Patent Publication Number: US-10767493-B2

Title: Turbine vane assembly with ceramic matrix composite vanes

Description:
FIELD OF THE DISCLOSURE 
     The present disclosure relates generally to vane assemblies for gas turbine engines, and more specifically to vanes that comprise ceramic-containing materials. 
     BACKGROUND 
     Gas turbine engines are used to power aircraft, watercraft, power generators, and the like. Gas turbine engines typically include a compressor, a combustor, and a turbine. The compressor compresses air drawn into the engine and delivers high pressure air to the combustor. In the combustor, fuel is mixed with the high pressure air and is ignited. Products of the combustion reaction in the combustor are directed into the turbine where work is extracted to drive the compressor and, sometimes, an output shaft. Left-over products of the combustion are exhausted out of the turbine and may provide thrust in some applications. 
     Products of the combustion reaction directed into the turbine flow over airfoils included in stationary vanes and rotating blades of the turbine. The interaction of combustion products with the airfoils heats the airfoils to temperatures that require the airfoils to be made from high-temperature resistant materials and/or to be actively cooled by supplying relatively cool air to the vanes and blades. To this end, some airfoils for vanes and blades are incorporating composite materials adapted to withstand very high temperatures. Design and manufacture of vanes and blades from composite materials presents challenges because of the geometry and strength required for the parts. 
     SUMMARY 
     The present disclosure may comprise one or more of the following features and combinations thereof. 
     A turbine section adapted for use in a gas turbine engine may include a case made from metallic materials, a turbine wheel, and a turbine vane assembly. The case may be shaped to extend around a central reference axis. The turbine wheel may be housed in the case. The turbine vane assembly may be fixed to the case and may be configured to smooth and redirect air moving along a primary gas path of the turbine section ahead of interaction with the turbine wheel. 
     In some embodiments, the turbine wheel may include a disk, a plurality of blades, and a rotatable seal. The disk may be mounted for rotation about the central reference axis relative to the case. The plurality of blades may be coupled to the disk for rotation with the disk. The rotatable seal element may be coupled to the disk for rotation with the disk. 
     In some embodiments, the turbine vane assembly may include a plurality of composite aero vanes made from ceramic matrix composite material, a plurality of structural vanes made from metallic materials, and a static seal element. The plurality of composite vanes may be shaped to provide inner and outer end walls defining the primary gas path as well as airfoils that extend across the primary gas path. The plurality of structural vanes may be shaped to provide airfoils that extend across the primary gas path. The static seal element may cooperate with the rotatable seal element of the turbine wheel to provide a seal for resisting movement of gasses across the seal when the turbine section is in use within a gas turbine engine. 
     In some embodiments, the static seal element may be fixed to the plurality of structural vanes so as to be in turn coupled to the case while remaining free for relative movement in relation to the composite aero vanes. Accordingly, the composite aero vanes may be substantially free from carrying mechanical loads applied by pressure on the static seal element to the case. In some embodiments, the seal provided by the rotatable seal element and the static seal element may be arranged radially inward of the primary gas path. 
     In some embodiments, the static seal element may include a seal panel and at least one seal land. The seal panel may divide axially adjacent compartments within the turbine section. The at least one seal land may be engaged by the rotatable seal element. 
     In some embodiments, the rotatable seal element may include a knife ring. The knife ring may engage the land of the static seal element. 
     In some embodiments, the plurality of structural vanes may each include an inner end wall, an outer end wall, and an airfoil. The inner end wall may face the primary gas path. The outer end wall may be spaced radially from the inner end wall and may face the primary gas path. The airfoil may extend from the inner end wall to the outer end wall across the primary gas path. 
     In some embodiments, the plurality of structural vanes may each include a seal mount. The seal mount may extend radially inwardly from the inner end wall away from the primary gas path. The static seal element may be fixed to the seal mount. 
     In some embodiments, the plurality of structural vanes may each include a case mount. The case mount may extend radially outwardly from the outer end wall away from the primary gas path. The case mount may engage the case to couple the structural vane and the static seal element to the case. 
     In some embodiments, the plurality of composite aero vanes may be each coupled to the case by a metallic spar. The spar may extend through the airfoils of the composite aero vanes. 
     In some embodiments, the seal element may be fixed to the seal mount by at least one fastener. In some embodiments, each of the plurality of composite aero vanes may be spaced apart from the static seal element. 
     According to an aspect of the present disclosure, a turbine vane assembly adapted for use in a gas turbine engine may include a plurality of composite aero vanes, a plurality of structural vanes, and a static seal element. The plurality of composite aero vanes may be made from ceramic matrix composite material. The plurality of structural vanes may be made from metallic materials. The static seal element may be fixed to the plurality of structural vanes and may be spaced apart from the composite aero vanes. 
     In some embodiments, each of the plurality of composite aero vanes may include an inner end wall, an outer end wall, and an airfoil. The inner end wall may extend partway around a central reference axis. The outer end wall may be spaced radially from the inner end wall to define a primary gas path therebetween. The airfoil may extend from the inner end wall to the outer end wall across the primary gas path. 
     In some embodiments, each of the plurality of structural vanes may include airfoils. The airfoils may extend across the primary gas path. In some embodiments, the static seal element may be arranged radially inward of the primary gas path. 
     In some embodiments, the plurality of structural vanes may each include a seal mount. The seal mount may extend radially inwardly from the inner end wall away from the primary gas path. The static seal element may be fixed to the seal mount. 
     In some embodiments, the plurality of structural vanes may each include a case mount. The case mount may extend radially outwardly from the outer end wall away from the primary gas path. The case mount may configured to engage a case so that the static seal element may be fixed to the case via the plurality of structural vanes. In some embodiments, the static seal element may be fixed to the seal mount by at least one fastener. 
     In some embodiments, the plurality of structural vanes may each include an inner end wall, an outer end wall, and an airfoil. The inner end wall may face the primary gas path. The outer end wall may spaced radially from the inner end wall and may face the primary gas path. The airfoil may extend from the inner end wall to the outer end wall across the primary gas path. 
     In some embodiments, the airfoil of each of the plurality of structural vanes may formed to include a cooling air passageway and a plurality of film cooling holes. The plurality of cooling holes may interconnect the cooling air passageway with the primary gas path. 
     In some embodiments, the plurality of composite aero vanes may be each mounted to a metallic spar. The spar may extend radially through the associated airfoil. 
     These and other features of the present disclosure will become more apparent from the following description of the illustrative embodiments. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cutaway of a gas turbine engine that includes a fan, a compressor, a combustor, and a turbine that includes a plurality of turbine wheel assemblies and turbine vane assemblies, the turbine vane assemblies being shown in further detail in  FIGS. 2-6 ; 
         FIG. 2  is a front elevation view of one turbine vane assembly included in the gas turbine engine of  FIG. 1  showing that the turbine vane assembly includes a plurality of composite aero vanes, a plurality of structural vanes made from metallic materials, and a static seal element mounted radially inward of the ceramic and structural vanes; 
         FIG. 3  is a front elevation view of a portion of the turbine vane assembly shown in  FIG. 2  illustrating that the structural vanes comprising metallic materials are spaced between and separate from the composite aero vanes and have cooling holes for discharging cooling air across the metallic materials of the structural vanes; 
         FIG. 4  is a cross-sectional view of a portion of the turbine section of the gas turbine engine of  FIG. 1  showing that the turbine vane assembly is mounted between a first stage turbine wheel and a second stage turbine wheel to redirect gas moving from the first stage turbine wheel toward the second stage turbine wheel and showing that the static seal element is fixed to the plurality of structural vanes for cooperation with a rotatable seal element of the second stage turbine wheel; 
         FIG. 5  is a detail view of the turbine section of  FIG. 4  showing that structural vanes are coupled directly with the static seal element to transfer axial loading through the structural vanes; 
         FIG. 6  is detail view of the turbine section similar to  FIG. 5  showing that the composite aero vanes are floating relative to the static seal element and that the composite aero vanes are only required to transfer aerodynamic loads acting on the composite aero vanes through a spar to a corresponding turbine case; 
         FIG. 7  is a cross-sectional view of the seal mount vane of  FIG. 5  showing that the structural vanes comprise metallic materials; and 
         FIG. 8  is a cross-sectional view of the self-supporting vane of  FIG. 6  showing that the composite aero vanes comprise ceramic matrix composite material. 
     
    
    
     DETAILED DESCRIPTION OF THE DRAWINGS 
     For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of illustrative embodiments illustrated in the drawings and specific language will be used to describe the same. 
     A turbine section  18  according to the present disclosure is adapted for use in a gas turbine engine  10  as suggested in  FIGS. 1 and 2 . The gas turbine engine  10  includes a fan  12 , a compressor  14 , a combustor  16 , and a turbine  18 . The fan  12  generates thrust for propelling an aircraft. The compressor  14  compresses and delivers air to the combustor  16 . The combustor  16  mixes fuel with the compressed air received from the compressor  14  and ignites the fuel. The hot, high-pressure gases from the burning fuel are directed into the turbine  18  where the turbine  18  extracts work from the gases to drive the compressor  14  and the fan  12 . In other embodiments, the gas turbine engine  10  may include a shaft, turboprop, or gearbox in place of the fan  12 . 
     The turbine section  18  includes a turbine case  20 , turbine wheels  22 ,  24 , and a turbine vane assembly  26  as shown in  FIGS. 1-3 . The turbine case  20  is made from metallic materials and shaped to extend around a central reference axis  11 . The turbine case  20  surrounds the other components of the turbine section  18 . The turbine wheels  22 ,  24  are mounted for rotating about the axis  11  in the case  20 . The turbine vane assembly  26  is fixed to the case  20  and is configured to smooth and redirect air moving along a primary gas path  25  of the turbine section  18  ahead of interaction with the turbine wheel  24  and downstream of turbine wheel  22 . 
     The turbine wheel  24  includes a disk  28 , a plurality of blades  30 , and a rotating seal element  32  as shown in  FIG. 4 . The disk  28  is mounted for rotation about the central reference axis  11  relative to the case  20 . The plurality of blades  30  are coupled to the disk for rotation with the disk  28 . The rotatable seal element  32  is coupled to the disk  28  for rotation with the disk  28 . 
     The turbine vane assembly  26  includes a plurality of composite aero vanes  34 , a plurality of structural vanes  36 , and a static seal element  38  as shown in  FIGS. 2-6 . Each of the composite aero vanes  34  comprises ceramic matrix material adapted to withstand high temperatures. However, the composite aero vanes  34  may have relatively low strength compared to the structural vanes  36 , which comprise metallic materials. The structural vanes  36  provide structural strength to the turbine vane assembly  26  by receiving the mechanical loads, such as the axial moment applied by pressure on the static seal element  38 . The static seal element  38  is fixed to the plurality of structural vanes  36  so as to be in turn coupled to the case  20  while remaining free for relative movement in relation to the composite aero vanes  34 . Accordingly, the composite aero vanes  34  are substantially free from carrying mechanical loads applied by pressure on the static seal element  38  to the case  20 . 
     The plurality of composite aero vanes  34  are made from integrally-formed ceramic matrix composite material as noted above. Each composite aero vane  34  is shaped to provide inner and outer end walls  48 ,  46  defining the primary gas path  25  and at least one airfoil  50  that extends across the primary gas path  25 . 
     The plurality of structural vanes  36  are made from metallic materials as noted above. Each structural vane  36  is shaped to provide inner and outer endwalls  74 ,  72  and at least one airfoil  76  that extend across the primary gas path  25 . 
     The static seal element  38  cooperates with the rotatable seal element  32  of the turbine wheel  24  to provide a compartment seal  35  for resisting movement of gasses across the compartment seal  35  when the turbine section  18  is in use within a gas turbine engine  10 . The compartment seal  35  formed by the static seal element  38  and the rotating seal element  32  seals between axially adjacent compartments  43 ,  44  resulting in a first pressure in the compartment  43  on the first stage turbine wheel  22  side and a second pressure in the compartment  44  on the second stage turbine wheel  24  side. The first pressure is greater than the second pressure resulting in a difference of pressure on either side  43 ,  44  of the static seal element  38 . The difference of pressure causes a pressure force to act on a seal panel  39  of the static seal component  38 . The pressure force results in an axial moment in the turbine vane assembly  26 . 
     In the illustrative embodiment, the static seal element  38  is a single integrally formed ring. In other embodiments, the static seal element  38  may include a plurality of static seal element segments to form the ring shape. The segmented static seal element  38  may also be provided with any appropriate seal apparatus to seal between each of the segments. 
     The static seal element  38  includes a seal panel  39 , seal lands  40 ,  41 , and a fastener  42  as shown in  FIGS. 4 and 5 . The seal panel  39  divides axially adjacent compartments  43 ,  44  within the turbine section  18 . The seal lands  40 ,  41  extend axially aft and away from the seal panel  39 . At least one seal land  40 ,  41  is engaged by the rotatable seal element  32  of the turbine wheel  24 . In the illustrative embodiment, both seal lands  40 ,  41  are engaged with the rotating seal element  32  to seal between the upstream turbine wheel  22  and the downstream turbine wheel  24 . The fastener  42  extends away from the seal panel  39  and engages a seal mount  78  of the structural vane  36 . The static seal element  38  is fixed to the seal mount  78  by at least one fastener  42 . In the illustrative embodiment, the fastener  42  is a bolt/nut combination. In other embodiments, the fastener  42  may be another suitable fastener  42  such as a pin, rivet, or integrated manufacturing retention (casting, welding, etc.). In other embodiments, the fastener  42  may be some other suitable mechanical joint such as a cross-key, a transition fit, or radial birdmouth with engaged rail. 
     The fastener  42  of the static seal element  38  is not fixed to the composite aero vane  34  in the illustrative embodiment. Each of the plurality of composite aero vanes  34  is spaced apart from the static seal element  38 . In the illustrative embodiment, the static seal element  38  is spaced apart from the inner end wall  48  of the composite aero vane  34  to leave a space  45  between the static seal element  38  and the inner end wall  48  of the composite aero vane  34 . In other embodiments, a seal may be arranged in the space  45  to seal between the static seal element  38  and the composite aero vane  34  and still allow relative movement of the static seal element  38  relative to the composite aero vane  34 . 
     Turning again to the plurality of composite aero vanes  34 , each of the plurality of composite aero vanes  34  includes an outer end wall  46 , an inner end wall  48 , an airfoil  50 , and a vane mount unit  52  as shown in  FIG. 6 . The inner end wall  48  is spaced radially inward of the outer end wall  46 . The airfoil  50  extends between and interconnects the outer end wall  46  and the inner end wall  48 . The airfoil  50  is shaped to redirect air moving along the primary gas path  25  of the turbine section  18  that extends radially from the outer end wall  46  to the inner end wall  48 . The airfoil  50  is also shaped to include a vane cavity  51  extending radially through the airfoil  50  and opens at the inner and outer end walls  46 ,  48 . The outer end wall  46  defines a radially outer boundary of the primary gas path  25  and the inner end wall  48  defines a radially inner boundary of the primary gas path  25 . The vane mount unit  52  mounts the composite aero vanes  34  to the turbine case  20  without engaging the static seal component  38 . 
     The airfoil  50  includes a radial outer end  54 , a radial inner end  55 , and a body  56  as shown in  FIG. 6 . The radial outer end  54  extends radially-outwardly past the outer end wall  46  outside the primary gas path  25  in the illustrative embodiment. The radial inner end  55  is spaced apart from the radial outer end  54  relative to the axis  11  and extends radially-inwardly past the inner end wall  48  outside the primary gas path  25 . The radial inner end  55  of the airfoil  50  of the composite aero vane  34  engages the vane mount unit  52 . The body  56  extends radially entirely between and interconnects the radial outer end  54  and the radial inner end  55 . 
     The vane mount unit  52  of the composite aero vanes  34  includes a carrier  60 , a spar  62 , and a clamp nut  64  as shown in  FIG. 6 . The spar  62  is made from metallic materials and the metallic spar  62  extends through the vane cavity  51  of the airfoils  50  of the composite aero vanes  34 . In some embodiments, the spar  62  may be hollow and include cooling holes to transmit cooling air to the composite aero vane  34  and/or into the inter-disk cavity between the turbine wheels  22 ,  24 . The spar  62  is configured to receive aerodynamic loads from the airfoil  50  of the composite aero vane  34  during use of the turbine section  18  in the gas turbine engine  10 . The carrier  60  is made from metallic materials and is coupled to the spar  62 . The carrier  60  engages the turbine case  20  to carry aerodynamic loads from the spar  62  to the turbine case  20 . The clamp nut  64  is located radially inward of the inner end wall  48  of the composite aero vane  34  and mates with the spar  62  to clamp the composite aero vane  34  blocking radial movement of the composite aero vane  34  relative to the axis  11 . 
     In the illustrative embodiment, the clamp nut radially retains the composite aero vane  34  relative to the spar  62 . In other embodiments, other methods to radially retain the composite aero vane  34  may be used, such as a pin, other fastener, or integrated manufacturing retention (casting, welding, etc.). In some embodiments, the spar  62  may couple to the static seal element  38  directly. 
     The carrier  60  of the vane mount unit  52  includes a body panel  66 , a forward mount hanger  68 , and an aft mount rail  70  as shown in  FIG. 6 . The forward hanger  68  extends radially outward from the carrier body panel  66  at a forward end of the body panel  66  and is engaged with a forward bracket  90  of the turbine case  20 . The aft mount rail  70  extends radially outward form the carrier body panel  66  at an aft end of the body panel  66  and is engaged with an aft bracket  92  of the turbine case  20 . The spar  62  couples to the body panel  66  of the carrier  60  in between the forward and aft attachment features  68 ,  70 . 
     The spar  62  of the vane mount unit  52  is shaped to engage the airfoil  50  of the composite aero vane  34  at a location radially outward or radially inward of the primary gas path  25 . The spar  62  engages the airfoil  50  of the composite aero vane  34  to transfer aerodynamic loads of the airfoil  50  to the spar  62  so that the spar  62  may carry the aerodynamic loads to the turbine case  20 . 
     Each of the structural vanes  36  includes an outer end wall  72 , an inner end wall  74 , an airfoil  76 , a case mount  77 , and a seal mount  78  as shown in  FIG. 5 . The inner end wall  74  faces the primary gas path  25 . The outer end wall  72  is spaced radially from the inner end wall  74  and faces the primary gas path  25 . The airfoil  76  extends from the inner end wall  74  to the outer end wall  72  across the primary gas path  25 . The case mount  77  extends radially outwardly from the outer end wall  72  away from the primary gas path  25 . The case mount  77  engages the case  20  to couple each structural vane  36  of the plurality of structural vanes  36  and static seal element  38  to the case  20 . The seal mount  78  extends radially inwardly from the inner end wall  74  away from the primary gas path  25  and the static seal element  38  is fixed to the seal mount  78 . 
     The airfoil  76  of the structural vanes  36  are formed to include a cooling air passageway  79  and a plurality of film cooling holes  80  as shown in  FIGS. 2-5 and 7 . The cooling air passageway  79  extends radially through the airfoil  76  of the structural vane  36  to transmit cooling air into the inter-disk cavity between the turbine wheels  22 ,  24 . The plurality of film cooling holes  80  interconnect the cooling air passageway  79  with the primary gas path  25 . In some embodiments, the airfoil  76  may include a metallic impingement tube to direct cooling air and/or locally increase cooling effectiveness. 
     In the illustrative embodiment, the airfoil  76  further includes a trailing edge cooling feature  83  as shown in  FIG. 7 . The trailing edge cooling feature  83  cools a trailing edge of the airfoil  76  of the structural vanes  36 . 
     The case mount  77  of the structural vanes  36  includes a forward hanger  84  and an aft rail  86  as shown in  FIG. 5 . The forward hanger  84  extends radially outward from the outer end wall  72  of the structural vane  36  relative to the axis  11 . The aft rail  86  is axially spaced apart from the forward hanger  84  and extends radially outward from the outer end wall  72  of the structural vane  36  relative to the axis  11 . The forward hanger  84  engages the case  20  at a location forward of the structural vane  36  and the aft rail  86  engages the case  20  at a location aft of the structural vane  36 . The forward hanger  84  and the aft rail  86  engage the case  20  to couple the structural vane  36  to the case  20  and transfer axial loads from the static seal element  38 , axial and/or circumferential loads from the structural vane  36 , and aerodynamic loads from structural vane  36  to the turbine case  20 . 
     Turing again to the turbine case  20 , the turbine case  20  may include an annular shell  88 , a forward bracket  90 , and an aft bracket  92  as shown in  FIGS. 4-6 . The annular shell  88  extends around the axis  11 . The forward bracket  90  extends radially inward from the annular shell  88 . The aft bracket  92  extends radially inward from the annular shell  88  at a location axially spaced from the forward bracket  90 . The forward and aft brackets  90 ,  92  also extend circumferentially at least partway around the overall circumferential length of the annular shell  88 . In the illustrative embodiment, the turbine case  20  only has two brackets  90 ,  92 . In other embodiments, the turbine case  20  may include two or more brackets. 
     In the illustrative embodiment, the forward bracket  90  provides a forward attachment feature for the case mount  77  of the structural vanes  36  and the vane mount unit  52  of the composite aero vanes  34  with a hook shape, while the aft bracket  92  provides an aft attachment feature for the case mount  77  and the vane mount unit  52  with a rail shape. In other embodiments, the forward and aft brackets  90 ,  92  may both be hook shaped. In other embodiments, the forward and aft attachment features  90 ,  92  may have another suitable shape (dovetail interface, T-shape interface, or other suitable interface shape). Additionally, seals may also be arranged between the brackets  90 ,  92  and the case mount  77  and/or the vane mount unit  52  to seal between the components. 
     In the illustrative embodiment, the forward and aft attachment features  90 ,  92  are axisymmetric about the axis  11 . The forward and aft attachment features  90 ,  92  use the same general attachment and load transfer method for both the composite aero and structural vanes  34 ,  36 . 
     In the illustrative embodiment, the rotatable seal element  32  includes a knife ring  33  as shown in  FIG. 4 . The knife ring  33  engages the seal lands  40 ,  41  of the static seal element  38  to form the seal  35  between the static seal element  38  and the rotatable seal element  32 . 
     The present disclosure is related to a turbine section  18  of a gas turbine engine  10  including a small number of metallic nozzle guide vanes  36  to carry structural loads that ceramic matrix composite vanes  34  cannot tolerate. In the illustrative embodiment, the ceramic matrix composite vanes  34  do not carry additional loading. As the ceramic matrix composite material does not need to carry additional loading, the stresses and/or design flexibility of the composite vanes  34  will improve. 
     The turbine vane assembly  26  may be configured to support other gas turbine engine components, such as an inter-stage seal  38 . Accordingly, an application of a metallic support structure is likely to be required to transmit the axial loading applied to the components to the high-pressure turbine casing  20 . 
     The present disclosure incorporates a combination of ceramic matrix composite and metallic vanes  34 ,  36 , whereby the inter-stage seal axial load is largely transmitted through the metallic vanes  36 . An optional metallic spar  62  may be installed inside the ceramic matrix composite vanes  34  and may in some embodiments, be configured to accommodate a portion of the inter-stage seal loading. 
     The number of metallic vanes e.g.  36  may be minimized as they can require greater cooling flows compared to the ceramic matrix composite vanes e.g.  34 . However, the number of metallic vanes  36  may be be greater than 1 to introduce an element of redundancy to the inter-stage seal support structure. The metallic vanes  36  may be circumferentially equally spaced. 
     The same aerodynamic definition or airfoil shape could be applied to both metal and ceramic matrix composite structures. However, consideration of the mismatch in thermal expansion is required, particularly at the ceramic matrix composite-metal vane interface. Achievement of equivalent aerodynamic performance between the two aerofoil styles could be a consideration to avoid introducing additional vibration forcing frequencies. 
     The ceramic matrix composite and metallic vanes  34 ,  36  may have improved aerodynamic performance when compared to a uniformly size set of ceramic matrix composite airfoils. The uniform ceramic matrix composite airfoils  34  may have a relatively large maximum thickness to increase and provide a sufficient second moment of area. However, a metallic aerofoil e.g.  36  can withstand larger mechanical loads so the mixed material set will have improved aerodynamic freedom i.e. option for reduced thickness and could result in an aerodynamically superior solution when compared to a uniformly size ceramic matrix composite vanes. 
     The stress of the metal spar (e.g.  62 ) is proportional to the loading (or number of metallic vanes) divided by the second moment of area of the vane, therefore an optimum aerodynamic solution may be to balance the number of metallic vanes (coolant consumption) against the size of the vanes (aerodynamic loss) that just provides an acceptable stress in the metallic vane. 
     The damage mechanisms associated with the ceramic matrix composite material are less studied than metallics and silicon carbide/silicon carbide ceramic matrix composites suffer from relatively low environmental durability. This concept takes advantage of the significant engine experience associated with metallic nickel vanes (e.g.  36 ) along with better understanding of the damage mechanisms, providing a robust and reliable support structure. It may be advantageous to only transmit the flow through the metallic vanes (e.g.  36 ). 
     While the disclosure has been illustrated and described in detail in the foregoing drawings and description, the same is to be considered as exemplary and not restrictive in character, it being understood that only illustrative embodiments thereof have been shown and described and that all changes and modifications that come within the spirit of the disclosure are desired to be protected.