Patent Publication Number: US-2022213798-A1

Title: Methods and Assemblies for Attaching Airfoils Within a Flow Path

Description:
FIELD 
     The present subject matter relates generally to gas turbine engines. More particularly, the present subject matter relates to outer and inner flow path boundary configurations for receipt of stator airfoils, as well as methods for assembling stator airfoils to a gas turbine flow path assembly. 
     BACKGROUND 
     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 and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere. 
     More particularly, 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 contain a plurality of stationary nozzle airfoils as well as a plurality of blade airfoils attached to a rotor that is driven by the flow of combustion gases against the blade airfoils. The turbine section may have other configurations as well. In any event, a flow path is defined by an inner boundary and an outer boundary, which both extend from the combustor through the stages of the turbine section. 
     Typically, the inner and outer boundaries defining the flow path comprise separate components. For example, an outer liner of the combustor, a separate outer band of a nozzle portion of a turbine stage, and a separate shroud of a blade portion of the turbine stage usually define at least a portion of the outer boundary of the flow path. However, utilizing separate components to form each of the outer boundary and the inner boundary requires a great number of parts, e.g., one or more seals may be required at each interface between the separate components to minimize leakage of fluid from the flow path, which can increase the complexity and weight of the gas turbine engine without eliminating leakage points between the separate components. Therefore, flow path assemblies may be utilized that have a unitary construction, e.g., a unitary outer boundary structure, where two or more components of the outer boundary are integrated into a single piece, and/or a unitary inner boundary structure, where two or more components of the inner boundary are integrated into a single piece. 
     A unitary construction of the flow path assembly may be furthered by various ways of assembling the turbine nozzle airfoils, which also may be referred to as stator vanes, with the outer boundary structure and the inner boundary structure. For example, the outer boundary structure and/or the inner boundary structure each may be constructed as a unitary structure or together may be constructed as a single unitary structure, with the nozzle airfoils inserted and secured during subsequent assembly. As another example, the nozzle airfoils may be integrally formed with one of the outer boundary structure or the inner boundary structure and attached to the other boundary structure during subsequent assembly. Separating the nozzle airfoils from the outer and/or inner boundary structures of the flow path assembly may simplify manufacturing, as well as reduce internal stresses compared to flow path assemblies comprising nozzle airfoils that are integral with both the outer and inner boundary structures. 
     Accordingly, improved flow path assemblies would be desirable. For example, a flow path assembly utilizing a unitary outer wall to define its outer boundary and having a plurality of nozzle airfoils received in pockets in the outer and inner boundaries would be beneficial. As another example, a flow path assembly utilizing a unitary inner wall to define its outer boundary and having a plurality of nozzle airfoils received in pockets in the outer and inner boundaries would be advantageous. Additionally, an inner wall that defines a plurality of slots for receipt of bayonets or projections from a plurality of nozzle airfoils to secure the nozzle airfoils to the inner wall would be useful. 
     BRIEF DESCRIPTION 
     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. 
     In one exemplary embodiment of the present disclosure, a flow path assembly for a gas turbine engine is provided. The flow path assembly comprises an inner wall; a unitary outer wall; and a plurality of nozzle airfoils, each nozzle airfoil having an inner end radially opposite an outer end. The unitary outer wall includes a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The combustor portion and the turbine portion are integrally formed as a single unitary structure. Further, the inner wall and the unitary outer wall define a combustor of the combustion section. Also, the unitary outer wall defines a plurality of outer pockets, each outer pocket configured for receipt of the outer end of one of the plurality of nozzle airfoils. The inner wall includes a forward segment and an aft segment and defines a plurality of inner pockets such that a portion of each inner pocket is defined by the forward segment and a remaining portion of each inner pocket is defined by the aft segment. Each inner pocket is configured for receipt of the inner end of one of the plurality of nozzle airfoils such that a nozzle airfoil extends from each inner pocket to a respective outer pocket. 
     In another exemplary embodiment of the present disclosure, a flow path assembly for a gas turbine engine is provided. The flow path assembly comprises an inner wall; an outer wall; and a plurality of nozzle airfoils, each nozzle airfoil having an inner end radially opposite an outer end. The inner wall and the unitary outer wall define a combustor of the combustion section. Moreover, the inner wall defines a plurality of inner pockets, each inner pocket configured for receipt of the inner end of one of the plurality of nozzle airfoils. The outer wall includes a forward segment and an aft segment and defines a plurality of outer pockets such that a portion of each outer pocket is defined by the forward segment and a remaining portion of each outer pocket is defined by the aft segment. Each outer pocket is configured for receipt of the outer end of one of the plurality of nozzle airfoils such that a nozzle airfoil extends from each inner pocket to a respective outer pocket. 
     In a further exemplary embodiment of the present disclosure, a flow path assembly for a gas turbine engine is provided. The flow path assembly comprises an inner wall defining a plurality of bayonet slots and a plurality of recesses along an aft surface of the inner wall. The flow path assembly also comprises a unitary outer wall including a combustor portion extending through a combustion section of the gas turbine engine and a turbine portion extending through at least a first turbine stage of a turbine section of the gas turbine engine. The turbine portion includes a plurality of nozzle airfoils, and the combustor portion and the turbine portion are integrally formed as a single unitary structure such that the plurality of nozzle airfoils is integral with the outer wall. The flow path assembly further comprises a first support member positioned radially inward of the inner wall to support the inner wall and a second support member positioned axially aft of the first support member. The second support member includes a plurality of tabs. An inner end of each nozzle airfoil is positioned against the inner wall, and each tab is received in a recess of the plurality of recesses defined in the inner wall such that the second support member fits against the aft surface of the inner wall to cover the plurality of bayonet slots. 
     These and other features, aspects and advantages of the present invention 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 invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  provides a schematic cross-section view of an exemplary gas turbine engine according to various embodiments of the present subject matter. 
         FIG. 2  provides a schematic exploded cross-section view of a combustion section and a high pressure turbine section of the gas turbine engine of  FIG. 1  according to an exemplary embodiment of the present subject matter. 
         FIG. 3A  provides a schematic cross-section view of the combustion section and high pressure turbine section of  FIG. 2  according to an exemplary embodiment of the present subject matter. 
         FIGS. 3B, 3C, 3D, and 3E  provide schematic cross-section views of the combustion section and high pressure turbine section of  FIG. 2  according to other exemplary embodiments of the present subject matter. 
         FIG. 3F  provides a partial perspective view of a portion of an integral outer boundary structure and inner boundary structure of the combustion section and high pressure turbine section of  FIG. 2  according to an exemplary embodiment of the present subject matter. 
         FIG. 4  provides a schematic cross-section view of a portion of a flow path assembly according to an exemplary embodiment of the present subject matter. 
         FIG. 5  provides a schematic cross-section view of a portion of a flow path assembly according to another exemplary embodiment of the present subject matter. 
         FIG. 6  provides a schematic cross-section view of the flow path assembly of  FIG. 5 , where the view is radially downward. 
     
    
    
     DETAILED DESCRIPTION 
     Reference will now be made in detail to present 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. Like or similar designations in the drawings and description have been used to refer to like 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 signify location or importance of the individual components. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows and “downstream” refers to the direction to which the fluid flows. 
     Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,  FIG. 1  is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of  FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine  10 , referred to herein as “turbofan engine  10 .” As shown in  FIG. 1 , the turbofan engine  10  defines an axial direction A (extending parallel to a longitudinal centerline  12  provided for reference) and a radial direction R. In general, the turbofan  10  includes a fan section  14  and a core turbine engine  16  disposed downstream from the fan section  14 . 
     The exemplary core turbine engine  16  depicted generally includes a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor  22  and a high pressure (HP) compressor  24 ; a combustion section  26 ; a turbine section including a high pressure (HP) turbine  28  and a low pressure (LP) turbine  30 ; and a jet exhaust nozzle section  32 . A high pressure (HP) shaft or spool  34  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 turbofan engine  10 , additional spools may be provided such that engine  10  may be described as a multi-spool engine. 
     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 depicted, fan blades  40  extend outward from disk  42  generally along the radial direction R. The fan blades  40  and disk  42  are together rotatable about the longitudinal axis  12  by LP shaft  36 . In some embodiments, a power gear box having a plurality of gears may be included for stepping down the rotational speed of the LP shaft  36  to a more efficient rotational fan speed. 
     Referring still to the exemplary embodiment of  FIG. 1 , disk  42  is covered by rotatable front nacelle  48  aerodynamically contoured to promote an airflow through the plurality of fan blades  40 . Additionally, the exemplary fan section  14  includes an annular fan casing 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 appreciated that nacelle  50  may be configured to be supported relative to the core turbine engine  16  by a plurality of circumferentially-spaced outlet guide vanes  52 . Moreover, a downstream section  54  of the nacelle  50  may extend over an outer portion of the core turbine engine  16  so as to define a bypass airflow passage  56  therebetween. 
     During operation of the turbofan engine  10 , a volume of air  58  enters turbofan  10  through an associated inlet  60  of the nacelle  50  and/or fan section  14 . As the volume of air  58  passes across fan blades  40 , a first portion of the air  58  as indicated by arrows  62  is directed or routed into the bypass airflow passage  56  and a second portion of the air  58  as indicated by arrows  64  is directed or routed into the LP compressor  22 . The ratio between the first portion of air  62  and the second portion of air  64  is commonly known as a bypass ratio. The pressure of the second portion of air  64  is then increased as it is routed through the high pressure (HP) compressor  24  and into the combustion section  26 , where it is mixed with fuel and burned to provide combustion gases  66 . 
     The combustion gases  66  are routed through the HP turbine  28  where a portion of thermal and/or kinetic energy from the combustion gases  66  is extracted via sequential stages of HP turbine stator vanes  68  that are coupled to the outer casing  18  and HP turbine rotor blades  70  that are coupled to the HP shaft or spool  34 , thus causing the HP shaft or spool  34  to rotate, thereby supporting operation of the HP compressor  24 . The combustion gases  66  are then routed through the LP turbine  30  where a second portion of thermal and kinetic energy is extracted from the combustion gases  66  via sequential stages of LP turbine stator vanes  72  that are coupled to the outer casing  18  and LP turbine rotor blades  74  that are coupled to the LP shaft or spool  36 , thus causing the LP shaft or spool  36  to rotate, thereby supporting operation of the LP compressor  22  and/or rotation of the fan  38 . 
     The combustion gases  66  are subsequently routed through the jet 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 substantially increased as the first portion of air  62  is routed 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 routing the combustion gases  66  through the core turbine engine  16 . 
     It will be appreciated that, although described with respect to turbofan  10  having 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 turboprops, turboshafts, turbojets, industrial and marine gas turbine engines, and/or auxiliary power units. 
     In some embodiments, components of turbofan engine  10 , particularly components within hot gas path  78 , such as components of combustion section  26 , HP turbine  28 , and/or LP turbine  30 , may comprise a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Of course, other components of turbofan engine  10 , such as components of HP compressor  24 , may comprise a CMC material. Exemplary CMC materials utilized for such components may include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron&#39;s SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon&#39;s NICALON®, Ube Industries&#39; TYRANNO®, and Dow Corning&#39;s SYLRAMIC®), alumina silicates (e.g., Nextel&#39;s  440  and  480 ), and chopped whiskers and fibers (e.g., Nextel&#39;s  440  and SAFFIL®), 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, bundles of the fibers, which may include a ceramic refractory material coating, are formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together (e.g., as plies) to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration or chemical vapor infiltration with silicon, to arrive at a component formed of a CMC material having a desired chemical composition. In other embodiments, the CMC material may be formed as, e.g., a carbon fiber cloth rather than as a tape. 
     As stated, components comprising a CMC material may be used within the hot gas path  78 , such as within the combustion and/or turbine sections of engine  10 . As an example, the combustion section  26  may include a combustor formed from a CMC material and/or one or more stages of one or more stages of the HP turbine  28  may be formed from a CMC material. However, CMC components may be used in other sections as well, such as the compressor and/or fan sections. Of course, in some embodiments, other high temperature materials and/or other composite materials may be used to form one or more components of engine  10 . 
       FIG. 2  provides an exploded view of a schematic cross-section of the combustion section  26  and the HP turbine  28  of the turbine section of the turbofan engine  10  according to an exemplary embodiment of the present subject matter.  FIG. 3A  provides an unexploded schematic cross-sectional view of the combustion section  26  and the HP turbine  28  of  FIG. 2  that focuses on an outer boundary of a flow path through the combustion section  26  and 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 particularly, for the depicted embodiment, HP turbine  28  includes a first turbine stage  82  and a second turbine stage  84 . In other embodiments, the HP turbine  28  may comprise 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 . Further, each turbine stage  82 ,  84  comprises a nozzle portion and a blade portion; the first turbine stage  82  includes nozzle portion  82 N and blade portion  82 B, and the second turbine stage  84  includes nozzle portion  84 N and blade portion  84 B. The nozzle portion  82 N of the first turbine stage  82  is located immediately downstream of the combustion section  26 , such that the nozzle portion  82 N of the first turbine stage  82  also may be referred to as a combustor discharge nozzle. Moreover, combustor  80  defines a generally annular combustion chamber  86  such that the combustor  80  may be described as a generally annular combustor. 
     Additionally, as described in greater detail below, a flow path  100  through the combustion section  26  and the HP turbine  28  is defined by an outer boundary and an inner boundary 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 HP turbine  28 ; thus, the flow path  100  may comprise at least a portion of the hot gas path  78  described above. Further, in other embodiments, the flow path  100  also may extend through LP turbine  30  and jet exhaust  32 ; in still other embodiments, the flow path  100  also may extend forward upstream of the combustion section  26 , e.g., into HP compressor  24 . As such, it will be appreciated that the discussion herein of the present subject matter with respect to combustion section  26  and HP turbine  28  is by way of example only and also may apply to different configurations of gas turbine engines and flow paths  100 . 
     As shown in the exploded view of  FIG. 2 , the outer and inner boundaries may be defined by an outer wall  102  and an inner wall  120 , respectively, which may include several portions of the combustion section  26  and HP turbine  28 . For instance, the combustor  80  includes an outer liner  108  defining an outer boundary of the flow path through the combustor  80 . Each nozzle portion  82 N,  84 N comprises an outer band defining an outer boundary of a flow path through the nozzle portion of each turbine stage, and each blade portion  82 B,  84 B comprises a shroud defining an outer boundary of a flow path through the blade portion of each turbine stage. More particularly, as shown in  FIG. 2 , the first turbine stage nozzle portion  82 N comprises outer band  110 , first turbine stage blade portion  82 B comprises shroud  112 , second turbine stage nozzle portion  84 N comprises outer band  114 , and second turbine stage blade portion  84 B comprises shroud  116 . These portions of the combustion section  26  and HP turbine  28  may comprise at least a portion of the outer wall  102 , as described in greater detail below. 
     Further, as illustrated in  FIG. 2 , the combustor  80  includes an inner liner  122  defining an inner boundary of the flow path through the combustor  80 . Each nozzle portion  82 N,  84 N comprises an inner band defining an inner boundary of the flow path through the nozzle portion of each turbine stage, and each blade portion  82 B,  84 B comprises one or more blade platforms that define an inner boundary of the flow path through the blade portion of each turbine stage. More particularly, as shown in  FIG. 2 , the first turbine stage nozzle portion  82 N comprises inner band  124 , first turbine stage blade portion  82 B comprises blade platforms  132 , second turbine stage nozzle portion  84 N comprises inner band  136 , and second turbine stage blade portion  84 B comprises blade platforms  132 . These portions of the combustion section  26  and HP turbine  28  may comprise at least a portion of the inner wall  122 , as described in greater detail below. 
     Moreover, in the depicted embodiment, a combustor dome  118  extends radially across a forward end  88  of the combustor  80 . The combustor dome  118  may be a part of outer wall  102 , may be a part of inner wall  120 , may be a part of both outer wall  102  and inner wall  120  (e.g., a portion of the combustor dome  118  may be defined by the outer wall  102  and the remainder may be defined by the inner wall  120 ), or may be a separate component from outer wall  102  and inner wall  120 . Additionally, a plurality of nozzle airfoils is positioned in each of the nozzle portions  82 N,  84 N. Each nozzle airfoil  126  within the first turbine stage nozzle portion  82 N extends radially from the outer band  110  to the inner band  124 , and the nozzle airfoils  126  are spaced circumferentially about the longitudinal centerline  12 . Each nozzle airfoil  128  within the second turbine stage nozzle portion  84 N extends radially from the outer band  114  to the inner band  136 , and the nozzle airfoils  128  are spaced circumferentially about the longitudinal centerline  12 . Further, a plurality of blade airfoils  130  are positioned in each of the blade portions  82 B,  84 B. Each blade airfoil  130  within the first turbine stage blade portion  82 B is attached to 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 spaced circumferentially about the longitudinal centerline  12 . Similarly, each blade airfoil  130  within the second turbine stage blade portion  84 B 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 spaced circumferentially about the longitudinal centerline  12 . Each blade airfoils  130  extends radially outward toward the outer wall  102 , i.e., the outer boundary of the flow path  100 , and a clearance gap is defined between a tip  140  of each blade airfoil  130  and the outer wall  102  such that each turbine rotor  134 ,  138  is free to rotate within its respective turbine stage. Although not depicted, each turbine rotor  134 ,  138  of the HP turbine  28  is connected to the HP shaft  34  ( FIG. 1 ). In such manner, rotor blade airfoils  130  may extract kinetic energy from the flow of combustion gases through the flow path  100  defined by the HP turbine  28  as rotational energy applied to the HP shaft  34 . 
     Accordingly, 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 and outer boundaries define the flow path for the combustion gases  66  through the combustion section  26  and HP turbine  28 . Portions of the outer boundary of the flow path assembly  101  may be integrated or unified into a single piece outer wall  102  that defines the radially outer boundary of the gas flow path  100 . For instance, the outer wall  102  may include a combustor portion  104  extending through a combustion section, such as combustion section  26 , and a turbine portion  106  extending through at least a first turbine stage of a turbine section, such as first turbine stage  82  of HP turbine  28 . The combustor portion  104  and 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 . 
     In the exemplary embodiment depicted in  FIG. 3A , the outer wall  102  includes a combustor portion  104  extending through the combustion section  26  and a turbine portion  106  extending 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 one turbine stage as just described) or through more stages (e.g., through one or more stages of the LP turbine  30  positioned downstream of 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, which is referred to herein as unitary outer wall  102 . 
     The term “unitary” as used herein denotes that the associated component, such as the outer wall  102 , is made as a single piece during manufacturing, i.e., the final unitary component is a single piece. Thus, a unitary component has a construction in which the integrated portions are inseparable and is different from a component comprising a plurality of 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 may be re-separated). The final unitary component may comprise a substantially continuous piece of material, or in other embodiments, may comprise a plurality of portions that are permanently bonded to one another. In any event, the various portions forming a unitary component are integrated with one another such that the unitary component is a single piece with inseparable portions. 
     As shown in  FIG. 3A , the combustor portion  104  of the unitary structure forming outer wall  102  includes the outer liner  108  of the combustor  80 . The turbine portion  106  includes the outer band  110  of the first turbine stage nozzle portion  82 N, the shroud  112  of the first turbine stage blade portion  82 B, the outer band  114  of the second turbine stage nozzle portion  84 N, and the shroud  116  of the second turbine stage blade portion  84 B. As stated, these outer boundary components are integrated into a single piece to form the unitary structure that is outer wall  102 . Thus, in the exemplary embodiment of  FIG. 2 , outer liner  108 , outer band  110 , shroud  112 , outer band  114 , and shroud  116  are integrally formed, i.e., constructed as a single unit or piece to form the integrated or unitary outer wall  102 . 
     In some embodiments, other portions of the flow path assembly  101  may be integrated into the unitary structure of outer wall  102 , and in still other embodiments, at least a portion of the outer boundary and the inner boundary are made as a single, unitary component such that the flow path assembly  101  may be referred to as an integrated flow path assembly. For example, referring to  FIG. 3B , the combustor portion  104  of unitary outer wall  102  also may include the combustor dome  118  that extends across the forward end  88  of combustor  80 . As such, in the exemplary embodiment of  FIG. 3B , the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , and combustor dome  118  are constructed as a single unit or piece to form the integrated or unitary outer wall  102 . That is, the outer liner  108 , outer bands  110 ,  114 , shrouds  112 ,  116 , and combustor dome  118  are integrally formed such that the outer liner  108 , outer bands  110 ,  114 , shrouds  112 ,  116 , and combustor dome  118  are a single unitary structure. 
     As another example, referring to  FIG. 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 integrated flow path assembly  101 . In the exemplary embodiment of  FIG. 3C , the combustor portion  104  further comprises the inner liner  122 , such that the inner liner  122  is integrated with the unitary structure of the outer wall  102  shown in  FIG. 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. In the exemplary embodiment of  FIG. 3D , the turbine portion  106  further includes the inner band  124  of the first turbine stage nozzle portion  82 N, such that the inner band  124  is integrated with the unitary structure of the flow path assembly  101  shown in  FIG. 3C . Accordingly, the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , combustor dome  118 , inner liner  122 , and inner band  124  are integrally formed such that the 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. In the exemplary embodiment of  FIG. 3E , the turbine portion  106  further includes the plurality of nozzle airfoils  126 , such that each nozzle airfoil  126  of the plurality of nozzle airfoils  126  of the first turbine stage nozzle portion  82 N is integrated with the unitary structure of the flow path assembly  101  shown in  FIG. 3D . Therefore, the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , combustor dome  118 , inner liner  122 , inner band  124 , and nozzle airfoils  126  are integrally formed such that the outer liner  108 , outer bands  110 ,  114 , shrouds  112 ,  116 , combustor dome  118 , inner liner  122 , inner band  124 , and nozzle airfoils  126  are a single unitary structure. 
     Of course, the nozzle airfoils  126  of the first turbine stage nozzle portion  82 N may be integrated with the outer wall  102  without being integrated with 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 , shrouds  112 ,  116 , and nozzle airfoils  126  are a single unitary structure, i.e., a unitary outer wall  102 . In other embodiments, the unitary outer wall  102  also may 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 nozzle airfoils  126  are integrally formed or constructed as a single unit or piece. In still other embodiments, the inner liner  122  also may be included, such that the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , combustor dome  118 , inner liner  122 , and nozzle airfoils  126  are integrally formed as a single unitary structure, i.e., a unitary outer wall  102 . 
     In yet other embodiments, the combustor dome  118  may not be integrated with either the outer wall  102  or the inner wall  120  in whole or in part. That is, the combustor dome  118  is a separate component from both the outer wall  102  and the inner wall  120 . As such, the flow path  100  may be discontinuous between the combustor dome  118  and outer wall  102 , as well as between combustor dome  118  and inner wall  120 . Further, in such embodiments, the combustor dome  118  is configured to move axially with respect to the inner wall  120  and the outer wall  102  but may be attached to, and accordingly supported by, one or more fuel nozzle assemblies  90 . More particularly, an axial slip joint may be formed between the combustor dome  118  and each of the outer wall  102  and the inner wall  120  such that the combustor dome  118  may move or float axially with respect to the inner wall  120  and outer wall  102 . Allowing the combustor dome  118  to float relative to the outer wall  102  and inner wall  120  can help control the position of the fuel nozzle assembly  90  with respect to the combustor dome  118  and combustor  80 . For example, the combustor dome  118 , outer wall  102 , and inner wall  120  may be made of a different material or materials than the fuel nozzle assembly  90 . As described in greater detail below, in an exemplary embodiment, the combustor dome  118 , outer wall  102 , and inner wall  120  are made from a ceramic matrix composite (CMC) material, and the fuel nozzle assembly  90  may be made from a metallic material, e.g., a metal alloy or the like. In such embodiment, the CMC material thermally grows or expands at a different rate than the metallic material. Thus, allowing the combustor dome  118  to move axially with respect to outer and inner walls  102 ,  120  may allow for tighter control of the immersion of swirler  92  of fuel nozzle assembly  90  within combustor dome  118 , as well as combustor  80 , than if the combustor dome  118  was attached to the outer and inner walls  102 ,  120 . Tighter control of the position of fuel nozzle assembly  90  and its components with respect to combustor  80  can reduce variation in operability and performance of engine  10 . 
     Additionally, in embodiments in which the combustor dome  118  is separate from the outer and inner walls  102 ,  120 , the outer wall  102  and inner wall  120  also may move axially and radially with respect to the combustor dome  118 . By decoupling the combustor dome  118  from the walls  102 ,  120  and allowing relative movement between the walls  102 ,  120  and the combustor dome  118 , stress coupling may be alleviated between the outer and inner walls  102 ,  120  and the combustor dome  118 . Moreover, any leakage between the uncoupled combustor dome  118  and outer and inner walls  102 ,  120  may be utilized as purge and/or film starter flow. 
       FIG. 3F  provides a partial perspective view of a portion of an integral flow path assembly  101 , having an outer wall  102  and inner wall  120  formed as a single piece component. As described with respect to  FIG. 3D  and shown in  FIG. 3F , in some embodiments of the combustion gas flow path assembly  101 , the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , combustor dome  118 , inner liner  122 , and inner band  124  are integrally formed such that the 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.  FIG. 3F  further illustrates that a plurality of openings  142  for receipt of fuel nozzle assemblies  90  and/or swirlers  92  may be defined in the forward end  88  of combustor  80  of the unitary flow path assembly  101 . Further, it will be appreciated that  FIG. 3F  illustrates only a portion of the integral flow path assembly  101  and that, although its entire circumference is not illustrated in  FIG. 3F , the flow path assembly  101  is a single, unitary piece circumferentially as well as axially. As such, the integral flow path assembly  101  defines a generally annular, i.e., generally ring-shaped, flow path between the outer wall  102  and inner wall  120 . 
     Integrating various components of the outer and inner boundaries of the flow path assembly  101  as described above can reduce the number of separate pieces or components within engine  10 , as well as reduce the weight, leakage, and complexity of the engine  10 , compared to known gas turbine engines. For instance, known gas turbine engines employ seals or sealing mechanisms at the interfaces between separate pieces of the flow path assembly to attempt to minimize leakage of combustion gases from the flow path. By integrating the outer boundary, for example, as described with respect to unitary outer wall  102 , split points or interfaces between the outer combustor liner and 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 as well as seals or sealing mechanisms required to prevent leakage. Similarly, by integrating components of the inner boundary, split points or interfaces between the integrated inner boundary components are eliminated, thereby eliminating leakage points and seals or sealing mechanisms required at the inner boundary. Accordingly, undesired leakage, as well as unnecessary weight and complexity, can be avoided by utilizing unitary components in the flow path assembly. Other advantages of unitary outer wall  102 , unitary inner wall  120 , and/or a unitary flow path assembly  101  will be appreciated by those of ordinary skill in the art. 
     As illustrated in  FIGS. 3A through 3F , the outer wall  102  and the inner wall  120  define a generally annular flow path therebetween. That is, the unitary outer wall  102  circumferentially surrounds the inner wall  120 ; stated differently, the unitary outer wall  102  is a single piece extending 360° degrees about the inner wall  120 , thereby defining a generally annular or ring-shaped flow path therebetween. As such, the combustor dome  118 , which extends across the forward end  88  of the combustor  80 , is a generally annular combustor dome  118 . Further, the combustor dome  118  defines an opening  142  for receipt of a fuel nozzle assembly  90  positioned at forward end  88 . The fuel nozzle assembly  90 , e.g., provides combustion chamber  86  with a mixture of fuel and compressed air from the compressor section, which is combusted within the combustion chamber  86  to generate a flow of combustion gases through the flow path  100 . The fuel nozzle assembly  90  may attach 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 embodiments, the fuel nozzle assembly  90  includes a swirler  92 , and in some embodiments, the swirler  92  may attach to the combustor dome  118 , but alternatively, the swirler  92  may float relative to the combustor dome  118  and flow path  100 . It will be appreciated that the fuel nozzle assembly  90  or swirler  92  may float relative to the combustor dome  118  and flow path  100  along both a radial direction R and an axial direction A or only along one or the other of the radial and axial directions R, A. Further, it will be understood that the combustor dome  118  may define a plurality of openings  142 , each opening receiving a swirler  92  or other portion of fuel nozzle assembly  90 . 
     As further illustrated in  FIGS. 3A through 3F , the flow path assembly  101  generally defines a converging-diverging flow path  100 . More particularly, the outer wall  102  and the inner wall  120  define a generally annular combustion chamber  86 , which forms a forward portion of the flow path  100 . Moving aft or downstream of combustion chamber  86 , the outer wall  102  and inner wall  120  converge toward one another, generally in the region of first turbine stage  82 . Continuing downstream of the first turbine stage  82 , the outer wall  102  and inner wall  120  then diverge, generally in the region of second turbine stage  84 . The outer wall  102  and inner wall  120  may continue to diverge downstream of the second turbine stage  84 . In exemplary embodiments, e.g., as shown in  FIG. 3A  and referring only to the unitary outer wall  102 , the first turbine stage nozzle outer band portion  110  and blade shroud portion  112  of the outer wall  102  converge toward the axial centerline  12 . The second turbine stage nozzle outer band portion  114  and blade shroud portion  116  of the outer wall  102  diverge away from the axial centerline  12 . As such, the outer boundary of flow path  100  formed by the unitary outer wall  102  defines a converging-diverging flow path  100 . 
     Turning to  FIG. 4 , a cross-sectional view is provided of a portion of the flow path assembly  101  according to an exemplary embodiment of the present subject matter. As shown in the depicted embodiment, the flow path assembly  101  includes an inner wall  120  and a unitary outer wall  102 . As described above, the unitary 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 a first turbine stage  82  of the turbine section  28 ; in the embodiment of  FIG. 4 , the turbine portion extends through the second turbine stage  84 . Further, the combustor portion  104  and the turbine portion  106  of the outer wall  102  are integrally formed as a single unitary structure and, thus, may be referred to as unitary outer wall  102 , and the inner wall  120  and the unitary outer wall  102  define the combustor  80 . 
     More particularly, in the illustrated embodiment, the combustor portion  104  of the unitary outer wall  102  comprises the outer liner  108  of the combustor  80 , and the turbine portion  106  comprises the outer band  110  of the first turbine stage nozzle portion  82 N, the shroud  112  of the first turbine stage blade portion  82 B, the outer band  114  of the second turbine stage nozzle portion  84 N, and the shroud  116  of the second turbine stage blade portion  84 B. The inner wall  120  also may include a unitary structure that may be referred to as unitary inner wall portion. For example, in  FIG. 4 , the inner wall  120  comprises a forward segment  120   a  and an aft segment  120   b , and the forward segment  120   a  is a unitary structure that may be referred to as unitary inner wall portion  120   a . The unitary inner wall portion  120   a  includes the inner liner  122  and a forward portion of the first turbine stage inner band  124 , which are integrally formed as unitary inner wall portion  120   a . The aft segment  120   b  comprises the remaining portion of the inner wall  120 , i.e., the remainder of the first turbine stage inner band  124 . In some embodiments, as previously described, the unitary outer wall  102  or unitary inner wall portion  120   a  also may include the combustor dome  118 , or the unitary outer wall  102  and the unitary inner wall portion  120   a  each may include a portion of the combustor dome  118 . In still other embodiments, the outer wall  102 , combustor dome  118 , and the forward segment  120   a  of inner wall  120  may be integrally formed as a single piece, unitary structure. 
     Referring still to the embodiment illustrated in  FIG. 4 , the flow path assembly  101  further includes a plurality of first turbine stage nozzle airfoils  126 . Each of the plurality of nozzle airfoils  126  extends between pockets defined in the inner boundary structure and the outer boundary structure of flow path assembly  101 . More particularly, each nozzle airfoil  126  has an inner end  126   a  radially opposite an outer end  126   b . The unitary outer wall  102  defines a plurality of outer pockets  170 , and each outer pocket  170  is configured for receipt of the outer end  126   b  of one of the plurality of first turbine stage nozzle airfoils  126 . Similarly, the inner wall  120  defines a plurality of inner pockets  174 . More specifically, as shown in  FIG. 4 , a portion of each inner pocket  174  is defined by the forward segment  120   a  of inner wall  120  and a remaining portion of each inner pocket  174  is defined by the aft segment  120   b . Further, each inner pocket  174  is configured for receipt of the inner end  126   a  of one of the plurality of first turbine stage nozzle airfoils  126 . Accordingly, a nozzle airfoil  126  extends from each inner pocket  174  to a respective outer pocket  170 . 
     Keeping with  FIG. 4 , the plurality of outer pockets  170  are defined along an area  178  of an inner surface  103  of the unitary outer wall  102 , and the unitary outer wall  102  is built up at the area  178  where the plurality of outer pockets  170  are defined. The built up area  178  may provide extra thickness in outer wall  102  such that the outer pockets  170  may be defined therein, without requiring the extra thickness throughout the entire unitary outer wall  102 . 
     Moreover, the forward segment  120   a  of the inner wall  120  comprises a forward flange  182 , and the aft segment  120   b  of the inner wall  120  comprises an aft flange  184 . Each flange  182 ,  184  extends radially inward from the inner wall  120 , i.e., toward the axial centerline  12 . The forward flange  182  is positioned adjacent the aft flange  184  when the forward segment  120   a  and the aft segment  120   b  are assembled in the flow path assembly  101 . Each flange  182 ,  184  defines at least one pin aperture  186  for receipt of a pin  188 . As such, at least one pin  188  secures the forward flange  182  to the aft flange  184 , e.g., to hold the forward and aft segments  120   a ,  120   b  of inner wall  120  in position with respect to one another. 
     It will be appreciated that the configuration illustrated in  FIG. 4  is by way of example only. For instance, although  FIG. 4  depicts only the first turbine stage nozzle portion  82 N of the flow path assembly  101 , the subject matter described with respect to the first turbine stage nozzle portion  82 N applies equally to the second turbine stage nozzle portion  84 N, and also may apply to any other nozzle portions of the turbine sections of the engine  10 . That is, while only the first turbine stage nozzle airfoils  126  and the adjacent portions of outer wall  102  and inner wall  120  are shown and described with respect to  FIG. 4 , the second turbine stage nozzle airfoils  128  and the adjacent portions of outer wall  102  and inner wall  120  may be similarly configured. More particularly, the outer wall  102  may define a plurality of second outer pockets for the receipt of the outer ends of nozzle airfoils  128 , and the second stage inner band  136  may comprise forward and aft segments that together define a plurality of second inner pockets for receipt of the inner ends of the nozzle airfoils  128 . The outer wall  102  may be built up in the area in which the second outer pockets are defined, e.g., to provide sufficient thickness for the second outer pockets to be defined in the outer wall  102 . Further, the forward and aft segments of the second stage inner band  136  may each include a flange, and the flanges of the forward and aft inner band segments may be bolted together, e.g., to attach the segments to one another. However, in other embodiments, the first turbine stage nozzle portion  82 N may be configured as shown in  FIG. 4  while the second turbine stage nozzle portion  84 N, as well as nozzle portions of other turbine stages, may be configured differently from the depicted embodiment. 
     In other exemplary embodiments, the outer wall  102  may comprise forward and aft segments and the inner wall  120  may be a unitary inner wall  120 . That is, some embodiments may utilize essentially an opposite configuration from the configuration illustrated in  FIG. 4 . More particularly, the unitary inner wall  120  may comprise the inner liner  122  of the combustor  80  and the inner band  124  of the first turbine stage nozzle portion  82 N. Further, the outer wall  102  may comprise a forward segment and an aft segment. The forward segment may be a unitary structure that may be referred to as unitary outer wall portion. The unitary outer wall portion includes the outer liner  108  and a forward portion of the first turbine stage outer band  110 , which are integrally formed as unitary outer wall portion. The aft segment comprises the remaining portion of the outer band  110  and may further include the first turbine stage shroud  112  and a forward portion of the second turbine stage outer band  114 . A second aft segment comprises at least the remainder of the second turbine stage outer band  114  and also may comprise the second turbine stage shroud  116 . In some embodiments, the forward segment of outer wall  102  or unitary inner wall  120  also may include the combustor dome  118 , or the outer wall forward segment and unitary inner wall  120  each may include a portion of the combustor dome  118 . In still other embodiments, the outer wall forward segment, combustor dome  118 , and inner wall  120  may be integrally formed as a single piece, unitary structure. 
     Moreover, similar to the embodiment shown in  FIG. 4 , each of the plurality of nozzle airfoils  126 ,  128  may extend between pockets defined in the inner boundary structure and the outer boundary structure of flow path assembly  101 . More specifically, the unitary inner wall  120  defines a plurality of inner pockets  174  and a plurality of second inner pockets. Each inner pocket  174  is configured for receipt of the inner end  126   a  of one of the plurality of first turbine stage nozzle airfoils  126 , and each second inner pocket is configured for receipt of the inner end of one of the plurality of second turbine stage nozzle airfoils  128 . 
     Similarly, the outer wall  102  defines a plurality of outer pockets  170 . A portion of each outer pocket  170  is defined by the forward segment of outer wall  102  and a remaining portion of each outer pocket  170  is defined by the aft segment. A plurality of second outer pockets also may be defined in the outer wall  102 , e.g., a portion of each second outer pocket may be defined by the forward portion of the second turbine stage outer band portion of the outer wall aft segment, and a remaining portion of each second outer pocket may be defined by the second aft segment, which includes the remainder of the second turbine stage outer band. Further, each outer pocket  170  is configured for receipt of the outer end  126   b  of one of the plurality of first turbine stage nozzle airfoils  126 , and each second outer pocket is configured for receipt of the outer end of one of the plurality of second turbine stage nozzle airfoils  128 . Accordingly, a first turbine stage nozzle airfoil  126  extends from each inner pocket  174  to a respective outer pocket  170 , and a second turbine stage nozzle airfoil  128  extends from each second inner pocket to a respective second outer pocket. 
     Further, similar to the embodiment shown in  FIG. 4 , the plurality of inner pockets  174  may be defined along an area of an inner surface  121  of the unitary inner wall  120 , and the unitary inner wall  120  may be built up at the area where the plurality of inner pockets  174  are defined. Likewise, the plurality of second inner pockets may be defined along an area of inner surface  121 , and the unitary inner wall  120  may be built up at the area where the plurality of second inner pockets are defined. The built up areas near the inner pockets may provide extra thickness in inner wall  120  such that the inner pockets may be defined therein, without requiring the extra thickness throughout the entirety of inner wall  120 . 
     Moreover, the forward and aft segments of the outer wall  102  may comprise flanges for securing the outer wall segments to one another. For example, the forward segment of the outer wall  102  may comprise a forward flange, the aft segment of the outer wall  102  may comprise a first aft flange and a second aft flange, and the second aft segment may comprise a third aft flange. Each flange may extend radially outward from the outer wall  102 , i.e., away from the axial centerline  12 . The forward flange may be positioned adjacent the first aft flange when the forward and aft segments are assembled in the flow path assembly  101 , and the second aft flange may be positioned adjacent the third aft flange when the aft segment and the second aft segment are assembled in the flow path assembly  101 . Each flange defines at least one pin aperture  186  for receipt of a pin  188 . As such, at least one pin  188  secures the forward flange to the first aft flange, e.g., to hold the forward and aft segments of outer wall  102  in position with respect to one another. Similarly, at least one pin  188  secures the second aft flange to the third aft flange, e.g., to hold the aft segment and the second aft segment of outer wall  102  in position with respect to one another. 
     It will be understood that each pocket  170 ,  172 ,  174 ,  176  may have a shape substantially similar to an axial cross-sectional shape of the respective airfoil  126 ,  128  received within the pocket. That is, each outer pocket  170 ,  172  may generally be described as an airfoil-shaped depression in the outer wall  102 , and each inner pocket  174 ,  176  may generally be described as an airfoil-shaped depression in the inner wall  120  and inner band  136 . More specifically, the first outer pockets  170  and first inner pockets  174  may have a shape corresponding to the first turbine stage nozzle airfoils  126 , the second outer pockets  172  and second inner pockets  176  may have a shape corresponding to the second turbine stage nozzle airfoils  128 . However, the pockets  170 ,  172 ,  174 ,  176  may have any appropriate shape for receiving the nozzle airfoils  126 ,  128 . 
     Utilizing an outer wall  102  or an inner wall  120  that comprises forward and aft segments may help simplify the assembly of the flow path assembly  101 . For example, referring to the exemplary embodiment of  FIG. 4 , the outer wall  102  and the forward segment  120   a  of the inner wall  120  may be positioned in the assembly. Next, the plurality of first turbine stage nozzle airfoils  126  may be installed, e.g., by inserting the outer ends  126   b  of the airfoils  126  into the first outer pockets  170  and positioning the inner ends  126   a  within the portion of the first inner pockets  174  defined by the forward segment  120   a  of inner wall  120 . Then, the aft segment  120   b  of the inner wall  120  may be slid into position, with the remainder of the first inner pockets  174  defined in the aft segment  120   b  enclosing the inner ends  126   a  of the first turbine stage nozzle airfoils  126  in the first inner pockets  174 . The first stage rotor  134  may then be installed, followed by the forward segment  136   a  of the second turbine stage inner band  136 . Next, the second turbine stage nozzle airfoils  128  may be installed, e.g., by inserting the outer ends  128   b  of the airfoils  128  into the second outer pockets  172  and positioning the inner ends  128   a  within the portion of the second inner pockets  176  defined by the forward segment  136   a  of inner band  136 . Then, the aft segment  136   b  of the inner band  136  may be slid into position, with the remainder of the second inner pockets  176  defined in the aft segment  136   b  enclosing the inner ends  128   a  of the second turbine stage nozzle airfoils  128  in the second inner pockets  176 . The second stage rotor  138  may be installed to complete the assembly of the flow path assembly  101 , as well as the combustion section  26  and HP turbine section  28  of the engine  10 . 
     Turning now to  FIGS. 5 and 6 , another exemplary embodiment of the flow path assembly  101  is provided. The depicted embodiment utilizes a bayonet type joint to secure the nozzle airfoils  126 ,  128 , which are integral with the outer wall  102 , to the inner wall  120  and inner band  136 . That is, referring particularly to  FIG. 5 , the first turbine stage nozzle airfoils  126  are integrated with the outer wall  102  to form a unitary structure with the outer boundary of the flow path assembly  101 . Thus, the outer wall  102  depicted in  FIG. 5  is similar to the embodiment of  FIG. 3E , although the airfoils  126  of  FIG. 5  are not integrated with the inner wall  120 , and the inner wall  120  and combustor dome  118  may, but need not, be integral with the unitary outer wall  102 . 
     Accordingly, as shown in  FIG. 5 , the inner end  126   a  of the nozzle airfoils  126  is positioned against the inner surface  121  of the inner wall  120 , and each inner end  126   a  includes a bayonet that is received in a bayonet slot defined in the inner wall  120  as described in greater detail below. Further, the inner wall  120  includes a built up portion or area  119  immediately upstream or forward of the nozzle airfoils  126 , such that the built up area  119  helps shield the interface between the nozzle airfoils  126  and the inner wall  120  from the hot combustion gases. It also will be appreciated that the built up area  119  of inner wall  120  may function similar to fillets that may be defined on nozzle airfoils  126  in other embodiments of the airfoils. The built up area  119  may be an area of built up CMC material that forms the inner wall  120 . 
     Keeping with  FIG. 5 , the inner end  126   a  of each nozzle airfoil  126  includes a projection or bayonet  190 , which is received in a bayonet slot  192  defined in the inner wall  120 . More particularly, referring to  FIG. 6 , the inner wall  120  defines a plurality of bayonet slots  192  such that each bayonet  190  is received in a bayonet slot  192  of the inner wall  120 . Each slot  192  generally has an L-shape, i.e., each bayonet slot  192  has a first leg  192   a  and a second leg  192   b  that extends generally perpendicular to the first leg  192   a . The first and second legs  192   a ,  192   b  are joined together to form a continuous pathway, i.e., bayonet slot  192 . Thus, when a bayonet  190  of a nozzle airfoil  126  is received within a bayonet slot  192 , the inner wall  120  may be slid such that the bayonet  190  moves along the first leg  192   a  and then the inner wall  120  may be rotated such that the bayonet  190  then moves along the second leg  192   b  until the bayonet  190  is seated within an end  192   c  of the bayonet slot  192 . Thus, the nozzle airfoils  126  are secured in the inner wall  120  via bayonet joints, which prevent axial movement of the nozzle airfoil bayonets  190  within the inner wall bayonet slots  192  to secure the airfoils  126  in the inner wall  120 . 
     Further, the bayonet slots  192  are defined such that the slots  192  are oriented to take any tangential and axial loads from the nozzle airfoils  126 . That is, the bayonet slots  192  are oriented such that a tangential and/or axial load on a nozzle airfoil  126  loads the airfoil into the inner wall  120  surrounding the bayonet  190  rather than causing the bayonet  190  to move along or within the slot  192 . Therefore, the bayonet slots  192  are defined in a manner that prevents the nozzle airfoils  126  from being dislodged from or moving from the slot ends  192   c  when the airfoils are subject to aerodynamic loads. 
     As also illustrated in  FIGS. 5 and 6 , a first support member  194  and a second support member  196  are positioned radially inward of the inner wall  120 , i.e., closer to the axial centerline  12  of engine  10 , to support the inner wall. The first support member  194  defines a groove  198  for receipt of a projection  200  of the inner wall  120 . The inner wall projection  200  provides an area for bayonet slots  192  to be defined in the inner wall  120  for the receipt of airfoil bayonets  190 . The first support member groove  198  fits against the projection  200  to support the inner wall  120  at the projection  200 . The first support member  194  further includes a housing  202  at a forward side  204  of the first support member. The housing  202  is configured for receipt of a seal  206 , which preferably is generally ring shaped to extend circumferentially against the inner wall  120 . For example, the seal  206  may be a piston ring seal or other suitable seal. The seal  206  helps prevent air leakage at the forward or upstream end of the inner wall  120 . 
     The second support member  196  is positioned axially aft of the first support member  194 . The second support member  196  includes a plurality of tabs  208 . As shown in  FIGS. 5 and 6 , the inner wall  120  defines a plurality of recesses  210  along an aft surface  123  of the inner wall. Each tab  208  of the second support member  196  is received in a recess  210  of the inner wall  120  to position or fit the second support member  196  against the aft surface  123  of inner wall  120  such that the second support member  196  covers the plurality of bayonet slots  192  defined in the inner wall  120 . As such, the second support member  196  helps prevent fluid leakage from or into the bayonet slots  192 . 
     As further depicted in  FIG. 5 , the first support member defines a flange  212  and the second support member defines a flange  214 . Each flange  212 ,  214  defines an aperture  216 , and a bolt  218  is received through the aperture  216  of the first support member  194  and the aperture  216  of the second support member  196  to secure the first support member to the second support member. Other ways of securing the first and second support members  194 ,  196  to one another also may be used, and in suitable embodiments, the first and second support members  194 ,  196  may be one integral, single piece component or may be divided into more than two components. 
     It will be appreciated that, although  FIGS. 5 and 6  depict only the first turbine stage nozzle portion  82 N of the flow path assembly  101 , the subject matter described with respect to the first turbine stage nozzle portion  82 N applies equally to the second turbine stage nozzle portion  84 N, and also may apply to any other nozzle portions of the turbine sections of the engine  10 . That is, while only the first turbine stage nozzle airfoils  126  are shown and described with respect to  FIGS. 5 and 6 , the second turbine stage nozzle airfoils  128  may be similarly configured, and the second stage inner band  136  may be configured similarly to the first stage inner band portion of the inner wall  120 . Further, first and second support members similar to those depicted in  FIGS. 5 and 6  also may be provided in the second turbine stage. However, in other embodiments, the first turbine stage nozzle portion  82 N may be configured as shown in  FIGS. 5 and 6  while the second turbine stage nozzle portion  84 N, as well as nozzle portions of other turbine stages, may be configured differently from the depicted embodiment. 
     As previously stated, the outer wall  102 , inner wall  120 , and combustor dome  118 , and in some embodiments, first and second turbine stage nozzle airfoils  126 ,  128 , may comprise a CMC material. More particularly, in exemplary embodiments, the combustor portion  104  and the turbine portion  106  of flow path assembly  101  are integrally formed from a CMC material such that the resulting unitary structure is a CMC component. For example, where the combustor portion  104  includes the outer liner  108  of the combustor  80  and the turbine portion  106  includes the outer band  110  of the first turbine stage nozzle portion  82 N, the shroud  112  of the first turbine stage blade portion  82 B, the outer band  114  of the second turbine stage nozzle portion  84 N, and the shroud  116  of the second turbine stage blade portion  84 B, the outer liner  108 , outer bands  110 ,  114 , and shrouds  114 ,  116  may be integrally formed from a CMC material to produce a unitary CMC outer wall  102 . As described above, in other embodiments, additional CMC components, such as the nozzle airfoils  126 ,  128 , 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 instance, where the inner wall  120  comprises separate components, e.g., inner liner  122 , inner bands  124 ,  136 , and blade platforms  132 , each component of the inner wall  120  may be formed from a CMC material. In embodiments in which two or more components are integrated to form a unitary inner wall  120 , the components may be integrally formed from a CMC material to construct a unitary CMC inner wall  120 . 
     Examples of CMC materials, and particularly SiC/Si—SiC (fiber/matrix) continuous fiber-reinforced ceramic composite (CFCC) materials and processes, are described in U.S. Pat. Nos. 5,015,540; 5,330,854; 5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. Such processes generally entail the fabrication of CMCs using multiple pre-impregnated (prepreg) layers, e.g., the ply material may include prepreg material consisting of ceramic fibers, woven or braided ceramic fiber cloth, or stacked ceramic fiber tows that has been impregnated with matrix material. In some embodiments, each prepreg layer is in the form of a “tape” comprising the desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and organic resin binders. Prepreg tapes can be formed by impregnating the reinforcement material with a slurry that contains the ceramic precursor(s) and binders. Preferred materials for the precursor will depend on the particular composition desired for the ceramic matrix of the CMC component, for example, SiC powder and/or one or more carbon-containing materials if the desired matrix material is SiC. Notable carbon-containing materials include carbon black, phenolic resins, and furanic resins, including furfuryl alcohol (C 4 H 3 OCH 2 OH). Other typical slurry ingredients include organic binders (for example, polyvinyl butyral (PVB)) that promote the flexibility of prepreg tapes, and solvents for the binders (for example, toluene and/or methyl isobutyl ketone (MIBK)) that promote the fluidity 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 component, for example, silicon and/or SiC powders in the case of a Si—SiC matrix. Chopped fibers or whiskers or other materials also may be embedded within the matrix as previously described. Other compositions and processes for producing composite articles, and more specifically, other slurry and prepreg tape compositions, may be used as well, such as, e.g., the processes and compositions described in U.S. Patent Application Publication No. 2013/0157037. 
     The resulting prepreg tape may be laid-up with other tapes, such that a CMC component formed from the tape comprises multiple laminae, each lamina derived from an individual prepreg tape. Each lamina contains a ceramic fiber reinforcement material encased in a ceramic matrix formed, wholly or in part, by conversion of a ceramic matrix precursor, e.g., during firing and densification cycles as described more fully below. In some embodiments, the reinforcement material is in the form of unidirectional arrays of tows, each tow containing continuous fibers or filaments. Alternatives to unidirectional arrays of tows may be used as well. Further, suitable fiber diameters, tow diameters, and center-to-center tow spacing will depend on the particular application, the thicknesses of the particular lamina and the tape from which it was formed, and other factors. As described above, other prepreg materials or non-prepreg materials may be used as well. 
     After laying up the tapes or plies to form a layup, the layup is debulked and, if appropriate, cured while subjected to elevated pressures and temperatures to produce a preform. The preform is then heated (fired) in a vacuum or inert atmosphere to decompose the binders, remove the solvents, and convert the precursor to the desired ceramic matrix material. Due to decomposition of the binders, the result is a porous CMC body that may undergo densification, e.g., melt infiltration (MI), to fill the porosity and yield the CMC component. Specific processing techniques and parameters for the above process will depend on the particular composition of the materials. For example, silicon CMC components may be formed from fibrous material that is infiltrated with molten silicon, e.g., through a process typically referred to as the Silcomp process. Another technique of manufacturing CMC components is the method known as the slurry cast melt infiltration (MI) process. In one method of manufacturing using the slurry cast MI method, CMCs are produced by initially providing plies of balanced two-dimensional (2D) woven cloth comprising silicon carbide (SiC)-containing fibers, having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers running in both directions of the weave. The term “silicon carbide-containing fiber” refers to a fiber having a composition that includes silicon carbide, and preferably is substantially silicon carbide. For instance, the fiber may have a silicon carbide core surrounded with carbon, or in the reverse, the fiber may have a carbon core surrounded by or encapsulated with silicon carbide. 
     Other techniques for forming CMC components include polymer infiltration and pyrolysis (PIP) and oxide/oxide processes. In PIP processes, silicon carbide fiber preforms are infiltrated with a preceramic polymer, such as polysilazane and then heat treated to form a SiC matrix. In oxide/oxide processing, aluminum or alumino-silicate fibers may be pre-impregnated and then laminated into a preselected geometry. Components may also be fabricated from a carbon fiber reinforced silicon carbide matrix (C/SiC) CMC. The C/SiC processing includes a carbon fibrous preform laid up on a tool in the preselected geometry. As utilized in the slurry cast method for SiC/SiC, the tool is made up of graphite material. The fibrous preform is supported by the tooling during a chemical vapor infiltration process at about 1200° C., whereby the C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D, and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes. For example, cut layers of 2D woven fabrics may be stacked in alternating weave directions as described above, or filaments may be wound or braided and combined with 3D weaving, stitching, or needling to form 2.5D or 3D preforms having multiaxial fiber architectures. Other ways of forming 2.5D or 3D preforms, e.g., using other weaving or braiding methods or utilizing 2D fabrics, may be used as well. 
     Thus, a variety of processes may be used to form a unitary structure, such as the outer wall  102  depicted in  FIG. 3A , as a unitary CMC component. More specifically, a plurality of plies of a CMC material may be used to form each unitary structure. The plurality of plies may be interspersed with one another to integrate the various portions forming the unitary structure. As an example, the unitary outer wall  102  of  FIG. 3A  may be made from a plurality of outer liner plies, a plurality of first turbine stage outer band plies, a plurality of first turbine stage shroud plies, a plurality of second turbine stage outer band plies, and a plurality of second turbine stage shroud plies. Where the outer liner plies meet the first turbine stage outer band plies, ends of the outer liner plies may be alternated with ends of the outer band plies to integrate the plies for forming the outer liner portion with the plies for forming the first turbine stage outer band portion of the unitary outer wall  102 . That is, any joints between the plies forming unitary outer wall  102  may be formed by alternating plies on one side of the joint with plies on the other side of the joint. As such, the plies for forming unitary outer wall  102  may be interspersed to integrate the plies and, thereby, each portion of the unitary outer wall  102 . Of course, the CMC plies may be laid up in other ways as well to form the unitary structure. In addition, laying up the plurality of CMC plies may include defining features of the unitary structure or other component (e.g., inner liner  122  when not integrated with inner band  124  to from a unitary inner wall  120  or separate combustor dome  118 ) such as openings  142  in combustor forward end  88 . 
     After the plurality of CMC plies are laid up to define a unitary CMC component preform, the preform is cured to produce a single piece, unitary CMC component, which is then fired and subjected to densification, e.g., silicon melt-infiltration, to form a final unitary CMC structure. Continuing with the above outer wall  102  example, the outer wall preform may be processed in an autoclave to produce a green state unitary outer wall  102 . Then, the green state unitary outer wall  102  may be placed in a furnace to burn out excess binders or the like and then placed in a furnace with a piece or slab of silicon and fired to melt infiltrate the unitary outer wall  102  with at least silicon. More particularly, for unitary outer wall  102  formed from CMC plies of prepreg tapes that are produced as described above, heating (i.e., firing) the green state component in a vacuum or inert atmosphere decomposes the binders, removes the solvents, and converts the precursor to the desired ceramic matrix material. The decomposition of the binders results in a porous CMC body; the body may undergo densification, e.g., melt infiltration (MI), to fill the porosity. In the foregoing example where the green state unitary outer wall  102  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 processes, and with any suitable materials including but not limited to silicon. In one embodiment, densification and firing may be conducted in a vacuum furnace or an inert atmosphere having an established atmosphere at temperatures above 1200° C. to allow silicon or other appropriate material or combination of materials to melt-infiltrate into the component. The densified CMC body hardens to a final unitary CMC outer wall  102 . In some embodiments, the final unitary structure may be finish machined, e.g., to bring the structure within tolerance or to define openings  142  in forward end  88 , and/or an environmental barrier coating (EBC) may be applied to the unitary structure, e.g., to protect the unitary structure from the hot combustion gases  66 . It will be appreciated that other methods or processes of forming CMC components, such as unitary CMC outer wall  102 , unitary CMC inner wall  120 , or the like may be used as well. 
     Additionally or alternatively, other processes for producing unitary components may be used to form unitary outer wall  102  and/or unitary inner wall  120 , and the unitary structure(s) may be formed from other materials. In some embodiments, an additive manufacturing process may be used to form unitary outer wall  102  and/or unitary inner wall  120 . For example, an additive process such as Fused Deposition Modeling (FDM), Selective Laser Sintering (SLS), Stereolithography (SLA), Digital Light Processing (DLP), Direct Metal Laser Sintering (DMLS), Laser Net Shape Manufacturing (LNSM), electron beam sintering or other known process may be used to produce a unitary outer wall  102  and/or a unitary inner wall  120 . Generally, an additive process fabricates components using three-dimensional information, for example, a three-dimensional computer model, of the component. The three-dimensional information is converted into a plurality of slices, each slice defining a cross section of the component for a predetermined height of the slice. The component is then “built-up” slice by slice, or layer by layer, until finished. Superalloy metallic materials or other suitable materials may be used in an additive process to form unitary outer wall  102  and/or a unitary inner wall  120 . In other embodiments, a unitary outer wall  102  and/or unitary inner wall  120  may be formed using a forging or casting process. Other suitable processes or methods may be used as well. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated 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. 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 equivalent structural elements with insubstantial differences from the literal language of the claims.