Patent Publication Number: US-2017370583-A1

Title: Ceramic Matrix Composite Component for a Gas Turbine Engine

Description:
FEDERALLY SPONSORED RESEARCH 
     This invention was made with government support under contact number FA8650-07-C-2802 of the United States Air Force. The government may have certain rights in the invention. 
    
    
     FIELD OF THE INVENTION 
     The present subject matter relates generally to ceramic matrix composite components and, more particularly, to ceramic matrix composite components for gas turbine engines. 
     BACKGROUND OF THE INVENTION 
     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 general 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. 
     Typically, the gas turbine engine includes a combustor having a combustion chamber defined by a combustor liner. The combustor liner includes an inner liner wall and an outer liner wall. Immediately downstream of the combustor is a turbine nozzle stage, including stationary guide vanes, stator vanes, etc., provided to direct therethrough the flow of combustion gases from the combustion section. The turbine nozzle stage usually includes a plurality of circumferentially spaced turbine nozzle sections. Similar to the combustor liner, each nozzle section usually has an inner endwall and an outer endwall, with a nozzle extending therebetween. Thus, typical gas turbine engines utilize a combustor liner that is separate from the turbine nozzle sections immediately downstream of the combustor, requiring multiple seals between the liner and nozzle stage to attempt to control parasitic leakage between the combustor and first turbine nozzle stage. The seals and their associate hardware add weight and complexity to the engine, which can negatively engine performance and assembly. 
     In addition, non-traditional high temperature materials, such as ceramic matrix composite (CMC) materials, are more commonly being used for various components within gas turbine engines. For example, because CMC materials can withstand relatively extreme temperatures, there is particular interest in replacing components within the flow path of the combustion gases with CMC materials. Combustor liners and turbine nozzle stages each have surfaces and/or features exposed to or within the flow path of the combustion gases. 
     Accordingly, a combustor and turbine nozzle stage assembly that essentially eliminates the need for sealing without adding unnecessary weight or complexity would be desirable. For example, an integral combustor liner and turbine nozzle stage, which eliminates the need for sealing between the liner and the nozzle stage, would be beneficial. In particular, an integral CMC combustor liner and turbine nozzle stage, i.e., a combustor liner and turbine nozzle stage integrally formed from a CMC material, would be advantageous. A method for forming an integral CMC combustor liner and turbine nozzle stage also would be useful. 
     BRIEF DESCRIPTION OF THE INVENTION 
     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 ceramic matrix composite component for a gas turbine engine is provided. The ceramic matrix composite component includes an inner wall defining a first inner surface; an outer wall defining a second inner surface; and a nozzle extending from the inner wall to the outer wall. The inner wall, outer wall, and nozzle are integrally formed from a ceramic matrix composite material such that the inner wall, outer wall, and nozzle are a single unitary component. 
     In another exemplary embodiment of the present disclosure, a method is provided for forming a ceramic matrix composite component of a gas turbine engine. The method includes laying up a plurality of plies of a ceramic matrix composite material; processing the plurality of plies to form a green state component; firing the green state component; and densifying the fired component to produce a final unitary component. The unitary component comprises a combustor liner portion and a combustor discharge nozzle stage portion. 
     In one exemplary aspect of the present disclosure, a method is provided for forming a ceramic matrix composite component of a gas turbine engine. The method includes laying up a plurality of plies of a ceramic matrix composite material; processing the plurality of plies to form a green state component; firing the green state component; and densifying the fired component to produce a final unitary component. Laying up the plurality of plies comprises interspersing a plurality of combustor liner plies with a plurality of combustor discharge nozzle stage plies. Further, the unitary component comprises an inner wall and an outer wall, and the inner and outer wall define a combustion chamber adjacent a forward end of the unitary component. The unitary component also comprises a nozzle extending from the inner wall to outer wall adjacent an aft end of the unitary component. 
     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  is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter. 
         FIG. 2  is a close-up, side view of a combustion section and a turbine section of the exemplary gas turbine engine of  FIG. 1 . 
         FIG. 3A  is a schematic view of a plurality of CMC plies of an integral combustor liner and combustor discharge nozzle stage in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 3B  is a schematic view of interspersed CMC plies of an integral combustor liner and combustor discharge nozzle stage in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 3C  is a schematic view of an integral combustor liner and combustor discharge nozzle stage after firing and densification in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 4  is a flow diagram of a method for forming an integral combustor liner and combustor discharge nozzle stage in accordance with an exemplary embodiment of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     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 turbomachine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of  FIG. 1 , the turbomachine is configured as a gas turbine engine, or rather as a high-bypass turbofan jet engine  12 , referred to herein as “turbofan engine  12 .” As shown in  FIG. 1 , the turbofan engine  12  defines an axial direction A (extending parallel to a longitudinal centerline  13  provided for reference), a radial direction R, and a circumferential direction C (extending about the longitudinal centerline  13 ) extending about the axial direction A. 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 and the core turbine engine  16  includes, 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 . Accordingly, the LP shaft  36  and HP shaft  34  are each rotary components, rotating about the axial direction A during operation of the turbofan engine  12 . 
     Referring still to the embodiment of  FIG. 1 , the fan section  14  includes a variable pitch fan  38  having a plurality of fan blades  40  coupled to a disk  42  in a spaced apart manner. As depicted, the fan blades  40  extend outwardly from disk  42  generally along the radial direction R. Each fan blade  40  is rotatable relative to the disk  42  about a pitch axis P by virtue of the fan blades  40  being operatively coupled to a suitable pitch change mechanism  44  configured to collectively vary the pitch of the fan blades  40  in unison. The fan blades  40 , disk  42 , and pitch change mechanism  44  are together rotatable about the longitudinal axis  12  by LP shaft  36  across a power gear box  46 . The power gear box  46  includes a plurality of gears for adjusting the rotational speed of the fan  38  relative to the LP shaft  36  to a more efficient rotational fan speed. More particularly, the fan section includes a fan shaft rotatable by the LP shaft  36  across the power gearbox  46 . Accordingly, the fan shaft may also be considered a rotary component, and is similarly supported by one or more bearings. 
     Referring still to the exemplary embodiment of  FIG. 1 , the disk  42  is covered by a rotatable front hub  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 . The exemplary nacelle  50  is 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  extends 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  12 , a volume of air  58  enters the turbofan  10  through an associated inlet  60  of the nacelle  50  and/or fan section  14 . As the volume of air  58  passes across the 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 arrow  64  is directed or routed into the core air flowpath  37 , or more specifically 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 . 
     In some embodiments, components of turbofan engine  12 , particularly components within hot gas path  78 , may comprise a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such components may include silicon carbide, 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). As further examples, the CMC materials may also include silicon carbide (SiC) or carbon fiber cloth. 
     Referring now to  FIG. 2 , a close-up, cross-sectional view is provided of the turbofan engine  12  of  FIG. 1  and particularly of the combustion section  26  and the HP turbine  28  of the turbine section. The depicted combustion section  26  generally includes an annular combustor  80 , and downstream of the combustion section  26 , the HP turbine  28  includes a plurality of turbine component stages. Each turbine component stage comprises a plurality of turbine components. More particularly, for the depicted embodiment, HP turbine  28  includes a plurality of turbine nozzle stages, such as first and second turbine nozzle stages  82 ,  84  shown in  FIG. 2 , as well as one or more stages of turbine rotor blades, such as turbine rotor blade stage  86 . 
     Typically, the combustor includes a combustion chamber defined by a combustor liner having an inner liner wall and an outer liner wall, and the HP turbine includes a first turbine nozzle stage located immediately downstream from the combustion section, such that the first turbine nozzle stage also may be referred to as a combustor discharge nozzle stage. The combustor discharge nozzle stage usually includes a plurality of circumferentially spaced turbine nozzle sections. Each nozzle section includes an inner endwall and an outer endwall, with a nozzle extending generally radially from the inner endwall to the outer endwall. Thus, typical turbofan engines utilize a combustor liner that is separate from the turbine nozzle sections immediately downstream of the combustor. 
     However, as illustrated in  FIG. 2 , turbofan engine  12  includes an integral combustor liner and combustor discharge nozzle stage  100 . The integral combustor liner and combustor discharge nozzle stage  100  depicted in  FIG. 2  has a forward end  102  and an aft end  104 . A combustor liner portion  106  is defined adjacent forward end  102 , and a combustor discharge nozzle stage portion  108  is defined adjacent aft end  104 . 
     Integral liner and nozzle stage  100  also includes an inner wall  110  defining a first inner surface  112  of integral liner and nozzle stage  100  and an outer wall  114  defining a second inner surface  116  of integral liner and nozzle stage  100 . In the depicted embodiment of  FIG. 2 , outer wall  114  extends generally circumferentially about inner wall  110 , i.e., outer wall  114  is spaced radially outward from inner wall  110 . A nozzle  118  extends generally radially, i.e., generally along the radial direction R, from inner wall  110  to outer wall  114  within the combustor discharge nozzle stage portion  108 . It will be appreciated that, while only one nozzle  118  is depicted in  FIG. 2 , integral liner and nozzle stage  100  includes a plurality of nozzles  118  spaced generally circumferentially about longitudinal centerline  13  within combustor discharge nozzle stage portion  108 . Each nozzle  118  of the plurality of nozzles extends generally radially from inner wall  110  to outer wall  114 . 
     The inner wall  110 , outer wall  114 , and nozzle  118  are integrally formed from a ceramic matrix composite material such that the inner wall  110 , outer wall  114 , and nozzle  118  are a single unitary component. More particularly, where integral liner and nozzle stage  100  includes a plurality of nozzles  118 , each nozzle  118  is integrally formed with inner wall  110  and outer wall  114  such that inner wall  110 , outer wall  114 , and the plurality of nozzles  118  are a single unitary component. As such, integral combustor liner and combustor discharge nozzle stage  100  also may be referred to as integral component  100  or unitary component  100 . In an exemplary embodiment, integral component  100  is formed from a CMC material. Methods and/or processes for forming an integral combustor liner and combustor discharge nozzle stage  100 , particularly an integral CMC combustor liner and combustor discharge nozzle stage, are described in greater detail below. 
     Further, the term “unitary” as used herein denotes that the associated component, particularly integral combustor liner and combustor discharge nozzle stage  100 , is made as a single piece during manufacturing, i.e., the unitary component is a continuous piece of material. Thus, a unitary component has a monolithic construction and is different from a component that has been made from a plurality of component pieces that have been joined together to form a single component. More specifically, in the exemplary embodiment of  FIG. 2 , inner wall  110 , outer wall  114 , and nozzle  118  are constructed as a single unit or piece to form unitary component  100 . 
     Referring still to  FIG. 2 , within combustor liner portion  106  of unitary component  100 , inner wall  110  and outer wall  114  define a combustion chamber  120  at or adjacent forward end  102  that extends generally along the axial direction A. Accordingly, a portion  110 C of inner wall  110  and a portion  114 C of outer wall  114  essentially define a combustor liner and, thus, form combustor liner portion  106  of unitary component  100 . At the aft end  104  of unitary component  100 , a portion  110 N of inner wall  110  and a portion  114 N of outer wall  114 , with nozzle  118  extending therebetween, essentially define a first nozzle stage of HP turbine  28  and, thus, form combustor discharge nozzle stage  108  of unitary component  100 . 
     A plurality of fuel nozzles  88  are positioned at forward end  102  of unitary component  100  for providing combustion chamber  120  with a mixture of fuel and compressed air from the compressor section. As discussed above, the fuel and air mixture is combusted within the combustion chamber  120  to generate a flow of combustion gases therethrough. As such, first inner surface  112  and second inner surface  116  generally define a hot side of unitary component  100 . The hot side is exposed to and defines in part a portion of the core air flowpath  37  extending through combustion chamber  120 , as well as combustor discharge nozzle stage portion  108  such that nozzle  118  is positioned within the core air flowpath  37 . Opposite the hot side is a cold side  122 , and although not depicted, inner wall  110  and/or outer wall  114  may include thermal management features, such as one or more cooling holes extending from the cold side to the hot side, to maintain a temperature of inner wall  110  and/or outer wall  114  within a desired operating temperature range. 
     Additionally, for the depicted exemplary embodiment of  FIG. 2 , turbofan engine  12  includes second turbine nozzle stage  84  downstream of integral combustor liner and combustor discharge nozzle stage  100 . That is, integral combustor liner and combustor discharge nozzle stage  100  extends from forward end  102  adjacent fuel nozzles  88  to aft end  104  adjacent second turbine nozzle stage  84  such that integral component  100  extends within combustion section  26  and HP turbine section  28 . Second turbine nozzle stage  84  includes a plurality of turbine nozzle sections  85  spaced along the circumferential direction C. Each second turbine nozzle section  85  includes a second stage turbine nozzle  87  positioned within the core air flowpath  37 , as well as an inner endwall  90  and an outer endwall  91 , with the second stage turbine nozzle  87  extending generally along the radial direction R from the inner endwall  90  to the outer endwall  91 . The inner endwall  90  and outer endwall  91  of the second nozzle section  85  each define a cold side  92   c  and an opposite hot side  92   h  exposed to and at least partially defining the core air flowpath  37 . 
     Located immediately downstream of the unitary component  100  and immediately upstream of the second turbine nozzle stage  84 , the HP turbine  28  includes a first stage  86  of turbine rotor blades  93 . First stage  86  of turbine rotor blades  93  includes a plurality of turbine rotor blades  93  spaced along the circumferential direction C and a first stage rotor  94 . The plurality of turbine rotor blades  93  are attached to first stage rotor  94 . Although not depicted, turbine rotor  94  is, in turn, connected to the HP shaft  34  ( FIG. 1 ). In such manner, turbine rotor blades  93  may extract kinetic energy from the flow of combustion gases through the core air flowpath  37  defined by the HP turbine  28  as rotational energy applied to the HP shaft  34 . Turbofan engine  12  additionally includes a shroud  95  exposed to and at least partially defining the core air flowpath  37 . Further, similar to inner wall  110  and outer wall  114  of unitary component  100  and inner endwall  90  and outer endwall  91  of second turbine nozzle stage  84 , each of the turbine rotor blades  93  includes a wall or platform  96 . Platform  96  of each of the turbine rotor blades  93  defines a cold side  97   c  and an opposite hot side  97   h  exposed to and at least in part defining the core air flowpath  37 . 
     As further illustrated in  FIG. 2 , aft end  104  of unitary component  100  includes a seal  98 , and each turbine nozzle section  85  of second turbine nozzle stage  84  includes a seal  98 . Additionally, platform  96  of each turbine rotor blade  93  includes a seal  99 . Seals  99  are configured to interact with the seals  98  of discharge nozzle stage portion  108  of unitary component  100  and turbine nozzle sections  85  forming second turbine nozzle stage  84 . The interaction of seals  98 ,  99  helps to prevent an undesired flow of combustion gases from the core air flowpath  37  between the first stage  86  of turbine rotor blades  93  and integral liner and nozzle stage  100 , as well as between first turbine blade stage  86  and second turbine nozzle stage  84 . However, as shown in  FIG. 2 , because combustor liner portion  106  is integrally formed with combustor discharge nozzle stage portion  108 , no seals are required to prevent undesired leakage of combustion gases between combustor  80  and the first stage  82  of turbine nozzles, i.e., combustor discharge nozzle stage portion  108  of unitary component  100 . As such, any leakage between the combustor and first turbine nozzle stage may be essentially eliminated, as well as any weight and complexity attributable to seals or sealing mechanisms that would be used between a combustor liner and combustor discharge nozzle stage when the combustor liner is separate from the combustor discharge nozzle stage. 
     Referring now to the schematic illustrations of  FIGS. 3A through 3C , integral combustor liner and combustor discharge nozzle stage  100  will be described in greater detail. Turning to  FIG. 3A , a plurality of plies  124  of a CMC material may be used to form the integral component  100 . In such embodiments, inner wall  110 , outer wall  114 , and nozzle  118  are formed from the CMC plies  124 . CMC plies  124  may be, e.g., plies pre-impregnated (pre-preg) with matrix material and may be formed from pre-preg tapes or the like. For example, the CMC plies may be formed from a prepreg tape comprising a desired ceramic fiber reinforcement material, one or more precursors of the CMC matrix material, and organic resin binders. According to conventional practice, prepreg tapes can be formed by impregnating the reinforcement material with a slurry that contains the ceramic precursor(s) and binders. The slurry also may contain solvents for the binders that promote the fluidity of the slurry to enable impregnation of the fiber reinforcement material, as well as one or more particulate fillers intended to be present in the ceramic matrix of the CMC component, e.g., silicon and/or SiC powders in the case of a Si—SiC matrix. Preferred materials for the precursor will depend on the particular composition desired for the ceramic matrix of the CMC component. For example, the precursor material may be 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). 
     As shown schematically in  FIG. 3B , the plurality of CMC plies  124  may include a plurality of CMC plies  126  for forming combustor liner portion  106  and a plurality of CMC plies  128  for forming combustor discharge nozzle stage portion  108 . Liner plies  126  may include plies for forming inner wall  110 C of combustor liner portion  106 , as well as plies for forming outer wall  114 C of combustor liner portion  106 . Similarly, nozzle stage plies  128  may include plies for forming inner wall  110 N of combustor discharge nozzle stage portion  108 , plies for forming outer wall  114 N of combustor discharge nozzle stage portion  108 , and plies for forming nozzles  118  of combustor discharge nozzle stage portion  108 . As such, nozzle stage plies  128  include plies for forming an inner endwall, an outer endwall, and a plurality of nozzles of a combustor discharge turbine nozzle stage. 
     In the exemplary embodiment depicted in  FIG. 3B , liner plies  126  and nozzle stage plies  128  are interspersed with one another. More specifically, where liner plies  126  meet nozzle stage plies  128 , plies  126  are alternated with plies  128  to integrate the plies for forming combustor liner portion  106  with the plies for forming combustor discharge nozzle stage portion  108 . That is, any joints between plies  126 ,  128  may be formed by alternating layers of plies  126 ,  128 . In some embodiments, single plies  126 ,  128  may be alternated to integrate plies  126  and  128  and thereby integrate combustor liner portion  106  with combustor discharge nozzle stage portion  108 . In other embodiments, one or more liner plies  126  may be formed in a stack that is alternated with a stack of one or more nozzle stage plies  128  to integrate plies  126  and  128  and thereby integrate combustor liner portion  106  with combustor discharge nozzle stage portion  108 . 
     Of course, integral combustor liner and combustor discharge nozzle stage  100  may be formed from a plurality of inner wall plies, a plurality of outer wall plies, and a plurality of nozzle plies, each ply made from a CMC material. The inner wall, outer wall, and nozzle plies may be interspersed, e.g., alternated where the plies meet as shown in  FIG. 3B , to form integral combustor liner and combustor discharge nozzle stage  100 . In this way, the plies forming the combustor liner portion  106  are interspersed, and thereby integrated, with the plies forming the combustor discharge nozzle stage portion  108 . 
     Further, it will be appreciated that any spacing between adjacent plies  126  and adjacent plies  128  shown in  FIG. 3B  is for purposes of illustration only. For example, in various embodiments, little to no space may be defined between adjacent plies  126  and adjacent plies  128  when plies  126 ,  128  are laid up during the process of forming the integral combustor liner and combustor discharge nozzle stage  100 . Rather, in exemplary embodiments, a ply  126  may be in contact with adjacent plies  126 , except where plies  126  are interspersed with plies  128  as described above. Of course, some spacing between adjacent plies  126  and/or adjacent plies  128  may result in the layup of plies  126 ,  128 , but not necessarily to the extent or between every adjacent ply as shown in the schematic representation of  FIG. 3B . 
     Referring now to  FIG. 3C , in an exemplary embodiment, the plurality of plies  124  defining inner wall  110 , outer wall  114 , and nozzle  118  are cured to produce a single piece component  100 , then fired and subjected to silicon melt-infiltration to form final unitary component  100 . For example, plies  124  may be processed in an autoclave to produce a green state integral liner and discharge nozzle stage  100 . Then, green state component  100  may be placed in a furnace with a piece or slab of silicon and fired to melt infiltrate the component  100  with silicon. More particularly, for unitary component  100  formed from CMC plies  124  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 component is fired with silicon, component  100  undergoes silicon melt-infiltration. The melt-infiltrated CMC body hardens to a final unitary CMC component  100 . 
       FIG. 4  provides a chart illustrating a method  400  for forming integral combustor liner and combustor discharge nozzle stage  100  according to an exemplary embodiment of the present subject matter. As shown at  402  in  FIG. 4 , a plurality of plies  124  of a CMC material for forming the unitary component  100  may be laid up to define a desired shape. During the layup generally shown at  402 , a desired component shape may be generally defined; the component shape may be finally defined after the plies are processed and machined as needed. Plies  124  may be laid up on a layup tool, mandrel, mold, or other appropriate device for supporting the plies and/or for defining the desired shape. Further, laying up plies  124  may comprise layering liner plies  126  and nozzle stage plies  128 , or inner wall, outer wall, and nozzle plies, by alternating layers of plies  126 ,  128  as previously described. That is, laying up plies  124  may include interspersing liner and nozzle stage plies  126 ,  128  or inner wall, outer wall, and nozzle plies. Interspersing plies  124  forming combustion liner portion  106  and combustor discharge nozzle stage portion  108  integrates portions  106 ,  108  such that the resultant component is integral combustor liner and combustor discharge nozzle stage  100 . 
     After the plies  124  are laid up, the plies may be processed, e.g., compacted and cured in an autoclave, as shown at  404  in  FIG. 4 . After processing, the plies form a green state component  100 , i.e., a green state integral liner and nozzle stage  100 . Green state component  100  is a single piece component, i.e., curing plies  124  produces a unitary component  100  formed from a continuous piece of CMC material. The green state component  100  then may undergo firing and densification, illustrated at  406  and  408  in  FIG. 4 , to produce a final unitary component  100 . As previously described, the unitary component  100  comprises inner wall  110  and outer wall  114 , which define combustor liner portion  106  adjacent the forward end  102  of component  100  and combustor discharge nozzle stage portion  108  adjacent the aft end  104  of component  100 . Nozzle  118  extends from inner  110  and outer wall  114  of unitary component  100 . 
     In an exemplary embodiment of method  400 , the green state component  100  is placed in a furnace with silicon to burn off any mandrel-forming materials and/or solvents used in forming the CMC plies  124 , to decompose binders in the solvents, and to convert a ceramic matrix precursor of the plies into the ceramic material of the matrix of the unitary CMC component  100 . The silicon melts and infiltrates any porosity created with the matrix as a result of the decomposition of the binder during burn-off/firing. 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. 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 materials to melt-infiltrate into the component  100 . After firing and densification, as shown at  410  in  FIG. 4 , the unitary component  100 , having combustor liner portion  106  and combustor discharge nozzle stage portion  108 , may be finish machined, if and as needed. Additionally or alternatively, an environmental barrier coating (EBC) may be applied to unitary component  100 . 
     Method  400  is provided by way of example only. For example, other processing cycles, e.g., utilizing other known methods or techniques for compacting and/or curing CMC plies, may be used. Further, unitary component  100  may be post-processed or densified using a melt-infiltration process or a chemical vapor infiltration process, or component  100  may be a matrix of pre-ceramic polymer fired to obtain a ceramic matrix. Alternatively, any combinations of these or other known processes 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.