Combustor assembly for turbine engine

Abstract

A combustor assembly for a gas turbine engine defining a radial direction and a circumferential direction includes a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one bushing defines a radial opening and an interface surface extending in a circumferential direction and in a radial direction; and a sealing member including a body defining a body surface extending in a radial direction and positioned adjacent to the interface surface of the at least one bushing, a flange extending forward from the body, and a radial element coupled to the flange and extending into a radial opening defined by the at least one bushing.

Description

[0001] FEDERALLY SPONSORED RESEARCH This invention was made with government support. The United States Government can have certain rights in the invention. TECHNICAL FIELD

[0002] The present subject matter relates generally to a combustor assembly for a gas turbine engine, or more particularly to a seal assembly for a combustor assembly and a method for manufacturing the same. BACKGROUND

[0003] A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the 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 burnt to provide combustion gases. The combustion gases are passed from the combustion section to the turbine section. The combustion gas flow through the turbine section drives the turbine section and is then passed through the exhaust section, e.g., to the atmosphere.

[0004] More generally, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are used as components within a gas turbine engine. For example, given the ability of CMC materials to withstand relatively extreme temperatures, particular attention has been directed to replacing components within the combustion section of a gas turbine engine with CMC materials. More particularly, inner and outer liners within the combustion section of a gas turbine engine are more generally formed of CMC materials.

[0005] In contrast, certain structural components that surround the inner and outer liners, as well as components positioned adjacent to such inner and outer liners, can be formed of metallic materials. However, the different coefficients of thermal expansion between the CMC liners and the metallic components can make it difficult to form a seal between the two components. Accordingly, a simplified assembly for forming a seal between a CMC component and a metallic component would be useful. SUMMARY

[0006] 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.

[0007] In one example embodiment of the present disclosure, a combustor assembly of a gas turbine engine is provided for defining a radial direction and a circumferential direction. The combustor assembly includes a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal component including a body defining a body surface extending along the radial direction and positioned adjacent the interface surface of the at least one liner, a flange extending forward from the body, and a radial element coupled to the flange and extending into the radial opening defined by the at least one liner.

[0008] In certain example embodiments, the at least one liner of the liner assembly is formed of a ceramic matrix composite material, and wherein the seal component is formed of a metallic material.

[0009] In certain example embodiments, the body surface of the body of the seal component is slidably engaged with the interface surface defined by the at least one liner of the liner assembly.

[0010] In certain example embodiments, the body of the seal component defines a downstream sealing surface configured for contacting a corresponding seal of an adjacent component within the gas turbine engine when installed.

[0011] In certain example embodiments, the radial opening is a first radial opening of a plurality of radial openings defined by the at least one liner at the downstream end of the at least one liner, and wherein the plurality of radial openings are spaced along the circumferential direction.

[0012] For example, in certain example embodiments, the radial element of the seal component is a first radial element of a plurality of radial elements of the seal component, and wherein each radial element of the plurality of radial elements is coupled to the flange of the seal component and extends into a respective radial opening of a plurality of radial openings defined by the at least one liner.

[0013] In certain example embodiments, the radial opening defines a length along the radial direction and a constant cross-sectional geometry along the length.

[0014] In certain example embodiments, the at least one liner includes an abradable coating positioned within the radial opening.

[0015] In certain example embodiments, the radial element of the seal component defines a cross-sectional geometry that is substantially equal to a cross-sectional geometry of the radial opening of the at least one liner.

[0016] In certain example embodiments, the seal component extends continuously along the circumferential direction to form a circumferential seal ring.

[0017] In certain example embodiments, the interface surface of the at least one liner extends continuously along the circumferential direction to form a complete loop.

[0018] In certain example embodiments, the at least one liner comprises a plurality of liners spaced along the circumferential direction, and wherein the plurality of liners together define the interface surface.

[0019] In another example embodiment of the present disclosure, a gas turbine engine defining a radial direction and a circumferential direction is provided. The gas turbine engine includes a compressor section, a combustor section, and a turbine section arranged in serial flow order, the combustor section including a combustor assembly including a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal component including a body defining a body surface extending along the radial direction and positioned adjacent to the interface surface of the at least one liner, a flange extending forward from the body, and a radial element coupled to the flange and extending into the radial opening defined by the at least one liner.

[0020] In certain example embodiments, the turbine section includes a first stage of airfoil components, wherein the first stage of airfoil components includes a base defining an upstream end, wherein the upstream end includes a seal plate, and wherein the body of the seal component contacts the seal plate to form a seal with the seal plate.

[0021] For example, in certain example embodiments, the body of the seal component defines a downstream seal surface, and wherein the downstream seal surface contacts the seal plate to form a seal with the seal plate.

[0022] In certain example embodiments, the at least one liner of the liner assembly is formed of a ceramic matrix composite material, and wherein the seal component is formed of a metallic material.

[0023] In another example aspect of the present disclosure, a method for manufacturing a combustor assembly of a gas turbine engine is provided, the combustor assembly including a liner assembly and a seal component, the liner assembly including at least one liner defining an interface surface at a downstream end. The method includes providing the seal component extending continuously along the circumferential direction to form a seal loop; and moving the seal component onto the at least one liner such that a body surface of a body of the seal component contacts the interface surface defined by the at least one liner.

[0024] In certain example aspects, the seal component includes a flange, and wherein moving the seal component onto the at least one liner includes moving the flange of the seal component over a radial opening defined by the at least one liner.

[0025] For example, in certain example aspects, the seal component further includes a radial element, and wherein the method further includes: moving the radial element of the seal component into a radial opening defined by the at least one liner; and coupling the radial element to the flange of the seal component.

[0026] In certain example aspects, the at least one liner of the liner assembly is formed of a ceramic matrix composite material, and wherein the seal component is formed of a metallic material.

[0027] Technical Solution 1. A combustor assembly of a gas turbine engine for defining a radial direction and a circumferential direction, the combustor assembly comprising: a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal component including a body, a flange, and a radial element, the body defining a body surface extending along the radial direction and positioned adjacent the interface surface of the at least one liner, the flange extending forwardly from the body, and the radial element coupled to the flange and extending into the radial opening defined by the at least one liner.

[0028] Technical Solution 2. The combustor assembly of Technical Solution 1, wherein the at least one liner of the liner assembly is formed of a ceramic matrix composite material, and wherein the seal component is formed of a metallic material.

[0029] Technical Solution 3. The combustor assembly of Technical Solution 1, wherein the body surface of the body of the seal component is slidably engaged with the interface surface defined by the at least one liner of the liner assembly.

[0030] Technical Solution 4. The combustor assembly of Technical Solution 1, wherein the body of the seal component defines a downstream sealing surface configured for contacting a corresponding seal of an adjacent component within the gas turbine engine when installed.

[0031] Technical Solution 5. The combustor assembly of Technical Solution 1, wherein the radial opening is a first radial opening of a plurality of radial openings defined by the at least one liner at the downstream end of the at least one liner, and wherein the plurality of radial openings are spaced along the circumferential direction.

[0032] TECHNICAL SOLUTION 6. The combustor assembly of TECHNICAL SOLUTION 5, wherein the radial element of the seal component is a first radial element of a plurality of radial elements of the seal component, and wherein each radial element of the plurality of radial elements is coupled to a flange of the seal component and extends into a respective radial opening of the plurality of radial openings defined by the at least one liner.

[0033] TECHNICAL SOLUTION 7. The combustor assembly of TECHNICAL SOLUTION 1, wherein the radial opening defines a length along the radial direction and a constant cross-sectional geometry along the length.

[0034] TECHNICAL SOLUTION 8. The combustor assembly of TECHNICAL SOLUTION 1, wherein the at least one liner includes a wear coating positioned within the radial opening.

[0035] TECHNICAL SOLUTION 9. The combustor assembly of TECHNICAL SOLUTION 1, wherein the radial element of the seal component defines a cross-sectional geometry that is substantially equal to a cross-sectional geometry of the radial opening of the at least one liner.

[0036] TECHNICAL SOLUTION 10. The combustor assembly of TECHNICAL SOLUTION 1, wherein the seal component extends continuously along the circumferential direction to form a circumferential seal ring.

[0037] TECHNICAL SOLUTION 11. The combustor assembly of TECHNICAL SOLUTION 1, wherein the interface surface of the at least one liner extends continuously along the circumferential direction to form a complete annulus.

[0038] TECHNICAL SOLUTION 12. The combustor assembly of TECHNICAL SOLUTION 1, wherein the at least one liner includes a plurality of liners spaced along the circumferential direction, and wherein the plurality of liners together define the interface surface.

[0039] TECHNICAL SOLUTION 13. A gas turbine engine defining a radial direction and a circumferential direction, the gas turbine engine comprising: a compressor section, a combustor section, and a turbine section arranged in serial flow order, the combustor section including a combustor assembly, the combustor assembly including a liner assembly at least partially defining a combustion chamber and including at least one liner extending between a downstream end and an upstream end, the downstream end of the at least one liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and A seal component comprising a body, a flange, and a radial element, the body defining a body surface extending along the radial direction and positioned adjacent to an interface surface of the at least one bushing, the flange extending forward from the body, and the radial element coupled to the flange and extending into the radial opening defined by the at least one bushing.

[0040] Technical solution 14. The gas turbine engine of technical solution 13, wherein the turbine section includes a first stage of airfoil components, wherein the first stage of airfoil components includes a base defining an upstream end, wherein the upstream end includes a seal plate, and wherein the body of the seal component contacts the seal plate to form a seal with the seal plate.

[0041] Technical solution 15. The gas turbine engine of technical solution 14, wherein the body of the seal component defines a downstream seal surface, and wherein the downstream seal surface contacts the seal plate to form the seal with the seal plate.

[0042] Technical solution 16. The gas turbine engine of technical solution 13, wherein at least one bushing of the bushing assembly is formed of a ceramic matrix composite material, and wherein the seal component is formed of a metallic material.

[0043] Technical solution 17. A method for manufacturing a combustor assembly of a gas turbine engine, the combustor assembly including a bushing assembly and a seal component, the bushing assembly including at least one bushing defining an interface surface at a downstream end, the method comprising: providing the seal component, the seal component continuously extending along a circumferential direction to form a seal ring; and moving the seal component onto the at least one bushing such that a body surface of a body of the seal component contacts the interface surface defined by the at least one bushing.

[0044] Technical solution 18. The method of technical solution 17, wherein the seal component includes a flange, and wherein moving the seal component onto the at least one bushing includes moving a flange of the seal component over a radial opening defined by the at least one bushing.

[0045] Technical solution 19. The method of technical solution 18, wherein the seal component further includes a radial element, and wherein the method further comprises: moving a radial element of the seal component into the radial opening defined by the at least one bushing; and coupling the radial element to the flange of the seal component.

[0046] Technical Solution 20. The method of Technical Solution 17, wherein at least one bushing of the bushing assembly is formed of a ceramic matrix composite material, and wherein the sealing component is formed of a metallic material.

[0047] 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. BRIEF DESCRIPTION OF DRAWINGS

[0048] A complete and open-ended disclosure of the application, including its best mode, to those of ordinary skill in the art, is set forth in the specification, which makes reference to the appended drawings, in which: Figure 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter.

[0049] Figure 2 is a schematic cross-sectional view of a combustor assembly according to an exemplary embodiment of the present disclosure.

[0050] Figure 3 is a partial enlarged schematic cross-sectional view of a downstream end of an outer bushing of the exemplary combustor assembly of Figure 2

[0051] Figure 4 is a partial enlarged schematic cross-sectional view of a downstream end of an outer bushing of the exemplary combustor assembly of Figure 3

[0052] Figure 5 is a partial enlarged schematic cross-sectional view of a downstream end of an outer bushing of a combustor assembly according to another exemplary embodiment of the present disclosure.

[0053] Figure 6 is a schematic view of a downstream end of an outer bushing of the exemplary combustor assembly of Figure 3

[0054] Figure 7 is a schematic view of a downstream end of a bushing assembly of a combustor assembly according to another exemplary embodiment of the present disclosure as viewed along a radial direction of the gas turbine engine.

[0055] Figure 8 is a flowchart of a method for manufacturing a combustor assembly according to an exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0056] ​​​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 refer to like or similar parts of the application.

[0057] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations.

[0058] As used herein, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another and do not necessarily have to refer to a position or order of importance of the individual components.

[0059] The terms "forward" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, forward refers to a position closer to the engine inlet and aft refers to a position closer to the engine nozzle or exhaust.

[0060] The terms "upstream" and "downstream" refer to relative directions 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.

[0061] The terms "coupled," "fixed," "attached" and the like, mean either a direct coupling, fixation, or attachment, as well as an indirect coupling, fixation or attachment by way of one or more intermediate constituent or features unless specifically specified herein, otherwise with the context clearly indicating otherwise.

[0062] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0063] Approximating language as used herein with respect to a given quantity should be interpreted in the context of the specific application in which the quantity is used and that the various sizing demands on such quantities will result in modifying the given language. For example, depending on the circumstances, "about" can include plus or minus ten percent from the indicated value, or in some instances, larger margins from the indicated value. Further, the use of "about" with respect to a given quantity should be interpreted as not being limited to the exact value of the given quantity, but rather to a range of values around the given quantity. For example, "about 10" means a range of values from about 5 to about 15, or in some instances, a range of values from about 9 to about 11. Language of the type "at least A, B, or C, and at most A, B or C" or "at least A, B, or C but not more than A, B or C" or the like, defines the ranges of valid numbers for A, B, or C. For example, "at least 1, 2, or 3, and at most 1, 2, or 3" means 1, 2, or 3.

[0064] Here, and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein, unless context or language indicates otherwise. For example, a range of "between A and B" includes an end value A or an end value B. Unless otherwise stated, the use of ordinal terms such as "first", "second", "third", etc. to describe a common but distinct member of a group will treat such terms as modifiers merely to distinguish that reference between groups, but not necessarily describing a position or importance of the member relative to the group.

[0065] Reference will now be made to the drawings wherein like numerals refer to like components throughout the several figures, Figure 1 is a schematic cross-sectional view of a gas turbine engine in accordance with example embodiments of the present disclosure. More particularly, for embodiments of Figure 1 is a high-bypass turbofan jet engine 10, referred to herein as "turbofan engine 10." As shown in Figure 1 The turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). Generally, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14.

[0066] The depicted example core turbine engine 16 generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encloses (in flow series 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.

[0067] For the depicted embodiment, the fan section 14 includes a variable-pitch fan 38 having a plurality of fan blades 40 coupled in spaced relation to a disc 42. As depicted, the fan blades 40 extend generally outwardly along the radial direction R from the disc 42. Since the fan blades 40 are operably coupled to suitable actuation members 44 configured to collectively vary the pitch of the fan blades 40 in unison, each fan blade 40 is rotatable relative to the disc 42 about a pitch axis P. The fan blades 40, disc 42, and actuation members 44 are rotatable together about the longitudinal axis 12 by the LP shaft 36 that traverses a power gear box 46. The power gear box 46 includes a plurality of gears for reducing the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.

[0068] Still referring to Figure 1In exemplary embodiments of the present application, the disk 42 is covered by a rotatable front nacelle 48 that aerodynamically profiles the airflow 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 the fan 38 and / or at least a portion of 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. A downstream section 54 of the nacelle 50 extends over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

[0069] During operation of the turbofan engine 10, an amount of air 58 enters the turbofan 10 through an associated inlet 60 of the nacelle 50 and / or fan section 14. As the amount of air 58 traverses the fan blades 40, a first portion of the air 58 is directed or passed into the bypass airflow passage 56 as indicated by arrow 62, and a second portion of the air 58 is directed or passed 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 generally referred to as the bypass ratio. The second portion of air 64 then passes through the high pressure (HP) compressor 24 and into the combustion section 26 where the second portion of air 64 is mixed with fuel and combusted to provide combustion gases 66.

[0070] 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 successive stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft or spool 34, thereby 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 is extracted from the combustion gases 66 via successive stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft or spool 36, thereby causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0071] The combustion gases 66 then pass through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the first portion of air 62 is also increased in pressure as it passes through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for passing the combustion gases 66 through the core turbine engine 16.

[0072] It should be appreciated, however, Figure 1 The exemplary turbofan engine 10 depicted is by way of example only, and in other exemplary embodiments, the turbofan engine 10 can have any other suitable configuration. For example, in other exemplary embodiments, the turbofan engine 10 can be any other suitable aeronautical gas turbine engine, such as a turboshaft engine, a turbojet engine, a turboprop engine, etc. Moreover, in yet other exemplary embodiments, aspects of the present disclosure can be incorporated into any other suitable gas turbine engine, e.g., including any suitable number or configuration of shafts, compressors, turbines, etc. Also, while depicted as an aeronautical gas turbine engine, in other embodiments, aspects of the present disclosure can be incorporated into a land-based gas turbine engine, an aeroderivative gas turbine engine, etc.

[0073] Reference is now made to Figure 2 , a partial enlarged cross-sectional view of a combustor assembly 100 according to exemplary embodiments of the present disclosure is provided. For example, Figure 2 The combustor assembly 100 of Figure 1 may be positioned in a combustion section 26 of the exemplary turbofan engine 10. More particularly, Figure 2 A side cross-sectional view of the exemplary combustor assembly 100 of Figure 2 is provided.

[0074] As shown, the combustor assembly 100 generally includes a liner assembly including at least one liner. In particular, for the embodiment shown, the at least one liner of the liner assembly includes an inner liner 102 extending generally along an axial direction A between a downstream end 104 (or aft end for the embodiment shown) and an upstream end 106 (or forward end for the embodiment shown), and an outer liner 108 also extending generally along the axial direction A between a downstream end 110 and an upstream end 112. The inner liner 102 and the outer liner 108 together at least partially define a combustion chamber 114 therebetween. As will be appreciated, for the embodiment shown, the combustor assembly 100 is configured as an annular combustor, such that the inner liner 102 and the outer liner 108 each extend along a circumferential direction C (see below) to define a circular / annular shape about a central axis (e.g., axis 12), and likewise such that the combustion chamber 114 is an annular combustion chamber. For example, the outer liner 102 can extend continuously along the circumferential direction C, or alternatively can include a plurality of liners forming a continuous outer liner for the combustor assembly 100. Similarly, the inner liner 108 can extend continuously along the circumferential direction C, or alternatively can include a plurality of liners forming a continuous inner liner for the combustor assembly 100. Such configurations will be discussed in greater detail below.

[0075] Additionally, the inner liner 102 and the outer liner 108 are each attached to an annular dome. More particularly, the annular dome includes an inner dome section 116 attached to the upstream end 106 of the inner liner 102 and an outer dome section 118 attached to the upstream end 112 of the outer liner 108. The inner dome section 116 and the outer dome section 118 can be integrally formed (or alternatively can be formed from multiple members attached in any suitable manner) and can each extend along the circumferential direction C to define an annular shape. The inner dome section 116 and the outer dome section 118 also each at least partially define a slot 122 for receiving the upstream end 106 of the inner liner 102 and the upstream end 112 of the outer liner 108, respectively.

[0076] The combustor assembly 100 also includes a plurality of fuel air mixers 124 spaced along the circumferential direction C and positioned at least partially within the annular dome. More particularly, the plurality of fuel air mixers 124 are disposed at least partially along the radial direction R between the outer dome section 118 and the inner dome section 116. Compressed air from the compressor section of the turbofan engine 10 flows into or through the fuel air mixers 124, where the compressed air is mixed with fuel and ignited to produce combustion gases 66 within the combustion chamber 114. The inner dome section 116 and the outer dome section 118 are configured to help provide such a flow of compressed air from the compressor section into or through the fuel air mixers 124. For example, the outer dome section 118 includes an outer shroud 126 at an upstream end 128 and the inner dome section 116 similarly includes an inner shroud 130 at an upstream end 132. The outer shroud 126 and the inner shroud 130 can help direct the flow of compressed air from the compressor section 26 into or through one or more of the fuel air mixers 124.

[0077] Also, the inner dome section 116 and the outer dome section 118 each include an attachment portion configured to help mount the combustor assembly 100 within the turbofan engine 10. For example, the outer dome section 118 includes an attachment extension 134 configured to be mounted to an outer combustor case 136 and the inner dome section 116 includes a similar attachment extension 138 configured to be attached to an annular support member 140 within the turbofan engine 10. In certain example embodiments, the inner dome section 116 can be integrally formed as a single annular member and similarly, the outer dome section 118 can also be integrally formed as a single annular member.

[0078] It should be appreciated, however, that in other example embodiments, the inner dome section 116 and / or the outer dome section 118 can alternatively be formed from one or more members joined in any suitable manner. For example, with reference to the outer dome section 118, in certain example embodiments, the outer shroud 126 can be formed separately from the outer dome section 118 and attached to the upstream end 128 of the outer dome section 118 using, for example, a welding process. Similarly, the attachment extension 134 can also be formed separately from the outer dome section 118 and attached to the upstream end 128 of the outer dome section 118 using, for example, a welding process. Additionally or alternatively, the inner dome section 116 can have a similar construction.

[0079] For the depicted embodiment, at least one liner of the liner assembly (and more particularly the inner liner 102 and the outer liner 108) is formed from a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature properties. Example CMC materials for such liners 102, 108 can include silicon carbide, 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 like sapphire or silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarns including silicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicate (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), as well as 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). At least certain CMC materials can have a coefficient of thermal expansion in the range of about 1.3 x 10 −6 in / in / °F to about 3.5 x 10 −6 in / in / °F.

[0080] Still referring to Figure 2, the combustor assembly 100 further includes features for forming a seal with a component of a gas turbine engine positioned adjacent to the combustor assembly 100. More particularly, for the illustrated embodiment, the combustor assembly 100 is configured to form a seal with a first stage of an airfoil component 150 of a turbine section of a gas turbine engine. For the illustrated embodiment, the first stage of the airfoil component 150 is a first stage of a turbine nozzle and includes a base defining an upstream end. More particularly, the first stage of the airfoil component 150 includes an outer base 152 defining an upstream end 154 and an inner base 156 defining an upstream end 158. The upstream end 154 of the outer base 152 and the upstream end 158 of the inner base 156 each include a seal plate 160. As will be appreciated from the discussion herein, features of the combustor assembly 100 are configured to form a seal with the seal plates 160 of the upstream end 158 of the outer base 152 and the upstream end 154 of the inner base 156.

[0081] More particularly, now referring also to Figure 3 ( Figure 3 Providing a partial enlarged view of the downstream end 110 of the outer liner 108 of at least one liner of the liner assembly), as depicted, the downstream end 110 of the outer liner 108 defines a radial opening 162 and an interface surface 164 extending along the radial direction R and along the circumferential direction C (as will be explained in greater detail below, for example, with reference to Figure 6 ). More particularly, the outer liner 108 of the liner assembly includes a groove 166 formed at the downstream end 110 thereof, wherein the interface surface 164 forms a portion of the groove 166. It will be appreciated that, as used herein, the term "downstream end" of a particular liner refers to a downstream section of the liner that includes less than twenty percent of the axial length of the liner.

[0082] Further, the combustor assembly 100 includes a seal component 168 operable with the downstream end 110 of the outer liner 108 to form a seal with an adjacent component of a gas turbine engine. For the illustrated embodiment, the seal component 168 generally includes a body 170, a flange 172, and a radial element 174. The flange 172 extends from the body 170 of the seal component 168 toward the radial opening 162. In particular, for the illustrated embodiment, the flange 172 extends forward from the body 170 past / over the radial opening 162 defined by the outer liner 108. Further, the radial element 174 of the seal component 168 is coupled to the flange 172 and at least partially extends into the radial opening 162 defined by the outer liner 108. In certain example embodiments, the radial element 174 can be fixedly coupled to the flange 172, for example, by welding or some other fixed mechanical coupling (e.g., a screw connection, a glue or epoxy connection, etc.). Alternatively, in other embodiments, the radial element 174 of the seal component 168 can be movably coupled to the flange 172, such as slidably coupled to the flange 172.

[0083] Notably, for the illustrated embodiment, the radial opening 162 defined by the outer bushing 108 is positioned immediately forward of the interface surface 164 and defines a length 176 along the radial direction R and a constant cross-sectional geometry along the length 176. For example, for the illustrated embodiment, the radial opening 162 defines a constant circular cross-sectional geometry along its length 176. Also for the illustrated embodiment, the radial element 174 of the seal member 168 defines a cross-sectional geometry that is substantially equal to the cross-sectional geometry of the radial opening 162 of the outer bushing 108. For example, for the illustrated embodiment, the radial element 174 defines a circular cross-sectional geometry having a diameter that is substantially equal to the diameter of the circular cross-sectional geometry of the radial opening 162. For example, the radial element 174 can be configured as a pin or other suitable structure. However, in other embodiments, the radial opening 162 and the radial element 174 can have other suitable complementary shapes.

[0084] Also, the body 170 of the seal member 168 defines a body surface 178 that extends along the radial direction R and is positioned adjacent to the interface surface 164 of the outer bushing 108. The body surface 178 and the interface surface 164 can together form a seal between the seal member 168 and the outer bushing 108. It will be appreciated that, as used herein, the term "seal", as can be defined between two components, indicates that the two components define a relatively small measured gap or no gap therebetween to limit or prevent any air flow therebetween. Thus, in certain example embodiments, a seal can indicate that the two components contact each other and form a substantially air-tight seal, or alternatively can indicate that the two components define a relatively small measured gap therebetween to restrict air flow therebetween in a desired manner.

[0085] Also, it will be appreciated that, for the illustrated embodiment, the body 170 of the seal member 168 contacts the seal plate 160 of the upstream end 154 of the outer base 152 to form a seal with the seal plate 160. More particularly, for the illustrated embodiment, the body surface 178 is positioned at a front / upstream side of the seal member 168, and the body 170 also includes a seal surface 180 that is positioned at a rear / downstream side of the seal member 168 that contacts the seal plate 160 to form a seal with the seal plate 160.

[0086] Briefly, it will be further appreciated that the upstream end 154 of the outer base 152 also includes an assembly for retaining the seal plate 160 in place and allowing the seal plate 160 to further form a seal with the upstream end 154 of the outer base 152. More particularly, the assembly includes a flange 182 with a bolt 184 extending therethrough to press an inner end 185 of the seal plate 160 against a lip 186 of the upstream end 154 of the outer base 152 and an outer end against a sealing surface 180 of the base 170 of the seal member 168.

[0087] In this manner, the seal member 168 can form a seal between the downstream end 110 of the outer liner 108 and an aft adjacent component of the gas turbine engine (i.e., the first stage of the airfoil component 150 for the illustrated embodiment).

[0088] As will be described in greater detail below, it will be further appreciated that the seal member 168 is configured to form a seal ring 194 (see Figure 6 that continuously extends in the circumferential direction C to form a circular shape (at least partially) around the downstream end 110 of the outer liner 108, which as noted above similarly has an annular shape. As also noted above, the outer liner 108 can be formed of a CMC material. However, in contrast, for the illustrated embodiment, the seal member 168 can be formed of a metallic material. For example, the seal member 168 can be formed of a metal, such as a nickel-based superalloy (which can have a coefficient of thermal expansion of about 8.3-8.5 x 10 −6 in / in / °F at temperatures of about 1000-1200 °F) or a cobalt-based superalloy (which can have a coefficient of thermal expansion of about 7.8-8.1 x 10 −6 in / in / °F at temperatures of about 1000-1200 °F). Thus, during operation of the gas turbine engine, the seal member 168 can expand relative to the outer liner 108 such that an effective diameter along the radial direction R of the seal ring 194 formed by the seal member 168 increases relative to an effective diameter along the radial direction R of the downstream end 110 of the outer liner 108.

[0089] Now referring specifically to Figure 4 such example aspects are depicted. More particularly, Figure 4 during operation of the gas turbine engine are depicted. Figure 3exemplary outer liner 108 and seal component 168, where the increased operating temperature causes the seal component 168 to expand relative to the outer liner 108. Thus, as shown, the flange 172 of the seal component 168 now defines a radial gap 188 with the downstream end 110 of the liner 108 along the radial direction R, and similarly, the body 170 of the seal component 168 defines a similarly sized radial gap 190 with the radially inner surface of the groove 166. However, given the above-described configuration of the outer liner 108 and seal component 168, the seal component 168 can maintain a seal between the downstream end 110 of the outer liner 108 and the upstream end 154 of an adjacent component of the gas turbine engine.

[0090] More particularly, as depicted, the body surface 178 of the body 170 of the seal component 168 slidably engages with the interface surface 164 defined by the outer liner 108 of the liner assembly. In this manner, despite the fact that the seal component 168 expands relative to the outer liner 108, the body surface 178 and the interface surface 164 can maintain a seal therebetween. Further, as shown, the radial element 174 of the seal component 168 slidably engages with the radial opening 162 defined by the outer liner 108. In this manner, the radial element 174 can ensure that the seal component 168 maintains its position relative to the outer liner 108 along the axial direction A, while allowing the seal component 168 to expand and / or contract along the radial direction R.

[0091] However, it will be appreciated that in other embodiments, any other suitable configuration of radial opening 162 and radial element 174 can provide for constraining the seal component 168 relative to the outer liner 108 along the axial direction A, while allowing relative movement along the radial direction R. For example, in at least certain exemplary embodiments, such as the alternative exemplary embodiment of Figure 5 , the outer liner 108 can include a wear coating 192 positioned at least partially within the radial opening 162. For example, as depicted in Figure 5 , the wear coating 192 can be effectively configured as a liner or other material configured to interface with the radial element 174 of the seal component 168 to reduce wear on the outer liner 108 (given the different materials - ceramic matrix composite versus metal material). However, in other embodiments, such a wear coating 192 can not be provided, or other suitable configurations can be provided.

[0092] Moreover, as noted above, the outer liner 108 extends about the circumferential direction C to define an annular shape, and similarly, the seal component 168 extends in the circumferential direction C to define an annular seal ring 194. Referring to Figure 6 , it is depicted that the outer liner 108 and the seal component 168 are configured to define a radial gap 196 between the downstream end 110 of the outer liner 108 and the upstream end 154 of the adjacent component of the gas turbine engine. In this manner, the seal component 168 can maintain a seal between the downstream end 110 of the outer liner 108 and the upstream end 154 of the adjacent component of the gas turbine engine. Figures 2 to 4a downstream end 110 of the outer liner 108 of the example combustor assembly 100, with the sealing component 168 positioned thereon. As shown, the sealing component 168 extends continuously along the circumferential direction C to form a circumferential seal ring 194, and the outer liner 108 similarly extends continuously along the circumferential direction C to form a similar annular shape. In this manner, it will be appreciated that the interface surface 164 of the outer liner 108 (which is masked by the sealing component 168 in Figure 6 ) extends continuously along the circumferential direction C to form a complete annulus, and the body surface 178 of the body 170 of the sealing component 168 similarly extends continuously along the circumferential direction C to form a complete annulus. The body surface 178 and the interface surface 164 can together define a 360 degree seal for the combustor assembly 100.

[0093] Moreover, as schematically depicted, it will be appreciated that the radial opening 162 described above and depicted in Figure 3 and Figure 4 is a first radial opening 162 of a plurality of radial openings 162 defined by the outer liner 108 at the downstream end 110 of the outer liner 108. For the illustrated embodiment, the plurality of radial openings 162 defined by the outer liner 108 are spaced along the circumferential direction C. Furthermore, as also schematically depicted, it will be appreciated that the radial element 174 of the sealing component 168 is a first radial element 174 of a plurality of radial elements 174 of the sealing component 168. Each radial element 174 of the plurality of radial elements 174 is coupled to the flange 172 of the sealing component 168 and extends into a respective radial opening 162 of the plurality of radial openings 162 of the outer liner 108. In this manner, the plurality of radial elements 174 positioned in the plurality of radial openings 162 can constrain the sealing component 168 relative to the outer liner 108 along the axial direction A during operation of the gas turbine engine, while still allowing for relative radial expansion between the sealing component 168 and the outer liner 108 during operation of the gas turbine engine.

[0094] However, it will be appreciated that in other example embodiments, any other suitable configuration can be provided for the combustor assembly 100, the gas turbine engine, etc. For example, now briefly referring to Figure 7 , a plan view of a downstream end of a liner assembly of a combustor assembly 100 is provided in accordance with another example embodiment of the present disclosure. For the illustrated embodiment, the liner assembly includes an outer liner 108 and a sealing component 168 positioned thereon. As shown, the sealing component 168 extends continuously along the circumferential direction C to form a circumferential seal ring 194, and the outer liner 108 similarly extends continuously along the circumferential direction C to form a similar annular shape. In this manner, it will be appreciated that the interface surface 164 of the outer liner 108 (which is masked by the sealing component 168 in Figure 7In the embodiment depicted, at least one outer bushing 108 of the bushing assembly is not a single outer bushing extending continuously along the circumferential direction C, but rather a plurality of outer bushings 108. The plurality of outer bushings 108 are arranged along the circumferential direction C to form an effective continuous outer bushing along the circumferential direction C. Each of the outer bushings 108 may be constructed in a manner similar to that described above. For example, each of the outer bushings 108 may define a downstream end 110, and one or more of such bushings define a radial opening 162 at the downstream end 110. Furthermore, each bushing defines an interface surface 164 extending along the radial direction R and along the circumferential direction C at the downstream end 110, such that the plurality of outer bushings 108 together define the interface surface 164, which extends continuously along the circumferential direction C to form a complete annulus.

[0095] Moreover, it will be understood that, although the above text is about Figures 3 to 6 The exemplary sealing member 168 described is configured to form a seal between the outer base 152 of the first stage of the airfoil member 150 and the outer bushing 108. In other embodiments, additional or alternative sealing members 168 may be provided to form a seal between the downstream end 104 of the inner bushing 102 and the inner base 156 of the first stage of the airfoil member 150 (see [link to documentation]). Figure 2 For example, in such embodiments, the downstream end 104 of the inner liner 102 may define one or more radial openings and interface surfaces, and the sealing member may include: a body having a body surface positioned adjacent to the interface surface; a flange; and a radial member coupled to the flange and extending at least partially into the radial opening of the inner liner. However, it is noteworthy that the sealing member may be dimensionally configured to define a gap (similar to) under cold / non-operating conditions. Figure 4 The gaps 188 and 190 in the middle are designed to allow the sealing components to expand during hot / operating conditions (e.g., operating conditions).

[0096] Now refer to Figure 8 A flowchart of a method 300 for manufacturing a combustor assembly for a gas turbine engine, according to an exemplary aspect of this disclosure, is provided. The combustor assembly manufactured according to method 300 can be used in accordance with the methods described above for... Figures 1 to 7 One or more of the other exemplary burner assemblies are constructed in a similar manner. Therefore, it will be understood that a burner assembly may include a bushing assembly and a sealing element, wherein the bushing assembly includes at least one bushing defining an interface surface at a downstream end.

[0097] Method 300 generally includes providing a sealing element at (302). The sealing element extends continuously in the circumferential direction to form a sealing ring. It is noteworthy that, as used herein, the term “providing” means only to make the component available and does not explicitly require the manufacture or formation of such component.

[0098] The method 300 also includes moving the seal component onto the at least one bushing at (304) such that the body surface of the body of the seal component contacts the interface surface defined by the at least one bushing of the bushing assembly of the combustor assembly. For the depicted example aspect, it will be appreciated that the seal component includes a flange, and moving the seal component onto the at least one bushing at (304) includes moving the flange of the seal component over the radial opening defined by the at least one bushing at (306).

[0099] Further, for the depicted example aspect, the seal component includes a radial element, and the method 300 also includes moving the radial element of the seal component into the radial opening defined by the at least one bushing at (308), and coupling the radial element to the flange of the seal component at (310).

[0100] As with the example embodiments described above, it will be appreciated that the at least one bushing of the bushing assembly can be formed of a ceramic matrix composite material, and the seal component can be formed of a metallic material. However, forming a combustor assembly according to one or more example aspects of the present disclosure can allow the seal component to move relative to the at least one bushing in a radial direction during operation of the gas turbine engine while maintaining a seal between the two components.

[0101] 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 elements that do not differ from the literal language of the claims, or if they include other structural elements that are insubstantial changes from the literal language of the claims.

Claims

1. A combustor assembly of a gas turbine engine defining a central axis, a radial direction, and a circumferential direction, the combustor assembly comprising: a liner assembly at least partially defining a combustion chamber and including at least one radial outer liner extending between a downstream end and an upstream end, the downstream end of the at least one radial outer liner defining a radial opening and an interface surface extending along the circumferential direction and along the radial direction; and a seal member including a body, a flange, and a radial element, the body extending outward relative to the central axis beyond the flange along the radial direction adjacent the interface surface, the body defining a body surface extending along the radial direction and positioned adjacent the interface surface of the at least one radial outer liner, the flange extending forward from the body, and the radial element coupled to the flange and extending into the radial opening defined by the at least one radial outer liner, the body surface slidably engaged with the interface surface to maintain a seal between the seal member and the outer liner in response to a radial gap formed between the seal member and the outer liner due to radial movement of the seal member relative to the outer liner. The body surface of the body of the seal member is slidably engaged with the interface surface defined by the at least one radial outer liner of the liner assembly.

2. The combustor assembly of claim 1, wherein, The body of the seal member defines a downstream sealing surface configured to contact a corresponding seal of an adjacent component within the gas turbine engine when installed.

3. The combustor assembly of claim 1, wherein, The radial opening is a first radial opening of a plurality of radial openings defined by the at least one radial outer liner at the downstream end of the at least one radial outer liner, and wherein the plurality of radial openings are spaced along the circumferential direction.

4. The combustor assembly of claim 1, wherein, The radial element of the seal member is a first radial element of a plurality of radial elements of the seal member, and wherein each radial element of the plurality of radial elements is coupled to the flange of the seal member and extends through a respective radial opening of the plurality of radial openings defined by the at least one radial outer liner.

5. The combustor assembly of claim 4, wherein, The radial opening defines a length along the radial direction and a constant cross-sectional geometry along the length.

6. The combustor assembly of claim 1, wherein, The radial opening extends only partially into the at least one radial outer liner.

7. The burner assembly of claim 1, wherein, The at least one radial outer liner includes a wear coating positioned within the radial opening.

8. The burner assembly of claim 1, wherein, The radial element of the seal member defines a cross-sectional geometry that is substantially equal to a cross-sectional geometry of the radial opening of the at least one radial outer liner.

9. The burner assembly of claim 1, wherein, The seal member extends continuously along the circumferential direction to form a circumferential seal ring.

10. The burner assembly of claim 1, wherein, ​

Images