Patent Publication Number: US-11391171-B2

Title: Methods and features for positioning a flow path assembly within a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of and claims priority to U.S. application Ser. No. 15/440,258, filed Feb. 23, 2017, the contents of which are incorporated herein by reference. 
    
    
     FIELD 
     The present subject matter relates generally to gas turbine engines. More particularly, the present subject matter relates to flow path assemblies of gas turbine engines and features for positioning a flow path assembly within a gas turbine engine. 
     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 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 forming the flow path assembly from a ceramic matrix composite (CMC) material. CMC materials are high temperature materials that are more commonly being used for various components within gas turbine engines. As such, CMC materials have a different rate of thermal expansion than, e.g., metallic materials such as metals or metal alloys. Therefore, where components supporting the CMC flow path assembly are made from one or more non-CMC materials, the CMC flow path assembly and the support components may thermally expand at different rates, which could affect the positioning of the flow path assembly within the gas turbine engine. 
     Accordingly, improved flow path assemblies would be desirable. For example, a flow path assembly utilizing a hub and spoke configuration to position the flow path assembly within a gas turbine engine would be useful. In particular, a flow path assembly utilizing positioning members to position the flow path assembly within a gas turbine engine and maintain the flow path assembly in a proper position while allowing for thermal growth of the flow path assembly and components that support the flow path assembly would be beneficial. 
     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 and a unitary outer wall that 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. The flow path assembly further comprises at least two positioning members for radially centering the flow path assembly within the gas turbine engine. The positioning members extend to the flow path assembly from one or more structures external to the flow path assembly. The positioning members constrain the flow path assembly tangentially and allow radial and axial movement of the flow path assembly. 
     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 and a unitary outer wall that 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. The flow path assembly also comprises a plurality of axial positioning members for positioning the flow path assembly within the gas turbine engine. The plurality of axial positioning members extends through a portion of the outer wall and a portion of the inner wall. The outer wall and the inner wall are configured to move axially along the plurality of positioning members to allow axial movement of the flow path assembly. 
     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 and a unitary outer wall that 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. The flow path assembly further comprises a plurality of radial positioning members for positioning the flow path assembly within the gas turbine engine. A plurality of openings are defined above the first turbine stage of the turbine portion, and each of the plurality of radial positioning members extends into one of the plurality of openings. 
     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. 
         FIGS. 4A, 4B, 4C, 5A, 5B, and 5C  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. 6  provides a cross-sectional view of a portion of a flow path assembly according to an exemplary embodiment of the present subject matter. 
         FIG. 7  provides a cross-sectional view of a flow path assembly according to an exemplary embodiment of the present subject matter. 
         FIG. 8  provides a perspective view of the flow path assembly of  FIG. 7 . 
         FIG. 9  provides a cross-sectional view of a flow path assembly according to an exemplary embodiment of the present subject matter. 
     
    
    
     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 SYLRAIVIIC®), 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 radially outer 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 . 
       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 , as well as  FIGS. 4A through 4C  and  FIGS. 5A and 5B  discussed in greater detail below, 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  FIGS. 4A and 4B , other exemplary embodiments of the present subject matter are illustrated.  FIG. 4A  provides a schematic cross-sectional view of the combustion section  26  and the HP turbine  28  of the turbine section according to one exemplary embodiment.  FIG. 4B  provides a schematic cross-sectional view of the combustion section  26  and the HP turbine  28  of the turbine section according to another exemplary embodiment.  FIG. 4C  provides a schematic cross-sectional view of the combustion section  26  and the HP turbine  28  of the turbine section according to yet another exemplary embodiment. 
     In the embodiments shown in  FIGS. 4A, 4B, and 4C , the outer wall  102  is formed as a single unitary structure and the inner wall  120  is formed as another single unitary structure, and together, the unitary outer wall  102  and the unitary inner wall  120  define the flow path  100 . However, it should be appreciated that the inner wall  120  need not be a single unitary structure. For example, in the embodiments shown in  FIGS. 4A, 4B, and 4C , the inner wall  120  could comprise an inner liner  122  formed separately from inner band  124 . 
     As described with respect to  FIGS. 3A through 3F , the unitary outer wall  102  of  FIGS. 4A, 4B, and 4C  defines an outer boundary and the inner wall  120  defines an inner boundary of the flow path  100 . Together, the unitary outer wall  102  and the inner wall  120  form a flow path assembly  101 . The unitary outer wall  102  extends from the forward end  88  of combustor  80  of the combustion section  26  through at least the first turbine stage  82  of the HP turbine  28 , and in the depicted embodiments, the unitary outer wall  102  extends from forward end  88  to an aft end of the second turbine stage  84  of HP turbine  28 . The inner wall  120  includes at least the inner liner  122 , and in embodiments in which the inner wall  120  is a unitary inner wall, the unitary inner wall  120  extends from the forward end  88  of the combustor  80  through the first turbine stage nozzle portion  82 N. Accordingly, as shown in  FIGS. 4A, 4B, and 4C , the outer wall  102  and inner wall  120  define the combustion chamber  86  of the combustor  80 . 
     Like the embodiments described with respect to  FIGS. 3A through 3F , the unitary outer wall  102  of the embodiments shown in  FIGS. 4A, 4B, and 4C  includes the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , and shroud  116 . Further, in the exemplary embodiment of  FIG. 4A , the unitary outer wall  102  includes the combustor dome  118  defined at the forward end  88  of the combustor  80 . Thus, the outer liner  108 , outer bands  110 ,  114 , shrouds  112 ,  116 , and 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 combustor dome  118 . Alternatively, as shown in the exemplary embodiment of  FIG. 4B , the unitary outer wall  102  includes a radially outer portion of the combustor dome  118 , such that the outer liner  108 , outer band  110 , shroud  112 , outer band  114 , shroud  116 , and a portion of the 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 combustor dome  118 . 
     Moreover, like the embodiments described with respect to  FIGS. 3A through 3F , the inner wall  120  of the embodiments shown in  FIGS. 4A, 4B, and 4C  at least includes the inner liner  122  of the combustor  80 . In some embodiments, such as illustrated in  FIGS. 4A and 4B , the inner wall  120  also includes the inner band  124  of the first turbine stage nozzle portion  82 N. In such embodiments, the inner liner  122  and inner band  124  are integrally formed as a single unitary structure, i.e., as a single unit or piece that may be referred to as unitary inner wall  120 . In other embodiments, as illustrated in  FIG. 4B , the unitary inner wall  120  may include a radially inner portion of the combustor dome  118  such that the inner liner  122  and the portion of the combustor dome  118  are integrally formed or constructed as a single unitary structure or such that the inner liner  122 , inner band  124 , and the portion of the combustor dome  118  are integrally formed or constructed as a single unitary structure. That is, in some embodiments, the unitary 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 still other embodiments, as shown in  FIG. 4C , the unitary inner wall  120  includes the combustor dome  118  defined at the forward end  88  of the combustor  80 . Thus, the combustor dome  118  and inner liner  122  (as well as inner band  124  in some embodiments) are integrally formed or constructed as a single unitary structure, i.e., inner wall  102  is a single unit or piece that includes combustor dome  118 . 
     Further, the first turbine stage nozzle airfoils  126  may be integrated with the outer wall  102  and/or with the inner wall  120 . As previously described, the first turbine stage nozzle airfoils  126  may be integrated with the outer wall  102 , but in other embodiments, the first turbine stage nozzle airfoils  126  may be integrated with the inner wall  120  and not the outer wall  102  or may be integrated with both the outer and inner walls  102 ,  120 . Whether formed separately from the walls  102 ,  120 , integrated with the inner wall  120  to form a single unitary structure with the inner wall  120 , integrated with the outer wall  102  to form a single unitary structure with the outer wall  102 , or integrated with both the outer and inner walls  102 ,  120  to form a single unitary structure with the outer and inner walls  102 ,  120 , a plurality of nozzle airfoils  126  extend from the inner wall  120  to the outer wall  102  within the first turbine stage nozzle portion  82 N. Additionally, as described above, the first turbine stage  82  includes a first stage rotor  134  having a plurality of rotor blade airfoils  130  attached thereto. Downstream of the first turbine stage  82 , a plurality of nozzle airfoils  128  extend from the inner band  136  to the outer wall  102  within the second turbine stage nozzle portion  84 N, and the second turbine stage blade portion  84 B includes a second stage rotor  138  having a plurality of rotor blade airfoils  130  attached thereto. 
     In the embodiments of  FIGS. 4A, 4B and 4C , the integrated or unitary outer wall  102  extends circumferentially about the integrated or unitary inner wall  120 . That is, the outer wall  102  circumferentially surrounds the inner wall  120  or the unitary outer wall  102  is a single piece extending 360° degrees about the inner wall  120 . As such, the outer wall  102  and the inner wall  120  define a generally annular flow path therebetween. Further, the combustor dome  118  extends across the forward end  88  of the combustor  80 , and whether integrated into the unitary outer wall  102  in whole or in part or integrated into the unitary inner wall  120  in whole or in part, the combustor dome  118  is a generally annular combustor dome  118 . 
     In addition, the flow path assembly  101  illustrated in the embodiments of  FIGS. 4A, 4B, and 4C  includes at least one opening  142  for receipt of a fuel nozzle assembly  90 . As described with respect to  FIGS. 3A through 3F , in some embodiments, the fuel nozzle assembly  90  may attach to the combustor dome  118 , which may be integrated with the outer wall  102  in whole as in the embodiment of  FIG. 4A  or in part as shown in  FIG. 4B , where the remainder is integrated with the inner wall  120 . As also described, the combustor dome  118  may be integrated with the inner wall  120  in whole as illustrated in  FIG. 4C , such that the fuel nozzle assembly  90  may attach to the combustor dome portion of unitary inner wall  120 . In other embodiments, the fuel nozzle assembly  90  does not attach to the combustor dome  118  but floats relative to the combustor dome  118  and the flow path  100 . As depicted, the fuel nozzle assembly  90  includes swirler  92 , which may be the portion of fuel nozzle assembly  90  that attaches to the combustor dome  118  or the portion that floats relative to the combustor dome  118  and flow path  100 . As previously described, the fuel nozzle assembly  90  or swirler  92  may float relative to the combustor dome  118  and flow path  100  along both the radial direction R and the axial direction A or only along one or the other of the radial and axial directions R, A. Moreover, as shown in  FIG. 3F , the combustor dome  118  may define a plurality of openings  142 , and each opening may receive a swirler  92  or other portion of fuel nozzle assembly  90 . 
     Referring still to  FIGS. 4A, 4B, and 4C , the unitary outer wall  102  and the inner wall  120  may define one or more features where the walls  102 ,  120  meet up with one another and, in some embodiments, may be attached to one another. For instance, in the embodiment of  FIG. 4A , the outer wall  102  defines a flange  144  along a radially inner edge of the outer wall  102  at the forward end  88  of the combustor  80 , and the inner wall  120  defines a flange  146  along a forward edge at the combustor forward end  88 . In the embodiment of  FIG. 4B , the outer wall flange  144  is defined along an edge of the combustor dome portion of the unitary outer wall  102 , and similarly, the inner wall flange  146  is defined along an edge of the combustor dome portion of the unitary inner wall  120 . As shown in  FIG. 4C , the outer wall  102  may define the outer wall flange  144  along a forward edge of the outer wall  102 , and the inner wall  120 , which includes combustor dome  118  in the illustrated embodiment, may define the inner wall flange  146  along a radially outer edge of the inner wall  120 .  FIGS. 4A, 4B, and 4C  illustrate that the flow path  100  may be discontinuous between the inner wall  120  and the outer wall  102 , i.e., formed from a separate inner and outer boundaries rather than integral inner and outer boundaries as shown in  FIGS. 3C through 3F . More particularly, the flow path  100  may be discontinuous where the outer wall flange  144  and the inner wall flange  146  are defined. 
     Thus, in the embodiment of  FIG. 4A , the outer wall  102  may be secured to the inner wall  120  at flanges  144 ,  146  near a radially inner, forward portion of the combustor  80 . Alternatively, the flanges  144 ,  146  as shown in  FIG. 4A  may define an area where the walls  102 ,  120  align or meet up with one another, e.g., flanges  144 ,  146  may define a slip joint between walls  102 ,  120 . In the embodiment of  FIG. 4B , the outer wall  102  may be secured to the inner wall  120  at flanges  144 ,  146  near a radial centerline of the combustor dome  118 . In other embodiments, the flanges  144 ,  146  as illustrated in  FIG. 4B  may define an area where the walls  102 ,  120  align or meet up with one another, e.g., flanges  144 ,  146  may define a slip joint between walls  102 ,  120 . In alternative embodiments, such as the embodiment of  FIG. 4C , the outer wall  120  may be secured to the inner wall  120  at flanges  144 ,  146  near a radially outer, forward portion of the combustor  80 , or the flanges  144 ,  146  as shown in  FIG. 4C  may define an area where the walls  102 ,  120  align or meet up with one another, e.g., flanges  144 ,  146  may define a slip joint between walls  102 ,  120  at a radially outer, forward portion of combustor  80 . In still other embodiments, the flanges  144 ,  146  may be defined in other locations such that the outer wall  102  and inner wall  120  are secured to, align, or meet up with one another at a location different from those depicted in  FIGS. 4A, 4B, and 4C . 
     Any suitable fastener or other attachment means may be used to secure the outer and inner walls  102 ,  120  at the flanges  144 ,  146 . For example, a plurality of apertures may be defined in each flange  144 ,  146 , and each aperture of the outer wall flange  144  may align with an aperture of the inner wall flange  146  for receipt of a fastener in each pair of aligned apertures. It will be appreciated that the outer wall  102  and the inner wall  120  may be attached to one another in other ways as well. Of course, in other embodiments as described above, the outer wall  102  and inner wall  120  may not be secured to one another but may move radially and/or axially with respect to one another. 
     Turning now to  FIGS. 5A, 5B, and 5C , schematic cross-sectional views are provided of the combustion section  26  and the HP turbine  28  of the turbine section of turbofan engine  10  according to other exemplary embodiments of the present subject matter. Unlike the embodiments of  FIGS. 3B through 3F  and  FIGS. 4A through 4C , the combustor dome  118  of the embodiments shown in  FIGS. 5A, 5B , and  5 C is not 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 . 
     Accordingly, as shown in  FIGS. 5A, 5B, and 5C , the outer wall  102  is a unitary outer wall including a combustor portion  104 , which extends through the combustion section  26  of engine  10 , and a turbine portion  106 , which extends through at least a first turbine stage of the turbine section of engine  10 . In the embodiments shown in  FIGS. 5A through 5C , the unitary outer wall  102  extends through the combustion section  26  to an aft end of HP turbine  28 , which includes two turbine stages  82 ,  84 . The combustor portion  104  and turbine portion  106  are integrally formed as a single unitary structure, i.e., unitary outer wall  102 . For example, as shown and described with respect to  FIG. 3A , the combustor portion  104  of the unitary outer wall  102  comprises the outer liner  108  of combustor  80 . The turbine portion  106  of unitary outer wall  102  comprises outer band  110  of 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 turbine portion  106  of unitary outer wall  102  also may include a plurality of nozzle airfoils  126 , which are integrally formed or constructed with the outer liner  108 , outer bands  110 ,  114 , and shrouds  112 ,  116  to form a single unitary structure, i.e., as a single unit or piece. 
     Further, as depicted in  FIGS. 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 instance, the inner wall  120  may comprise separate components defining the inner boundary of the flow path  100 . In other embodiments, the inner wall  120  may be a unitary inner wall  120  including an inner liner  122  and inner band  124  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 unitary inner wall  120  including inner liner  122 , inner band  124 , and first turbine stage nozzle airfoils  126  integrally formed as a single unitary structure, i.e., as a single unit or piece. Further, in the depicted embodiments of  FIGS. 5A, 5B, and 5C , the flow path  100  may be discontinuous between the inner wall  120  and the outer wall  102 , i.e., formed from a separate inner and outer boundaries rather than integral inner and outer boundaries as shown in  FIGS. 3C through 3F . More particularly, 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 . 
     Referring particularly to  FIG. 5A , the combustor dome  118  is positioned at forward end  88  of combustor  80  of 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 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  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  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 . 
     Further, 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. 
     As illustrated in  FIG. 5A , the combustor dome  118  includes an outer wing  152  and an inner wing  154 . The outer wing  152  extends aft along the outer wall  102 , and the inner wing  154  extends aft along the inner wall  120 . The wings  152 ,  154  may help guide the combustor dome  118  as it moves with respect to the outer wall  102  and inner wall  120 , and the wings  152 ,  154  also may help maintain the radial position or alignment of the combustor dome  118  as it moves axially. The wings may provide a consistent gap between the dome  118  and walls  102 ,  120  for purge and/or film starter flow as previously described. 
     Turning to  FIG. 5B , in other embodiments, each wing  152 ,  154  may extend forward from the combustor dome body  156 , rather than aft as shown in  FIG. 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 , e.g., to a metal dome supporting fuel nozzle assembly  90  and/or to either or both of the outer wall  102  and inner wall  120  at the forward end  88  of combustor  80 . In some embodiments, the forward-extending wings  152 ,  154  of combustor dome  118  may be pinned or otherwise attached to the outer wall  102  and the inner wall  120  as shown in  FIG. 5B . In still other embodiments, one of the wings  152 ,  154  may extend forward and the other wing  152 ,  154  may extend aft with respect to body  156 , and the combustor dome  118  may be attached to the fuel nozzle assembly  90  or to another component. 
     Referring now to  FIG. 5C , another exemplary embodiment of a separate combustor dome  118  and outer and inner walls  102 ,  120  is illustrated. In the embodiment illustrated in  FIG. 5C , the combustor dome  118  includes a forward-extending inner wing  154  but no outer wing  152 ; rather, an outer end  158  of the combustor dome  118  extends to the outer wall  102 . To retain the combustor dome  118  and seal against combustion gas leakage around the dome, the inner wing  154  is pinned with the inner wall  120  at the forward end  88  of the combustor  80 , and the outer end  158  is preloaded against the outer wall  102 . More particularly, a spring element  160  is pinned with the outer wall  102  at the combustor forward end  88 , and the spring element  160  presses against the body  156  of the combustor dome  118  to preload the outer end  158  of the combustor dome  118  into a lip  162  defined in the outer wall  102 . By utilizing the mounting configuration illustrated in  FIG. 5C , positive definite retention and sealing of the combustor dome  118  may be provided while minimizing thermal stresses in the dome, which is particularly useful when the combustor dome  118  is made from a CMC material. 
     Turning to  FIGS. 6, 7, and 8 , cross-sectional views are provided of a portion of the flow path assembly  101  according to exemplary embodiments of the present subject matter. As shown in the depicted embodiments, 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 . For example, the turbine portion  106  may be a high pressure turbine section  28  that extends through the first turbine stage  82  and a 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 embodiments, 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 be a unitary structure that may be referred to as unitary inner wall  120 ; for example, as shown in  FIGS. 7 and 8 , the inner wall  120  may be a unitary structure comprising the inner liner  122  and first turbine stage inner band  124 , which are integrally formed as unitary inner wall  120 . The second turbine stage inner band  136  also may be part of the inner wall  120 , although the inner band  136  is not integral with the inner liner  122  and first turbine stage inner band  124 . In some embodiments, e.g., as illustrated in  FIG. 7 , the unitary outer wall  102  or unitary inner wall  120  also may include the combustor dome  118 , or as illustrated in  FIG. 8 , the combustor dome  118  may be separate from both the outer wall  102  and the inner wall  120 . In other embodiments, the unitary outer wall  102  and the unitary inner wall  120  each may include a portion of the combustor dome  118 . In still other embodiments, the outer wall  102 , combustor dome  118 , and inner wall  120  may be integrally formed as a single piece, unitary structure. 
       FIGS. 6, 7, 8, and 9  depict various features for positioning the flow path assembly  101  using one or more positioning members, e.g., in a hub and spoke configuration, where the flow path assembly  101  is a hub, one or more spokes center and/or constrain the flow path assembly  101  in one or more directions while allowing for different thermal growth rates between different materials. As will be appreciated from the foregoing description of the gas turbine engine  10  and flow path assembly  101 , a positioning system including one or more positioning members may be used to position or center the flow path assembly  101  within the outer casing  18  of the engine  10 . Further, the positioning system may help position the flow path assembly  101  downstream of the compressor section  24  of the engine  10 . Moreover, the positioning system may help position the flow path assembly  101  with respect to the one or more fuel nozzle assemblies  90  of the engine  10 , e.g., such that the fuel nozzles are located at a proper depth with respect to the combustor  80 . Additionally, as previously described, the flow path assembly  101  comprises a unitary outer wall  102 , which at least forms a single piece outer boundary of the flow path  100  but also may integrate other portions of the flow path assembly  101 . The unitary outer wall  102  extends through the combustion section  26  and at least the first turbine stage  82  of the HP turbine section  28  but also may extend through additional turbine stages. Thus, the positioning members illustrated in  FIGS. 6, 7, 8, and 9  position the entire flow path assembly  101  within engine  10 , rather than several separate pieces of a flow path through the combustion and turbine sections (such as separate outer and inner liners, outer and inner bands, shrouds, etc.). Accordingly, methods of positioning the flow path assembly  101  within the engine  10  include positioning the single piece outer boundary of the flow path assembly  101  using the one or more positioning members as more fully described below. 
     Moreover, the positioning configurations described herein allow positive radial and angular positioning of the inner boundary of flow path  100 , e.g., inner wall  120  and inner band  136 , and any related inner boundary hardware. However, the positioning systems described herein do not over-constrain the inner boundary and related hardware or inhibit movement of the inner boundary and its hardware as the components thermally expand as engine temperatures increase. Thus, the positioning systems allow the inner boundary and its hardware to relatively freely thermally expand, thereby allowing relative radial growth between components having different coefficients of thermal expansion, such as CMC components and metallic components. 
     Referring particularly to the embodiment illustrated in  FIG. 6 , at least two positioning members or spokes center the flow path assembly  101 , with each positioning member extending from a structure external to the flow path assembly  101  to the outer wall  102 . More specifically, a first positioning member  162  extends from a first mounting component  164  attached to the outer casing  18  to a first receptacle  166  on the outer wall  102 , and a second positioning member  172  extends from a second mounting component  174  attached to the outer casing  18  to a second receptacle  176  on the outer wall  102 . As shown in  FIG. 6 , each of the first positioning member  162  and the second positioning member  172  is a generally cylindrical pin that passes through the respective mounting component to the respective receptacle. That is, the pin forming first positioning member  162  passes through an aperture in the first mounting component  164  and is received in an opening  165  in the first receptacle  166 . A first grommet  168  and a first bushing  167  in the first receptacle  166  help position and retain the first positioning member  162  in the first receptacle  166 , and a first nut  170  helps retain the first positioning member  162  within the first mounting component  164 . Similarly, the pin forming second positioning member  172  passes through an aperture in the second mounting component  174  and is received in an opening  175  in the second receptacle  176 . A second grommet  178  and a second bushing  177  in the second receptacle  176  help position and retain the second positioning member  172  in the second receptacle  176 , and a second nut  180  helps retain the second positioning member  172  within the second mounting component  174 . Each of the first mounting component  164  and the second mounting component  174  may be configured as, e.g., a hanger or the like. 
     As depicted in  FIG. 6 , the first receptacle  166  is located on an outer surface  103  of the outer wall  102  at the first turbine stage shroud portion  112  of the outer wall  102 , and the second receptacle  176  is located on the outer surface  103  at the second turbine stage shroud portion  116  of the outer wall  102 . As such, the first receptacle  166  is located over, or radially outward from, the first turbine stage blade portion  82 B, and the second receptacle  176  is located over, or radially outward from, the second turbine stage blade portion  84 B, such that the first and second positioning members  162 ,  172  are positioned over the blade portions of the turbine stages  82 ,  84 . Thus, any distortion of the flow path assembly  101  or surrounding structures that results from, e.g., thermal expansion, circumferential out-of-roundness, or the like may be controlled in the area of the positioning members  162 ,  172 , particularly in the radial direction R because the positioning members extend radially toward the flow path assembly  101 . More particularly, as described in greater detail herein, the flow path assembly  101 , or at least the outer wall  102  and inner wall  120 , may be made from a CMC material while the mounting components  164 ,  174  and other components of the engine  10  are metallic components, such that the CMC components have a different rate of thermal expansion than the metallic components. During operation of the engine  10 , the metallic components generally thermally expand at a greater rate than the CMC components, which may affect the position of the flow path assembly  101  within the engine  10 . By locating the positioning members  162 ,  172  over the blade portions  82 B,  84 B, the distortions or shifting position of flow path assembly  101  may be locally controlled in the radial direction R to help preserve the clearance gap G between the blade tips  140  and the outer wall  102 , which may be a critical area compared to other portions of the flow path assembly  101  because of, e.g., the impact on the performance of the engine  10 . More specifically, the thermal profile may not be uniform about the circumference of the flow path assembly  101 ; thus, the positioning members  162 ,  172  may help keep the flow path assembly  101  in round or enforce the profile of the flow path assembly  101  as the various components expand or contract with thermal changes in the engine. 
     Further, one or both of the positioning members  162 ,  172  may be allowed some axial movement within their respective receptacles  166 ,  176 , e.g., to account for some variation in part positioning along the axial direction A or to allow for axial thermal expansion. For instance, the opening  175  in the second receptacle  176  may have an oblong or slot shape in the axial direction A, with the grommet  178  and bushing  177  having complementary oblong shapes. As such, the generally cylindrical pin or second positioning member  172  may move along the axial direction A within the opening  175 , e.g., to absorb variations or changes in axial positioning of the flow path assembly  101 . On the other hand, the first and second positioning members  162 ,  172  may constrain the flow path assembly radially and tangentially, although some variation in positioning or in size may be allowed along the radial direction R, e.g., to account for thermal expansion or the like. 
     It will be appreciated that the first and second positioning members  162 ,  172  shown in  FIG. 6  are by way of example only, and a plurality of first and second positioning members  162 ,  172  may be used to center and/or constrain the flow path assembly  101 . That is, the plurality of positioning members  162 ,  172  may be circumferentially spaced apart from one another, and each of the plurality of positioning members  162 ,  172  may generally extend from one or more structures external to the flow path assembly  101  to the outer wall  102 , e.g., to one of a plurality of receptacles  166 ,  176  positioned on the outer wall  102 . The number of each positioning member  162 ,  172  may be optimized, e.g., to contribute a minimal amount of weight to the engine  10  through the positioning member  162 ,  172  while adequately positioning the flow path assembly  101  within the engine  10 . In some embodiments, an optimal number of first positioning members  162  will correspond to the number of airfoils, nozzle or blade airfoils, in the first turbine stage  82 , and an optimal number of second positioning members  172  will correspond to the number of airfoils, nozzle or blade airfoils, in the second turbine stage  84 . In other embodiments, the number of each positioning member  162 ,  172  may be less than or more than the number of airfoils in the respective turbine stage. 
     Turning now to  FIGS. 7 and 8 , another exemplary embodiment is provided of a hub and spoke configuration for positioning the flow path assembly  101  using one or more spokes or positioning members.  FIG. 7  provides a schematic cross-section view of the flow path assembly  101 , and  FIG. 8  provides a forward end view of the flow path assembly  101  and the positioning members with an external mounting component, which supports the positioning members, removed. Similar to the embodiment shown in  FIG. 6 , the embodiment depicted in  FIGS. 7 and 8  utilizes at least two, and preferably at least three, positioning members to center the flow path assembly  101 , with each positioning member extending adjacent to the inner wall  120  from a structure external to the flow path assembly  101 . More particularly, a positioning member  182  extends axially from a mounting component  184  into a slot  186  defined at an inner forward end  88   a  of the flow path assembly  101 . The mounting component  184  is external to and separate from the flow path assembly  101  and, e.g., helps support the fuel nozzle assembly  90 . Still more particularly, as shown in  FIG. 8 , a plurality of positioning members  182  extends axially aft from the external mounting component  184 , and each positioning member  182  is configured as a generally cylindrical pin. Each of the plurality of positioning members  182  is received in one of a plurality of slots  186  defined in the outer and inner wall flanges  144 ,  146  of the inner forward end  88   a  of the flow path assembly  101 . That is, as shown in  FIGS. 7 and 8  and as described above with respect to  FIG. 4A , in some embodiments, the combustor dome  118  is integrally formed with the outer wall  102 , and the unitary outer wall  102  including the combustor dome  118  defines an outer wall flange  144  along a radially inner edge at the forward end  88  of the flow path assembly  101 . Similarly, the inner wall  120  defines an inner wall flange  146  along a forward edge at the forward end  88 . The slots  186  may be defined through the outer and inner wall flanges  144 ,  146  such that the slots through each flange  144 ,  146  align when the flow path assembly  101  is assembled and such that a positioning member  182  may be received in each slot  186 . 
     Further, as illustrated in  FIG. 7 , the outer wall  102  defines an outer lip  188  that contacts an aft outer axial support  190  and the inner wall  120  defines an inner lip  192  that contacts an aft inner axial support  194  to limit the axial movement of the flow path assembly  101 . More specifically, the flow path assembly  101  may “float” axially, i.e., be allowed some axial movement due to thermal expansion, variable positions of components due to tolerance variations, etc., by sliding along the positioning members  182 . The outer and inner axial supports  190 ,  194  limit the aft axial movement of the flow path assembly  101  by preventing aft axial movement once the outer lip  188  contacts the outer axial support  190  and the inner lip  192  contacts the inner axial support  194 . The outer and inner axial supports  190 ,  194  may be attached to a structural component of the gas turbine engine  10 , such as a casing or other structural member. As described in greater detail herein, the flow path assembly  101 , or at least the outer wall  102  and inner wall  120  may be made from a CMC material while the mounting component  184 , outer axial support  190 , and inner axial support  194  are metallic components, such that the CMC components have a different rate of thermal expansion than the metallic components. During operation of the engine  10 , the metallic components generally thermally expand at a greater rate than the CMC components, which may affect the position of the flow path assembly  101  within the engine  10 . Thus, the axial positioning member(s)  182  and the interface between the outer lip  188  and outer axial support  190  and the inner lip  192  and inner axial support  194  help control the position of the flow path assembly  101  within the engine  10  while allowing for the differing rates of thermal expansion between the components. 
     Accordingly, similar to the embodiment described with respect to  FIG. 6 , the configuration of  FIGS. 7 and 8  positions the flow path assembly  101  axially, tangentially, and radially while generally constraining the assembly  101  tangentially and radially but allowing some relative movement axially. That is, the positioning members  182  located within the slots  186  as shown in  FIGS. 7 and 8  substantially prevent tangential movement of the flow path assembly  101  such that the assembly  101  does not rotate with respect to the axial centerline  12 . Further, the positioning members  182  and axial supports  190 ,  194  substantially limit radial movement of the flow path assembly  101 , although the assembly  101  may be allowed some radial movement, e.g., the slots  186  may be longer along the radial direction R than the radial height of the positioning members  182  such that the positioning members  182  can move radially within the slots  186 , to account for some variation in part positioning along the radial direction R or to allow for radial thermal expansion. Additionally, as previously described, the flow path assembly  101  may be allowed to move along the axial direction A, with the axial movement limited by the outer and inner axial supports  190 ,  194 , to allow for any variations in part tolerances or positioning, for thermal expansion, or the like. 
     Moreover, it will be appreciated that the embodiment shown in  FIGS. 7 and 8  is by way of example only and other similar configurations of the hub and spoke system for positioning the flow path assembly  101  may be used. For instance, as described with respect to  FIGS. 4A, 4B, and 4C , the combustor dome  118  may be integrally formed with the outer wall  102  or the inner wall  120 , or a portion of the combustor dome  118  may be integrally formed with the outer wall  102  while the remaining portion is integrally formed with the inner wall  120 . Thus, in some embodiments, the flanges  144 ,  146  may not be defined at the inner forward end  88   a  of the flow path assembly  101  but, for example, may be defined at an outer forward end  88   b  of the flow path assembly  101 . In such embodiments, the positioning members  182  may extend adjacent to the outer wall  102  from a component external to the flow path assembly  101 , such as mounting component  184 , rather than adjacent to the inner wall  120  as shown in  FIGS. 7 and 8 . 
     Referring now to  FIG. 9 , another exemplary embodiment is provided of a hub and spoke configuration for positioning the flow path assembly  101  using one or more spokes or positioning members. Similar to the embodiment shown in  FIGS. 6, 7, and 8 , the embodiment depicted in  FIG. 9  utilizes at least two, and preferably at least three, positioning members to center the flow path assembly  101 , with each positioning member extending to the outer wall  102  or inner wall  120  from a structure external to the flow path assembly  101 . More particularly, as described above with respect to  FIG. 5A , in some embodiments of flow path assembly  101 , the combustor dome  118  may be separate from the outer wall  102  and the inner wall  120 . As such, the positioning members for positioning the flow path assembly  101 , e.g., the hub in a hub and spoke configuration, may extend to both the outer wall  102  and the inner wall  120 . 
     As depicted in  FIG. 9 , a first positioning member  202  extends radially from a first mounting component  204  into an axial slot  206  defined in the outer wall  102  at the outer forward end  88   b  of the flow path assembly  101 . The first mounting component  204  is external to and separate from the flow path assembly  101  and, e.g., helps support the fuel nozzle assembly  90 . The first positioning member  202  is configured as a generally cylindrical pin, but other positioning member configurations may be used as well. Additionally, a plurality of first positioning members  202  may be provided, the first positioning members  202  circumferentially spaced apart from one another about the outer forward end  88   b . In such embodiments, the outer wall  102  may define a plurality of slots  206  such that each of the plurality of first positioning members  202  may be received in a separate slot  206 , or several first positioning members  202  may be received in the same slot  206 , i.e., fewer slots  206  than first positioning members  202  may be defined in the outer wall  102 . 
     Further, a second positioning member  212  extends radially from the first mounting component  204  into an aperture  214  defined in the inner wall  120  at the inner forward end  88   a  of the flow path assembly  101 . As shown in  FIG. 9 , the second positioning member  212  may be configured as a bolt and nut, such that the second positioning member  212  bolts the inner wall  120  to the first mounting component  204 . However, other configurations of positioning member  212  may be used as well, e.g., generally cylindrical pins or the like, and in some embodiments, the second positioning member  212  may be substantially similar to the first positioning member  202 . In addition, a plurality of second positioning members  212  may be provided, the second positioning members  212  circumferentially spaced apart from one another about the inner forward end  88   a . In such embodiments, the inner wall  120  may define a plurality of apertures  214  such that each of the plurality of second positioning members  212  may be received in an aperture  214 , or where second positioning members  212  are configured similar to the first positioning members  202 , several second positioning members  212  may be received in the same aperture  214 , i.e., fewer apertures  214  than second positioning members  212  may be defined in the inner wall  120 . 
     Moreover, a third positioning member  222  extends radially through a second mounting component  224  into an opening  226  defined in the first turbine stage shroud portion  112  of the outer wall  102 . The second mounting component  224  may be configured as a hanger or the like and may be attached to the outer casing  18  or another structural component of the gas turbine engine  10  and, as shown, is separate from and external to the flow path assembly  101 . A slot  228  is defined in the second mounting component  224 , through which the third positioning member  222  is inserted such that the third positioning member  222  extends radially toward the opening  226 . Further, a bushing  230  in the opening  226  helps position and retain the third positioning member  222  in the opening  226 . The third positioning member  222  may be generally cylindrical or may have any another suitable shape. Additionally, it will be appreciated that the third positioning member  222  depicted in  FIG. 9  is by way of example only, and a plurality of third positioning members  222  may be used to center and/or constrain the flow path assembly  101 . That is, the plurality of third positioning members  222  may be circumferentially spaced apart from one another, and each of the plurality of positioning members  222  may generally extend from one or more structures external to the flow path assembly  101  to the outer wall  102 , e.g., each third positioning member  222  may extend into one of a plurality of openings  226  defined in the outer wall  102 . Similar to the positioning members  162 ,  172  of  FIG. 6 , the number of third positioning members  222  may be optimized, e.g., to contribute a minimal amount of weight to the engine  10  through the positioning members  222  while adequately positioning the flow path assembly  101  within the engine  10 . In some embodiments, an optimal number of third positioning members  222  will correspond to the number of blades  130  in the first turbine stage blade portion  82 B, but in other embodiments, the number of third positioning members  222  may be less than or more than the number of blades  130  in the first turbine stage blade portion  82 B. 
     As illustrated in  FIG. 9 , the outer wall  102  is built up, or has an increased thickness, in the area of the first turbine stage shroud portion  112  of the outer wall  102 , which is above the first turbine stage blades  130 . The opening  226  is defined through the outer surface  103  of the outer wall  102  at the shroud  112  such that the third positioning member  222  is positioned over the blade portion of the first turbine stage  82 . As described with respect to  FIG. 6 , the flow path assembly  101 , or at least the outer wall  102  and inner wall  120 , may be made from a CMC material while the mounting components  204 ,  224  and other components of the engine  10  are metallic components, such that the CMC components have a different rate of thermal expansion than the metallic components. During operation of the engine  10 , the metallic components generally thermally expand at a greater rate than the CMC components, which may affect the position of the flow path assembly  101  within the engine  10 . Using the system illustrated in  FIG. 9 , any distortion of the flow path assembly  101  or surrounding structures that results from, e.g., thermal expansion or the like may be controlled in the area of the positioning members  202 ,  212 ,  222 . In particular, by locating the third positioning member(s)  222  over the blade portion  82 B, the distortions may be locally controlled in the radial direction R to help preserve the clearance gap G between the blade tips  140  and the outer wall  102 , which may be a critical area compared to other portions of the flow path assembly  101  because of, e.g., the impact on the performance of the engine  10 . More particularly, the thermal profile may not be uniform about the circumference of the flow path assembly  101 ; thus, the third positioning members(s)  222  may help keep the flow path assembly  101  in round or enforce the profile of the flow path assembly  101  as the various components expand or contract with thermal changes in the engine. The third positioning member(s)  222  also generally constrain the flow path assembly  101  axially and tangentially. Further, as illustrated in  FIG. 9 , the slot(s)  206  in which first positioning member(s)  202  is received may be sized to allow for some axial movement of the first positioning member(s)  202  within the slot(s)  206 , e.g., to account for some variation in part positioning along the axial direction A or to allow for axial thermal expansion. However, the second positioning member(s)  212  may constrain the inner wall  120  radially, tangentially, and axially by bolting the inner wall  120  to the first mounting component  204 . In other embodiments, the second positioning member(s)  212  may allow some relative movement of the inner wall  120 , e.g., to properly position the inner boundary of flow path  100  within the engine  10  or to compensate for thermal expansion of various components during engine operation. 
     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 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  as shown in the embodiments of  FIGS. 5A and 5B ) such as openings  142  in combustor forward end  88 , outer wall flange  144 , and inner wall flange  146 . 
     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.