Abstract:
Devices and methods are disclosed for making ceramic matrix composite (CMC) nozzles that limit thermal stresses from expansion and contraction, maintain tolerance on critical engineering dimensions, and reduces parasitic leakage associated with split line gaps in the CMC components. Cantilevered and herringbone patterns are formed by the split line gaps in the endwalls of the nozzles.

Description:
FIELD OF THE INVENTION 
       [0001]    The present subject matter relates generally to nozzles of gas turbine engines, and more particularly to devices and methods for making nozzles with split line gaps configured to reduce thermal stresses in the ceramic matrix composite (CMC) components and reduce parasitic leakage associated with the split line gaps. 
       BACKGROUND OF THE INVENTION 
       [0002]    A 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 enters 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 that route from the combustion section through a hot gas path defined within the turbine section, and then exhausted from the turbine section via the exhaust section. 
         [0003]    In particular configurations, the turbine section includes, in serial flow order, a high pressure (HP) turbine and a low pressure (LP) turbine. The HP turbine and the LP turbine each include various rotatable turbine components such as turbine rotor blades, rotor disks and retainers, and various stationary turbine components such as stator vanes or nozzles, turbine shrouds and engine frames. The rotatable and the stationary turbine components at least partially define the hot gas path through the turbine section. As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to the rotatable turbine components and the stationary turbine components. 
         [0004]    Nozzles utilized in gas turbine engines, and in particular HP turbine nozzles, are often arranged as an array of airfoil-shaped vanes extending between annular inner and outer endwalls which define the primary flowpath through the nozzles. Nozzles having integral inner and outer endwalls experience thermal stress concentration due to the closed structure of the nozzle assembly. The thermal stress and leakage of the components of neighboring nozzles arranged in an annular array is of particular concern for optimal gas turbine engine performance. Expansion and contraction of nozzle materials affects dimensions between features of neighboring nozzles, and in particular the airfoils. It is generally desirable that these engineering dimensions remain within desired predetermined tolerances for optimal gas turbine engine performance when the nozzles experience many cycles of thermal stress. If some of these dimensions are smaller than a predetermined optimal range, the gas turbine engine compressor can stall. If larger than the predetermined optimal range, the efficiency of the gas turbine engine can be lowered. 
         [0005]    Accordingly, improved devices and methods for making CMC nozzles is desired. In particular, methods and devices for making nozzles that limit thermal stresses from expansion and contraction, maintain tolerance on critical engineering dimensions, and reduces parasitic leakage associated with split line gaps in the CMC components would be advantageous. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    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. 
         [0007]    A cantilevered device is generally provided that limits both thermal stresses in the CMC components and leakage associated with split line gaps, along with methods for making such nozzles. 
         [0008]    In accordance with one embodiment, the cantilevered nozzle includes at least two airfoils configured in a cantilevered pattern, each airfoil having an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge. An outer endwall is disposed radially outward of each airfoil, the outer endwall defining a leading edge face, a trailing edge face, and a radially outwardly-facing end surface. An inner endwall is disposed radially inward of each airfoil, the inner endwall defining a leading edge face, a trailing edge face, and a radially inwardly-facing end surface. Only one of the outer endwall and the inner endwall is segmented and the other endwall is integral. At least one split line gap is disposed on the segmented endwall adjacent to an endwall side surface. The at least one split line gap is positioned in a generally axial direction between each airfoil and extends between the leading edge face and trailing edge face of the segmented endwall. 
         [0009]    In accordance with another embodiment, the cantilevered nozzle includes at least two airfoils configured in a herringbone pattern, each airfoil having an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge. An outer endwall is disposed radially outward of each airfoil, the outer endwall comprising a leading edge face, a trailing edge face, and a radially outwardly-facing end surface. An inner endwall is disposed radially inward of each airfoil, the inner endwall comprising a leading edge face, a trailing edge face, and a radially inwardly-facing end surface. At least two split line gaps are disposed alternately on the outer endwall and the inner endwall adjacent to an endwall side surface. The at least two split line gaps are positioned in a generally axial direction between the airfoils and extending between the leading edge face and trailing edge face of the outer endwall or the inner endwall. 
         [0010]    In accordance with another embodiment, a device and method of making a nozzle assembly is disclosed. The nozzle assembly includes at least two airfoils, each airfoil having an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge. An outer endwall is disposed radially outward of each airfoil, the outer endwall comprising a leading edge face, a trailing edge face, and a radially outwardly-facing end surface. An inner endwall is disposed radially inward of each airfoil, the inner endwall comprising a leading edge face, a trailing edge face, and a radially inwardly-facing end surface. At least one split line gap is disposed adjacent a side surface on a segmented endwall selected from at least one of the group consisting of the outer endwall and the inner endwall. At least one split line gap is positioned in a generally axial direction between each airfoil and extends between the leading edge face and trailing edge face of said segmented endwall. A nozzle support structure includes a strut extending through each airfoil, the outer endwall of the nozzle and the inner endwall of the nozzle. An outer hanger is disposed radially outward of each airfoil, the outer hanger comprising a radially inwardly-facing end surface adjacent said outer endwall outwardly-facing end surface. An inner hanger is disposed radially inward of each airfoil, the inner hanger comprising a radially outwardly-facing end surface adjacent said inner endwall inwardly-facing end surface. 
         [0011]    In some embodiments, the strut of the first nozzle assembly is joined to at least one of the inner hanger or the outer hanger of the first nozzle assembly and the strut of the second nozzle assembly is joined to at least one of the inner hanger or the outer hanger of the second nozzle assembly. In other embodiments, the strut of the first nozzle assembly is connected to at least one of the inner hanger or the outer hanger of the first nozzle assembly and the strut of the second nozzle assembly is connected to at least one of the inner hanger or the outer hanger of the second nozzle assembly. 
         [0012]    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 
         [0013]    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: 
           [0014]      FIG. 1  is a schematic cross-sectional view of a gas turbine engine in accordance with one embodiment of the present disclosure; 
           [0015]      FIG. 2  is an enlarged circumferential cross sectional side view of a high pressure turbine portion of a gas turbine engine in accordance with one embodiment of the present disclosure; 
           [0016]      FIG. 3  is a perspective view of an assembled nozzle assembly in accordance with one embodiment of the present disclosure; 
           [0017]      FIG. 4  is a perspective view of a fully segmented nozzle assembly with joined neighboring nozzles, without split line gaps of the present disclosure; 
           [0018]      FIG. 5  is a perspective view of a three airfoil segment of neighboring nozzles illustrating the outer endwall split line gaps between adjacent nozzles in accordance with the cantilevered embodiment of the present disclosure; 
           [0019]      FIG. 6  is a perspective view of joined neighboring nozzle array assembly in accordance with the cantilevered embodiment of the present disclosure; 
           [0020]      FIG. 7  is a perspective view of airfoils of neighboring nozzles illustrating the alternating outer and inner endwall split line gaps between adjacent nozzles in accordance with the herringbone embodiment of the present disclosure; 
           [0021]      FIG. 8  is a perspective view of joined neighboring nozzle array assembly in accordance with the herringbone embodiment of the present disclosure. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0022]    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 flow direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the flow direction from which the fluid flows, and “downstream” refers to the flow direction to which the fluid flows. 
         [0023]    Further, as used herein, the terms “axial” or “axially” refer to a dimension along a longitudinal axis of an engine. The term “forward” used in conjunction with “axial” or “axially” refers to a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “rear” used in conjunction with “axial” or “axially” refers to a direction toward the engine nozzle, or a component being relatively closer to the engine nozzle as compared to another component. The terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. 
         [0024]    Gas turbine nozzles having integral inner and outer endwalls experience thermal stress concentration due to the closed structure of the nozzle assembly. Splitting a single endwall, inner or outer, forms a cantilevered nozzle structure with split line gaps that allows the integral (non-split) endwall to drive the thermal response of the component without fighting stresses imposed by the opposite (split) endwall. Alternatively, splitting the inner and outer endwalls, to form a herringbone nozzle structure, with split line gaps that allows the integral (non-split) portion of the endwall to drive the thermal response of the component without fighting stresses imposed by the opposite (split) portion of the endwall. Additionally, these embodiments provide larger nozzle segments to be joined thereby reducing the number of joints, split line cuts and gaps. Balancing leakage from split line cuts as well as thermal stresses is a critical design optimization in turbine component design. The present disclosure increases nozzle design space and provides optimized leakage and stress designs. Partially combining components through integral endwalls provides leakage benefit over a fully segmented component that manifests as a reduction in parasitic flows in the turbine design. 
         [0025]    Referring now to the drawings,  FIG. 1  is a schematic cross-sectional view of an exemplary high-bypass turbofan type engine  10  herein referred to as “turbofan  10 ” as may incorporate various embodiments of the present disclosure. As shown in  FIG. 1 , the turbofan  10  has a longitudinal or axial centerline axis  12  that extends therethrough for reference purposes. In general, the turbofan  10  may include a core turbine or gas turbine engine  14  disposed downstream from a fan section  16 . 
         [0026]    The gas turbine engine  14  may generally include a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  may be formed from multiple casings. The outer casing  18  encases, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor  22 , a high pressure (HP) compressor  24 , a combustion section  26 , a turbine section including a high pressure (HP) turbine  28 , 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 . The (LP) spool  36  may also be connected to a fan spool or shaft  38  of the fan section  16 . In particular embodiments, the (LP) spool  36  may be connected directly to the fan spool  38  such as in a direct-drive configuration. In alternative configurations, the (LP) spool  36  may be connected to the fan spool  38  via a speed reduction device  37  such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within engine  10  as desired or required. 
         [0027]    As shown in  FIG. 1 , the fan section  16  includes a plurality of fan nozzles  40  that are coupled to and that extend radially outwardly from the fan spool  38 . An annular fan casing or nacelle  42  circumferentially surrounds the fan section  16  and/or at least a portion of the gas turbine engine  14 . It should be appreciated by those of ordinary skill in the art that the nacelle  42  may be configured to be supported relative to the gas turbine engine  14  by a plurality of circumferentially-spaced outlet guide vanes  44 . Moreover, a downstream section  46  of the nacelle  42  (downstream of the guide vanes  44 ) may extend over an outer portion of the gas turbine engine  14  so as to define a bypass airflow passage  48  therebetween. 
         [0028]      FIG. 2  provides an enlarged cross sectioned view of the HP turbine  28  portion of the gas turbine engine  14  as shown in  FIG. 1 , as may incorporate various embodiments of the present invention. As shown in  FIG. 2 , the HP turbine  28  includes, in serial flow relationship, a first stage  50  which includes an annular array  52  of stator vane nozzles  54  (only one shown) axially spaced from an annular array  56  of turbine rotor nozzles  58  (only one shown). The HP turbine  28  further includes a second stage  60  which includes an annular array  62  of stator vane nozzles  64  (only one shown) axially spaced from an annular array  66  of turbine rotor nozzles  68  (only one shown). The turbine rotor nozzles  58 ,  68  extend radially outwardly from and are coupled to the HP spool  34  ( FIG. 1 ). As shown in  FIG. 2 , the stator vane nozzles  54 ,  64  and the turbine rotor nozzles  58 ,  68  at least partially define a hot gas path  70  for routing combustion gases from the combustion section  26  ( FIG. 1 ) through the HP turbine  28 . 
         [0029]    As further shown in  FIG. 2 , the HP turbine may include one or more shroud assemblies, each of which forms an annular ring about an annular array of rotor nozzles. For example, a shroud assembly  72  may form an annular ring around the annular array  56  of rotor nozzles  58  of the first stage  50 , and a shroud assembly  74  may form an annular ring around the annular array  66  of turbine rotor nozzles  68  of the second stage  60 . In general, shrouds of the shroud assemblies  72 ,  74  are radially spaced from nozzle tips  76 ,  78  of each of the rotor nozzles  68 . A radial or clearance gap CL is defined between the nozzle tips  76 ,  78  and the shrouds. The shrouds and shroud assemblies generally reduce leakage from the hot gas path  70 . 
         [0030]    It should be noted that shrouds and shroud assemblies may additionally be utilized in a similar manner in the low pressure compressor  22 , high pressure compressor  24 , and/or low pressure turbine  30 . Accordingly, shrouds and shrouds assemblies as disclosed herein are not limited to use in HP turbines, and rather may be utilized in any suitable section of a gas turbine engine. 
         [0031]    The position and condition of stator vane nozzles  54 ,  64  in an engine  10  is of particular concern, especially as affected by expansion and contraction of the nozzles due to the thermal stress and leakage of the nozzle assembly as it experiences numerous hot gas operation cycles. Accordingly, and referring now to  FIG. 3 through 8 , the present disclosure is further directed to devices and methods for assembling neighboring nozzles  102  of a gas turbine engine  10  to include endwall split line gaps. The neighboring nozzles  102  in accordance with the present disclosure are nozzles which are or will be next to one another in an annular array in engine  10 . Nozzles  102  as disclosed herein may be utilized in place of stator vanes  54 , stator vanes  64 , or any other suitable stationary airfoil-based assemblies in an engine. 
         [0032]    As shown for example in  FIG. 3 , a nozzle  102  in accordance with the present disclosure includes an airfoil  110 , which has outer surfaces defining a pressure side  112 , a suction side  114 , a leading edge  116  and a trailing edge  118 . The pressure side  112  and suction side  114  extend between the leading edge  116  and the trailing edge  118 , as is generally understood. In typical embodiments, airfoil  110  is generally hollow, thus allowing cooling fluids to be flowed therethrough and structural reinforcement components to be disposed therein. 
         [0033]    Nozzle  102  can further include an inner endwall  120  and an outer endwall  130 , each of which is connected to the airfoil  110  at radially outer ends thereof generally along a radial direction  104 . For the cantilever embodiment ( FIGS. 5 and 6 ), adjacent nozzles  102  in an array of nozzles may be situated side by side along a circumferential direction  106 , as shown, and positioned or cut such that the inner endwall  120  is integral, or contiguous, and neighboring side surfaces of the segmented outer endwall  130  contain split line gaps and are not in contact thereby cantilevering each nozzle from its inner endwall. Similarly, the nozzles can cantilever from the outer endwall  130  with the split line gaps positioned on the inner endwall  120 . 
         [0034]    For the herringbone embodiment ( FIGS. 7 and 8 ), adjacent nozzles  102  in an array of nozzles may be situated side by side along a circumferential direction  106 , as shown, and positioned or cut such that every other neighboring nozzle of the inner endwall  120  contains a split line gap disposed at the nozzle side surface and are not in contact. Additionally, every other neighboring nozzle of the outer endwall  130  contains a split line gap disposed at the nozzle side surface and are not in contact, thereby forming a herringbone interconnecting pattern for the nozzle assembly. Inner endwall  120  may be disposed radially inward of the airfoil  110 , while outer endwall  130  may be disposed radially outward of the airfoil  110 . Inner endwall  120  may include, for example, a radially inwardly-facing end surface  121  and a radially outwardly-facing end surface  122  which are spaced apart radially from each other. Inner endwall  120  may further include various side surfaces, including a pressure side slash face  124 , suction side slash face  125 , leading edge face  126  and trailing edge face  127 . Similarly, outer endwall  130  may include, for example, a radially inwardly-facing end surface  131  and a radially outwardly-facing end surface  132  which are spaced apart radially from each other. Outer endwall  130  may further include various side surfaces, including a pressure side slash face  134 , suction side slash face  135 , leading edge face  136  and trailing edge face  137 . 
         [0035]    In exemplary embodiments, the airfoil  110 , inner endwall  120  and outer endwall  130  may be formed from ceramic matrix composite (“CMC”) materials. Alternatively, however, other suitable materials, such as suitable plastics, composites, metals, etc., may be utilized. 
         [0036]    Nozzles  102  may be subjected to various loads during operation of the engine  10 , including loads along an axial direction (as defined along the centerline  12 ). Further, differences in the materials utilized to form a nozzle  102  and associated support structure  108  (i.e. CMC and metal, respectively, in exemplary embodiments) may cause undesirable relative movements of the nozzle  102  and/or support structure  108  during engine operation, in particular along the radial direction  104 . It is generally desirable to improve the load transmission between the associated nozzle  102  and support structure  108  and reduce the risk of damage to the component of the nozzle  102  that interface with the support structure  108  due to such loading and relative movement. The split line gaps arranged in a cantilevered or herringbone pattern as described in the present disclosure provide space for relative movement within design dimensional tolerances thereby reducing thermal stress on the nozzle assembly components. 
         [0037]    As seen in  FIGS. 3 and 4 , neighboring nozzles  102  are referred to respectively as a first nozzle  210  and a second nozzle  212 . Neighboring nozzle assemblies  100  are referred to respectively as a first nozzle assembly  200  and a second nozzle assembly  202 . Neighboring nozzles support structures  108  are referred to respectively as a first nozzle support structure  220  and a second nozzle support structure  222 . First nozzle assembly  200  includes first nozzle  210  and first nozzle support structure  220 , and second nozzle assembly  202  includes second nozzle  212  and second nozzle support structure  222 . It should be understood that first and second nozzle assemblies  200 ,  202 , nozzles  210 ,  212 , and nozzle support structures  220 ,  222  may be any two neighboring nozzle assemblies  100 , nozzles  102 , and nozzle support structures  108 , respectively, within or to be utilized within an engine  10 . 
         [0038]    In  FIG. 3 , a nozzle  102  in accordance with the present disclosure includes an airfoil  110 , which has outer surfaces defining a pressure side  112 , a suction side  114 , a leading edge  116  and a trailing edge  118 . The pressure side  112  and suction side  114  extend between the leading edge  116  and the trailing edge  118 , as is generally understood. In typical embodiments, airfoil  110  is generally hollow, thus allowing cooling fluids to be flowed therethrough and structural reinforcement components to be disposed therein. 
         [0039]    As further illustrated in  FIG. 3 , nozzle  102  may be a component of a nozzle assembly  100 , which may additionally include a nozzle support structure  108 . Each support structure  108  may be coupled to a nozzle  102  to support the nozzle  102  in engine  10 . Further support structure  108  may transmit loads from the nozzle  102  to various other components within the engine  10 . 
         [0040]    Support structure  108  may include, for example, a strut  140 . Strut  140  may generally extend through the airfoil  110 , such as generally radially through the interior of the airfoil  110 . Strut  140  may further extend through the inner endwall  120  and the outer endwall  130 , such as through bore holes (not labeled) therein. In general, strut  208  may carry loads between the radial ends of the nozzle  102  to other components of the support structure  108 . The loads may be transferred through these components to other components of the engine  10 , such as the engine casing, etc. 
         [0041]    For example, support structure  108  may include an inner hanger  150  and an outer hanger  160 , each of which is connected to strut  140  at radially outer ends thereof generally along radial direction  104 . Adjacent support structures  108  in an array of support structures  108  may be situated side by side along circumferential direction  106 , as shown, with neighboring surfaces of the inner hangers  150  in contact and neighboring surfaces of the outer hangers  150 . Inner hanger  150  may be disposed radially inward of the strut  140 , while outer hanger  160  may be disposed radially outward of the strut  140 . Further, inner hanger  150  may be positioned generally radially inward of the airfoil  110  and inner endwall  120 . Outer hanger  160  may be positioned generally radially outward of the airfoil  110  and outer endwall  130 . Inner hanger  150  may include, for example, a radially inwardly-facing end surface  151  and a radially outwardly-facing end surface  152  which are spaced apart radially from each other. Inner hanger  150  may further include various side surfaces, including a pressure side slash face  154 , suction side slash face  155 , leading edge face  156  and trailing edge face  157 . Similarly, outer hanger  160  may include, for example, a radially inwardly-facing end surface  161  and a radially outwardly-facing end surface  162  which are spaced apart radially from each other. Outer hanger  160  may further include various side surfaces, including a pressure side slash face  164 , suction side slash face  165 , leading edge face  166  and trailing edge face  167 . 
         [0042]    In exemplary embodiments, the strut  140 , inner hanger  150  and outer hanger  160  are formed from metals. Alternatively, however, other suitable materials, such as suitable plastics, composites, etc., may be utilized. 
         [0043]    Accordingly, and referring now to  FIG. 5 , a three-airfoil nozzle segment  300  cantilevered from the inner endwall  120  in accordance with the present disclosure may further include one or more endwall split line gaps  200 ,  202  which are used to control nozzle material expansion and contraction loads between the associated nozzle  102  and support structure as well as between neighboring nozzles  102 . Each split line gap  200 ,  202  of a nozzle  102  is saw cut through the outer endwall or dimensionally formed on each nozzle segment. The split line gaps extend generally axially through the endwall from the leading edge face  136  to the trailing edge face  137 . The inner endwall  120  is integral, or contiguous, with no split line gaps. Alternatively, the nozzles  102  can be cantilevered from the outer endwall  130  with the split line gaps  200 ,  202  positioned on the inner endwall  120 . 
         [0044]      FIG. 6  is a perspective view of joined neighboring nozzle  102  array assembly in accordance with the cantilevered embodiment of the present disclosure and  FIG. 5 . The embodiment shown is cantilevered from the integral or contiguous inner endwall  120  with split line gaps  200 ,  202  positioned full perimeter on the outer endwall  130 . 
         [0045]      FIG. 7  is a perspective view of five airfoils  102  with two segments of neighboring nozzles illustrating the alternating outer endwall  408  and inner endwall  410  split line gaps  400 ,  402 ,  404 ,  406  between adjacent nozzle segments in accordance with the herringbone embodiment of the present disclosure. This embodiment may require additional connection joints at the interface between some of the airfoils and the endwalls. For example, one airfoil (of the two airfoil segment) may have no endwall, either outer or inner depending on the relative position of the segment in the array, until the neighboring segment is joined to provide the missing endwall. The connection may nest the airfoil inside of an endwall cavity that matches the airfoil profile. 
         [0046]      FIG. 8  is a perspective view of joined neighboring nozzle  102  array assembly in accordance with the herringbone embodiment of the present disclosure and  FIG. 7 . This embodiment shows the alternating split line gaps  400 ,  402 ,  404 , and  406  positioned between every other airfoil around the full perimeter on the outer endwall  130  and inner endwall  120 . 
         [0047]    Methods in accordance with the present disclosure may include, for example, assembling a first nozzle assembly  200  and a second nozzle assembly  202 .  FIGS. 3 and 4  illustrate one embodiment of a nozzle assembly, which may be a first nozzle assembly  200  or a second nozzle assembly  202 , which has been assembled in accordance with the present disclosure. In the embodiment of  FIG. 4 , the steps of assembling the first and second nozzle assemblies  200 ,  202  are performed before other steps of the present method, including a joining step as discussed herein. 
         [0048]    An assembled first or second nozzle assembly  200 ,  202  includes a nozzle  210 ,  212  and a nozzle support structure  220 ,  222 . The strut  140  of the nozzle support structure  220 ,  222  generally extends through the nozzle  210 ,  212 , such as through the airfoil  110 , inner endwall  120  and outer endwall  130  thereof. In exemplary embodiments, the step of assembling a first nozzle assembly  200  and/or second nozzle assembly  202  includes, for example, the step of inserting the strut  140  of the first or second nozzle support structure  220 ,  222  through the first or second nozzle  210 ,  222 , such as through the airfoil  210 , inner endwall  120  and outer endwall  130  thereof. The step of assembling the first nozzle assembly  200  and/or second nozzle assembly  202  may further include, for example, the step of joining the strut  140  of the first or second nozzle support structure  220 ,  222  to one or both of the inner hanger  150  or outer hanger  160  of the first or second nozzle support structure  220 ,  222 . In some embodiments, the strut  140  may be integral with one of the inner hanger  150  or outer hanger  160 , and thus not require joining to this hanger. In other embodiments, the strut  140  may require joining to both hangers  150 ,  160 . For example, in the embodiment of  FIG. 3 , the strut  140  is integral with the outer hanger  160  and joined to inner hanger  150 . 
         [0049]    Joining of components in accordance with the present disclosure may form a joint  230  between the components. In exemplary embodiments, joining is accomplished by brazing the components, such as the strut  140  and inner and/or outer hangers  150 ,  160 , together. Alternatively, joining may be accomplished by welding or another suitable joining technique. Joining techniques in accordance with the present disclosure generally utilized a melted and then solidified filler material and/or melted and then solidified surfaces of the components to fix the subject components together. Connecting of components in accordance with the present disclosure may be accomplished via, for example, a suitable mechanical fastener or another suitable technique that generally results in a removable connection. 
         [0050]    A method in accordance with the present disclosure may further include, for example, the step of joining the first nozzle support structure  210  and the second nozzle support structure  212  together. For example, the joining step may include joining the inner hangers  150  of the first nozzle support structure  210  and second nozzle support structure  222  together and joining the outer hangers  160  of the first nozzle support structure  210  and second nozzle support structure  212  together. In particular, and as shown for example in  FIG. 4 , the suction side slash face  155  of the inner hanger  150  of the first nozzle support structure  210  and the pressure side slash face  154  of the inner hanger  150  of the second nozzle support structure  212  may be joined together, and the suction side slash face  165  of the outer hanger  160  of the first nozzle support structure  210  and the pressure side slash face  164  of the outer hanger  160  of the second nozzle support structure  212  may be joined together. Connecting of components in accordance with the present disclosure may be accomplished via, for example, a suitable mechanical fastener or another suitable technique that generally results in a removable connection. 
         [0051]    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 languages of the claims.