Patent Publication Number: US-11391158-B2

Title: Composite airfoil assembly with separate airfoil, inner band, and outer band

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of and claims priority to U.S. application Ser. No. 15/922,263, filed Mar. 15, 2018, the contents of which are incorporated herein by reference. 
    
    
     FEDERALLY SPONSORED RESEARCH 
     This invention was made with government support under contact number FA8650-15-D-2501 awarded by the Department of the Air Force. The U.S. government may have certain rights in the invention. 
    
    
     FIELD 
     The present subject matter relates generally to gas turbine engines. More particularly, the present subject matter relates to composite airfoil assemblies for gas turbine engines, such as composite turbine nozzle fairings for gas turbine engines. 
     BACKGROUND 
     More commonly, non-traditional high temperature composite materials, such as ceramic matrix composite (CMC) materials, are being used in applications such as gas turbine engines. Components fabricated from such materials have a higher temperature capability compared with typical components, e.g., metal components, which may allow improved component performance and/or increased engine temperatures. Composite components may provide other advantages as well, such as an improved strength to weight ratio. 
     Typically, a CMC turbine nozzle fairing comprises an airfoil, an inner band, and an outer band that are integrally formed as one single component, with curved transition zones between the airfoil and each of the inner band and outer band. However, the transition from the airfoil to the band sections in the CMC turbine nozzle fairing generally comprises complex shapes in the vicinity of the curvature such that the nozzle fairings are difficult to lay up, resulting in a long manufacturing cycle time and low yield, and also are difficult to compact, often resulting in poor compaction. Additionally, thermal differences, i.e., a thermal fight, between the airfoil and bands produce high stresses in the nozzle fairings, which limits the acceptability of part defects and results in tighter inspection limits for non-destructive examination of the parts. Moreover, known CMC nozzle fairings typically are singlets and can allow leakage between each separate nozzle fairing. 
     Accordingly, improved airfoil assemblies would be useful. In particular, an airfoil assembly comprising an airfoil that is separate from each of the inner band and outer band would be advantageous. Further, an airfoil assembly having a separate airfoil, inner band, and outer band that is simply supported, with a positively located airfoil, would be desirable. 
     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 subject matter, an airfoil assembly for a gas turbine engine is provided. The airfoil assembly comprises an airfoil having a concave pressure side opposite a convex suction side and an inner end radially spaced apart from an outer end. The pressure side and the suction side extend axially from a leading edge to a trailing edge. The airfoil assembly further comprises an inner band defining an inner opening shaped complementary to the inner end of the airfoil and an outer band defining an outer opening shaped complementary to the outer end of the airfoil. The inner end of the airfoil is received with the inner opening and the outer end of the airfoil is received within the outer opening. The airfoil assembly also comprises a strut extending radially through a cavity defined by the airfoil, as well as a first pad defined at a first radial location within the cavity and a second pad defined at a second radial location within the cavity. The first radial location is different from the second radial location. 
     In another exemplary embodiment of the present subject matter, an airfoil assembly for a gas turbine engine is provided. The airfoil assembly comprises an airfoil having a concave pressure side opposite a convex suction side and an inner end radially spaced apart from an outer end. The pressure side and the suction side extend axially from a leading edge to a trailing edge. The airfoil assembly also comprises an inner band defining an inner opening shaped complementary to the inner end of the airfoil and an outer band defining an outer opening shaped complementary to the outer end of the airfoil. The inner end of the airfoil is received with the inner opening and the outer end of the airfoil is received within the outer opening. The inner band includes a first flowpath surface, a first non-flowpath surface opposite the first flowpath surface, and a first inner flange extending radially from the first non-flowpath surface. The outer band includes a second flowpath surface, a second non-flowpath surface opposite the second flowpath surface, and a first outer flange extending radially from the second non-flowpath surface. The inner band is secured to an inner support structure by a first inner fastener extending through the first inner flange, and the outer band is secured to an outer support structure by a first outer fastener extending through the first outer flange. 
     In a further exemplary embodiment of the present subject matter, an airfoil assembly for a gas turbine engine is provided. The airfoil assembly comprises an airfoil having a concave pressure side opposite a convex suction side and an inner end radially spaced apart from an outer end. The pressure side and the suction side extend axially from a leading edge to a trailing edge. The airfoil assembly further comprises an inner band defining an inner opening shaped complementary to the inner end of the airfoil and an outer band defining an outer opening shaped complementary to the outer end of the airfoil. The inner end of the airfoil is received with the inner opening and the outer end of the airfoil is received within the outer opening. The inner band includes a first flowpath surface, a first non-flowpath surface opposite the first flowpath surface, and a first inner flange extending radially from the first non-flowpath surface. The outer band includes a second flowpath surface, a second non-flowpath surface opposite the second flowpath surface, and a first outer flange extending radially from the second non-flowpath surface. The inner band is secured to an inner support structure by a first inner fastener extending through the first inner flange, and the outer band is secured to an outer support structure by a first outer fastener extending through the first outer flange. Moreover, a strut extends radially through a cavity defined by the airfoil, and a first pad is defined at a first radial location within the cavity and a second pad is defined at a second radial location within the cavity. The first radial location is different from the second radial location. Each of the inner band, outer band, and airfoil are formed from a ceramic matrix composite material. 
     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 perspective view of a doublet airfoil assembly, having two airfoils separate from an inner band and outer band, according to an exemplary embodiment of the present subject matter. 
         FIG. 3  provides an axial cross-section view of one airfoil of the exemplary airfoil assembly of  FIG. 2 . 
         FIG. 4  provides a side schematic view of an airfoil assembly having a single pinned flange on each of the inner band and outer band, according to an exemplary embodiment of the present subject matter. 
         FIG. 5  provides a side schematic view of an airfoil assembly having a double pinned flange on each of the inner band and outer band, 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 “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust. 
     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. 
     The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein. 
     The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. 
     Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 10 percent margin. 
     Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. 
     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 an axial 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 . 
     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 axial centerline  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 . 
     In some embodiments, components of turbofan engine  10 , particularly components within or defining the hot gas path  78 , may comprise a composite material, such as a ceramic matrix composite (CMC) material having high temperature capability. Composite materials generally comprise a fibrous reinforcement material embedded in matrix material, e.g., a ceramic matrix material. The reinforcement material serves as a load-bearing constituent of the composite material, while the matrix of a composite material serves to bind the fibers together and act as the medium by which an externally applied stress is transmitted and distributed to the fibers. 
     Exemplary CMC materials may include silicon carbide (SiC), silicon, silica, or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron&#39;s SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon&#39;s NICALON®, Ube Industries&#39; TYRANNO®, and Dow Corning&#39;s SYLRAMIC®), alumina silicates (e.g., Nextel&#39;s 440 and 480), and chopped whiskers and fibers (e.g., Nextel&#39;s 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, bundles of the fibers, which may include a ceramic refractory material coating, are formed as a reinforced tape, such as a unidirectional reinforced tape. A plurality of the tapes may be laid up together (e.g., as plies) to form a preform component. The bundles of fibers may be impregnated with a slurry composition prior to forming the preform or after formation of the preform. The preform may then undergo thermal processing, such as a cure or burn-out to yield a high char residue in the preform, and subsequent chemical processing, such as melt-infiltration 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. 
     Turning to  FIG. 2 , an exemplary airfoil assembly  100 , e.g., a nozzle fairing assembly for turbofan engine  10 , is illustrated. The airfoil assembly  100  comprises two airfoils  102 , an inner band  104 , and an outer band  106 . Because the airfoil assembly  100  includes two airfoils  102  received within a single inner band  104  and a single outer band  106 , the assembly may be referred to as a doublet airfoil assembly  100 . In the depicted exemplary embodiment, each airfoil  102 , the inner band  104 , and the outer band  106  is formed from a CMC material. As shown in  FIG. 2 , the exemplary CMC airfoils  102  each include a concave pressure side  108  opposite a convex suction side  110 . Opposite pressure and suction sides  108 ,  110  of the airfoil  102  radially extend between an inner end  112  and an outer end  114  along an airfoil span S. Moreover, pressure and suction sides  108 ,  110  of the airfoil  102  extend axially between a leading edge  116  and an opposite trailing edge  118 , and the pressure and suction sides  108 ,  110  define an outer surface  120  of the airfoil  102 . Further, each illustrated airfoil  102  includes an inner parapet  122  that extends about the airfoil  102  at its inner end  112 , and an outer parapet  124  that extends about the airfoil  102  at its outer end  114 . Additionally, referring to  FIG. 3 , each airfoil  102  includes a trailing edge portion  126  that defines its trailing edge  118 . The trailing edge portion  126  is located aft of a cavity  128  defined by the airfoil  102 . The cavity  128  extends the radial length, i.e., the span S, of the airfoil  102 . 
     As further shown in  FIG. 2 , the inner and outer bands  104 ,  106  are relatively thin CMC structures that are separate from the airfoil  102 . That is, each of the airfoil  102 , inner band  104 , and outer band  106  are separately formed from a CMC material such that each component is an individual piece. In the depicted embodiment, the airfoil assembly  100  is a turbine nozzle fairing assembly, and a plurality of airfoil assemblies  100  may be positioned circumferentially adjacent one another to form an annular turbine nozzle stage, e.g., a plurality of turbine nozzles positioned circumferentially about the axial centerline  12  of the engine  10 . As such, each of the inner band  104  and outer band  106  form a liner along the hot gas path  78 , protecting metallic components and the like from the heat of the combustion gases  66 . 
     As illustrated in  FIG. 2 , the inner band  104  defines two inner openings  130  that are shaped complementary to the inner end  112  of each airfoil  102 . As such, the inner end  112  of each airfoil  102  is received within a respective inner opening  130 . Similarly, the outer band  106  defines two outer openings  132  shaped complementary to the outer end  114  of each airfoil  102 , such that the outer end  114  of each airfoil  102  is received within a respective outer opening  132 . An inner seal  134  extends around the inner end  112  of each airfoil  102  such that the inner seal  134  is positioned between the inner end  112  and the inner band  104  to seal against leakage through the inner opening  130 . Likewise, an outer seal  136  extends around the outer end  114  of each airfoil  102  such that the outer seal  136  is positioned between the outer end  114  and the outer band  106  to seal against leakage through the outer opening  132 . Moreover, the inner and outer seals  134 ,  136  are positioned to engage the inner band  104  and outer band  106 , respectively. The inner and outer seals  134 ,  136  are illustrated for only one airfoil  102  in  FIG. 2 ; the seals  134 ,  136  for the other airfoil  102  in  FIG. 2  are omitted for clarity. 
     Turning to  FIGS. 3, 4, and 5 , the airfoil assembly  100  will be described in greater detail.  FIG. 3  provides an axial cross-section view of one airfoil  102  of the exemplary airfoil assembly  100  of  FIG. 2 .  FIG. 4  provides a side schematic view of the airfoil assembly  100  having a single pinned flange on each of the inner band  104  and outer band  106 , according to one exemplary embodiment of the present subject matter.  FIG. 5  provides a side schematic view of the airfoil assembly  100  having a double pinned flange on each of the inner band  104  and outer band  106 , according to another exemplary embodiment of the present subject matter. 
     As illustrated in  FIGS. 3 and 5 , a strut  140  extends radially through each airfoil  102 , the inner band  104 , and the outer band  106  of the airfoil assembly  100 . More particularly, the strut  140  extends through each airfoil  102  within the cavity  128  defined in the airfoil  102 . The strut  140  includes a first pad  142  at a first radial location R 1  within the cavity  128  and a second pad  144  at a second radial location R 2  within the cavity  128 . As depicted in  FIG. 5 , the first radial location R 1  is different from the second radial location R 2 , and the first and second radial locations R 1 , R 2  are determined from the axial centerline  12  of the engine  10 . In other embodiments, the first and second pads  142 ,  144  may be defined on the airfoil  102  rather than the strut  140 . Although not illustrated in  FIG. 4 , it will be appreciated that a strut  140  could extend through the airfoil  102  as shown in  FIGS. 3 and 5 . 
     Keeping with  FIGS. 3 and 5 , the airfoil  102  is also constrained axially such that axial loading of the airfoil  102  transfers the load to an inner support structure  174  and an outer support structure  178 . As shown, a radially extending first slot  146  is defined in the trailing edge portion  126  of each airfoil  102 . Each first slot  146  is configured for receipt of a first pin  148 . Further, a radially extending second slot  150  is defined in each airfoil  102  for receipt of a second pin  152 . The depicted airfoil assembly  100  could be configured as a first stage or a second stage turbine nozzle assembly, i.e., when installed within a gas turbine engine. If configured as a second stage assembly  100 , the first slot  146  is defined in the outer end  114  of each airfoil  102 , and the second slot  150  is defined in the inner end  112  of each airfoil  102 . A first aperture  154  is defined in the outer support structure  178  adjacent each first slot  146  such that each first pin  148  passes through the first aperture  154  into the first slot  146 , and a second aperture  156  is defined in the inner support structure  174  adjacent each second slot  150  such that each second pin  152  passes through the second aperture  156  into the second slot  150 . Thus, each airfoil  102  is constrained axially by a pair of pins  148 ,  152  extending radially into the airfoil  102 , the first pin  148  at the outer end  114  of the trailing edge portion  126 , and the second pin  152  at the inner end  112  of the airfoil  102  just aft of the cavity  128 . Further, the pins  148 ,  152  prevent the airfoil  102  from bottoming out within the openings  130 ,  132 , as well as from pulling out of the openings  130 ,  132 . 
     For a second stage nozzle assembly, the depicted pin configuration supports and locates the respective airfoil  102 , with the axial loads passing into the inner and outer support structures  174 ,  178 . However, it will be appreciated that where the airfoil assembly  100  is a first stage nozzle assembly (e.g., as illustrated in  FIG. 4 ), the first slots  146  are typically defined in the inner end  112  of the airfoil  102 , and the second slots  150  are typically defined in the outer end  114  of the airfoil  102 . Likewise, in such embodiments, the first apertures  154  are defined in the inner support structure  174  such that the first pins  148  pass through the first aperture  154  into the first slot  146  at the inner end  112  of the trailing edge portion  126  of each airfoil  102 , and the second apertures  156  are defined in the outer support structure  178  such that the second pins  152  pass through the outer support structure  178  into the second slot  150  just aft of the cavity  128  of each airfoil  102 . Such a pin configuration adequately axially constrains the first stage nozzle assembly  100 , supporting and locating each airfoil  102  of the assembly  100  such that the axial loads pass into the inner and outer support structures  174 ,  178 . As such, the second stage nozzle is cantilevered from the outer casing  18 , with an inter-stage seal attached to the inner support structure  174 . The cantilevered configuration may be applied to nozzles of any turbine stage that requires the nozzles to be cantilevered, e.g., third, fourth, etc. stage nozzles. In contrast, the first stage nozzle is non-cantilevered, with the inner and outer support structures  174 ,  178  being supported by the engine&#39;s static structure, e.g., outer casing  18 . 
     Further, in either the first or second stage embodiment of the airfoil assembly  100 , the first and second pins  148 ,  152  radially constrain each airfoil  102 . Moreover, referring to  FIG. 3 , each airfoil  102  is configured to contact the first and second pads  142 ,  144  of the strut  140  extending through the airfoil  102  and load the slot  146  into the pin  148  when the airfoil  102  is tangentially loaded. More specifically, when each airfoil  102  experiences a tangential load, the airfoil  102  is configured to load into the pads  142 ,  144  such that an interior surface  138  of the airfoil  102 , which defines the cavity  128  of the airfoil  102 , contacts the first and second pads  142 ,  144 . More generally, e.g., for embodiments in which the pads  142 ,  144  are not part of or attached to the strut  140 , the first and second pads  142 ,  144  provide a loading surface or stop for the airfoil  102  to load into the strut  140  when the airfoil  102  is tangentially loaded. As such, the first and second pads  142 ,  144  and the first pin  148  transfer the tangential loads to the strut  140  and inner support structure  174 . The first and second pins  148 ,  152  support and locate each airfoil  102  while transferring radial and axial loads to the inner and outer support structures  174 ,  178 . 
     Further, in exemplary embodiments of the airfoil assembly  100 , the strut  140  is formed from a metallic material, such as a metal or metal alloy. Accordingly, the CMC airfoil  102  and the metallic strut  140  have different coefficients of thermal expansion α. As previously described, each airfoil  102  is separate from the inner band  104  and outer band  106 , and is not attached or fastened to the strut  140 , and as such, the airfoil  102  is free to float radially to accommodate the difference in coefficients of thermal expansion α between the airfoil  102  and the strut  140 . That is, the strut  140  will begin to expand at a lower temperature than each airfoil  102 , and the airfoil  102  has freedom to move radially to accommodate the thermal expansion of the strut  140 . 
     Turning now to  FIGS. 4 and 5 , the inner band  104  includes a first flowpath surface  160  and a first non-flowpath surface  162  opposite the first flowpath surface  160 . Similarly, the outer band  106  includes a second flowpath surface  164  and a second non-flowpath surface  166  opposite the second flowpath surface  164 . The first and second flowpath surfaces  160 ,  164  help define the hot gas path  78 , while the first and second non-flowpath surfaces  162 ,  166  are positioned outside of the hot gas path  78 . 
     Referring particularly to  FIG. 4 , in one exemplary embodiment, a first inner flange  168  extends radially from the first non-flowpath surface  162 , and a first outer flange  170  extends radially from the second non-flowpath surface  166 . A first inner fastener  172  extends through an aperture  173  in the first inner flange  168  and into the inner support structure  174  to secure the inner band  104  to the inner support structure  174 . Likewise, a first outer fastener  176  extends through an aperture  177  in the first outer flange  170  and into the outer support structure  178  to secure the outer band  106  to the outer support structure  178 . Thus, the exemplary embodiment of  FIG. 4  comprises a single pinned flange on each of the inner band  104  and outer band  106  to attach the bands  104 ,  106  to their respective support structures  174 ,  178 , e.g., metallic hangers or the like. It will be appreciated that more than one inner fastener  172  and outer fastener  176 , e.g., each inner fastener  172  and each outer fastener  176  circumferentially spaced apart from one another, may pass through the inner band  104  and outer band  106 , respectively, to secure the inner and outer bands  104 ,  106  to their respective support structures  174 ,  178 . In such embodiments, the aperture  173  for an additional inner fastener  172  and the aperture  177  for an additional outer fastener  176  extend in the circumferential direction C to account for the thermal expansion difference between the support structures  174 ,  178  (e.g., metal hangers) and the CMC bands  104 ,  106 . 
     As shown in  FIG. 4 , the first inner flange  168  extends from the first non-flowpath surface  162  near a mid-portion  180  of the inner band  104 , and the first outer flange  170  extends from the second non-flowpath surface  166  near an aft end  182  of the outer band  106 . Preferably, each of the first inner flange  168  and the first outer flange  170  are defined on the respective band  104 ,  106  near a resultant pressure force to minimize moment into the pin joint between the first inner flange  168  and the inner support structure  174  and the first outer flange  170  and the outer support structure  178 . That is, the various forces formed by a pressure distribution across the bands  104 ,  106  produce a moment at each pin joint. The position of each flange  168 ,  178  on its respective band  104 ,  106  is selected to minimize the moment created by the pressure forces that is taken through each pin  148 ,  152 . 
     Referring now to  FIG. 5 , in another exemplary embodiment of the airfoil assembly  100 , each of the inner band  104  and outer band  106  includes a double pinned flange attaching the band  104 ,  106  to its respective support structure  174 ,  178 . More particularly, in addition to the first inner flange  168 , the inner band  104  includes a second inner flange  184  extending radially from the first non-flowpath surface  162 , and in addition to the first outer flange  170 , the outer band  106  includes a second outer flange  186  extending radially from the second non-flowpath surface  166 . A second inner fastener  188  extends through an aperture  189  ( FIG. 2 ) in the second inner flange  184  and into the inner support structure  174 , and a second outer fastener  190  extends through an aperture  191  ( FIG. 2 ) in the second outer flange  186  and into the outer support structure  178 . As such, the first and second inner fasteners  172 ,  188  secure the inner band  104  to the inner support structure  174 , and the first and second outer fasteners  176 ,  190  secure the outer band  106  to the outer support structure  178 . As described above, more than one inner fastener  188  and outer fastener  190 , e.g., each inner fastener  188  and each outer fastener  190  circumferentially spaced apart from one another, may pass through the inner band  104  and outer band  106 , respectively, to secure the inner and outer bands  104 ,  106  to their respective support structures  174 ,  178 . In such embodiments, the aperture  189  for an additional inner fastener  188  and the aperture  191  for an additional outer fastener  190  extend in the circumferential direction C to account for the thermal expansion difference between the support structures  174 ,  178  (e.g., metal hangers) and the CMC bands  104 ,  106 . 
     As further depicted in  FIG. 5 , in the double pinned embodiment, the first inner flange  168  extends from the first non-flowpath surface  162  of the inner band  104  near the mid-portion  180  of the inner band  104 . The second inner flange  184  extends from the first non-flowpath surface  162  near a forward end  192  of the inner band  104 . Moreover, the first outer flange  170  extends from the second non-flowpath surface  166  of the outer band  106  near the aft end  182  of the outer band  106 , while the second outer flange  186  extends from the second non-flowpath surface  166  near a forward end  194  of the outer band  106 . As a result, the first inner flange  168  and the second inner flange  184  extend from the first non-flowpath surface  162  of the inner band  104  such that the first and second inner flanges  168 ,  184  straddle the inner support structure  174 . Similarly, the first outer flange  170  and the second outer flange  186  extend from the second non-flowpath surface  166  of the outer band  106  such that the first and second outer flanges straddle the outer support structure  178 . As such, the first and second inner flanges  168 ,  184  wrap around the inner support structure  174  and the first and second outer flanges  170 ,  186  wrap around the outer support structure  178 , thereby forming a CMC liner around each of the first and outer support structures  174 ,  178  out of the direct flow path  78  of the combustion gases  66 . As previously described, the support structures  174 ,  178  may be hangers formed from a metallic material, e.g., a metal or metal alloy, such that the CMC bands  104 ,  106  with their flanges  168 ,  184 ,  170 ,  186  help protect the support structures  174 ,  178  from the high temperature combustion gases  66 . 
     Additionally, although the airfoil assembly  100  depicted in the exemplary embodiments includes two airfoils, the airfoil assembly  100  described herein could be formed as a singlet, triplet, etc. More particularly, for a singlet assembly  100 , the inner band  104  defines one inner opening  130  and the outer band  106  defines one outer opening  132 . A single airfoil  102  extends from the inner band  104  to the outer band  106  with the inner end  112  of the airfoil  102  positioned in the inner opening  130  and the outer end  114  of the airfoil  102  positioned in the corresponding outer opening  132 . A triplet airfoil assembly  100  would include three airfoils  102 , with each airfoil  102  extending between an inner opening  130  and an outer opening  132  such that the triplet assembly  100  includes three inner openings  130  and three outer openings  132 . An airfoil assembly  100  including any appropriate number of airfoils  102  extending from the inner band  104  and outer band  106  may be formed. It will be appreciated that, by reducing the number of inner and outer band segments, split line leakage, i.e., leakage between adjacent inner and outer band segments, can be reduced. More specifically, split line leakage is eliminated where splits in the inner band  104  and outer band  106  are eliminated to form a doublet, triplet, etc. airfoil assembly  100 . However, as shown in  FIGS. 2, 4, and 5 , seals  196  also may be included between adjacent inner and outer bands  104 ,  106  to help minimize leakage between band segments. 
     As described herein, the airfoils  102 , inner band  104 , and outer band  106  may be formed from a CMC material. After laying up the CMC plies as previously described to form a layup or preforms for each of the airfoils  102 , inner band  104 , and outer band  106 , the layups are debulked and, if appropriate, cured while subjected to elevated pressures and temperatures to produce cured preforms, e.g., the layups may be cured in an autoclave to form autoclaved bodies. In exemplary embodiments, the autoclaved bodies are 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 for each preform is a porous CMC fired body that may undergo densification, e.g., melt infiltration (MI), to fill the porosity and yield the respective 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 a tool 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. 
     Optionally, after processing, the CMC component (i.e., CMC airfoil  102 , CMC inner band  104 , and CMC outer band  106 ) may be finish machined, if and as needed, and coated with one or more coatings, such as an environmental barrier coating (EBC). Moreover, the method described above is provided by way of example only. As an example, other known methods or techniques for curing composite plies, as well as for densifying a CMC component, may be utilized. Alternatively, any combinations of these or other known processes may be used. Further, although in the exemplary embodiments described herein the airfoil assembly  100  as a turbine nozzle fairing assembly, it will be appreciated that the airfoil assembly  100  could be any nozzle fairing assembly. Additionally, although described herein with respect to CMC airfoils  102 , CMC inner bands  104 , and CMC outer bands  106 , the present subject matter could be applied to an airfoil assembly  100  formed from any suitable material, such as a polymer matrix composite (PMC) material or other composite material. Moreover, the airfoil assembly  100  may comprise any appropriate combination of materials, e.g., in some embodiments, at least one of the airfoil  102 , inner band  104 , or outer band  106  may be formed from a different material than the remainder of the components of the airfoil assembly. 
     Accordingly, as described herein, an airfoil assembly having airfoils separate from an inner band and an outer band may be constructed. By forming each component of the airfoil assembly as a separate piece, complicated CMC ply layups may be avoided, which may reduce the complexity of the manufacturing process for the airfoil assembly while increasing part yield and maintaining suitable compaction of each component part of the airfoil assembly. Further, separating each airfoil of the assembly from the inner and outer bands eliminates stresses due to the thermal fight between the airfoil and bands. By eliminating such stresses, the airfoil assemblies described herein are more defect tolerant than known integral airfoil and band designs, which opens up the non-destructive examination limits of the assemblies. Separation of the airfoil and bands also enables doublet, triplet, etc. airfoil assemblies, which eliminate split line leakage. Therefore, compared to known airfoil assembly designs, the airfoil assemblies described herein have improved producibility and reduced stresses, which reduces defects and increases acceptability of the airfoil assemblies. Other advantages of the subject matter described herein also may be realized by those of ordinary skill in the art. 
     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.