects the LP turbine 30 to the LP compressor 22. In other embodiments of the turbofan engine 10, additional spools can be provided.

[0104] The fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in spaced relation. As shown, the fan blades 40 extend generally outwardly in the radial direction R from the disk 42. The fan blades 40 and disk 42 can rotate together about the longitudinal axis 12 by the LP shaft 36. In some embodiments, a power gearbox having a plurality of gears can be included to step down the rotational speed of the LP shaft 36 to a more efficient fan rotational speed.

[0105] Referring still to the exemplary embodiment of Figure 1 the disk 42 is covered by a rotatable hub cap 48 that is aerodynamically contoured to facilitate the flow of gases through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially encircles at least a portion of the fan 38 and / or the core turbine engine 16. The nacelle 50 is supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Further, as shown in Figure 1 the downstream section 54 of the nacelle 50 can extend over the exterior of the core turbine engine 16 so as to define a bypass air flow passage 56 therebetween.

[0106] During operation of the turbofan engine 10, a volume of air 58 enters the turbofan 10 via the associated inlet 60 of the nacelle 50 and / or fan section 14. As this volume of air 58 passes across the fan blades 40, a first portion of the air 58 is directed or channeled into the bypass airflow passage 56 as indicated by arrow 62, and a second portion of the air 58 is directed or channeled into the LP compressor 22 as indicated by arrow 64. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the bypass ratio. The pressure of the second portion of air 64 is then increased as it passes through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to provide combustion gases 66.

[0107] The combustion gases 66 pass through the HP turbine 28, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 and HP turbine rotor blades 70 coupled to the HP shaft or spool 34, thus causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 then pass through the LP turbine 30, where a second portion of the thermal and kinetic energy from the combustion gases 66 is extracted via sequential stages of LP turbine stator vanes 72 and LP turbine rotor blades 74 coupled to the LP shaft or spool 36, thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0108] The combustion gases 66 subsequently pass through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 passes through the bypass airflow passage 56 before it is exhausted from the fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, LP turbine 30, and jet exhaust nozzle section 32 at least partially define a hot gas path 78 to pass the combustion gases 66 through the core turbine engine 16.

[0109] It will be appreciated that, although described with reference to a turbofan 10 having a core turbine engine 16, the present subject matter can be applicable to other types of turbomachinery. For example, the present subject matter can be suitable for use in connection with or in turboprops, turboshafts, turbojets, industrial and marine gas turbine engines, and / or auxiliary power units.

[0110] In some embodiments, components of the turbofan engine 10, particularly components within the hot gas path 78, such as components of the combustion section 26, the HP turbine 28, and / or the LP turbine 30, can be formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Other components of the turbofan engine 10, such as components of the HP compressor 24, can also be formed of a CMC material. Exemplary CMC materials for such components can include silicon carbide (SiC), silicon, silica, or alumina matrix materials, and combinations thereof. Ceramic fibers can be embedded within the matrix, such as oxidation-stable reinforcing fibers including monofilaments such as sapphire or silicon carbide (e.g., Textron's SCS-6), and short fibers including silicon carbide (e.g., Nippon Carbon's NICALON® and KYOVAC®), alumina (e.g., Nextel's Ube Industries' UBE® and Dow Corning's ), aluminum silicate (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and ), and optional ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) grit and yarn.

[0111] In certain embodiments, fiber tows that can include a coating of ceramic refractory material are formed into reinforcing tapes, such as unidirectional reinforcing tapes. Multiple tapes can be laminated together (e.g., as a ply) to form a preform. The fiber tows can be impregnated with a slurry composition prior to forming the preform or after forming the preform. The preform can then undergo a heat treatment, such as curing or burn-out, to produce a high char residue in the preform, and subsequently a chemical treatment, such as silicon melt infiltration or chemical vapor infiltration, to achieve a component formed of a CMC material having a desired chemical composition. In other embodiments, the CMC material can be formed into, for example, carbon fiber cloth rather than tapes.

[0112] As noted above, components formed of CMC materials can be used in the hot gas path 78, such as within the combustion section and / or turbine section of the engine 10. For example, the combustion section 26 can include a combustor formed of a CMC material, and / or one or more stages of the HP turbine 28 can be formed of a CMC material. However, as noted above, CMC components can also be used in other sections of the turbofan engine 10, such as the compressor and / or fan sections. In some embodiments, additionally or alternatively, other high temperature materials and / or other composite materials can be used to form one or more components of the engine 10.

[0113] Figure 2 Provided are Figure 1A schematic cross-sectional view of the combustion section 26 and the HP turbine 28 of the turbine section of the turbofan engine 10. As shown, the combustion section 26 includes a generally annular combustor 80, and downstream of the combustion section 26, the HP turbine 28 includes multiple turbine stages. More specifically, for the illustrated embodiment, the HP turbine 28 includes a first turbine stage 82 and a second turbine stage 84. In other embodiments, the HP turbine 28 may include a different number of turbine stages; for example, the HP turbine 28 may include one turbine stage or more than two turbine stages. The first turbine stage 82 is located immediately downstream of the combustion section 26, while the second turbine stage 84 is located immediately downstream of the first turbine stage 82. Furthermore, each turbine stage 82, 84 includes a nozzle assembly and a blade assembly. The first turbine stage 82 includes a nozzle assembly 82N and a blade assembly 82B, and the second turbine stage 84 includes a nozzle assembly 84N and a blade assembly 84B. The nozzle assembly 82N of the first turbine stage 82 is positioned immediately downstream of the combustion section 26, such that the nozzle assembly 82N of the first turbine stage 82 can also be referred to as a burner exhaust nozzle. Furthermore, the burner 80 defines a generally annular combustion chamber 86, such that the burner 80 can be described as a generally annular burner. That is, the burner 80 extends annularly around the axial centerline 12 in the circumferential direction C.

[0114] Additionally, as described in detail below, the flow path 100 passing through the combustion section 26 and the HP turbine 28 is defined by the outer and inner boundaries of the flow path assembly 101. The outer and inner boundaries form the flow path 100 for the combustion gas 66 to flow through the combustion section 26 and the HP turbine 28; therefore, the flow path 100 may include the aforementioned hot gas path 78. Figure 1 At least a portion of the flow path 100. Furthermore, in other embodiments, the flow path 100 may extend through the LP turbine 30 and the jet exhaust port 32. Figure 1 In other embodiments, the flow path 100 may extend upstream of the combustion section 26, for example, to the HP compressor 24. Figure 1 Therefore, it will be appreciated that the inventive aspects disclosed herein can also be applied to different sections, configurations and / or components of the gas turbine engine and flow path 100, and are not limited to components within combustion section 26 and HP turbine 28.

[0115] for Figure 2In the illustrated embodiment, the outer and inner boundaries of the flow path 100, which defines at least a portion of the combustion section 26 and the HP turbine 28, are defined by an outer wall 102 and an inner wall 120, respectively. Generally, the outer wall 102 is spaced apart from the inner wall 120 in a radial direction R. As shown, the outer wall 102 and the inner wall 120 define portions of the outer and inner boundaries of the flow path 100. For example, the outer wall 102 defines an outer bushing portion 108 that forms the outer boundary of the flow path 100 through the burner 80. The outer wall 102 also defines an outer band portion 110 that forms the outer boundary of the flow path 100 through the nozzle assembly 82N of the first turbine stage 82 and an outer band portion 114 that forms the outer boundary of the flow path 100 through the nozzle assembly 84N of the second turbine stage 84. Furthermore, the outer wall 102 defines a shroud portion 112 forming the outer boundary of the flow path 100 passing through the blade assembly 82B of the first turbine stage 82 and a shroud portion 116 forming the outer boundary of the flow path 100 passing through the blade assembly 84B of the second turbine stage 84. The shroud portion 112 of the outer wall 102 is positioned between the outer belt portion 110 and the outer belt portion 114 in the axial direction A.

[0116] In addition, such as Figure 2 As shown, the inner wall 120 defines an inner bushing portion 122 that forms the inner boundary of the flow path 100 through the burner 80. The inner wall portion 120 also defines an inner band portion 124 that forms the inner boundary of the flow path 100 through the nozzle assembly 82N of the first turbine stage 82. Although in Figure 2 In the illustrated embodiment, the blade platform 132 is not shown integrally with the inner wall 120, but the blade platform 132 forms the inner boundary of the flow path 100 through the blade assembly 82B of the first-stage turbine 82, the inner band 136 forms the inner boundary of the flow path 100 through the nozzle assembly 84N of the second-stage turbine 84, and the blade platform 132 forms the inner boundary of the flow path 100 through the blade assembly 84B of the second-stage turbine 84.

[0117] Furthermore, the burner dome 118 extends radially across the front end 88 of the burner 80. The burner dome 118 may be a portion of the outer wall 102, the inner wall 120, or both (e.g., a portion of the burner dome 118 may be defined by the outer wall 102, while the remainder may be defined by the inner wall 120), or it may be a component separate from the outer wall 102 and the inner wall 120. The burner dome 118 defines an opening 142 for receiving a fuel nozzle assembly 90 positioned at the front end 88. The fuel nozzle assembly 90, for example, supplies a mixture of fuel and compressed air from the compressor section to the combustion chamber 86, which is burned within the combustion chamber 86 to produce a flow of combustion gas 66 through the flow path 100. The fuel nozzle assembly 90 may be attached to the burner dome 118, or it may “float” relative to the burner dome 118 and the flow path 100, i.e., the fuel nozzle assembly 90 may not need to be attached to the burner dome 118. In the illustrated embodiment, the fuel nozzle assembly 90 includes a swirler 92, and in some embodiments, the swirler 92 may be attached to the burner dome 118; alternatively, the swirler 92 may float relative to the burner dome 118 and the flow path 100. It will be appreciated that the fuel nozzle assembly 90 or the swirler 92 may float relative to the burner dome 118 and the flow path 100 along both the radial direction R and the axial direction A, or only one or the other of the radial direction R and the axial direction A. Furthermore, it will be understood that the burner dome 118 may define a plurality of openings 142 spaced apart from each other in the circumferential direction C, each opening accommodating the swirler 92 or other portions of the fuel nozzle assembly 90.

[0118] like Figure 2 As further shown, nozzle assemblies 82N and 84N include multiple guide vanes. Specifically, nozzle assembly 82N includes multiple airfoils 126. Figure 2 Only one is shown in the image), and the nozzle assembly 84N includes multiple airfoils 128. Figure 2 (Only one is shown in the image). Each nozzle airfoil 126 within the nozzle assembly 82N of the first turbine stage 82 extends radially R from the outer band portion 110 of the outer wall 102 to the inner band portion 124 of the inner wall 120. The nozzle airfoils 126 are arranged in a ring array around the axial centerline 12 and spaced apart from each other in the circumferential direction C. Similarly, each nozzle airfoil 128 within the nozzle assembly 84N of the second turbine stage 84 extends radially R from the outer band portion 114 of the outer wall 102 to the inner band 136. The nozzle airfoils 128 are arranged in a ring array around the axial centerline 12 and spaced apart from each other in the circumferential direction C.

[0119] A plurality of blade airfoils 130 are positioned in each of the blade assemblies 82B, 84B. Each blade airfoil 130 within the blade assembly 82B of the first turbine stage 82 is attached to a blade platform 132, which in turn is attached to the first stage rotor 134. The blade airfoils 130 attached to the first stage rotor 134 via the blade platforms 132 are circumferentially spaced about the axial centerline 12. Similarly, each blade airfoil 130 within the blade assembly 84B of the second turbine stage 84 is attached to a blade platform 132, which in turn is attached to the second stage rotor 138. The blade airfoils 130 attached to the second stage rotor 138 via the blade platforms 132 are circumferentially spaced about the axial centerline 12. Each blade airfoil 130 extends radially outward from their respective blade platforms 132 toward the outer wall 102 (i.e., the outer boundary of the flowpath 100). A gap is defined between the radially outer tip of each blade airfoil 130 and the outer wall 102, such that each turbine rotor 134, 138 is free to rotate within its respective turbine stage. Although not shown, each turbine rotor 134, 138 of the HP turbine 28 is connected to the HP shaft 34 Figure 1 In this manner, the blade airfoils 130 can extract kinetic energy from the flow of combustion gases 66 passing through the flowpath 100 defined by the HP turbine 28 as rotational energy applied to the HP shaft 34.

[0120] Accordingly, the flowpath 100 through the combustion section 26 and the HP turbine 28 is defined by a flowpath assembly 101 having an inner boundary and an outer boundary, and the inner and outer boundaries define the flowpath 100 for the combustion gases 66 through the combustion section 26 and the HP turbine 28. Portions of the outer boundary of the flowpath assembly 101 can be unitized or integrated as a single piece outer wall 102 that defines the radially outer boundary of the gas flowpath 100. For example, the outer wall 102 can include a combustor portion that extends through a combustion section, such as the combustion section 26, and a turbine portion that extends through at least a first turbine stage of a turbine section, such as the first turbine stage 82 of the HP turbine 28. The combustor portion and the turbine portion can be integrally formed such that the combustor portion and the turbine portion are a single unitary structure, i.e., the unitary outer wall 102.

[0121] For example, the outer wall 102 can include a combustor portion that extends through a combustion section, such as the combustion section 26, and a turbine portion that extends through at least a first turbine stage of a turbine section, such as the first turbine stage 82 of the HP turbine 28. The combustor portion and the turbine portion can be integrally formed such that the combustor portion and the turbine portion are a single unitary structure, i.e., the unitary outer wall 102. Figure 2In the illustrated embodiments, for example, a portion of the outer wall 102 is integrally formed or integrated as a single piece. As shown, the outer wall 102 includes a burner portion 104 extending through the combustion section 26 and a turbine portion 106 extending through at least a first turbine stage 82 and a second turbine stage 84 of the turbine section. In other embodiments, the turbine portion 106 may extend through fewer stages (e.g., through one turbine stage as described above) or through more stages (e.g., through one or more stages of an LP turbine 30 positioned downstream of an HP turbine 28). The burner portion 104 and the turbine portion 106 are integrally formed such that the burner portion 104 and the turbine portion 106 are a single integral structure, referred to herein as the integral outer wall 102.

[0122] As used herein, the term "integral" refers to a related component, such as outer wall 102, which is made as a single piece during manufacturing; that is, the final integral component is a single piece. Therefore, an integral component has a construction in which the integrated parts are indivisible, and differs from a component comprising multiple individual components that have been joined together and are referred to as a single component once joined, even if the components remain distinct and the single component is not indivisible (i.e., the component can be reseparated). The final integral component may comprise substantially continuous material components, or in other embodiments, may comprise multiple parts permanently joined together. In any case, the various parts forming the integral component are integrated together such that the integral component is a single piece with indivisible parts.

[0123] like Figure 2 As shown, in this embodiment, the burner portion 104 forming the integral structure of the outer wall 102 includes an outer bushing portion 108 of the burner 80 and at least a portion of a burner dome 118 extending across the front end 88 of the burner 80. The turbine portion 106 of the outer wall 102 includes an outer band portion 110 of the nozzle assembly 82N of the first turbine stage 82, a shroud portion 112 of the blade assembly 82B of the first turbine stage 82, an outer band portion 114 of the nozzle assembly 84N of the second turbine stage 84, and a shroud portion 116 of the blade assembly 84B of the second turbine stage 84. These outer boundary members are integrated as a single piece to form the integral structure of the outer wall 102. Therefore, in Figure 2 In an exemplary embodiment, at least a portion of the outer bushing portion 108, the outer belt portion 110, the protective cover portion 112, the outer belt portion 114, the protective cover portion 116, and the burner dome 118 are integrally formed, i.e., configured as a single unit or piece to form an integral or monolithic outer wall 102.

[0124] In addition, for Figure 2In the illustrated embodiment, at least a portion of the inner wall 120 that bounds the inner boundary of the flowpath 100 is integral with the outer wall 102 to form an integrated flowpath assembly 101. As shown, the combustor portion 104 further includes an inner liner portion 122, and the turbine portion 106 further includes an inner band portion 124 of the nozzle assembly 82N of the first stage turbine 82. In this manner, the inner liner portion 122 and the inner band portion 124 of the inner wall 120 are integral with the monolithic outer wall 102. Thus, the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, the shroud portion 116, the combustor dome portion 118, the inner liner portion 122, and the inner band portion 124 are integrally formed as a single monolithic structure.

[0125] In alternative example embodiments, some portions of the outer boundary and the inner boundary need not be integrally formed as a single monolithic structure. For example, in some embodiments, the combustor dome 118 can be a separate member from the outer wall 102 (formed by the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, and the shroud portion 116) and the inner wall 120 (formed by the inner liner portion 122 and the inner band portion 124). In yet other embodiments, only the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, and the shroud portion 116 are integrally formed as a single monolithic structure, while the combustor dome 118, the inner liner portion 122, and the inner band portion 124 are separate members. In yet other embodiments, the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, the shroud portion 116, and the combustor dome 118 are integrally formed as a single monolithic structure, while the inner liner portion 122 and the inner band portion 124 are separate members. In some embodiments, the outer liner portion 108 and the outer band portion 110 are integrally formed as a single monolithic structure, while the remaining members are separate members. In further embodiments, the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, the shroud portion 116, the combustor dome 118, and the inner liner portion 122 are integrally formed as a single monolithic structure, while the inner band portion 124 is a separate member. In some further embodiments, the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, the shroud portion 116, and the combustor dome 118 are integrally formed as a single monolithic structure, while the inner liner portion 122 and the inner band portion 124 are integrally formed as a single monolithic structure, and wherein the two monolithic structures can be attached or otherwise connected to one another. It will be recognized that other combinations of portions of the flowpath assembly 101 can be integrally formed as a single monolithic piece with one another, or can be separate members.

[0126] Compared to known gas turbine engines, the integration of various components of the outer and inner boundaries of the flow path assembly 101, as described above, reduces the number of individual parts or components within the engine 10, and reduces the weight, leakage, and complexity of the engine 10. For example, known gas turbine engines employ seals or sealing mechanisms at the mating points between individual parts of the flow path assembly to attempt to minimize leakage of combustion gases from the flow path. By integrating the outer boundary, for example, as described with respect to the integral outer wall 102, split points or mating points between the outer combustor bushing and the first turbine stage outer belt, the first turbine stage outer belt and the first turbine stage shroud, etc., can be eliminated, thereby eliminating leakage points and the seals or sealing mechanisms required to prevent leakage. Similarly, by integrating the components of the inner boundary, split points or mating points between the integrated inner boundary components are eliminated, thereby eliminating leakage points and the seals or sealing mechanisms required at the inner boundary. Accordingly, by utilizing integral components in the flow path assembly, undesirable leakage and unnecessary weight and complexity can be avoided. Other advantages of the integral outer wall 102, integral inner wall 120 and / or integral flow path assembly 101 will be recognized by those skilled in the art.

[0127] In some embodiments, nozzle guide vanes may be inserted through mounting openings defined in the outer wall 102 and inner wall 120 to form a nozzle assembly of flow path assembly 100, such as nozzle assembly 82N of a first turbine stage 82, nozzle assembly 84N of a second turbine stage 84, etc. An exemplary manner in which nozzle guide vanes may be inserted through the outer wall 102 and inner wall 120 and positioned to form a nozzle assembly of flow path assembly 101 is provided below. Such methods and the resulting nozzle assemblies are suitable for flow path assemblies with continuous CMC boundaries (e.g., ...). Figure 2 The flow path component 101 is particularly useful.

[0128] Figure 3 Provided Figure 2 A schematic axial cross-sectional view of a portion of the first-stage nozzle assembly 82N of the flow path assembly 101. As shown, the outer wall 102 is spaced apart from the inner wall 120 along the radial direction R. That is, the outer wall 102 is positioned relative to the axial centerline 12 ( Figure 2) positioned radially outward of the inner wall 120. The outer wall 102 has an inner surface 152 and an outer surface 154 spaced apart from the inner surface 152 along the radial direction R. The outer wall 102 defines a plurality of mounting openings 150. The mounting openings 150 extend through a radial thickness of the outer wall 102. That is, the mounting openings 150 extend between the inner surface 152 and the outer surface 154 of the outer wall 102 along the radial direction R. Each mounting opening 150 is shaped to complement a radial cross-section of the airfoil 126 of a given vane 68, e.g., such that the airfoil 126 of the vane 68 can be inserted therethrough. The mounting openings 150 are spaced apart from one another along the circumferential direction C. In particular, the outer wall 102 defines an array of circumferentially spaced apart mounting openings 150. The mounting openings 150 defined by the outer wall 102 are generally aligned with one another along the axial direction A for the nozzle assembly 82N. Additionally, for this embodiment, the mounting openings 150 are defined by the outer wall 102 along the turbine portion 106 of the outer wall 102.

[0129] The inner wall 120 has an inner surface 162 and an outer surface 164 spaced apart from the inner surface 162 along the radial direction R. Generally, the flowpath 100 is defined between the outer surface 164 of the inner wall 120 and the inner surface 152 of the outer wall 102. The inner wall 120 defines a plurality of mounting openings 160. The mounting openings 160 extend through a radial thickness of the inner wall 120. That is, the mounting openings 160 extend between the inner surface 162 and the outer surface 164 of the inner wall 120 along the radial direction R. The mounting openings 160 are spaced apart from one another along the circumferential direction C. More particularly, the inner wall 120 defines an array of circumferentially spaced apart mounting openings 160. Each mounting opening 160 is shaped to complement a radial cross-section of the airfoil 126 of a given vane 68, e.g., such that the airfoil 126 of the vane 68 can be inserted therethrough. Moreover, the mounting openings 160 defined by the inner wall 120 are generally aligned with one another along the axial direction A. Additionally, the mounting openings 160 of the inner wall 120 and the mounting openings 150 of the outer wall 102 are aligned along the axial direction A. That is, for this embodiment, the mounting openings 160 are defined by the inner wall 120 along the turbine portion 106 of the inner wall 120. Figure 3 The mounting openings 160 of the inner wall 120 and the mounting openings 150 of the outer wall 102 are aligned along the axial direction A for the nozzle assembly 82N of the first turbine stage 82 shown in FIG. 1. Thus, for this embodiment, the mounting openings 160 are defined by the inner wall 120 along the turbine portion 106 of the inner wall 120. Figure 2 It will be appreciated that the outer wall 102 and the inner wall 120 can define other arrays of mounting openings along different stages of the turbine section of the turbine engine, e.g., as shown in FIG. 2. Figure 6

[0130] ​Notably, for this embodiment, each mounting opening 160 of the inner wall 120 is positioned between adjacent mounting openings 150 of the outer wall 102 along the circumferential direction C. Thus, each mounting opening 150 of the outer wall 102 is thus positioned between adjacent mounting openings 160 of the inner wall 120 along the circumferential direction C. For example, as shown in FIG. 1 1, a first mounting opening 160A of the inner wall 120 is positioned between a first mounting opening 150A of the outer wall 102 and a second mounting opening 150B of the outer wall 102 along the circumferential direction C. This pattern continues around the entirety of the annular shape of the nozzle assembly 82N. Moreover, in some embodiments, the mounting openings 160 of the inner wall 120 are positioned intermediate between the mounting openings 150 of the outer wall 102 along the circumferential direction C. Figure 3

[0131] As further shown in FIG. 1 1, the outer wall 102 defines positioning grooves 156 that are spaced apart from one another along the circumferential direction C. More specifically, the outer wall 102 defines an array of circumferentially spaced apart positioning grooves 156. The positioning grooves 156 defined by the outer wall 102 are generally aligned with one another for the nozzle assembly 82N along the axial direction A. The positioning grooves 156 of the outer wall 102 are positioned opposite the mounting openings 160 of the inner wall 120 along the radial direction R. In other words, each mounting opening 160 of the inner wall 120 has an associated radially aligned positioning groove 156 defined by the outer wall 102. The positioning grooves 156 are defined by the outer wall 102 such that they extend radially from the inner surface 152 to a location radially inward of the outer surface 154 of the outer wall 102. Thus, the positioning grooves 156 of the outer wall 102 do not extend through the outer wall 102. Each positioning groove 156 has a radial cross-section that is shaped to complement and sized to accommodate a positioning end of a nozzle vane inserted through the inner wall 120, as will be explained more fully below. Figure 3

[0132] The inner wall 120 defines positioning grooves 166 that are spaced apart from one another along the circumferential direction C. Specifically, the inner wall 120 defines an array of circumferentially spaced apart positioning grooves 166. The positioning grooves 166 defined by the inner wall 120 are generally aligned with one another for the nozzle assembly 82N along the axial direction A. The positioning grooves 166 of the inner wall 120 are positioned opposite the mounting openings 150 of the outer wall 102 along the radial direction R. In other words, each mounting opening 150 of the outer wall 102 has an associated radially aligned positioning groove 166 defined by the inner wall 120. The positioning grooves 166 are defined by the inner wall 120 such that they extend radially from the outer surface 164 to a location radially outward of the inner surface 162 of the inner wall 120. Thus, the positioning grooves 166 of the inner wall 120 do not extend through the inner wall 120. Each positioning groove 166 has a radial cross-section that is shaped to complement and sized to accommodate a positioning end of a nozzle vane inserted through the outer wall 102, as will be explained more fully below.​​

[0133] Nozzle assembly 82N includes a plurality of vanes 68. Each vane 68 extends in the radial direction R between a mounting end 170 and a positioned end 172. Further, each vane 68 has an airfoil 126 positioned at least partially within flowpath 100, a mounting flange 174 positioned at the mounting end 170 of vane 68, and a seal portion 176 connecting airfoil 126 and mounting flange 174. For each vane 68, seal portion 176 is positioned between airfoil 126 and mounting flange 174 in the radial direction R. As shown in Figure 3 Nozzle assembly 82N includes a plurality of first vanes and a plurality of second vanes. In an alternating fashion, the first vanes are inserted inward into flowpath 100 through mounting openings 150 of outer wall 102 relative to axial centerline 12 in the radial direction R, and the second vanes are inserted outward into flowpath 100 through mounting openings 160 of inner wall 120 relative to axial centerline 12 in the radial direction R. More specifically, for this embodiment, every other vane 68 positioned along the circumferential direction C is a first vane that is inserted inward into flowpath 100 through one of mounting openings 150 of outer wall 102 in the radial direction R, and every other vane 68 positioned along the circumferential direction C is a second vane that is inserted outward into flowpath 100 through one of mounting openings 160 of inner wall 120 in the radial direction R. Thus, the plurality of first vanes and the plurality of second vanes are arranged in an alternating fashion along the circumferential direction C.

[0134] For example, as shown in Figure 3 first vane 68A is inserted through mounting opening 150A of outer wall 102. When first vane 68A is fully inserted in place, mounting flange 174 of first vane 68A is positioned outside of outer wall 102 relative to axial centerline 12 (as shown by the dashed line) in the radial direction R, and airfoil 126 of first vane 68A is positioned at least partially within flowpath 100. Second vane 68B is positioned adjacent to first vane 68A along the circumferential direction C and is inserted through mounting opening 160B of inner wall 120. When second vane 68B is fully inserted in place, mounting flange 174 of second vane 68B is positioned inside of inner wall 120 relative to axial centerline 12 (as shown by the dashed line) in the radial direction R, and airfoil 126 of second vane 68B is positioned at least partially within flowpath 100. Continuing the alternating arrangement, first vane 68C is positioned adjacent to second vane 68B along the circumferential direction C and is inserted through mounting opening 150C of outer wall 102. When first vane 68C is fully inserted in place, mounting flange 174 of first vane 68C is positioned outside of outer wall 102 relative to axial centerline 12 (as shown by the dashed line) in the radial direction R, and airfoil 126 of first vane 68C is positioned at least partially within flowpath 100. Figure 2 Figure 2 Figure 2 ​​The first guide vane 68C is positioned radially R outside the outer wall 102, and the airfoil 126 is at least partially positioned within the flow path 100. The second guide vane 68D is positioned circumferentially C adjacent to the first guide vane 68C and inserted through the mounting opening 160D of the inner wall 120. When the second guide vane 68D is fully inserted, the mounting flange 174 of the second guide vane 68D is positioned relative to the axial centerline 12( Figure 2 The airfoil 126 of the second guide vane 68D is located inside the inner wall 120 in the radial direction R, and is at least partially located within the flow path 100.

[0135] like Figure 3 As shown, the guide vanes 68A, 68B, 68C, and 68D are inserted through their respective mounting openings 150A, 160B, 150C, and 160D in a radial direction R that alternates between inward and outward. That is, the first guide vanes 68A and 68C are inserted radially inward relative to the axial centerline 12, and the second guide vanes 68B and 68D are inserted radially outward relative to the axial centerline 12. Furthermore, in this embodiment, each of the plurality of guide vanes 68 extends radially R between the outer wall 102 and the inner wall 120. However, in… Figure 3 In the illustrated embodiment, no guide vane 68 extends through both the inner wall 120 and the outer wall 102.

[0136] Advantageously, for example, compared to a design in which all blades are inserted radially inward through the outer wall 102, the alternating radial insertion direction of the guide vanes 68 between the outer wall 102 and the inner wall 120 provides more space for securing and sealing the guide vanes 68 in place. That is, the alternating insertion direction of the guide vanes provides circumferential space between the mounting flanges 174 of the guide vanes 68. Therefore, the mounting flanges 174 can have a larger size without overlapping or contacting each other. Increasing the size of the mounting flanges 174 and / or providing sufficient circumferential space between them provides a more stable nozzle assembly and prevents unwanted wear on the guide vanes and on the outer wall 102 and inner wall 120, among other benefits.

[0137] Each mounting opening 150, 160 of the outer wall 102 and inner wall 120 is sized to receive one of the guide vanes 68 and is shaped to be complementary to the radial section of the guide vane 68 through which it is inserted. Specifically, the mounting openings 150, 160 are shaped to be complementary to the radial section of the airfoil 126 of each guide vane 68 and are sized to receive that radial section. Furthermore, the mounting openings 150, 160 are sized and shaped such that the sealing portion 176 of the guide vane 68 (and therefore the mounting flange 174) cannot pass through its insertion. For example, as... Figure 3As shown in FIG. 2, the seal portion 176 of each vane 68 wedges the vane 68 into engagement with the outer wall 102 for each vane 68 inserted through one of the mounting openings 150 of the outer wall 102. Similarly, the seal portion 176 of each vane 68 wedges the vane 68 into engagement with the inner wall 120 for each vane 68 inserted through one of the mounting openings 160 of the inner wall 120. The seal portion 176 of each vane 68 prevents or inhibits fluid leakage to or from the flowpath 100. The vanes 68 can be press- or interference- fit within their respective mounting openings 150, 160 such that the seal portion 176 of each vane 68 seals the mounting openings 150, 160 about their entire circumference. In some preferred embodiments, the seal portion 176 of each vane 68 is sized such that the seal portion 176 is flush or radially aligned with the inner surface 152 of the outer wall 102 or the outer surface 164 of the inner wall 120 depending on the insertion direction of the particular vane 68.

[0138] For this embodiment, as in Figure 4 and Figure 5 As best shown in the partial enlarged view of FIG. 2, each mounting opening 150, 160 includes an insertion end 180 and a flowpath end 182 radially spaced from the insertion end 180. For the mounting openings 150 defined by the outer wall 102, the insertion end 180 of each mounting opening 150 is positioned radially outward of the flowpath end 182 along the radial direction R. For the mounting openings 160 defined by the inner wall 120, the insertion end 180 of each mounting opening 160 is positioned radially inward of the flowpath end 182 along the radial direction R. Further, as shown, each mounting opening 150, 160 is defined in part by a sidewall 158, 168, respectively. Notably, the insertion end 180 of each mounting opening 150, 160 has a greater radial cross-sectional area than the flowpath end 182. Accordingly, the sidewall 158, 168 of each mounting opening 150, 160 is angled relative to the radial direction R such that the sidewall 158, 168 converges as it extends along the radial direction R toward its respective flowpath end 182. Conversely, the sidewall 158, 168 of each mounting opening 150, 160 is angled relative to the radial direction R such that the sidewall 158, 168 diverges as it extends along the radial direction R toward its respective insertion end 180. The seal portion 176 of each vane 68 can include one or more angled surfaces 178 shaped to complement the sidewall 158, 168 of the mounting opening 150, 160. In this manner, the mechanical advantage of the complementary angled surfaces wedges the seal portion 176 into sealing engagement with the sidewall 158, 168 of the mounting opening 150, 160 when the vane 68 is inserted into its respective mounting opening 150, 160.

[0139] As Figure 3As further shown in FIG. 1, when the vanes 68 are inserted through their respective mounting openings 150, 160, the mounting flanges 174 of the vanes 68 inserted through the mounting openings 150 of the outer wall 102 engage the outer surface 154 of the outer wall 102, and the mounting flanges 174 of the vanes 68 inserted through the mounting openings 160 of the inner wall 120 engage the inner surface 162 of the inner wall 120. Accordingly, the mounting flanges 174 of the vanes 68 provide stops when the vanes 68 are inserted in their respective mounting openings 150, 160 and prevent the vanes 68 from sliding through their respective mounting openings 150, 160. In some embodiments, the mounting flanges 174 of each vane 68 extend at least twice the distance from the arc of the airfoil 126 to the side wall 158, 168 of the outer or inner wall 102, 120 that defines the mounting opening 150 through which the vane 68 is inserted.

[0140] In addition, when the vanes 68 are inserted through their respective mounting openings 150, 160, the mounting end 170 of each vane 68 is positioned in one of the positioning grooves 166 of the inner wall 120 or one of the positioning grooves 156 of the outer wall 102, depending on the direction of insertion of the vane 68. For example, as shown in FIG. 1, for vanes 68 inserted radially inward through the outer wall 102, the mounting end 170 of each vane 68 is positioned in the associated positioning groove 166 defined by the inner wall 120. For vanes 68 inserted radially outward through the inner wall 120, the mounting end 170 of each vane 68 is positioned in the associated positioning groove 156 defined by the outer wall 102. Figure 3 By positioning the positioning end 172 of each vane 68 in one of the positioning grooves 156, 166, the vanes 68 are constrained in place and prevented from moving about during operation of the turbofan engine 10 (FIG. 1). Figure 1

[0141] Figure 6 A perspective cross-sectional view of a portion of the flowpath assembly 101 of Figure 2 is provided, showing the outer wall 102 and the inner wall 120 formed as a single piece member. For this embodiment, the outer liner portion 108, the outer band portion 110, the shroud portion 112, the outer band portion 114, the shroud portion 116, the combustor dome 118, the inner liner portion 122, and the inner band portion 124 (see FIG. 1) are integrally formed as a single unitary structure. In addition, as shown in FIG. 1, the vanes 68 are inserted radially inward through the outer wall 102 (indicated by solid arrows) alternate with the vanes 68 inserted radially outward through the inner wall 120 (indicated by dashed arrows) along the circumferential direction C. In addition, as shown in FIG. 1, the vanes 69 of the second stage nozzle assembly 84N can likewise be inserted through the outer wall 102 and the inner wall 120 in an alternating fashion. It will be appreciated that the vanes 68 of the first stage nozzle assembly 84N and the vanes 69 of the second stage nozzle assembly 84N can be inserted through the outer wall 102 and the inner wall 120 in any suitable alternating fashion. Figure 2 Figure 6 Figure 6 Figure 6 ​​​​Only a portion of the integral flowpath assembly 101 is shown, and although not the entire circumference thereof is shown in Figure 6 the flowpath assembly 101 is a single integral piece both circumferentially and axially. As such, the integral flowpath assembly 101 defines a generally annular (i.e., generally annular shape) flowpath between the outer wall 102 and the inner wall 120. Additionally, it will be appreciated that the vanes of the nozzle assemblies 82N, 84N can be annularly inserted in an alternating fashion around the assembly.

[0142] In some embodiments, the vanes of the nozzle assembly can be cooled by a suitable fluid, such as by cooling air supplied by a suitable cooling source. For example, Figure 7 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly is provided, showing various cooling features of the assembly. For example, as shown in Figure 7 the nozzle assembly can be Figure 3 the nozzle assembly 82N of FIG. 1. As shown, for this embodiment, cooling fluid is supplied to the first vanes 68A, 68C from the outer diameter of the assembly, and to the second vanes 68B, 68D from the inner diameter of the assembly. Specifically, as shown, cooling fluid C OD is supplied to the first vanes 68A and to the first vanes 68C. Cooling fluid C OD may flow through one or more internal passages defined by the first vanes 68A and the first vanes 68C. The mounting flange 174 and / or the outer wall 102 can define one or more metering holes operable to contain cooling fluid C OD . The airfoils 126 of the vanes 68A, 68C can define a plurality of openings such that cooling fluid C OD may be discharged into the flowpath 100. Further, cooling fluid C ID is supplied to the second vanes 68B and to the second vanes 68D. Cooling fluid C ID may flow through one or more internal passages defined by the second vanes 68B and the second vanes 68D. The mounting flange 174 and / or the inner wall 120 can define one or more metering holes operable to contain cooling fluid C ID . The airfoils 126 of the vanes 68B, 68D can define a plurality of openings such that cooling fluid C ID may be discharged into the flowpath 100. Notably, for this embodiment, every other vane 68 is supplied with coolant from the outer diameter of the assembly, while every other vane 68 is supplied with coolant from the inner diameter of the assembly.

[0143] Although Figure 7 the nozzle assembly 82N of FIG. 1 is shown as a first stage nozzle assembly, it will be appreciated that the nozzle assemblies of other stages can likewise include cooling features, such as described above. For example, the nozzle assemblies of Figure 7 and Figure 2 may be cooled in the same or similar manner as the first stage nozzle assembly 82N of FIG. 1.Figure 6 For example, in some example embodiments, the vanes can each define one or more internal cooling cavities. For at least some vanes, the one or more internal cooling cavities can be supplied with cooling fluid (e.g., compressor discharge air) via a supply plenum positioned at an outer diameter of the nozzle assembly. The cooling fluid can be discharged from the internal cooling cavities via a discharge plenum positioned at an inner diameter. In this manner, the cooling fluid can flow radially inward through the vanes. The cooling fluid can be supplied in a radially inward direction to all of the vanes, or alternatively, at least some of the vanes can be cooled by cooling fluid flowing in a radially outward direction. For example, the one or more internal cooling cavities can be supplied with cooling fluid via a supply plenum positioned at an inner diameter of the nozzle assembly. The cooling fluid can be discharged from the internal cooling cavities via a discharge plenum positioned at an outer diameter. In this manner, the cooling fluid can flow radially outward through the vanes. In some embodiments, the cooling fluid can be supplied in a radially outward direction to all of the vanes, or as described above, where the direction of cooling fluid flow through the internal cooling cavities can alternate depending on the direction of vane insertion. In yet some further embodiments, the cooling fluid can be supplied through metallic tubes extending through the CMC vanes of the nozzle assembly.

[0144] Figure 8 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine in accordance with example embodiments of the present subject matter is provided. For the illustrated embodiment of Figure 8 For the illustrated embodiment, the vanes 68 of the nozzle assembly 82N are radially inserted inwardly and outwardly in an alternating manner as described above. However, for this embodiment, the outer wall 102 defines a plurality of through openings 190 instead of the positioning grooves 156( Figure 3 ) As shown, the through openings 190 are positioned opposite the mounting openings 160 defined by the inner wall 120 in the radial direction R. Additionally, the inner wall 120 defines a plurality of through openings 192 instead of the positioning grooves 166( Figure 3 ) As shown, the through openings 192 are positioned opposite the mounting holes 150 defined by the outer wall 102 in the radial direction R.

[0145] As Figure 8 As further shown, the positioning end 172 of each vane 68 is inserted through one of the through openings 190, 192. Specifically, the positioning end 172 of the first vane 68A is inserted through the through opening 192 defined by the inner wall 120 such that at least a portion of the first vane 68A is positioned radially inward of the inner wall 120 relative to the axial centerline 12( Figure 2 ) in the radial direction R. The positioning end 172 of the second vane 68B is inserted through the through opening 190 defined by the outer wall 102 such that at least a portion of the second vane 68B is positioned radially outward of the outer wall 102 relative to the axial centerline 12( Figure 2) positioned radially outward of the outer wall 102 along the radial direction R. The positioning end 172 of the first vane 68C is inserted through a through opening 192 defined by the inner wall 120 such that at least a portion of the first vane 68C is positioned relative to the axial centerline 12 Figure 2 ) positioned radially inward of the inner wall 120 along the radial direction R. Further, the positioning end 172 of the second vane 68D is inserted through a through opening 190 defined by the outer wall 102 such that at least a portion of the second vane 68D is positioned relative to the axial centerline 12 Figure 2 ) positioned radially outward of the outer wall 102 along the radial direction R.

[0146] For this embodiment, each vane 68 is held at its positioning end 172 by a holding device 195. The holding device 195 can be any suitable mechanical holding device, such as a ring clamp, a retaining pin, a nut, etc. As shown, the first vane 68A is held at its positioning end 172 by a holding device 195. The holding device 195 can fixedly position the positioning end 172 relative to the outer wall 102. In some embodiments, a radial space in which no component is positioned can be defined between the holding device 195 and the outer surface 154 of the outer wall 102, for example, to allow for thermal growth. In still other embodiments, a bushing or the like can be positioned therebetween to protect the CMC outer wall 102, for example, from delamination. The first vane 68C is held by a holding device 195 in a similar manner as the first vane 68A.

[0147] Further, the second vane 68B is held at its positioning end 172 by a holding device 195. The holding device 195 can fixedly position the positioning end 172 relative to the inner wall 120. In some embodiments, a radial space in which no component is positioned can be defined between the holding device 195 and the inner surface 162 of the inner wall 120, for example, to allow for thermal growth. In still other embodiments, a bushing or the like can be positioned therebetween to protect the CMC inner wall 120, for example, from delamination. The second vane 68D is held by a holding device 195 in a similar manner as the second vane 68B.

[0148] Figure 9 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine according to yet another exemplary embodiment of the present subject matter is provided. For the illustrated embodiment of Figure 9 the illustrated embodiment of, the vanes 68 of the nozzle assembly 82N are inserted radially inward and outward in an alternating manner as described above. For this embodiment, the outer wall 102 defines a plurality of through openings 190 and the inner wall 120 defines a plurality of through openings 192, for example, similar to Figure 8 the illustrated embodiment of. Further, as shown in Figure 9 the positioning ends 172 of the vanes 68A, 68B, 68C, and 68D are each inserted through their associated through openings 190, 192. the positioning ends 172 of the vanes 68A, 68B, 68C, and 68D are each inserted through their associated through openings 190, 192.

[0149] For the illustrated embodiment of Figure 9 , each vane 68 is held at its location end 172 by one or more plies that are laminated and heat treated as described herein. As shown, the first vane 68A is held at its location end 172 to the cold side or outer surface 154 of the outer wall 102 by one or more plies 200. The plies 200 can secure the location end 172 to the outer wall 102. The first vane 68C is held at its location end 172 to the outer surface 154 of the outer wall 102 by one or more plies 200 in a similar manner as the first vane 68A. In addition, the second vane 68B is held at its location end 172 to the cold side or inner surface 162 of the inner wall 120 by one or more plies 202. The plies 202 can secure the location end 172 to the inner wall 120. The second vane 68D is held at its location end 172 to the inner surface 162 of the inner wall 120 by one or more plies 202 in a similar manner as the second vane 68B.

[0150] Figure 10 A schematic axial cross-sectional view of a portion of a first stage nozzle assembly of a flowpath assembly for a turbine engine in accordance with further example embodiments of the present subject matter is provided. For the illustrated embodiment of Figure 10 , the vanes 68 of the nozzle assembly 82N are radially inserted inwardly and outwardly in an alternating manner as described above. For this embodiment, the outer wall 102 defines a plurality of through openings 190 and the inner wall 120 defines a plurality of through openings 192, e.g., similar to the illustrated embodiments of Figure 8 and Figure 9 . In addition, as shown in Figure 10 , the location ends 172 of the vanes 68A, 68B, 68C, and 68D are each inserted through their associated through openings 190, 192.

[0151] For the illustrated embodiment of Figure 10 , each vane 68 is held at its location end 172 by a bonding material. For example, the bonding material can be a braze-type bond or other infiltrating-type bonding material. As shown, the first vane 68A is held at its location end 172 to the cold side or outer surface 154 of the outer wall 102 by a bonding material 206. The bonding material 206 can secure the location end 172 to the outer wall 102. The first vane 68C is held at its location end 172 to the outer surface 154 of the outer wall 102 by a bonding material 206 in a similar manner as the first vane 68A. In addition, the second vane 68B is held at its location end 172 to the cold side or inner surface 162 of the inner wall 120 by a bonding material 208. The bonding material 208 can secure the location end 172 to the inner wall 120. The second vane 68D is held at its location end 172 to the inner surface 162 of the inner wall 120 by a bonding material 208 in a similar manner as the second vane 68B.

[0152] Figure 11 A schematic axial section view of a portion of a first-stage nozzle assembly for a flow path assembly of a turbine engine, according to another further exemplary embodiment of this subject matter, is provided. Figure 11 In the illustrated embodiment, the guide vanes 68 of the nozzle assembly 82N are inserted radially inward and outward in an alternating manner as described above. For this embodiment, the outer wall 102 defines a plurality of through openings 190, while the inner wall 120 defines a plurality of through openings 192, for example, similar to... Figure 8 , Figure 9 and Figure 10 The illustrated embodiment. Furthermore, as... Figure 11 As shown, the positioning ends 172 of guide vanes 68A, 68B, 68C and 68D are each inserted through their respective through openings 190, 192.

[0153] for Figure 11 In the illustrated embodiment, each guide vane 68 may define a mounting end recess 210 at or near the mounting end 170. Each guide vane 68 may also define a positioning end recess 212 at or near the positioning end 172. The mounting end recess 210 of each guide vane 68 is operable to receive a mounting barb ring 214, and the positioning end recess 212 of each guide vane 68 is operable to receive a positioning barb ring 216. Rings 214, 216 may be formed of any suitable material capable of being compressed when the guide vane is inserted into place. For example, rings 214, 216 may be formed of CMC material. When the guide vane 68 is inserted into its respective position, the mounting barb ring 214 is received within the mounting end recess 210 and compressed / deformed to form a barb seal at or near the mounting end 170 of the guide vane. Similarly, the positioning barb ring 216 is received within the positioning end recess 212 and compressed / deformed to form a barb seal at or near the positioning end 172. The formed barbed seal reduces or, in some cases, eliminates leakage in the flow path 100. Although barbed seals formed at the mounting end 170 and positioning end 172 of each guide vane 68 by engagement of rings 214, 216 with the guide vane 68 and the corresponding walls 102, 120 are shown, it will be appreciated that in some embodiments, the nozzle assembly 82N may have only a formed positioning end barbed seal or a mounting end barbed seal.

[0154] Figure 12 A flowchart is provided of an exemplary method 300 for assembling a nozzle assembly for a turbine engine, defined in axial, radial, and circumferential directions, according to exemplary embodiments of this subject matter. For example, method 300 can be used to assemble a first turbine stage nozzle assembly 82N and / or a second turbine stage nozzle assembly 84N as described herein. Method 300 can also be used to assemble other nozzle assemblies, such as nozzle assemblies within the compression section of a gas turbine engine or within the LP turbine of a gas turbine engine. Additionally, Figure 12Steps are shown performed in a particular order for purposes of illustration and discussion. Those of ordinary skill in the art, using the disclosures provided herein, will understand that the various steps of any of the methods disclosed herein can be changed, modified, extended, rearranged, and / or omitted in various ways without departing from the scope of the present disclosure.

[0155] At 302, the method 300 includes inserting a plurality of vanes radially inward through an outer wall of a nozzle assembly. For example, the nozzle assembly can be the nozzle assembly 82N, the outer wall can be the outer wall 102, and the vanes can be the vanes 68 of Figure 3 FIG. 1. As shown in FIG. 1, the vanes 68 are inserted radially inward through the mounting openings 150 defined by the outer wall 102. The inserted vanes 68 are inserted radially inward such that the vanes 68 extend between the outer wall 102 and the inner wall 120 along the radial direction R, and such that the airfoils 126 of the vanes 68 are positioned within the flowpath 100. Further, the mounting flanges 174 of each vane 68 remain positioned radially outward of the outer wall 102, and can engage the outer surface 154 of the outer wall 102. The positioned end 172 of the inserted vanes 68 is positioned within the positioned recess 166 defined by the inner wall 120. Further, the sealing portion 176 of each vane 68 is wedged into place with the outer wall 102 to seal the mounting openings 150 in the outer wall 102. Figure 3

[0156] In some embodiments, the outer wall 102 can be a monolithic outer wall formed as a monolithic structure having at least an outer liner portion 108 Figure 2 and an outer band portion 110 Figure 2 . In such embodiments, the mounting openings 150 are defined by the outer liner portion 108 of the outer wall 102. In some embodiments, the outer wall 102 can simply be an outer band separate from the combustion liner and downstream shroud.

[0157] At 304, the method 300 includes inserting a plurality of vanes radially outward through an inner wall of a flowpath assembly. Further, in such embodiments, the plurality of vanes inserted radially inward through the outer wall are circumferentially alternating with the plurality of vanes inserted radially outward through the inner wall. For example, the inner wall can be the inner wall 120, and the vanes can be the vanes 68 of Figure 3 FIG. 1. As shown in FIG. 1, the vanes 68 are inserted radially inward through the mounting openings 150 defined by the outer wall 102. The inserted vanes 68 are inserted radially inward such that the vanes 68 extend between the outer wall 102 and the inner wall 120 along the radial direction R, and such that the airfoils 126 of the vanes 68 are positioned within the flowpath 100. Further, the mounting flanges 174 of each vane 68 remain positioned radially outward of the outer wall 102, and can engage the outer surface 154 of the outer wall 102. The positioned end 172 of the inserted vanes 68 is positioned within the positioned recess 166 defined by the inner wall 120. Further, the sealing portion 176 of each vane 68 is wedged into place with the outer wall 102 to seal the mounting openings 150 in the outer wall 102. Figure 3 ​As shown in FIG. 6, the vanes 68 are inserted radially outward through the mounting openings 160 defined by the inner wall 120. Inserting the vanes 68 radially outward such that the vanes 68 extend in the radial direction R between the inner wall 120 and the outer wall 102, and such that the airfoils 126 of the vanes 68 are positioned within the flowpath 100. Further, the mounting flanges 174 of each vane 68 remain positioned radially inward of the inner wall 120 and can engage the inner surface 162 of the inner wall 120. The positioned end 172 of the inserted vanes 68 is seated within and positioned by the positioned recess 156 defined by the outer wall 102. Further, the sealing portion 176 of each vane 68 is wedged into place with the inner wall 120 to seal the mounting openings 160 in the inner wall 120. As best shown in FIG. 6, the mounting openings 150, 160 defined by the outer wall 102 and the inner wall 120 can be axially aligned. Figure 6

[0158] Notably, as shown in FIG. 3, in some embodiments, the plurality of vanes 68 inserted inward through the outer wall 102 in the radial direction R at 302 alternate with the plurality of vanes 68 inserted outward through the inner wall 120 in the radial direction R at 304 in the circumferential direction C. That is, for the illustrated embodiment of FIG. 3, every other vane 68 around the annulus of the nozzle assembly 82N is inserted radially inward through the outer wall 102, and every other vane 68 around the annulus of the nozzle assembly 82N is inserted radially outward through the inner wall 102. In this manner, as described previously, for example, alternating the radial insertion direction of the vanes 68 between the outer wall 102 and the inner wall 120 provides more space for securing and sealing the vanes 68 into place than a design in which all of the vanes are inserted radially inward through the outer wall 102, among other benefits. Figure 3 Figure 3

[0159] In some embodiments, the inner wall 102 can be of unitary construction with the outer wall 102 (e.g., as shown in FIGS. 5A and 5B). The inner wall 120 can comprise at least an inner band portion 124 (e.g., as shown in FIGS. 5A and 5B). In such embodiments, the mounting openings 160 are defined by the inner sleeve portion 124 of the inner wall 120. In some embodiments, the inner wall 120 can simply be an inner band separate from the inner combustion sleeve and downstream platform. Figure 2 Figure 6 Figure 2

[0160] ​​​​​​In some embodiments, such as when there is an odd number of guide vanes in the nozzle assembly, not all guide vanes of the nozzle assembly alternate between inwardly inserted and outwardly inserted guide vanes. For example, a method of assembling a nozzle assembly for defining axial, radial, and circumferential directions of a turbine engine may include: inserting a first guide vane radially inward through an outer wall of the nozzle assembly, adjacent to the first guide vane in the circumferential direction, and inserting a second guide vane radially outward through an inner wall of the nozzle assembly, spaced radially from the outer wall, adjacent to the first guide vane in the circumferential direction. In some embodiments, the method further includes inserting a third guide vane radially inward through the outer wall of the nozzle assembly, adjacent to the first guide vane and opposite the second guide vane in the circumferential direction. In this manner, the first guide vane is positioned circumferentially between the second and third guide vanes. Both the first and third guide vanes are inserted radially inward through the outer wall of the nozzle assembly, while the second guide vane is inserted radially outward through the inner wall of the nozzle assembly.

[0161] For example, Figure 13 A schematic axial cross-sectional view of a portion of a nozzle assembly 82N according to an exemplary embodiment of the present subject is provided. For this embodiment, the nozzle assembly 82N includes a plurality of triplet segments spaced apart from each other in the circumferential direction C. Figure 13 The figure shows two complete segments, including a first segment 94 and an adjacent second segment 96. As shown, both the first segment 94 and the second segment 96 have three guide vanes; therefore, these segments are triplet segments.

[0162] For the first segment 94, moving from left to right, the first guide vane 68-1 is inserted radially R through the outer wall 102 of the nozzle assembly 82N, such that its airfoil 126 extends through the mounting opening 150 of the outer wall 102 and is at least partially positioned within the flow path 100. It is noteworthy that the mounting flange 174 of the first guide vane 68-1 is positioned relative to the axial centerline (e.g., Figure 1 The axial centerline 12) is located on the outside of the outer wall 102 along the radial direction R. It is adjacent to the first guide vane 68-1 along the circumferential direction C (from...). Figure 13 From a certain perspective, immediately to the right of the first guide vane 68-1, the second guide vane 68-2 is inserted radially R through the inner wall 120 of the nozzle assembly, such that its airfoil 126 extends through the mounting opening 160 of the inner wall 120 and is at least partially positioned within the flow path 100. It is noteworthy that the mounting flange 174 of the second guide vane 68-2 is positioned relative to the axial centerline (e.g., ...). Figure 1 The axial centerline 12) is located inside the inner wall 120 along the radial direction R. It is adjacent to the second guide vane 68-2 along the circumferential direction C (from...). Figure 13from the perspective of the viewer looking immediately to the right of the second vane 68-2 of the first segment 94), the other first vane 68-1 (corresponding to the third vane described above) is inserted inwardly through the outer wall 102 of the nozzle assembly 82N in the radial direction R such that its airfoil 126 extends through the mounting opening 150 of the outer wall 102 and is at least partially positioned within the flowpath 100. As shown, the mounting flange 174 of the first vane 68-1 is positioned outside of the outer wall 102 in the radial direction R relative to the axial centerline (e.g., the axial centerline 12 of Figure 1 the first segment 94. Thus, the vanes of the first segment 94 alternate between first vanes and second vanes.

[0163] As Figure 13 further shown in Figure 13 , the second segment 96 adjacent to the first segment 94 has the same vane configuration as the first segment 94. Accordingly, the first vane 68-1 of the first segment 94 (i.e., the first vane 68-1 furthest to the right in Figure 13 ) is adjacent to the first vane 68-1 of the second segment 96 (i.e., the first vane 68-1 furthest to the left in ).

[0164] In yet other embodiments, a method of assembling a nozzle assembly of a turbine engine defining an axial direction, a radial direction, and a circumferential direction can include inserting a first vane inwardly through an outer wall of the nozzle assembly in the radial direction and adjacent to the first vane in the circumferential direction, inserting a second vane outwardly through an inner wall of the nozzle in the radial direction, the inner wall being spaced apart from the outer wall in the radial direction. The method can further include inserting a third vane outwardly through the inner wall of the nozzle assembly in the radial direction adjacent to the second vane and opposite the first vane in the circumferential direction. In this manner, the second vane is positioned between the first vane and the third vane in the circumferential direction. Both the second vane and the third vane are inserted outwardly through the inner wall of the nozzle assembly in the radial direction, while the first vane is inserted inwardly through the outer wall of the nozzle assembly in the radial direction. For example, in this embodiment with reference to Figure 13 , the nozzle segment 82N can be configured such that the second vane 68-2 is positioned in the location shown for the first vane 68-1 and the first vane 68-1 is positioned in the location shown for the second vane 68-2.

[0165] In some embodiments, the mounting flange of a vane of the nozzle assembly can be at least partially positioned within its respective mounting opening, as shown in Figure 14 and Figure 15 . Indeed, the mounting flange of the vane can be wedgedly engaged with its respective wall.

[0166] In some example embodiments, with reference to Figure 14 , a nozzle assembly 82N for a turbine engine is provided. The turbine engine can define an axial direction A( Figure 2), a radial direction R, a circumferential direction C, and an axial centerline 12 (A) extending along an axial direction A. Figure 2 The nozzle assembly 82N includes an outer wall 102 defining a mounting opening 150 and an inner wall 120 defining a mounting opening 160. The outer wall 102 and the inner wall 120 are spaced apart from one another along the radial direction R. The inner wall 120 and the outer wall 102 define a flowpath 100. The nozzle assembly 82N includes a first vane 68A having an airfoil 126 and a mounting flange 174. The airfoil 126 of the first vane 68A is positioned at least partially within the flowpath 100, and the mounting flange 174 of the first vane 68A is positioned at least partially within the mounting opening 150, and more specifically the mounting opening 150A, of the outer wall 102. The nozzle assembly 82N includes a second vane 68B positioned adjacent to the first vane 68A along the circumferential direction C. The second vane 68B has an airfoil 126 and a mounting flange 174. The airfoil 126 of the second vane 68B is positioned at least partially within the flowpath 100, and the mounting flange 174 of the second vane 68B is positioned at least partially within the mounting opening 160, and more specifically the mounting opening 160B, of the inner wall 120.

[0167] In some embodiments, the mounting flange 174 of the first vane 68A is positioned entirely within the mounting opening 150A of the outer wall 102. Also, in some embodiments, the mounting flange 174 of the second vane 68B is positioned entirely within the mounting opening 160B of the inner wall 120. Further, in some embodiments, the outer wall 102 has an outer surface 154 and an inner surface 152. A distance between the outer surface 154 and the inner surface 152 defines a radial thickness of the outer wall 102. In such embodiments, the mounting flange 174 of the first vane 68A extends along the radial direction R between the outer surface 154 and the inner surface 152 of the outer wall 102. Additionally or alternatively, in some embodiments, the inner wall 120 has an outer surface 164 and an inner surface 162. A distance between the outer surface 164 and the inner surface 162 of the inner wall 120 defines a radial thickness of the inner wall 120. In such embodiments, the mounting flange 174 of the second vane 68B extends along the radial direction R between the outer surface 164 and the inner surface 162 of the inner wall 120.

[0168] However, in other embodiments, the mounting flange 174 of the first vane 68A need not extend along the radial direction R between the outer surface 154 and the inner surface 152 of the outer wall 102, and the mounting flange 174 of the second vane 68B need not extend along the radial direction R between the outer surface 164 and the inner surface 162 of the inner wall 120. For example, as shown in FIG. 6, the mounting flange 174 of the first vane 68A can extend along the radial direction R between the outer surface 154 and the inner surface 152 of the outer wall 102, and the mounting flange 174 of the second vane 68B can extend along the radial direction R between the outer surface 164 and the inner surface 162 of the inner wall 120. Figure 15As shown, the mounting flanges 174 of the first guide vane 68A and the second guide vane 68B are positioned within their respective mounting openings 150A and 160B, and respectively placed on their respective mounting seats 188A and 188b. That is, the mounting flange 174 of the first guide vane 68A is placed on the mounting seat 188A, and the mounting flange 174 of the second guide vane 68B is placed on the mounting seat 188B. The mounting flanges 174 placed on their respective mounting seats 188A and 188B provide additional retention for the guide vanes and serve as positioning aids when the guide vanes 68A and 68B are inserted through their respective outer walls 102 and inner walls 120.

[0169] like Figure 14 and Figure 15 As shown, in some embodiments, the sidewall of the mounting flange 174 of the first guide vane 68A wedgably engages the outer wall 102. In some embodiments, the sidewall of the mounting flange 174 of the second guide vane 68B wedgably engages the inner wall 120. The wedging of the first guide vane 68A with the outer wall 102 provides the mechanical advantage of holding the first guide vane 68A in place. Similarly, the wedging of the second guide vane 68B with the inner wall 120 provides the mechanical advantage of holding the second guide vane 68B in place.

[0170] In some embodiments, reference Figure 14 The outer wall 102 has an outer surface 154 and an inner surface 152 spaced apart from each other in the radial direction R. The outer wall 102 defines an outer periphery of a mounting opening 150A at the outer surface 154 and an inner periphery of a mounting opening 150A at the inner surface 152. Figure 14 As shown, the outer periphery of the mounting opening 150A on the outer wall 102 is larger than its inner periphery. Furthermore, as shown, the inner wall 120 has an outer surface 164 and an inner surface 162 spaced apart from each other in the radial direction R. The inner wall 120 defines the outer periphery of the mounting opening 160B at the outer surface 164 and the inner periphery of the mounting opening 160B at the inner surface 162. As shown, the inner periphery of the mounting opening 160B on the inner wall 120 is larger than its outer periphery. The geometry of the mounting openings 150A and 160B facilitates the wedging of the mounting flange relative to the outer wall 102 and the inner wall 120.

[0171] In some embodiments, the mounting opening 150A of the outer wall 102 is one of a plurality of mounting openings 150 of the outer wall 102, and the mounting opening 160A of the inner wall 120 is one of a plurality of mounting openings 160 of the inner wall 120. In such embodiments, the first vane 68A is one of a plurality of first vanes, and the second vane 68B is one of a plurality of second vanes. Further, in some embodiments, the airfoil 126 of each first vane of the plurality of first vanes is positioned at least partially within the flowpath 100, and the mounting flange 174 of each first vane of the plurality of first vanes is positioned at least partially in one of the plurality of mounting openings 150 of the outer wall 102. In some further embodiments, the airfoil 126 of each second vane of the plurality of second vanes is positioned at least partially within the flowpath 100, and the mounting flange 174 of each second vane of the plurality of second vanes is positioned at least partially in one of the plurality of mounting openings 160 of the inner wall 120. In some further embodiments, the plurality of first vanes and the plurality of second vanes are arranged in an alternating fashion along the circumferential direction C, for example, as shown in FIGS. 1-2. Figure 14 and Figure 15

[0172] In some embodiments, the inner wall 120 defines positioning recesses 166 that are spaced apart from one another along the circumferential direction C and are positioned opposite the mounting openings 160 of the outer wall 102 along the radial direction R, and the outer wall 102 defines positioning recesses 156 that are spaced apart from one another along the circumferential direction C and are positioned opposite the mounting openings 160 of the inner wall 120 along the radial direction R. Further, in some embodiments, each vane of the plurality of first vanes and the plurality of second vanes has a positioning end 172 opposite the mounting flange 174 of the vane, and wherein the positioning end 172 of each vane of the plurality of first vanes is received within one of the positioning recesses 166 defined by the inner wall 120, and wherein the positioning end 172 of each vane of the plurality of second vanes is received within one of the positioning recesses 156 defined by the outer wall 102. In some embodiments, the mounting openings 160 of the inner wall 120 are positioned intermediate or substantially intermediate between the mounting openings 150 of the outer wall 102 along the circumferential direction C.

[0173] In some embodiments, the outer wall 102, the inner wall 120, the first vanes 68A, and the second vanes 68B are each formed of a CMC material. The turbine engine in which the nozzle assembly is positioned can be a gas turbine engine, such as an aircraft gas turbine engine.

[0174] ​In some example embodiments, a turbine engine is provided. The turbine engine defines an axial direction, a radial direction, a circumferential direction, and an axial centerline extending along the axial direction. The turbine engine includes a combustion section and a turbine section positioned downstream of the combustion section. The turbine engine has an inner wall defining a mounting opening. The turbine engine also has an outer wall having a combustor portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section. The combustor portion and the turbine section are integrally formed as a single monolithic structure. The turbine portion of the outer wall defines the mounting opening. The turbine engine also has a first vane having an airfoil and a mounting flange. The mounting flange of the first vane is at least partially positioned in the mounting opening defined by the outer wall. The turbine engine also has a second vane positioned adjacent the first vane along the circumferential direction. The second vane has an airfoil and a mounting flange. The mounting flange of the second vane is at least partially positioned in the mounting opening of the inner wall. In some embodiments, the outer wall, the inner wall, the first vane, and the second vane are formed of a CMC material.

[0175] In other embodiments, the inner wall defines a positioning recess positioned opposite the mounting opening of the outer wall along the radial direction, and the outer wall defines a positioning recess positioned opposite the mounting opening of the inner wall along the radial direction. In such embodiments, the first vane and the second vane can each extend along the radial direction between a mounting end and a positioning end. The positioning end of the first vane is positioned in the positioning recess of the inner wall, and the positioning end of the second vane is positioned in the positioning recess of the outer wall. In some embodiments, the mounting opening of the inner wall and the mounting opening of the outer wall are aligned along the axial direction.

[0176] A method for assembling a nozzle assembly for a turbine engine defining an axial direction, a radial direction, and a circumferential direction will now be provided. For example, Figure 14 and Figure 15 The nozzle assembly of can be assembled according to the following example method. The method includes inserting a first vane inwardly through an outer wall of the nozzle assembly along the radial direction such that a mounting flange of the first vane engages the outer wall and is at least partially positioned within a mounting opening defined by the outer wall. The method also includes, adjacent the first vane along the circumferential direction, inserting a second vane outwardly through an inner wall spaced apart from the outer wall along the radial direction such that a mounting flange of the second vane engages the inner wall and is at least partially positioned in a mounting opening defined by the inner wall.

[0177] In some embodiments, the method further includes inserting a third vane adjacent the first vane and opposite the second vane in the circumferential direction, inward through the outer wall of the nozzle assembly in the radial direction, such that the mounting flange of the third vane engages the outer wall and is at least partially positioned within the second mounting opening defined by the outer wall. In other embodiments, the method further includes inserting a third vane adjacent the second vane and opposite the first vane in the circumferential direction, outward through the inner wall of the nozzle assembly in the radial direction, such that the mounting flange of the third vane engages the inner wall and is at least partially positioned within the second mounting opening defined by the inner wall.

[0178] In yet other embodiments, the first vane is one of a plurality of first vanes and the second vane is one of a plurality of second vanes, and wherein the method further includes: inserting the plurality of first vanes inward through the outer wall of the nozzle assembly in the radial direction, such that the mounting flange of each of the plurality of first vanes engages the outer wall and is at least partially positioned within a respective mounting opening defined by the outer wall; and inserting the plurality of second vanes outward through the inner wall of the nozzle assembly in the radial direction, such that the mounting flange of each of the plurality of second vanes engages the inner wall and is at least partially positioned within a respective mounting opening defined by the inner wall. In some embodiments, the plurality of first vanes inserted inward through the outer wall in the radial direction alternate in the circumferential direction with the plurality of second vanes inserted outward through the inner wall in the radial direction.

[0179] In some embodiments, the plurality of vanes inserted inward through the outer wall in the radial direction extend between the outer wall and the inner wall in the radial direction, and the plurality of vanes inserted outward through the inner wall in the radial direction extend between the outer wall and the inner wall in the radial direction.

[0180] In some further embodiments, each of the plurality of vanes inserted inward through the outer wall of the nozzle assembly in the radial direction extends between a positioning end and a mounting end, and wherein the positioning end of each vane inserted inward through the outer wall in the radial direction is positioned in one of a plurality of positioning recesses defined by the inner wall, and wherein each of the plurality of vanes inserted outward through the inner wall of the nozzle assembly in the radial direction extends between a positioning end and a mounting end, and wherein the positioning end of each vane inserted outward through the inner wall in the radial direction is positioned in one of a plurality of positioning recesses defined by the outer wall.

[0181] In some embodiments, each of the plurality of vanes inserted inward through the outer wall of the nozzle assembly in the radial direction is wedged into a mounting opening defined by the outer wall. In other embodiments, each of the plurality of vanes inserted outward through the inner wall of the nozzle assembly in the radial direction is wedged into a mounting opening defined by the inner wall.

[0182] Notably, Figure 14 and Figure 15The embodiments disclosed in the detailed description can include one or more embodiments disclosed in Figure 7 , Figure 8 , Figure 9 , Figure 10 , Figure 11 and Figure 13 .

[0183] As previously mentioned, the outer wall 102, inner wall 120, and combustor dome 118, and in some embodiments, the vanes 68, 69, are formed of a CMC material. Examples of CMC materials, and in particular, SiC / Si-SiC (fiber / matrix) continuous fiber reinforced ceramic composite (CFCC) materials and processes are described in U.S. Patents Nos. 5,015,540; 5,330,854; 5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441 and U.S. Patent Application Publication No. 2004 / 0067316. Such processes generally require the use of multiple pre-impregnated (pre-preg) layers to fabricate the CMC, for example, the ply material can comprise a pre-preg material composed of ceramic fibers, woven or braided ceramic fiber cloth, or stacked ceramic fiber tows that have been impregnated with a matrix material. In some embodiments, each pre-preg layer is in the form of a "tape" that includes a desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and an organic resin binder. The pre-preg tape can be formed by impregnating the reinforcement material with a slurry that includes the ceramic precursors and the binder. The preferred materials for the precursors will depend on the particular composition of the ceramic matrix of the CMC component that is desired, for example, SiC powder and / or one or more carbon-containing materials, if the desired matrix material is SiC. Notable carbon-containing materials include carbon black, phenolic resin, and furan resin, including furfuryl alcohol (C4H3OCH2OH). Other typical slurry ingredients include an organic binder (e.g., polyvinyl butyral (PVB)) that promotes the flexibility of the pre-preg tape and a solvent for the binder (e.g., toluene and / or methyl isobutyl ketone (MIBK)) that can promote the flowability of the slurry, enabling impregnation of the fiber reinforcement material. The slurry can further contain one or more particulate fillers that are intended to be present in the ceramic matrix of the CMC component, for example, silicon and / or SiC powder, in the case of a Si-SiC matrix. As previously mentioned, chopped fibers or whiskers or other materials can also be embedded within the matrix. Other ingredients and methods for producing composite articles can also be used, and more specifically, other slurries and pre-preg tape ingredients, such as the processes and ingredients described in U.S. Patent Application Publication No. 2013 / 0157037.

[0184] The resulting prepreg tape can be laminated with other tapes such that the CMC component formed from the tapes includes a plurality of plies, each ply derived from an individual prepreg tape. Each ply includes a ceramic fiber reinforcement material encased in a ceramic matrix formed, in whole or in part, by the transformation of ceramic matrix precursors during a sintering and densification cycle, for example, as more fully described below. In some embodiments, the reinforcement material is in the form of unidirectional arrays of tows, each tow including continuous fibers or filaments. Alternative arrangements of unidirectional arrays of tows can also be used. Further, the appropriate fiber diameter, tow diameter, and center-to-center tow spacing will depend on the particular application, the particular ply, and the thickness of the tape from which the ply is formed, among other factors. As noted above, other prepreg or non-prepreg materials can also be used.

[0185] After the tapes or plies are laminated to form a layup, the layup is thinned and, if appropriate, cured while being subjected to elevated pressure and temperature to produce a preform. The preform is then heated (fired) in a vacuum or inert atmosphere to decompose the binder, remove solvents, and transform the precursors into the desired ceramic matrix material. As a result of the decomposition of the binder, a porous CMC body results, which can be subjected to densification, such as melt infiltration (MI), to fill the pores and produce a CMC component. The specific processing techniques and parameters used for the above-described process will depend on the specific composition of the material. For example, silicon CMC components can be formed from a fibrous material infiltrated with molten silicon, such as by a process commonly referred to as the Silcomp process. Another technique for manufacturing CMC components is a method referred to as the slurry cast melt infiltration (MI) process. In one method of manufacturing using the slurry cast MI process, a CMC is produced by first providing plies of a balanced two-dimensional (2D) woven cloth including silicon carbide (SiC) containing fibers, the woven cloth having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers extending in both directions of the fabric. The term "silicon carbide containing fibers" means fibers having a composition that includes silicon carbide, and preferably is substantially silicon carbide. For example, the fibers can have a silicon carbide core surrounded by carbon, or conversely, the fibers can have a carbon core surrounded or encapsulated by silicon carbide.

[0186] Other techniques for forming CMC components include polymer infiltration and pyrolysis (PIP) and oxide / oxide processes. In a PIP process, silicon carbide fiber preforms are infiltrated with a preceramic polymer, such as polysilazane, and then heat treated to form a SiC matrix. In an oxide / oxide process, aluminum or aluminosilicate fibers can be pre-impregnated and then laminated into a preselected geometry. Components can also be made from carbon fiber reinforced silicon carbide matrix (C / SiC) CMC. The C / SiC process includes carbon fiber preforms placed in a tool in a preselected geometry. The tool is made of a graphite material, as used in the slurry cast method for SiC / SiC. The fiber preforms are supported by the tool during a chemical vapor infiltration process at about 1200°C, thereby forming a C / SiC CMC component. In other embodiments, 2D, 2.5D, and / or 3D preforms can be used in the MI, CVI, PIP, or other processes. For example, cut layers of 2D woven fabric can be stacked in alternating weave directions as described above, or the filaments can be wound or woven and combined with 3D braiding, stitching, or needling to form a 2.5D or 3D preform having a multi-axial fiber architecture. Other ways of forming 2.5D or 3D preforms can also be used, for example, using other weaving or braiding methods or utilizing 2D fabrics.

[0187] Accordingly, various processes can be used to form the integral structures, such as the outer wall 102, as an integral CMC component. More specifically, multiple plies of CMC material can be used to form each integral structure. The multiple plies can be interspersed with one another to integrally form various portions of the integral structure. For example, the integral outer wall 102 can be made from a plurality of outer liner plies, a plurality of first turbine stage outer band plies, a plurality of first turbine stage shroud plies, a plurality of second turbine stage outer band plies, and a plurality of second turbine stage shroud plies. At the point where the outer liner plies meet the first turbine stage outer band plies, the ends of the outer liner plies can alternate with the ends of the outer band plies to integrally form the plies used to form the outer liner portion with the plies used to form the first turbine stage outer band portion of the integral outer wall 102. That is, any joint between plies forming the integral outer wall 102 can be formed by alternating the plies on one side of the joint with the plies on the other side of the joint. In other embodiments, the CMC plies can also be layered in other ways to form the integral structure. Additionally, layering the multiple CMC plies can include defining features of the integral structure or other components, such as the inner liner portion 122 when not integrally formed with the inner band portion 124 to form the integral inner wall 120 or the separate combustor dome 118, such as the openings 142 in the combustor forward end 88.

[0188] After the plurality of CMC plies are laminated to define the monolithic CMC component preform, the preform is cured to produce a single-piece monolithic CMC component, which is then fired and subjected to densification, such as silicon melt infiltration, to form the final monolithic CMC structure. Continuing with the example of the outer wall 102 described above, the outer wall preform can be processed in an autoclave to produce a green-state monolithic outer wall 102. The green-state monolithic outer wall 102 can then be placed in a furnace to burn off excess binder and the like, and then placed in a furnace with a silicon piece or silicon plate and fired to at least melt infiltrate the monolithic outer wall 102 with silicon. More specifically, for a monolithic outer wall 102 formed from CMC plies produced from prepreg tapes as described above, heating (i.e., firing) the green-state component in a vacuum or inert atmosphere decomposes the binder, removes solvents, and converts the precursors to the desired ceramic matrix material. Decomposition of the binder results in a porous CMC body; the body can undergo densification, such as melt infiltration (MI), to fill the pores. In the foregoing example of firing the green-state monolithic outer wall 102 with silicon, the outer wall 102 undergoes silicon melt infiltration. However, densification can be performed using any known densification technique, including but not limited to Silcomp, melt infiltration (MI), chemical vapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), and oxide / oxidation processes, and with any suitable material, including but not limited to silicon. In one embodiment, densification and firing can be performed in a vacuum furnace or inert atmosphere with an established atmosphere at a temperature above 1200 °C to allow silicon or other suitable material or combination of materials to melt infiltrate into the component. The densified CMC body hardens into the final monolithic CMC outer wall 102. In some embodiments, optionally, the final monolithic structure can be finished, such as to bring the structure within tolerance or to define the openings 142 in the forward end 88, and / or an environmental barrier coating (EBC) can be applied to the monolithic structure, such as to protect the monolithic structure from hot combustion gases 66. It will be recognized that other methods or processes of forming CMC components can also be used, such as the monolithic CMC outer wall 102, monolithic CMC inner wall 120, etc.

[0189] Additionally or alternatively, other processes for producing monolithic components can be used to form the monolithic outer wall 102 and / or the monolithic inner wall 120, and the monolithic structure can be formed from other materials. In some embodiments, an additive manufacturing process can be used to form the monolithic outer wall 102 and / or the monolithic inner wall 120. For example, an additive process, such as fused deposition modeling (FDM), selective laser sintering (SLS), stereolithography (SLA), digital light processing (DLP), direct metal laser sintering (DMLS), laser net shape manufacturing (LNSM), electron beam sintering, or other known processes can be used to produce the monolithic outer wall 102 and / or the monolithic inner wall 120. Generally, an additive process uses three-dimensional information (e.g., a three-dimensional computer model) of a component to manufacture the component. The three-dimensional information is converted into a plurality of slices, each slice defining a cross-section of the component for a predetermined height of the slice. The component is then "built up" slice-by-slice or layer-by-layer until completion. A superalloy metal material or other suitable material can be used in the additive process to form the monolithic outer wall 102 and / or the monolithic inner wall 120. In other embodiments, a forging or casting process can be used to form the monolithic outer wall 102 and / or the monolithic wall 120. Other suitable processes or methods can also be used.

[0190] This written description uses examples to disclose the application, including the best mode, and also to enable any person skilled in the art to practice the application, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the application is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural

[0191] Other aspects of the application are provided by the subject matter of the following clauses:

[0192] 1. A nozzle assembly for a turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending in the axial direction, the nozzle assembly comprising: an outer wall defining a mounting opening; an inner wall defining a mounting opening, the inner wall and the outer wall defining a flowpath; a first vane having an airfoil and a mounting flange, wherein the airfoil of the first vane extends through the mounting opening of the outer wall and is at least partially positioned within the flowpath, and wherein the mounting flange of the first vane is positioned radially outward of the outer wall relative to the axial centerline; a second vane positioned adjacent the first vane in the circumferential direction and having an airfoil and a mounting flange, wherein the airfoil of the second vane extends through the mounting opening of the inner wall and is at least partially positioned within the flowpath, and wherein the mounting flange of the second vane is positioned radially inward of the inner wall relative to the axial centerline.

[0193] 2. The nozzle assembly of any preceding clause, wherein the mounting opening of the outer wall is one of a plurality of mounting openings, and the mounting opening of the inner wall is one of a plurality of mounting openings, and wherein the first vane is one of a plurality of first vanes, and the second vane is one of a plurality of second vanes, and wherein the airfoil of each of the plurality of first vanes extends through one of the plurality of mounting openings of the outer wall and is at least partially positioned within the flowpath, and wherein the mounting flange of each of the plurality of first vanes is positioned radially outward of the outer wall relative to the axial centerline, and wherein the airfoil of each of the plurality of second vanes extends through one of the plurality of mounting openings of the inner wall and is at least partially positioned within the flowpath, and wherein the mounting flange of each of the plurality of second vanes is positioned radially inward of the inner wall relative to the axial centerline.

[0194] 3. The nozzle assembly of any preceding clause, wherein at least two pairs of adjacent vanes of the nozzle assembly include one of the plurality of first vanes and one of the plurality of second vanes.

[0195] 4. The nozzle assembly of any preceding clause, wherein the plurality of first vanes and the plurality of second vanes are arranged in an alternating fashion along the circumferential direction.

[0196] 5. The nozzle assembly of any preceding clause, wherein the inner wall defines positioning recesses spaced apart from one another along the circumferential direction and positioned radially opposite the mounting openings of the outer wall, and the outer wall defines positioning recesses spaced apart from one another along the circumferential direction and positioned radially opposite the mounting openings of the inner wall, and wherein each vane of the plurality of first vanes and the plurality of second vanes has a positioning end opposite the mounting flange of the vane, and wherein the positioning end of each vane of the plurality of first vanes is received within one of the positioning recesses defined by the inner wall, and wherein the positioning end of each vane of the plurality of second vanes is received within one of the positioning recesses defined by the outer wall.

[0197] 6. The nozzle assembly of any preceding clause, wherein the mounting opening of the inner wall is positioned intermediate the mounting openings of the outer wall along the circumferential direction.

[0198] 7. The nozzle assembly of any preceding clause, wherein the mounting opening of the outer wall is sized to accommodate the airfoil of the first vane and shaped to complement a radial cross-section of the airfoil of the first vane, and the mounting opening of the inner wall is sized to accommodate the airfoil of the second vane and shaped to complement a radial cross-section of the second vane.

[0199] 8. The nozzle assembly of any preceding clause, wherein the outer wall has an inner surface and an outer surface spaced apart from the inner surface in the radial direction, and the mounting flange of the first vane engages the outer surface of the outer wall, and wherein the inner wall has an inner surface and an outer surface spaced apart from the inner surface in the radial direction, and the mounting flange of the second vane engages the inner surface of the inner wall.

[0200] 9. The nozzle assembly of any preceding clause, wherein the first vane has a seal portion connecting the airfoil and the mounting flange, and the second vane has a seal portion connecting the airfoil and the mounting flange, and wherein the seal portion of the first vane wedges the first vane into engagement with the outer wall, and the seal portion of the second vane wedges the second vane into engagement with the inner wall.

[0201] 10. The nozzle assembly of any preceding clause, wherein the nozzle assembly defines an outer diameter and an inner diameter, and wherein the cooling fluid is delivered to the first vane from the outer diameter of the nozzle assembly, and wherein the cooling fluid is delivered to the second vane from the inner diameter of the nozzle assembly.

[0202] 11. A turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending in the axial direction, the turbine engine comprising: a combustion section; a turbine section positioned downstream of the combustion section; an inner wall defining a mounting opening; an outer wall comprising a combustor portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section, the combustor portion and the turbine portion being integrally formed as a single unitary structure, wherein the turbine portion of the outer wall defines the mounting opening; a first vane having an airfoil and a mounting flange, wherein the airfoil of the first vane extends through the mounting opening of the outer wall, and wherein the mounting flange of the first vane is positioned radially outward of the outer wall relative to the axial centerline; and a second vane positioned adjacent the first vane in the circumferential direction and having an airfoil and a mounting flange, wherein the airfoil of the second vane extends through the mounting opening of the inner wall, and wherein the mounting flange of the second vane is positioned radially inward of the inner wall relative to the axial centerline.

[0203] 12. The turbine engine of any preceding clause, wherein the outer wall, the inner wall, the first vane, and the second vane are formed of a ceramic matrix composite (CMC) material.

[0204] 13. The turbine engine of any preceding clause, wherein the inner wall defines a positioning groove positioned radially opposite the mounting opening of the outer wall, and the outer wall defines a positioning groove positioned radially opposite the mounting opening of the inner wall.

[0205] 14. The turbine engine of any preceding clause, wherein the first vane and the second vane each extend in the radial direction between a mounting end and a positioning end, and wherein the positioning end of the first vane is positioned in the positioning groove of the inner wall and the positioning end of the second vane is positioned in the positioning groove of the outer wall.

[0206] 15. The turbine engine of any preceding clause, wherein the mounting opening of the inner wall and the mounting opening of the outer wall are aligned in the axial direction.

[0207] 16. A method for assembling a nozzle assembly of a turbine engine for defining an axial direction, a radial direction, and a circumferential direction, the method comprising: inserting a first vane inward through an outer wall of the nozzle assembly in the radial direction and adjacent the first vane in the circumferential direction, inserting a second vane outward through an inner wall of the nozzle in the radial direction, the inner wall being spaced apart from the outer wall in the radial direction.

[0208] 17. The method of any preceding clause, further comprising: inserting a third vane inward through the outer wall of the nozzle assembly in the radial direction adjacent the first vane in the circumferential direction and opposite the second vane.

[0209] 18. The method of any preceding clause, further comprising: inserting a third vane outward through the inner wall of the nozzle assembly in the radial direction adjacent the second vane in the circumferential direction and opposite the first vane.

[0210] 19. The method of any preceding clause, wherein the first vane is one of a plurality of first vanes and the second vane is one of a plurality of second vanes, and wherein the method further comprises: inserting the plurality of first vanes inward through the outer wall of the nozzle assembly in the radial direction; and inserting the plurality of second vanes outward through the inner wall of the nozzle assembly in the radial direction, wherein the plurality of first vanes inserted inward through the outer wall in the radial direction alternate in the circumferential direction with the plurality of second vanes inserted outward through the inner wall in the radial direction.

[0211] 20. The method of any preceding clause, wherein the plurality of vanes inserted inward through the outer wall in the radial direction extend in the radial direction between the outer wall and the inner wall, and the plurality of vanes inserted outward through the inner wall in the radial direction extend in the radial direction between the outer wall and the inner wall.

[0212] 21. The method of any preceding paragraph, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted outwardly through the inner wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the outer wall, and wherein each of the plurality of vanes inserted outwardly through the inner wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted outwardly through the inner wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the outer wall.

[0213] 22. The method of any preceding paragraph, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction has a sealing portion that wedges into a mounting opening defined by the outer wall, and wherein each of the mounting openings has an insertion end and a flowpath end spaced apart from the insertion end in the radial direction, the insertion end positioned radially outward of the flowpath end, and wherein the insertion end of each of the mounting openings has a greater radial cross-sectional area than the flowpath end.

[0214] 23. A nozzle assembly for a turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending in the axial direction, the nozzle assembly comprising: an outer wall defining a mounting opening; an inner wall defining the mounting opening, the inner wall and the outer wall defining a flowpath; a first vane having an airfoil and a mounting flange, wherein the airfoil of the first vane is positioned at least partially within the flowpath and the mounting flange of the first vane is positioned at least partially within the mounting opening of the outer wall; a second vane positioned adjacent the first vane in the circumferential direction and having an airfoil and a mounting flange, wherein the airfoil of the second vane is positioned at least partially within the flowpath and the mounting flange of the second vane is positioned at least partially within the mounting opening of the inner wall.

[0215] 24. The nozzle assembly of any preceding paragraph, wherein the mounting flange of the first vane is positioned completely within the mounting opening of the outer wall.

[0216] 25. The nozzle assembly of any preceding paragraph, wherein the mounting flange of the second vane is positioned completely within the mounting opening of the inner wall.

[0217] 26. The nozzle assembly of any preceding paragraph, wherein the outer wall extends in the radial direction between an outer surface and an inner surface, and wherein the mounting flange of the first vane extends between the outer surface and the inner surface of the outer wall.

[0218] 27. The nozzle assembly of any preceding paragraph, wherein the inner wall extends in the radial direction between an outer surface and an inner surface, and wherein the mounting flange of the second vane extends between the outer surface and the inner surface of the inner wall.

[0219] 28. The nozzle assembly of any preceding clause, wherein the side wall of the mounting flange of the first vane wedges against the outer wall.

[0220] 29. The nozzle assembly of any preceding clause, wherein the side wall of the mounting flange of the second vane wedges against the inner wall.

[0221] 30. The nozzle assembly of any preceding clause, wherein the outer wall has an outer surface and an inner surface spaced apart from one another in the radial direction, and wherein the outer wall defines an outer perimeter of the mounting opening at the outer surface and an inner perimeter of the mounting opening at the inner surface, and wherein the outer perimeter of the mounting opening of the outer wall is greater than the inner perimeter.

[0222] 31. The nozzle assembly of any preceding clause, wherein the inner wall has an outer surface and an inner surface spaced apart from one another in the radial direction, and wherein the inner wall defines an outer perimeter of the mounting opening at the outer surface and an inner perimeter of the mounting opening at the inner surface, and wherein the inner perimeter of the mounting opening of the inner wall is greater than the outer perimeter.

[0223] 32. The nozzle assembly of any preceding clause, wherein the mounting opening of the outer wall is one of a plurality of mounting openings of the outer wall, and the mounting opening of the inner wall is one of a plurality of mounting openings of the inner wall.

[0224] 33. The nozzle assembly of any preceding clause, wherein the first vane is one of a plurality of first vanes, and the second vane is one of a plurality of second vanes.

[0225] 34. The nozzle assembly of any preceding clause, wherein the airfoil of each first vane of the plurality of first vanes is positioned at least partially within the flowpath, and the mounting flange of each first vane of the plurality of first vanes is positioned at least partially in one of the plurality of mounting openings of the outer wall.

[0226] 35. The nozzle assembly of any preceding clause, wherein the airfoil of each second vane of the plurality of second vanes is positioned at least partially within the flowpath, and the mounting flange of each second vane of the plurality of second vanes is positioned at least partially in one of the plurality of mounting openings of the inner wall.

[0227] 36. The nozzle assembly of any preceding clause, wherein the plurality of first vanes and the plurality of second vanes are arranged in an alternating manner in the circumferential direction.

[0228] 37. The nozzle assembly of any preceding clause, wherein the inner wall defines positioning recesses spaced apart from one another in the circumferential direction and positioned opposite the mounting openings of the outer wall in the radial direction, and the outer wall defines positioning recesses spaced apart from one another in the circumferential direction and positioned opposite the mounting openings of the inner wall in the radial direction.

[0229] 38. The nozzle assembly of any preceding clause, wherein each vane of the first plurality of vanes and the second plurality of vanes has a locating end opposite the mounting flange of the vane, and wherein the locating end of each vane of the first plurality of vanes is received within one of the locating grooves defined by the inner wall, and wherein the locating end of each vane of the second plurality of vanes is received within one of the locating grooves defined by the outer wall.

[0230] 39. The nozzle assembly of any preceding clause, wherein the mounting openings of the inner wall are positioned intermediate between the mounting openings of the outer wall in the circumferential direction.

[0231] 40. The nozzle assembly of any preceding clause, wherein the outer wall, the inner wall, the first vanes, and the second vanes are formed of a ceramic matrix composite (CMC) material.

[0232] 41. The nozzle assembly of any preceding clause, wherein the turbine engine is a gas turbine engine.

[0233] 42. The nozzle assembly of any preceding clause, wherein the gas turbine engine is an aero gas turbine engine.

[0234] 43. A turbine engine defining an axial direction, a radial direction, a circumferential direction, and an axial centerline extending in the axial direction, the turbine engine comprising: a combustion section; a turbine section positioned downstream of the combustion section; an inner wall defining a mounting opening; an outer wall comprising a combustor portion extending through the combustion section and a turbine portion extending through at least a portion of the turbine section, the combustor portion and the turbine portion being integrally formed as a single unitary structure, wherein the turbine portion of the outer wall defines the mounting opening; a first vane having an airfoil and a mounting flange, wherein the mounting flange of the first vane is positioned at least partially in the mounting opening defined by the outer wall; and a second vane positioned adjacent the first vane in the circumferential direction and having an airfoil and a mounting flange, wherein the mounting flange of the second vane is positioned at least partially in the mounting opening of the inner wall.

[0235] 44. The turbine engine of any preceding clause, wherein the outer wall, the inner wall, the first vanes, and the second vanes are formed of a ceramic matrix composite (CMC) material.

[0236] 45. The turbine engine of any preceding clause, wherein the inner wall defines a locating groove positioned opposite the mounting opening of the outer wall in the radial direction, and the outer wall defines a locating groove positioned opposite the mounting opening of the inner wall in the radial direction.

[0237] 46. The turbine engine of any preceding clause, wherein the first vane and the second vane each extend in the radial direction between a mounting end and a positioned end, and wherein the positioned end of the first vane is positioned in a positioned recess of the inner wall and the positioned end of the second vane is positioned in a positioned recess of the outer wall.

[0238] 47. The turbine engine of any preceding clause, wherein the mounting opening of the inner wall and the mounting opening of the outer wall are aligned in the axial direction.

[0239] 48. A method for assembling a nozzle assembly for a turbine engine, the turbine engine defining an axial direction, a radial direction, and a circumferential direction, the method comprising: inserting a first vane inward through an outer wall of the nozzle assembly in the radial direction such that a mounting flange of the first vane engages the outer wall and is at least partially positioned within a mounting opening defined by the outer wall; and adjacent the first vane in the circumferential direction, inserting a second vane outward through an inner wall spaced apart from the outer wall in the radial direction such that a mounting flange of the second vane engages the inner wall and is at least partially positioned in a mounting opening defined by the inner wall.

[0240] 49. The method of any preceding clause, further comprising: adjacent the first vane in the circumferential direction and opposite the second vane, inserting a third vane inward through the outer wall of the nozzle assembly in the radial direction such that a mounting flange of the third vane engages the outer wall and is at least partially positioned within a second mounting opening defined by the outer wall.

[0241] 50. The method of any preceding clause, further comprising: adjacent the second vane in the circumferential direction and opposite the first vane, inserting a third vane outward through the inner wall of the nozzle assembly in the radial direction such that a mounting flange of the third vane engages the inner wall and is at least partially positioned in a second mounting opening defined by the inner wall.

[0242] 51. The method of any preceding clause, wherein the first vane is one of a plurality of first vanes and the second vane is one of a plurality of second vanes, and wherein the method further comprises: inserting the plurality of first vanes inward through the outer wall of the nozzle assembly in the radial direction such that the mounting flange of each of the plurality of first vanes engages the outer wall and is at least partially positioned within a respective mounting opening defined by the outer wall; and inserting the plurality of second vanes outward through the inner wall of the nozzle assembly in the radial direction such that the mounting flange of each of the plurality of second vanes engages the inner wall and is at least partially positioned within a respective mounting opening defined by the inner wall.

[0243] 52. The method of any preceding clause, wherein the plurality of first vanes inserted inward through the outer wall in the radial direction alternate in the circumferential direction with the plurality of second vanes inserted outward through the inner wall in the radial direction.

[0244] 53. The method of any preceding clause, wherein the plurality of vanes inserted inwardly through the outer wall in the radial direction extend between the outer wall and the inner wall in the radial direction, and the plurality of vanes inserted outwardly through the inner wall in the radial direction extend between the outer wall and the inner wall in the radial direction.

[0245] 54. The method of any preceding clause, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted inwardly through the outer wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the inner wall, and wherein each of the plurality of vanes inserted outwardly through the inner wall of the nozzle assembly in the radial direction extends between a locating end and a mounting end, and wherein the locating end of each vane inserted outwardly through the inner wall of the nozzle assembly in the radial direction is positioned in one of a plurality of locating recesses defined by the outer wall.

[0246] 55. The method of any preceding clause, wherein each of the plurality of vanes inserted inwardly through the outer wall of the nozzle assembly in the radial direction is wedged into a mounting opening defined by the outer wall.

[0247] 56. The method of any preceding clause, wherein each of the plurality of vanes inserted outwardly through the inner wall of the nozzle assembly in the radial direction is wedged into a mounting opening defined by the inner wall.

Claims

1. A nozzle assembly for a turbine engine, comprising: a first wall defining a first mounting opening; a second wall spaced from the first wall, the first wall and the second wall defining a flowpath, the second wall defining a second mounting opening; a first vane positioned at least partially within the flowpath and at least partially within the first mounting opening, the first vane having a first mounting flange that is larger than the first mounting opening to secure the first vane to the first wall, the first vane further having a first retention feature that secures the first vane to the second wall; and a second vane positioned at least partially within the flowpath and at least partially within the second mounting opening, the second vane having a second mounting flange that is larger than the second mounting opening to secure the second vane to the second wall, the second vane further having a second retention feature that secures the second vane to the first wall. the first wall is spaced from the second wall along a radial direction defined by the turbine engine, and wherein the first wall is positioned interior to the second wall along the radial direction relative to an axial centerline of the turbine engine.

2. The nozzle assembly of claim 1, wherein, the first wall is spaced from the second wall along a radial direction defined by the turbine engine, and wherein the first wall is positioned exterior to the second wall along the radial direction relative to an axial centerline of the turbine engine.

3. The nozzle assembly of claim 1, wherein, the second wall defines a first through opening through which the first vane extends, the first vane being secured exterior to the flowpath by the first retention feature.

4. The nozzle assembly of claim 1, wherein, the second wall has a first surface facing the flowpath and a second surface facing away from the flowpath, and wherein a radial space is defined between the first retention feature and the second surface of the second wall.

5. The nozzle assembly of claim 1, wherein, the first retention feature and / or the second retention feature is one of an annular clip, a retaining pin, and a nut.

6. The nozzle assembly of claim 1, wherein, the first retention feature and / or the second retention feature is one or more plies.

7. The nozzle assembly of claim 1, wherein, the one or more plies are laminated and secured exterior to the first vane and / or the second vane.

8. The nozzle assembly of claim 7, wherein, the first retention feature and / or the second retention feature is a bonding material.

9. The nozzle assembly of claim 1, wherein, the bonding material is a braze-type bond.

10. The nozzle assembly of claim 9, wherein, the first vane defines a locating end recess, and wherein the first retention feature is a locating barb ring that is received within the locating end recess and compressed between the first vane and the second wall.

11. The nozzle assembly of claim 1, wherein, the first vane defines a mounting end recess, and wherein a mounting barb ring is received within the mounting end recess and compressed between the first vane and the first wall.

12. The nozzle assembly of claim 1, wherein, the second vane defines a locating end recess, and wherein the second retention feature is a locating barb ring that is received within the locating end recess and compressed between the second vane and the first wall.

13. The nozzle assembly of claim 1, wherein, the second vane defines a mounting end recess, and wherein a mounting barb ring is received within the mounting end recess and compressed between the second vane and the second wall.

14. The nozzle assembly of claim 1, wherein, ​ 15. The nozzle assembly of claim 1, wherein, The first wall defines a second through opening through which the second vane extends, the second vane being secured by the second retention feature outside of the flowpath.

16. A nozzle assembly for a turbine engine, comprising: a first wall defining a first mounting opening; a second wall spaced from the first wall, the first wall and the second wall defining a flowpath, the second wall defining a second mounting opening; a first vane positioned at least partially within the flowpath and extending between a first end and a second end, the first end of the first vane having a first mounting flange positioned within the first mounting opening and wedgingly engaging the first wall, the second end of the first vane being secured to the second wall; and a second vane positioned at least partially within the flowpath and extending between a first end and a second end of the second vane, the first end of the second vane having a second mounting flange positioned within the second mounting opening and wedgingly engaging the second wall, the second end of the second vane being secured to the first wall.

17. The nozzle assembly of claim 16, wherein, The first mounting flange of the first vane is positioned entirely within the first mounting opening of the first wall.

18. The nozzle assembly of claim 16, wherein, The first wall has a first surface and a second surface, and wherein the first mounting flange extends between the first surface and the second surface.

19. The nozzle assembly of claim 16, wherein, The first wall has a radial thickness and the first mounting flange has a radial thickness, and wherein the radial thickness of the first mounting flange is less than the radial thickness of the first wall.

20. The nozzle assembly of claim 16, wherein, The first wall has a mounting seat, and wherein the first mounting flange is seated on the mounting seat.

21. A method for assembling a nozzle assembly for a turbine engine, the turbine engine defining a radial direction, the method comprising: inserting a first vane inwardly along the radial direction such that a locating end of the first vane is inserted through a locating opening defined by an inner wall of the nozzle assembly and such that a mounting flange of the first vane engages an outer wall of the nozzle assembly, the inner wall being spaced from the outer wall along the radial direction; and inserting a second vane outwardly along the radial direction such that a locating end of the second vane is inserted through a locating opening defined by the outer wall and such that a mounting flange of the second vane engages the inner wall of the nozzle assembly.

Images