Integrated flow path structure

Abstract

A flow path assembly and a gas turbine engine are provided. The flow path assembly may include a burner dome positioned at a forward end of a combustor of a combustion section of the gas turbine engine and a unitary outer wall, the unitary outer wall including a burner portion and a turbine portion, the burner portion extending through the combustion section, and the turbine portion extending through at least a first turbine stage of the turbine section of the gas turbine engine. The burner portion and the turbine portion are integrally formed as a single, unitary structure. The flow path assembly also includes an inner wall extending from the forward end of the burner through at least the combustion section. The burner dome extends radially from the unitary outer wall to the inner wall and is configured to move axially relative to the inner wall and the unitary outer wall. Other flow path assembly and gas turbine engine configurations are provided.

Description

[0001] This application is a divisional application of the invention patent application with application number 201780084777.2 filed on December 6, 2017, and invention name “Integrated flow path structure”. Technical Field

[0002] The present subject matter generally relates to gas turbine engines. More particularly, the present subject matter relates to an integral structure for defining a flow path within a gas turbine engine. Background Art

[0003] A gas turbine engine typically includes a fan and a core arranged in flow communication with each other. Furthermore, the core of a gas turbine engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section in serial flow order. During operation, air is supplied from the fan to the 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 and combusted within the combustion section to provide combustion gases. From the combustion section, the combustion gases are transported to the turbine section. The combustion gas flow through the turbine section drives the turbine section and then passes through the exhaust section, for example, to the atmosphere.

[0004] More specifically, the combustion section includes a combustor having a combustion chamber defined by a combustor liner. Downstream of the combustor, the turbine section includes one or more stages. For example, each stage may have multiple stationary nozzle airfoils and multiple blade airfoils attached to a rotor that is driven by the flow of combustion gases against the blade airfoils. The turbine section may also have other configurations. For example, the turbine may be a counter-rotating turbine without stationary nozzle airfoils. In any case, the flow path is defined by inner and outer boundaries, both of which extend from the combustor through the stages of the turbine section.

[0005] Typically, the inner and outer boundaries that define the flow path include separate components. For example, the outer liner of the combustor, the separate outer band of the nozzle portion of the turbine stage, and the separate shroud of the blade portion of the turbine stage typically define at least a portion of the outer boundary of the flow path. Using separate components to form each of the outer and inner boundaries may require one or more seals at each interface between the separate components to minimize leakage of fluid from the flow path. As a result, a large number of components may be required to construct the flow path assembly, which can increase the complexity and weight of the gas turbine engine. In addition, although seals can be provided, the use of separate components in the flow path assembly provides several points for fluid to leak from the flow path. The increased weight, complexity, and leakage can negatively impact engine performance and the assembly of the engine during manufacturing.

[0006] Therefore, there is a need for improved flow path assemblies. For example, an outer boundary structure extending through the combustion section and at least the first stage of the turbine section would be beneficial. Furthermore, a flow path assembly comprising an inner boundary structure and an outer boundary structure would be useful. Furthermore, a flow path assembly comprising an integral combustor dome, an inner boundary structure, and an outer boundary structure would be helpful. Furthermore, a gas turbine engine having a flow path assembly having an outer boundary structure would be advantageous. Summary of the Invention

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

[0008] In an exemplary embodiment of the present disclosure, a flow path assembly for a gas turbine engine is provided. The flow path assembly includes a combustor dome positioned at a forward end of a combustor in a combustion section of the gas turbine engine. The flow path assembly also includes a one-piece outer wall comprising a combustor portion extending through the combustion section and a turbine portion extending through at least a first turbine stage of the turbine section of the gas turbine engine. The combustor portion and the turbine portion are integrally formed as a single, integral structure. The flow path assembly also includes an inner wall extending from the forward end of the combustor through at least the combustion section. The combustor dome extends radially from the one-piece outer wall to the inner wall, and the combustor dome is configured to move axially relative to the inner wall and the one-piece outer wall.

[0009] In another exemplary embodiment of the present disclosure, a flow path assembly for a gas turbine engine is provided. The flow path assembly extends from a combustor through a turbine section of the gas turbine engine. The turbine section includes a first turbine stage immediately downstream of the combustor and a second turbine stage immediately downstream of the first turbine stage. The flow path assembly includes a burner dome positioned at a forward end of the burner and an inner wall defining an inner boundary of a flow path. The inner wall includes an inner liner of the burner. The flow path assembly further includes an outer wall defining an outer boundary of the flow path. The outer wall extends from the forward end of the burner through at least the first turbine stage. The burner dome is a separate component from the inner wall and the outer wall.

[0010] In another exemplary embodiment of the present disclosure, a gas turbine engine is provided. The gas turbine engine includes a combustion section including a combustor. The gas turbine engine also includes a turbine section including a first turbine stage positioned immediately downstream of the combustion section and a second turbine stage positioned immediately downstream of the first turbine stage. The combustion section and the turbine section define a flow path. The combustion section includes an inner liner, an outer liner, and a combustor dome. The inner liner defines an inner boundary of the flow path through the combustion section, and the outer liner defines an outer boundary of the flow path through the combustion section. The combustor dome is positioned at a forward end of the combustor and extends between the inner and outer liner. Each of the first and second turbine stages of the turbine section includes a nozzle portion and a blade portion. Each nozzle portion includes an inner band and an outer band. The inner band defines an inner boundary of the flow path through the nozzle portion, and the outer band defines an outer boundary of the flow path through the nozzle portion. Each blade portion includes a shroud that defines an outer boundary of the flow path through the blade portion. The outer liner, outer band, and shroud are integrally formed such that the outer liner, outer band, and shroud form a single, integral outer wall. Furthermore, the combustor dome is a separate component from the liner and the integral outer wall.

[0011] These and other features, aspects and advantages of the present invention will be 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 present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] A full and enabling disclosure of the invention, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification, which refers to the accompanying drawings, in which:

[0013] Figure 1 Schematic cross-sectional views of exemplary gas turbine engines according to various embodiments of the present subject matter are provided.

[0014] Figure 2 According to exemplary embodiments of the present subject matter, Figure 1 Schematic exploded cross-sectional view of the combustion section and high-pressure turbine section of a gas turbine engine.

[0015] Figure 3A According to exemplary embodiments of the present subject matter, Figure 2 Schematic cross-sectional view of the combustion section and the high-pressure turbine section.

[0016] Figure 3B , 3C, 3D and 3E provide other exemplary embodiments according to the present subject matter Figure 2 Schematic cross-sectional view of the combustion section and the high-pressure turbine section.

[0017] Figure 3FAccording to exemplary embodiments of the present subject matter, Figure 2 A partial perspective view of the overall outer boundary structure and a portion of the inner boundary structure of the combustion section and the high-pressure turbine section.

[0018] Figure 4A , 4B, 4C, 5A, 5B and 5C provide other exemplary embodiments according to the present subject matter Figure 2 Schematic cross-sectional view of the combustion section and the high-pressure turbine section. DETAILED DESCRIPTION

[0019] Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar designations in the drawings and the description have been used to refer to the same or similar parts of the invention. As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of the various components. The terms "upstream" and "downstream" refer to relative directions with respect to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which a fluid is flowing and "downstream" refers to the direction toward which a fluid is flowing.

[0020] Referring now to the drawings, in which like reference numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference) and a radial direction R. Generally, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14 .

[0021] The exemplary core turbine engine 16 shown generally includes a substantially tubular casing 18 defining an annular inlet 20. Casing 18 encloses, in serial flow relationship: a compressor section including a supercharger 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 an exhaust nozzle section 32. A high-pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22. In other embodiments of the turbofan engine 10, additional spools may be provided, such that the engine 10 may be described as a multi-spool engine.

[0022] For the depicted embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, fan blades 40 extend generally outwardly from disk 42 in a radial direction R. Fan blades 40 and disk 42 are rotatable together about longitudinal axis 12 via LP shaft 36. In some embodiments, a power gearbox having a plurality of gears may be included for stepping down the rotational speed of LP shaft 36 to a more efficient rotational fan speed.

[0023] Still refer to Figure 1 In an exemplary embodiment, the disk 42 is covered by a rotatable forward nacelle 48 that is aerodynamically shaped to facilitate 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 surrounds the fan 38 and / or at least a portion of the core turbine engine 16. It should be understood that the nacelle 50 can be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Furthermore, a downstream section 54 of the nacelle 50 can extend over an outer portion of the core turbine engine 16 to define a bypass airflow passage 56 therebetween.

[0024] During operation of turbofan engine 10, a volume of air 58 enters turbofan 10 through nacelle 50 and / or associated inlet 60 of fan section 14. As volume of air 58 passes through fan blades 40, a first portion of air 58, as indicated by arrow 62, is directed or channeled into bypass airflow passage 56, and a second portion of air 58, as indicated by arrow 64, is directed or channeled into LP compressor 22. The ratio between first portion of air 62 and second portion of air 64 is generally referred to as the bypass ratio. The pressure of second portion of air 64 then increases as it is directed through high pressure (HP) compressor 24 and into combustion section 26, where it mixes with fuel and combusts to provide combustion gases 66.

[0025] The combustion gases 66 are directed 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 blades 68 coupled to the 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 are then directed 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 blades 72 coupled to the 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.

[0026] The combustion gases 66 are then directed through the exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is significantly increased as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the core turbine engine 16.

[0027] It should be understood that although described with respect to a turbofan 10 having a core turbine engine 16, the present subject matter may be applicable to other types of turbomachinery. For example, the present subject matter may be suitable for use with or in turboprop engines, turboshafts, turbojets, industrial and marine gas turbine engines, and / or auxiliary power units.

[0028] 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, HP turbine 28, and / or LP turbine 30, may include a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capabilities. Of course, other components of the turbofan engine 10, such as components of the HP compressor 24, may include CMC materials. Exemplary CMC materials for these components may include silicon carbide (SiC), silicon, silica, or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded in the matrix, such as oxidation-stable reinforcing fibers, including monofilaments such as sapphire and silicon carbide (e.g., Textron's SCS-6), as well as rovings and yarns, including silicon carbide (e.g., Nippon Carbon's Ube Industries and Dow Corning ), aluminum silicates (e.g., Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's 440 and ), and optionally 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). For example, in certain embodiments, the fiber bundles that may include a ceramic refractory coating are formed into reinforcement tapes, such as unidirectional reinforcement tapes. Multiple tapes may be laid together (e.g., as layers) to form a preform component. The fiber bundles may be impregnated with the slurry composition before or after forming the preform. The preform may then undergo a thermal treatment, such as curing or burnout, to produce a high coke residue in the preform, and subsequent chemical treatment, such as melt infiltration or chemical vapor infiltration of silicon, to obtain a component formed of a CMC material having the desired chemical composition. In other embodiments, the CMC material may be formed, for example, as carbon fiber cloth instead of a tape.

[0029] As described above, components comprising CMC materials may be used within 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 may include a combustor formed from CMC materials, and / or one or more stages of the HP turbine 28 may be formed from CMC materials. However, CMC components may also be used in other sections, such as the compressor section and / or the fan section. Of course, in some embodiments, other high-temperature materials and / or other composite materials may be used to form one or more components of the engine 10.

[0030] Figure 2 An exploded view of a schematic cross section of a combustion section 26 and an HP turbine 28 of the turbine section of a turbofan engine 10 is provided, according to an exemplary embodiment of the present subject matter. Figure 3A Provided Figure 2FIG2 is an unexploded schematic cross-sectional view of the combustion section 26 and HP turbine 28, focusing on the outer boundaries of the flow path through the combustion section 26 and the HP turbine 28. The depicted combustion section 26 includes a generally annular combustor 80, and downstream of the combustion section 26, the HP turbine 28 includes a plurality of turbine stages. More specifically, for the depicted 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 positioned immediately downstream of the combustion section 26, and the second turbine stage 84 is positioned immediately downstream of the first turbine stage 82. Furthermore, each turbine stage 82, 84 includes a nozzle portion and a blade portion; the first turbine stage 82 includes a nozzle portion 82N and a blade portion 82B, and the second turbine stage 84 includes a nozzle portion 84N and a blade portion 84B. The nozzle portion 82N of the first turbine stage 82 is located immediately downstream of the combustion section 26, so that the nozzle portion 82N of the first turbine stage 82 can also be referred to as a combustor exhaust nozzle. In addition, the combustor 80 defines a generally annular combustion chamber 86, so that the combustor 80 can be described as a generally annular combustor.

[0031] Furthermore, as described in greater detail below, a flow path 100 through the combustion section 26 and the HP turbine 28 is defined by the outer and inner boundaries of a flow path assembly 101. The outer and inner boundaries form a flow path for the combustion gases 66 through the combustion section 26 and the HP turbine 28; thus, the flow path 100 may include at least a portion of the hot gas path 78 described above. Furthermore, in other embodiments, the flow path 100 may also extend through the LP turbine 30 and the jet exhaust 32; in still other embodiments, the flow path 100 may also extend forward upstream of the combustion section 26, for example, into the HP compressor 24. Thus, it should be understood that the discussion of the present subject matter herein with respect to the combustion section 26 and the HP turbine 28 is merely exemplary and may also be applied to different configurations of gas turbine engines and flow paths 100.

[0032] like Figure 2 As shown in the exploded view of FIG, the outer and inner boundaries may be defined by outer wall 102 and inner wall 120, respectively, which may include portions of combustion section 26 and HP turbine 28. For example, combustor 80 includes an outer liner 108 that defines the outer boundary of the flow path through combustor 80. Each nozzle section 82N, 84N includes an outer band that defines the outer boundary of the flow path through the nozzle section of each turbine stage, and each blade section 82B, 84B includes a shroud that defines the outer boundary of the flow path through the blade section of each turbine stage. More specifically, as shown in FIG. Figure 2As shown, first turbine stage nozzle portion 82N includes outer band 110, first turbine stage blade portion 82B includes shroud 112, second turbine stage nozzle portion 84N includes outer band 114, and second turbine stage blade portion 84B includes shroud 116. These portions of combustion section 26 and HP turbine 28 may include at least a portion of outer wall 102, as described in more detail below.

[0033] In addition, if Figure 2 As shown, the combustor 80 includes a liner 122 that defines the inner boundary of the flow path through the combustor 80. Each nozzle section 82N, 84N includes an inner band that defines the inner boundary of the flow path through the nozzle section of each turbine stage, and each blade section 82B, 84B includes one or more blade platforms that define the inner boundary of the flow path through the blade section of each turbine stage. More specifically, as shown in FIG. Figure 2 As shown, first turbine stage nozzle portion 82N includes inner band 124, first turbine stage blade portion 82B includes blade platform 132, second turbine stage nozzle portion 84N includes inner band 136, and second turbine stage blade portion 84B includes blade platform 132. These portions of combustion section 26 and HP turbine 28 may include at least a portion of inner wall 122, as described in more detail below.

[0034] Furthermore, in the depicted embodiment, a combustor dome 118 extends radially beyond the forward end 88 of the combustor 80. The combustor dome 118 may be a portion of the outer wall 102, a portion of the inner wall 120, a portion of both the outer wall 102 and the inner wall 120 (e.g., a portion of the combustor dome 118 may be defined by the outer wall 102 and the remaining portion may be defined by the inner wall 120), or a separate component from the outer wall 102 and the inner wall 120. Additionally, a plurality of nozzle airfoils are located in each nozzle section 82N, 84N. Each nozzle airfoil 126 in the first turbine stage nozzle section 82N extends radially from an outer band 110 to an inner band 124, and the nozzle airfoils 126 are circumferentially spaced about the longitudinal centerline 12. Each nozzle airfoil 128 in the second turbine stage nozzle section 84N extends radially from an outer band 114 to an inner band 136, and the nozzle airfoils 128 are circumferentially spaced about the longitudinal centerline 12. Furthermore, a plurality of blade airfoils 130 are positioned in each blade section 82B, 84B. Each blade airfoil 130 within the first turbine stage blade section 82B is attached to a blade platform 132, which in turn is attached to a first stage rotor 134. The blade airfoils 130 attached to the first stage rotor 134 are circumferentially spaced about the longitudinal centerline 12. Similarly, each blade airfoil 130 within the second turbine stage blade section 84B is attached to a blade platform 132, which in turn is attached to a second stage rotor 138. The blade airfoils 130 attached to the second stage rotor 138 are circumferentially spaced about the longitudinal centerline 12. Each blade airfoil 130 extends radially outward toward the outer wall 102 (i.e., the outer boundary of the flow path 100) and defines a gap between a tip 140 of each blade airfoil 130 and the outer wall 102, allowing each turbine rotor 134, 138 to freely 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, rotor blade airfoils 130 may extract kinetic energy from the flow of combustion gases through flow path 100 defined by HP turbine 28 as rotational energy that is applied to HP shaft 34 .

[0035] Thus, a flow path 100 through the combustion section 26 and the HP turbine 28 is defined by a flow path assembly 101 having an inner boundary and an outer boundary, and the inner boundary and the outer boundary define the flow path of the combustion gases 66 through the combustion section 26 and the HP turbine 28. Portions of the outer boundary of the flow path assembly 101 can be integrated or unified into a single outer wall 102 that defines the radial outer boundary of the gas flow path 100. For example, the outer wall 102 can include a combustor portion 104 that extends through the combustion section (e.g., the combustion section 26) and a turbine portion 106 that extends through at least the first turbine stage of the turbine section (e.g., the first turbine stage 82 of the HP turbine 28). The combustor portion 104 and the turbine portion 106 are integrally formed such that the combustor portion and the turbine portion are a single, unitary structure, i.e., a unitary outer wall 102.

[0036] exist Figure 3A In the exemplary embodiment shown, the outer wall 102 includes a combustor portion 104 that extends through the combustion section 26 and a turbine portion 106 that extends through at least the first turbine stage 82 and the second turbine stage 84 of the turbine section. In other embodiments, the turbine portion 106 may extend through fewer stages (e.g., through the one turbine stage just described) or through more stages (e.g., through one or more stages of the LP turbine 30 located downstream of the HP turbine 28). The combustor portion 104 and the turbine portion 106 are integrally formed such that the combustor portion 104 and the turbine portion 106 are a single, unitary structure, referred to herein as a unitary outer wall 102.

[0037] As used herein, the term "integral" means that the associated components, such as the outer wall 102, are made into a single piece during manufacture, i.e., the final integral component is a single piece. Thus, an integral component has a structure in which the integrated parts are inseparable and is different from a component comprising multiple separate component pieces that have been joined together and, once joined, are referred to as a single component even though the component pieces remain distinct and the single component is not inseparable (i.e., the pieces can be separated again). The final integral component may comprise a substantially continuous piece of material or, in other embodiments, may comprise multiple parts that are permanently bonded to one another. In any case, the various parts that form the integral component are integrated with one another so that the integral component is a single piece having inseparable parts.

[0038] like Figure 3AAs shown, the combustor section 104 of the integral structure forming the outer wall 102 includes the outer liner 108 of the combustor 80. The turbine section 106 includes the outer band 110 of the first turbine stage nozzle section 82N, the shroud 112 of the first turbine stage blade section 82B, the outer band 114 of the second turbine stage nozzle section 84N, and the shroud 116 of the second turbine stage blade section 84B. As described above, these outer boundary components are integrated into a single piece to form an integral structure as the outer wall 102. Therefore, in Figure 2 In the exemplary embodiment, outer liner 108 , outer band 110 , shroud 112 , outer band 114 , and shroud 116 are integrally formed, ie, constructed as a single unit or piece, to form unitary or unitary outer wall 102 .

[0039] In some embodiments, other portions of the flow path assembly 101 may be integrated into the integral structure of the outer wall 102, and in still other embodiments, at least a portion of the outer boundary and the inner boundary are made into a single integral component, such that the flow path assembly 101 may be referred to as a unitary flow path assembly. Figure 3B The burner portion 104 of the integral outer wall 102 may also include a burner dome 118 extending beyond the front end 88 of the burner 80. Figure 3B In the exemplary embodiment, outer liner 108, outer bands 110, shrouds 112, outer bands 114, shrouds 116, and combustor dome 118 are constructed as a single unit or piece to form a unitary or monolithic outer wall 102. That is, outer liner 108, outer bands 110, 114, shrouds 112, 116, and combustor dome 118 are integrally formed such that outer liner 108, outer bands 110, 114, shrouds 112, 116, and combustor dome 118 are a single, unitary structure.

[0040] As another example, refer to Figure 3C At least a portion of the inner wall 120 defining the inner boundary of the flow path 100 may be integrated with the outer wall 102 to form an integral flow path assembly 101. Figure 3C In the exemplary embodiment of the present invention, the combustor section 104 further includes a liner 122 such that the liner 122 is aligned with the Figure 3B Thus, the outer liner 108, outer band 110, shroud 112, outer band 114, shroud 116, combustor dome 118, and inner liner 122 are integrally formed such that the outer liner 108, outer bands 110, 114, shrouds 112, 116, combustor dome 118, and inner liner 122 are a single, unitary structure. Figure 3D In the exemplary embodiment of the present invention, the turbine section 106 further includes an inner band 124 of the first turbine stage nozzle section 82N such that the inner band 124 is aligned with the inner band 124. Figure 3C16, combustor dome 118, inner liner 122, and inner band 124 are integrally formed such that outer liner 108, outer bands 110, 114, shrouds 112, 116, combustor dome 118, inner liner 122, and inner band 124 are a single, unitary structure. Figure 3E In the exemplary embodiment of the present invention, the turbine section 106 further includes a plurality of nozzle airfoils 126 such that each of the plurality of nozzle airfoils 126 of the first turbine stage nozzle section 82N is aligned with the nozzle airfoil 126 of the first turbine stage nozzle section 82N. Figure 3D 14 , 116 , combustor dome 118 , inner liner 122 , inner band 124 , and nozzle airfoil 126 are integrally formed such that outer liner 108 , outer bands 110 , 114 , shrouds 112 , 116 , combustor dome 118 , inner liner 122 , inner band 124 , and nozzle airfoil 126 are a single, unitary structure.

[0041] Of course, the nozzle airfoil 126 of the first turbine stage nozzle portion 82N may be integrated with the outer wall 102 rather than the inner wall 120. For example, the plurality of nozzle airfoils 126 may be formed as a single unit or piece with the outer liner 108, outer band 110, shroud 112, outer band 114, shroud 116, such that the outer liner 108, outer bands 110, 114, shroud 112, 116, and the nozzle airfoil 126 are a single, unitary structure, i.e., the unitary outer wall 102. In other embodiments, the unitary outer wall 102 may further include the combustor dome 118, such that the outer liner 108, outer band 110, shroud 112, outer band 114, shroud 116, combustor dome 118, and the nozzle airfoil 126 are integrally formed or constructed as a single unit or piece. In other embodiments, a liner 122 may also be included such that the outer liner 108 , outer band 110 , shroud 112 , outer band 114 , shroud 116 , combustor dome 118 , liner 122 , and nozzle airfoil 126 are integrally formed as a single, unitary structure, i.e., a unitary outer wall 102 .

[0042] Figure 3F A partial perspective view of a portion of an integral flow path assembly 101 is provided having an outer wall 102 and an inner wall 120 formed as a single piece. Figure 3D Described and Figure 3FAs shown, in some embodiments of the combustion gas flow path assembly 101, the outer liner 108, the outer band 110, the shroud 112, the outer band 114, the shroud 116, the combustor dome 118, the inner liner 122, and the inner band 124 are integrally formed such that the outer liner 108, the outer bands 110, 114, the shrouds 112, 116, the combustor dome 118, the inner liner 122, and the inner band 124 are a single, unitary structure. Figure 3F It is also shown that a plurality of openings 142 for receiving the fuel nozzle assembly 90 and / or the swirler 92 may be defined in the front end 88 of the combustor 80 of the integrated flow path assembly 101. Furthermore, it should be understood that Figure 3F Only a portion of the integral flow path assembly 101 is shown, and although its entire periphery is not shown Figure 3F 1 , but the flow path assembly 101 is a single integral piece in the circumferential direction and the axial direction. Thus, the integral flow path assembly 101 defines a generally annular flow path between the outer wall 102 and the inner wall 120, ie, a generally ring-shaped flow path.

[0043] Integrating the various components of the outer and inner boundaries of flow path assembly 101, as described above, can reduce the number of separate components or parts within engine 10, as well as reduce the weight, leakage, and complexity of engine 10, compared to known gas turbine engines. For example, known gas turbine engines employ seals or sealing mechanisms at the interfaces between separate components of the flow path assembly in an attempt to minimize leakage of combustion gases from the flow path. By integrating the outer boundaries, for example, as described with respect to integral outer wall 102, the split points or interfaces between the outer combustor liner and the first turbine stage outer band, the first turbine stage outer band 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 boundaries, the split points or interfaces between the integral inner boundary components are eliminated, thereby eliminating leakage points at the inner boundaries and the seals or sealing mechanisms required at the inner boundaries. Thus, by utilizing integral components within the flow path assembly, unwanted leakage, as well as unnecessary weight and complexity, can be avoided. Those skilled in the art will appreciate additional advantages of integral outer wall 102, integral inner wall 120, and / or integral flow path assembly 101.

[0044] like Figures 3A to 3FAs shown, the outer wall 102 and the inner wall 120 define a generally annular flow path therebetween. That is, the integral outer wall 102 circumferentially surrounds the inner wall 120; in other words, the integral outer wall 102 is a single piece that extends 360° around the inner wall 120, thereby defining a generally annular or ring-shaped flow path therebetween. Thus, the combustor dome 118 extending beyond the forward end 88 of the combustor 80 is a generally annular combustor dome 118. Furthermore, the combustor dome 118 defines an opening 142 for receiving the fuel nozzle assembly 90 located at the forward end 88. The fuel nozzle assembly 90 provides the combustion chamber 86 with a mixture of fuel and compressed air from the compressor section, for example, which is combusted within the combustion chamber 86 to produce a flow of combustion gases through the flow path 100. The fuel nozzle assembly 90 may be attached to the combustor dome 118 or may "float" relative to the combustor dome 118 and the flow path 100, i.e., the fuel nozzle assembly 90 may not be attached to the combustor 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 combustor dome 118, but alternatively, the swirler 92 may float relative to the combustor dome 118 and the flow path 100. It should be understood that the fuel nozzle assembly 90 or the swirler 92 may float relative to the combustor dome 118 and the flow path 100 in both the radial direction R and the axial direction A, or in only one or the other of the radial direction R and the axial direction A. Furthermore, it should be understood that the combustor dome 118 may define a plurality of openings 142, each of which receives a swirler 92 or other portion of the fuel nozzle assembly 90.

[0045] like Figures 3A to 3F and discussed in more detail below Figures 4A to 4C as well as Figure 5A and Figure 5B As further shown in FIG, flow path assembly 101 generally defines a converging-diverging flow path 100. More specifically, outer wall 102 and inner wall 120 define a generally annular combustion chamber 86, which forms a forward portion of flow path 100. Moving aft or downstream of combustion chamber 86, outer wall 102 and inner wall 120 generally converge toward each other in the region of first turbine stage 82. Continuing downstream of first turbine stage 82, outer wall 102 and inner wall 120 generally then diverge in the region of second turbine stage 84. Outer wall 102 and inner wall 120 may continue to diverge downstream of second turbine stage 84. In an exemplary embodiment, for example, as Figure 3AAs shown and referring only to the unitary outer wall 102, a first turbine stage nozzle outer band portion 110 and a blade shroud portion 112 of the outer wall 102 converge toward the axial centerline 12. A second turbine stage nozzle outer band portion 114 and a blade shroud portion 116 of the outer wall 102 diverge away from the axial centerline 12. Thus, the outer boundary of the flow path 100 formed by the unitary outer wall 102 defines a converging-diverging flow path 100.

[0046] Go to Figure 4A and 4B , showing other exemplary embodiments of the present subject matter. Figure 4A A schematic cross-sectional view of a combustion section 26 and an HP turbine 28 of a turbine section is provided according to an exemplary embodiment. Figure 4B A schematic cross-sectional view of a combustion section 26 and an HP turbine 28 of a turbine section is provided according to another exemplary embodiment. Figure 4C A schematic cross-sectional view of a combustion section 26 and an HP turbine 28 of a turbine section is provided according to yet another exemplary embodiment.

[0047] exist Figure 4A 4B and 4C, the outer wall 102 is formed as a single integral structure, and the inner wall 120 is formed as another single integral structure, and the integral outer wall 102 and the integral inner wall 120 together define the flow path 100. However, it should be understood that the inner wall 120 does not need to be a single integral structure. For example, in Figure 4A In the embodiments shown in Figures 4B and 4C, the inner wall 120 may include an inner lining 122 formed separately from the inner belt 124.

[0048] As about Figures 3A to 3F As described, Figure 4A , 4B and 4C, the integral outer wall 102 defines the outer boundary of the flow path 100, and the inner wall 120 defines the inner boundary of the flow path 100. The integral outer wall 102 and the inner wall 120 together form the flow path assembly 101. The integral outer wall 102 extends from the front end 88 of the combustor 80 of the combustion section 26 through at least the first turbine stage 82 of the HP turbine 28, and in the illustrated embodiment, the integral outer wall 102 extends from the front end 88 to the rear end of the second turbine stage 84 of the HP turbine 28. The inner wall 120 includes at least an inner liner 122, and in the embodiment where the inner wall 120 is an integral inner wall, the integral inner wall 120 extends from the front end 88 of the combustor 80 through the first turbine stage nozzle portion 82N. Thus, as shown Figure 4A 4B and 4C , the outer wall 102 and the inner wall 120 define the combustion chamber 86 of the combustor 80 .

[0049] About Figures 3A to 3F The described embodiment is similar, Figure 4A, the one-piece outer wall 102 of the embodiment shown in 4B and 4C includes an outer lining 108, an outer band 110, a shield 112, an outer band 114 and a shield 116. In addition, in Figure 4A In the exemplary embodiment of the present invention, the integral outer wall 102 includes a combustor dome 118 defined at the forward end 88 of the combustor 80. Thus, the outer liner 108, the outer bands 110, 114, the shrouds 112, 116, and the combustor dome 118 are integrally formed or constructed as a single, unitary structure, i.e., the outer wall 102 is a single unit or piece including the combustor dome 118. Alternatively, as Figure 4B In the exemplary embodiment shown, unitary outer wall 102 includes a radially outer portion of combustor dome 118 such that outer liner 108, outer band 110, shroud 112, outer band 114, shroud 116, and a portion of combustor dome 118 are integrally formed or constructed as a single, unitary structure, i.e., outer wall 102 is a single unit or piece that includes a portion of combustor dome 118.

[0050] In addition, with regard to Figures 3A to 3F The described embodiment is similar, Figure 4A , the inner wall 120 of the embodiments shown in 4B and 4C includes at least the liner 122 of the combustor 80. In some embodiments, for example Figure 4A and 4B As shown, the inner wall 120 also includes an inner band 124 of the first turbine stage nozzle portion 82N. In such an embodiment, the inner liner 122 and the inner band 124 are integrally formed as a single, unitary structure, i.e., as a single unit or piece that can be referred to as a unitary inner wall 120. In other embodiments, such as Figure 4B As shown, the integral inner wall 120 may include a radially inner portion of the combustor dome 118 such that the inner liner 122 and a portion of the combustor dome 118 are integrally formed or constructed as a single integral structure or such that the inner liner 122, the inner band 124, and a portion of the combustor dome 118 are integrally formed or constructed as a single integral structure. That is, in some embodiments, the integral inner wall 120 is a single unit or piece that includes a portion of the combustor dome 118 (and may or may not include the inner band 124). In other embodiments, such as Figure 4C As shown, the integral inner wall 120 includes a combustor dome 118 defined at the forward end 88 of the combustor 80. Thus, the combustor dome 118 and the inner liner 122 (and, in some embodiments, the inner band 124) are integrally formed or constructed as a single, unitary structure, i.e., the inner wall 102 is a single unit or piece including the combustor dome 118.

[0051] Additionally, the first turbine stage nozzle airfoil 126 may be integrated with the outer wall 102 and / or the inner wall 120. As previously described, the first turbine stage nozzle airfoil 126 may be integrated with the outer wall 102, but in other embodiments, the first turbine stage nozzle airfoil 126 may be integrated with the inner wall 120 instead of the outer wall 102, or may be integrated with both the outer wall 102 and the inner wall 120. Whether formed separately from the walls 102, 120, integrated with the inner wall 120 to form a single, unitary structure therewith, integrated with the outer wall 102 to form a single, unitary structure therewith, or integrated with both the outer wall 102, 120 to form a single, unitary structure therewith, a plurality of nozzle airfoils 126 extend within the first turbine stage nozzle portion 82N from the inner wall 120 to the outer wall 102. Additionally, as described above, first turbine stage 82 includes a first-stage rotor 134 having a plurality of rotor blade airfoils 130 attached thereto. Downstream of first turbine stage 82, a plurality of nozzle airfoils 128 extend from an inner band 136 to outer wall 102 within second turbine stage nozzle portion 84N, and second turbine stage blade portion 84B includes a second-stage rotor 138 having a plurality of rotor blade airfoils 130 attached thereto.

[0052] exist Figure 4A In the embodiments of Figures 4B and 4C, the integral or one-piece outer wall 102 extends circumferentially around the integral or one-piece inner wall 120. That is, the outer wall 102 circumferentially surrounds the inner wall 120, or the one-piece outer wall 102 is a single piece that extends 360° around the inner wall 120. Thus, the outer wall 102 and the inner wall 120 define a generally annular flow path therebetween. Furthermore, the burner dome 118 extends beyond the front end 88 of the combustor 80 and, whether wholly or partially integrated into the one-piece outer wall 102 or wholly or partially integrated into the one-piece inner wall 120, is a generally annular burner dome 118.

[0053] in addition, Figure 4A , 4B and 4C, the flow path assembly 101 includes at least one opening 142 for receiving the fuel nozzle assembly 90. Figures 3A to 3F As described, in some embodiments, the fuel nozzle assembly 90 may be attached to the combustor dome 118, which may be as shown. Figure 4A In the embodiment of the present invention, the outer wall 102 is integrally integrated, or as Figure 4B 1 and 2. As shown in FIG, the combustor dome 118 is partially integrated with the outer wall 102 and the remaining portion is integrated with the inner wall 120. As also described, the combustor dome 118 can be as shown in FIG. Figure 4CAs shown, the fuel nozzle assembly 90 is integrally integrated with the inner wall 120 so that the fuel nozzle assembly 90 can be attached to the combustor dome portion of the integral inner wall 120. In other embodiments, the fuel nozzle assembly 90 is not attached to the combustor dome 118, but floats relative to the combustor dome 118 and the flow path 100. As shown, the fuel nozzle assembly 90 includes a swirler 92, which can be the portion of the fuel nozzle assembly 90 that is attached to the combustor dome 118 or the portion that floats relative to the combustor dome 118 and the flow path 100. As previously described, the fuel nozzle assembly 90 or the swirler 92 can float relative to the combustor dome 118 and the flow path 100 in the radial direction R and the axial direction A, or in only one or the other of the radial direction R and the axial direction A. In addition, as Figure 3F As shown, the combustor dome 118 may define a plurality of openings 142 , and each opening may receive the swirler 92 or other portion of the fuel nozzle assembly 90 .

[0054] Still refer to Figure 4A , 4B and 4C, the outer wall 102 and the inner wall 120 of a unit can define one or more features wherein the walls 102, 120 intersect with each other and, in some embodiments, can be attached to each other. For example, in Figure 4A In the embodiment of FIG. 8 , the outer wall 102 defines a flange 144 along a radially inner edge of the outer wall 102 at the front end 88 of the combustor 80 , and the inner wall 120 defines a flange 146 along a forward edge at the combustor front end 88 . Figure 4B In the embodiment of FIG. 1 , the outer wall flange 144 is defined along the edge of the burner dome portion of the integral outer wall 102 , and similarly, the inner wall flange 146 is defined along the edge of the burner dome portion of the integral inner wall 120 . Figure 4C As shown, the outer wall 102 may define an outer wall flange 144 along a forward edge of the outer wall 102 , and the inner wall 120 , including the combustor dome 118 in the illustrated embodiment, may define an inner wall flange 146 along a radially outer edge of the inner wall 120 . Figure 4A 4B and 4C show that the flow path 100 can be discontinuous between the inner wall 120 and the outer wall 102, that is, formed by separate inner and outer boundaries rather than by a continuous flow path 100 as shown in FIG. Figures 3C to 3F The integral inner and outer boundaries are shown as forming a .More specifically, the flow path 100 may be discontinuous, wherein an outer wall flange 144 and an inner wall flange 146 are defined.

[0055] Therefore, in Figure 4A In the embodiment of FIG. 1 , the outer wall 102 may be secured to the inner wall 120 at flanges 144 , 146 proximate the radially inner front portion of the combustor 80 . Alternatively, as Figure 4AThe flanges 144, 146 shown may define areas where the walls 102, 120 align or intersect one another, for example, the flanges 144, 146 may define a sliding joint between the walls 102, 120. Figure 4B In some embodiments, the outer wall 102 may be secured to the inner wall 120 at flanges 144, 146 proximate the radial centerline of the combustor dome 118. In other embodiments, such as Figure 4B The flanges 144, 146 shown may define areas where the walls 102, 120 align or intersect one another, for example, the flanges 144, 146 may define a sliding joint between the walls 102, 120. Figure 4C In an alternative embodiment to this embodiment, the outer wall 120 may be secured to the inner wall 120 at flanges 144, 146 proximate the radially outer front of the combustor 80, or as Figure 4C The flanges 144, 146 shown may define areas where the walls 102, 120 align or intersect one another. For example, the flanges 144, 146 may define a sliding joint between the walls 102, 120 at the radially outer front of the combustor 80. In other embodiments, the flanges 144, 146 may be defined at other locations such that the outer wall 102 and the inner wall 120 are aligned or intersected at locations other than the radially outer front of the combustor 80. Figure 4A , the positions shown in 4B and 4C are fixed, aligned or intersecting with each other.

[0056] Any suitable fasteners or other attachment means may be used to secure the outer wall 102 and inner wall 120 to the flanges 144, 146. For example, a plurality of holes may be defined in each flange 144, 146, and each hole in the outer wall flange 144 may be aligned with a hole in the inner wall flange 146 to receive a fastener in each pair of aligned holes. It should be understood that the outer wall 102 and inner wall 120 may also be attached to each other in other ways. Of course, in other embodiments as described above, the outer wall 102 and inner wall 120 may not be fixed to each other, but may be movable radially and / or axially relative to each other.

[0057] Now go to Figure 5A 5B and 5C provide schematic cross-sectional views of the combustion section 26 and the HP turbine 28 of the turbine section of the turbofan engine 10 according to other exemplary embodiments of the present subject matter. Figures 3B to 3F and Figures 4A to 4C The embodiments are different. Figure 5A 5B and 5C are not integrally or partially integrated with the outer wall 102 or the inner wall 120. That is, the burner dome 118 is a separate component from the outer wall 102 and the inner wall 120.

[0058] Therefore, if Figure 5A5B and 5C, the outer wall 102 is an integral outer wall including a combustor portion 104 and a turbine portion 106. The combustor portion 104 extends through the combustion section 26 of the engine 10, and the turbine portion 106 extends through at least the first turbine stage of the turbine section of the engine 10. Figures 5A to 5C In the illustrated embodiment, the integral outer wall 102 extends through the combustion section 26 to the aft end of the HP turbine 28, which includes the two turbine stages 82, 84. The combustor section 104 and the turbine section 106 are integrally formed as a single, unitary structure, namely the integral outer wall 102. For example, as described with respect to Figure 3A As shown and described, the combustor portion 104 of the integral outer wall 102 includes an outer liner 108 of the combustor 80. The turbine portion 106 of the integral outer wall 102 includes an outer band 110 of the first turbine stage nozzle portion 82N, a shroud 112 of the first turbine stage blade portion 82B, an outer band 114 of the second turbine stage nozzle portion 84N, and a shroud 116 of the second turbine stage blade portion 84B. The turbine portion 106 of the integral outer wall 102 may also include a plurality of nozzle airfoils 126 that are integrally formed or constructed with the outer liner 108, the outer bands 110, 114, and the shrouds 112, 116 to form a single, unitary structure, i.e., as a single unit or piece.

[0059] In addition, if Figure 5A , 5B and 5C , the inner wall 120 extends from the forward end 88 of the combustor 80 through at least the combustion section 26. For example, the inner wall 120 may include separate components that define the inner boundary of the flow path 100. In other embodiments, the inner wall 120 may be a one-piece inner wall 120 including the inner liner 122 and the inner band 124 that are integrally formed as a single, unitary structure, i.e., as a single unit or piece. As another example, the inner wall 120 may be a one-piece inner wall 120 including the inner liner 122, the inner band 124, and the first turbine stage nozzle airfoil 126 that are integrally formed as a single, unitary structure, i.e., as a single unit or piece. Additionally, in Figure 5A 5B and 5C, the flow path 100 may be discontinuous between the inner wall 120 and the outer wall 102, that is, by separate inner and outer boundaries rather than as shown in FIG. Figures 3C to 3F More specifically, the flow path 100 may be discontinuous between the combustor dome 118 and the outer wall 102 and between the combustor dome 118 and the inner wall 120 .

[0060] Specific reference Figure 5AA combustor dome 118 is positioned at the forward end 88 of the combustor 80 of the combustion section 26 and extends radially from the outer wall 102 to the inner wall 120. The combustor dome 118 is configured to move axially relative to the inner wall 120 and the outer wall 102, but may be attached to and, therefore, supported by, one or more fuel nozzle assemblies 90. More specifically, an axial sliding joint 150 is formed between the combustor dome 118 and each of the outer wall 102 and the inner wall 120, such that the combustor dome 118 can move axially, or float, relative to the inner wall 120 and the outer wall 102. Allowing the combustor dome 118 to float relative to the outer wall 102 and the inner wall 120 may help control the position of the fuel nozzle assembly 90 relative to the combustor dome 118 and the combustor 80. For example, the combustor dome 118, the outer wall 102, and the inner wall 120 may be made of a different material than the fuel nozzle assembly 90. As described in more detail below, in an exemplary embodiment, the combustor dome 118, outer wall 102, and inner wall 120 are made of a ceramic matrix composite (CMC) material, and the fuel nozzle assembly 90 may be made of a metallic material, such as a metal alloy. In such an embodiment, CMC materials thermally grow or expand at a different rate than metallic materials. Therefore, allowing the combustor dome 118 to move axially relative to the outer wall 102 and inner wall 120 may allow for tighter control over the immersion of the swirler 92 of the fuel nozzle assembly 90 within the combustor dome 118 and within the combustor 80 than over the attachment of the combustor dome 118 to the outer wall 102 and inner wall 120. Tighter control over the position of the fuel nozzle assembly 90 and its components relative to the combustor 80 may reduce variability in the operability and performance of the engine 10.

[0061] Furthermore, the outer wall 102 and the inner wall 120 may also move axially and radially relative to the burner dome 118. By decoupling the burner dome 118 from the walls 102, 120 and allowing relative movement between the walls 102, 120 and the burner dome 118, the stress coupling between the outer and inner walls 102, 120 and the burner dome 118 may be relieved. Furthermore, any leakage between the uncoupled burner dome 118 and the outer and inner walls 102, 120 may be used as a purge and / or membrane initiator flow.

[0062] like Figure 5AAs shown, the burner dome 118 includes outer wings 152 and inner wings 154. The outer wings 152 extend rearwardly along the outer wall 102, and the inner wings 154 extend rearwardly along the inner wall 120. The wings 152, 154 can help guide the burner dome 118 as it moves relative to the outer wall 102 and the inner wall 120, and the wings 152, 154 can also help maintain the radial position or alignment of the burner dome 118 as the burner dome 118 moves axially. As previously described, the wings can provide a consistent gap between the dome 118 and the walls 102, 120 for purge and / or membrane initiator flow.

[0063] Go to Figure 5B In other embodiments, each wing 152, 154 may extend forwardly from the combustor dome body 156 rather than Figure 5A The forward-extending wings 152, 154 may be used to mount the combustor dome 118 to a component other than the fuel nozzle assembly 90 / swirler 92, such as a metal dome supporting the fuel nozzle assembly 90, and / or to either or both of the outer wall 102 and the inner wall 120 at the forward end 88 of the combustor 80. In some embodiments, the forward-extending wings 152, 154 of the combustor dome 118 may be pinned or otherwise attached to the outer wall 102 and the inner wall 120, such as Figure 5B In other embodiments, one of the wings 152 , 154 may extend forward and the other wing 152 , 154 may extend rearward relative to the body 156 , and the combustor dome 118 may be attached to the fuel nozzle assembly 90 or another component.

[0064] Now refer to Figure 5C , another exemplary embodiment of a separate combustor dome 118 and outer and inner walls 102, 120 is shown. Figure 5C In the illustrated embodiment, the burner dome 118 includes forwardly extending inner wings 154 but lacks outer wings 152; instead, an outer end 158 of the burner dome 118 extends to the outer wall 102. To retain the burner dome 118 and seal against combustion gas leakage around the dome, the inner wings 154 are pinned to the inner wall 120 at the front end 88 of the burner 80, and the outer ends 158 are preloaded against the outer wall 102. More specifically, a spring element 160 is pinned to the outer wall 102 at the burner front end 88 and presses against the body 156 of the burner dome 118 to preload the outer end 158 of the burner dome 118 into a lip 162 defined in the outer wall 102. By utilizing Figure 5C The mounting configuration shown in , may provide positive retention and sealing of the combustor dome 118 while minimizing thermal stresses in the dome, which is particularly useful when the combustor dome 118 is made of CMC material.

[0065] As previously mentioned, outer wall 102, inner wall 120, and combustor dome 118 may comprise CMC material. More specifically, in the exemplary embodiment, combustor section 104 and turbine section 106 of flow path assembly 101 are integrally formed from CMC material, such that the resulting unitary structure is a CMC component. For example, where combustor section 104 comprises outer liner 108 of combustor 80, and turbine section 106 comprises outer band 110 of first turbine stage nozzle section 82N, shroud 112 of first turbine stage blade section 82B, outer band 114 of second turbine stage nozzle section 84N, and shroud 116 of second turbine stage blade section 84B, outer liner 108, outer bands 110, 114, and shrouds 114, 116 may be integrally formed from CMC material to create a unitary CMC outer wall 102. As described above, in other embodiments, additional CMC components may be integrally formed with the outer liner 108, outer bands 110, 114, and shrouds 114, 116 to construct a unitary CMC outer wall 102. Similarly, the inner wall 120 may be formed from a CMC material. For example, where the inner wall 120 includes separate components, such as the inner liner 122, inner bands 124, 136, and blade platform 132, each component of the inner wall 120 may be formed from a CMC material. In embodiments where two or more components are integrated to form a unitary inner wall 120, these components may be integrally formed from a CMC material to construct a unitary CMC inner wall 120.

[0066] 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. Patent 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, as well as U.S. Patent Application Publication No. 2004 / 0067316. These processes typically require the use of multiple pre-impregnated (prepreg) layers to manufacture the CMC. For example, the layer material may include prepreg materials 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 prepreg layer is in the form of a "tape" and includes the desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and an organic resin binder. Prepreg tapes can be formed by impregnating reinforcement materials with a slurry containing a ceramic precursor and a binder. The preferred material for the precursor will depend on the specific composition desired for the ceramic matrix of the CMC part. For example, if the desired matrix material is SiC, the preferred materials are SiC powder and / or one or more carbonaceous materials. Notable carbonaceous materials include carbon black, phenolic resins, and furan resins, including furfuryl alcohol (C4H3OCH2OH). Other typical slurry ingredients include an organic binder (e.g., polyvinyl butyral (PVB)) to promote the flexibility of the prepreg tape and a solvent (e.g., toluene and / or methyl isobutyl ketone (MIBK)) to promote the flowability of the slurry to enable impregnation of the fiber reinforcement material. The slurry may further contain one or more particulate fillers intended to be present in the ceramic matrix of the CMC part, for example, in the case of a Si-SiC matrix, silicon and / or SiC powder. As previously mentioned, chopped fibers or whiskers or other materials may also be embedded in the matrix. Other compositions and processes for preparing composite articles, more specifically, other slurry and prepreg compositions, such as those described in US Patent Application Publication No. 2013 / 0157037, may also be used.

[0067] The resulting prepreg tape can be stacked with other tapes so that the CMC part formed from the tape comprises multiple laminates, each laminate derived from a separate prepreg tape. Each laminate comprises a ceramic fiber reinforcement encased in a ceramic matrix, which is formed, for example, in whole or in part, by conversion of a ceramic matrix precursor during a firing and densification cycle, as described more fully below. In some embodiments, the reinforcement is in the form of an array of unidirectional tows, each tow comprising continuous fibers or filaments. Alternatives to unidirectional tow arrays may also be used. Furthermore, the appropriate fiber diameter, tow diameter, and center-to-center tow spacing will depend on the specific application, the thickness of the particular laminate and the tape from which it is formed, as well as other factors. As described above, other prepreg materials or non-prepreg materials may also be used.

[0068] After the tapes or layers are laid up to form a stack, the stack is compacted and, if appropriate, cured under high pressure and temperature to produce a preform. The preform is then heated (fired) in a vacuum or inert atmosphere to decompose the binder, remove the solvent, and convert the precursor into the desired ceramic matrix material. As a result of the decomposition of the binder, a porous CMC body is produced, which can undergo densification, such as melt infiltration (MI), to fill the pores and produce a CMC part. The specific processing techniques and parameters used for the above-mentioned processing will depend on the specific composition of the material. For example, a silicon CMC part can be formed from a fiber material infiltrated with molten silicon, such as by a process commonly known as the Silcomp process. Another technique for manufacturing CMC parts is a method known as slurry casting melt infiltration (MI) processing. In one method of manufacturing using the slurry casting MI method, a CMC is produced by first providing a balanced two-dimensional (2D) woven fabric layer containing silicon carbide (SiC) fibers, the balanced two-dimensional (2D) woven fabric having two weave directions that are substantially at 90° angles to each other and having substantially the same number of fibers extending in both directions of the fabric. The term "silicon carbide-containing fiber" refers to a fiber having a composition that includes silicon carbide, and preferably is essentially silicon carbide. For example, the fiber can have a silicon carbide core surrounded by carbon, or conversely, the fiber can have a carbon core surrounded by or encapsulated by silicon carbide.

[0069] Other techniques for forming CMC parts include polymer infiltration and pyrolysis (PIP) and oxide / oxide processing. In the PIP process, a silicon carbide fiber preform is infiltrated with a pre-ceramic polymer, such as polysilazane, and then heat treated to form a SiC matrix. In the oxide / oxide process, aluminum or aluminosilicate fibers can be pre-impregnated and then laminated into a preselected geometry. Parts can also be made from carbon fiber reinforced silicon carbide matrix (C / SiC) CMCs. The C / SiC process includes a carbon fiber preform laid up on a tool in a preselected geometry. The tool is made of a graphite material, as used in a slurry casting method for SiC / SiC. In a chemical vapor infiltration process at approximately 1200°C, the fiber preform is supported by a tool, thereby forming a C / SiC CMC part. In other embodiments, 2D, 2.5D and / or 3D preforms can be used for MI, CVI, PIP or other processes. For example, cut layers of 2D fabric can be stacked in alternating weave directions as described above, or the filaments can be twisted or braided and combined with 3D weaving, stitching, or needle punching to form a 2.5D or 3D preform with a multiaxial fiber architecture. Other ways of forming a 2.5D or 3D preform can also be used, for example, using other weaving or weaving methods or utilizing 2D fabrics.

[0070] Therefore, various processes can be used to form a unitary structure, e.g. Figure 3A The outer wall 102 shown in FIG is a one-piece CMC component. More specifically, multiple layers of CMC material can be used to form each one-piece structure. Multiple layers can be interlaced with each other to integrate the various parts of the one-piece structure. As an example, Figure 3A The integral outer wall 102 can be made of a plurality of outer liners, a plurality of first turbine stage outer band layers, a plurality of first turbine stage shroud layers, a plurality of second turbine stage outer band layers and a plurality of second turbine stage shroud layers. Where the outer liners intersect with the first turbine stage outer band layer, the ends of the outer liners may alternate with the ends of the outer band layers so that the layers used to form the outer liner portion are integrated with the layers used to form the first turbine stage outer band portion of the integral outer wall 102. That is, any joints between the layers forming the integral outer wall 102 can be formed by alternating the layers on one side of the joint with the layers on the other side of the joint. In this way, the layers used to form the integral outer wall 102 can be interspersed with integrated layers, thereby integrating each portion of the integral outer wall 102. Of course, the CMC layers can also be stacked in other ways to form an integral structure. Additionally, stacking multiple CMC layers can include defining an integral structure or other components (e.g., when not integrated with the inner band 124 to form a portion such as Figure 5A and 5B The embodiment shown in FIG. 1 shows an integral inner wall 120 or a liner 122 separating the combustor dome 118 , such as openings 142 in the combustor front end 88 , the outer wall flange 144 , and the inner wall flange 146 .

[0071] After laying up multiple CMC layers to define a monolithic CMC part preform, the preform is cured to produce a single, monolithic CMC part, 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 the green, monolithic outer wall 102. The green, monolithic outer wall 102 can then be placed in a furnace to burn off excess binder, etc., and then placed in a furnace with a silicon wafer or frit and fired to infiltrate the monolithic outer wall 102 with at least silicon melt. More specifically, for the monolithic outer wall 102 formed from the CMC layers of the prepreg tape produced as described above, heating (i.e., firing) the green part in a vacuum or inert atmosphere decomposes the binder, removes the solvent, and converts the precursor to the desired ceramic matrix material. The decomposition of the binder produces a porous CMC body; the body can then undergo densification, such as melt infiltration (MI), to fill the pores. In the foregoing example, in which a single outer wall 102 in a green state is fired with silicon, the outer wall 102 undergoes silicon melt infiltration. However, densification may 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 / oxide treatment, and with any suitable material, including but not limited to silicon. In one embodiment, densification and firing may be performed in a vacuum furnace or an inert atmosphere having an atmosphere established at a temperature above 1200° C. to allow silicon or other suitable material or combination of materials to melt infiltrate the component. The densified CMC body hardens into the final, one-piece CMC outer wall 102. In some embodiments, the final one-piece structure may be finish-machined, for example, to bring the structure within tolerances or to define an opening 142 in the front end 88, and / or an environmental barrier coating (EBC) may be applied to the one-piece structure, for example, to protect the one-piece structure from the effects of the hot combustion gases 66. It should be understood that other methods or processes of forming CMC components (eg, the integral CMC outer wall 102 , the integral CMC inner wall 120 , etc.) may also be used.

[0072] Additionally or alternatively, other processes for producing integral components may be used to form the integral outer wall 102 and / or the integral inner wall 120, and the integral structure may be formed from other materials. In some embodiments, an additive manufacturing process may be used to form the integral outer wall 102 and / or the integral inner wall 120. For example, the integral outer wall 102 and / or the integral inner wall 120 may be produced using additive processes or other known processes, 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), and electron beam sintering. Typically, additive processes use three-dimensional information of a component (e.g., a three-dimensional computer model) to manufacture the component. The three-dimensional information is converted into a plurality of slices, each slice defining a cross-section of the component at a predetermined height for the slice. The component is then "built" slice by slice or layer by layer until it is complete. Superalloy metal materials or other suitable materials may be used in additive processes to form the integral outer wall 102 and / or the integral inner wall 120. In other embodiments, a forging or casting process may be used to form the integral outer wall 102 and / or the integral inner wall 120. Other suitable processes or methods may also be used.

[0073] This written description uses examples to disclose the invention, including the best mode, and also to enable those skilled in the art to practice the invention, including making and using any devices or systems and performing any combined methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that are insubstantially different from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

Claims

1. A flow path assembly for a gas turbine engine, characterized in that: The flow path assembly comprises: a combustor dome positioned at a forward end of a combustor of a combustion section of the gas turbine engine, the combustor dome including inner wings; an outer wall extending from the front end of the combustor through the combustion section; and an inner wall extending from the front end of the burner through the combustion section, wherein the burner dome is a separate component from the outer wall and the inner wall and extends radially from the outer wall to the inner wall, wherein the inner wing extends forward along the inner wall, wherein fasteners extend through the inner wings and the inner wall to attach the burner dome to the inner wall, and wherein the burner dome is attached only to the inner wall.

2. The flow path assembly according to claim 1, wherein: wherein the burner dome further includes an outer end radially opposite the inner wing and a body extending from the outer end to the inner wing, and wherein a spring element is fixed to the outer wall at the front end of the burner with a pin, the spring element pressing against the body of the burner dome to preload the outer end against the outer wall.

3. The flow path assembly according to claim 2, wherein: Wherein the outer wall defines a lip on a radially inner surface of the outer wall adjacent the combustor dome, and wherein the outer end is configured to contact the lip to preload the outer end against the outer wall.

4. The flow path assembly according to claim 1, wherein: wherein the outer wall comprises an outer liner of the combustor, an outer band of a nozzle portion of a first turbine stage of a turbine section of the gas turbine engine, and a shroud of a blade portion of the first turbine stage, and wherein the outer liner, the outer band, and the shroud are integrally formed as a single, unitary structure.

5. The flow path assembly according to claim 4, wherein: wherein the inner wall comprises an inner liner of the combustor and an inner band of a nozzle portion of the first turbine stage, and wherein the inner liner and the inner band are integrally formed as a single, unitary structure.

6. The flow path assembly according to claim 1, wherein: Wherein the outer wall, the inner wall and the combustor dome comprise ceramic matrix composite materials.

7. A flow path assembly for a gas turbine engine, characterized in that: The flow path assembly extends from a combustor through a turbine section of the gas turbine engine, the turbine section including a first turbine stage immediately downstream of the combustor and a second turbine stage immediately downstream of the first turbine stage, the flow path assembly defining a flow path, the flow path assembly comprising: an inner wall defining an inner boundary of the flow path, the inner wall comprising an inner liner of the combustor; an outer wall defining an outer boundary of the flow path, the outer wall comprising an outer liner of the combustor, and a burner dome positioned at a front end of the burner, the burner dome comprising a body and inner wings extending forward from the body along the inner wall, wherein fasteners extend through the inner wings and the inner wall to attach the burner dome to the inner wall, wherein the burner dome is attached only to the inner wall.

8. The flow path assembly according to claim 7, wherein: The burner dome further includes an outer end extending radially to the outer wall.

9. The flow path assembly according to claim 8, wherein: wherein a spring element is pinned to the outer wall at the front end of the burner, and wherein the spring element is configured to press against the body of the burner dome to preload the outer end of the burner dome into a lip defined in the outer wall.

10. The flow path assembly according to claim 9, wherein: wherein the burner dome is formed from a ceramic matrix composite material.

11. The flow path assembly according to claim 7, wherein: wherein the outer wall further comprises an outer band of a nozzle portion of the first turbine stage and a shroud of a blade portion of the first turbine stage, and wherein the outer liner, the outer band, and the shroud are integrally formed from a ceramic matrix composite material as a single, unitary structure.

12. The flow path assembly according to claim 7, wherein: Wherein the combustor dome defines an opening for receiving a fuel nozzle assembly.

13. A gas turbine engine, characterized in that: include: a combustion section, said combustion section including a burner; and a turbine section comprising a first turbine stage positioned immediately downstream of the combustion section and a second turbine stage positioned immediately downstream of the first turbine stage, wherein the combustion section and the turbine section define a flow path, wherein the combustion section includes an inner liner defining an inner boundary of the flow path through the combustion section, an outer liner defining an outer boundary of the flow path through the combustion section, and a combustor dome positioned at a forward end of the combustor and extending between the inner liner and the outer liner, wherein each of the first and second turbine stages of the turbine section includes a nozzle portion and a blade portion, each nozzle portion including an inner band defining an inner boundary of the flow path through the nozzle portion and an outer band defining an outer boundary of the flow path through the nozzle portion, each blade portion including a shroud defining an outer boundary of the flow path through the blade portion, wherein the combustor dome is a separate component from the inner liner and the outer liner, wherein the combustor dome comprises a body and inner wings extending forward from the body along the liner, wherein fasteners extend through the inner wings and the liner to attach the combustor dome to the liner, and wherein the combustor dome is attached only to the liner.

14. The gas turbine engine according to claim 13, wherein: Further including: A fuel nozzle assembly has a swirler extending through an opening in the combustor dome.

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