Patent Abstract:
One embodiment of the present invention is a gas turbine engine. Another embodiment is a gas turbine engine vane system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine vanes. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Full Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims the benefit of U.S. Provisional Patent Application 61/290,843, filed Dec. 29, 2009, and is incorporated herein by reference. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to gas turbine engines, and more particularly, to gas turbine engine vanes. 
     BACKGROUND 
     Gas turbine engine vanes remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
     SUMMARY 
     One embodiment of the present invention is a gas turbine engine. Another embodiment is a gas turbine engine vane system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine vanes. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
         FIG. 1  schematically depicts a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 2  depicts a non-limiting example of a vane assembly for a gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 3  depicts a non-limiting example of an inner ring, an outer ring and spokes coupling the inner and outer ring of the vane assembly of  FIG. 2 . 
         FIG. 4  depicts a non-limiting example of an airfoil with a cooling air tube and spokes disposed in the cooling air tube of the vane assembly of  FIG. 2 . 
         FIG. 5  is a cross-section through the airfoil depicted in  FIG. 4  that illustrates a bushing and a pad that transfer loads between the airfoil and a spoke of the vane assembly of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
     For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
     Referring now to the drawings, and in particular,  FIG. 1 , a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically depicted. In one form, gas turbine engine  10  is an axial flow machine, e.g., an aircraft propulsion power plant. In other embodiments, gas turbine engine  10  may be a centrifugal flow machine or a combination axial centrifugal flow machine. It will be understood that the present invention is equally applicable to various gas turbine engine configurations, for example, including turbojet engines, turbofan engines, turboprop engines, and turboshaft engines having axial, centrifugal and/or axi-centrifugal compressors and/or turbines. 
     In the illustrated embodiment, gas turbine engine  10  includes an engine core  12 . Engine core  12  includes a compressor  14  having a plurality of blades and vanes  16 , and outlet guide vanes (OGV)  18 , a diffuser  20 , a combustor  22  and a turbine  24 . Diffuser  20  and combustor  22  are fluidly disposed between OGV  18  of compressor  14  and turbine  24 . Turbine  24  is drivingly coupled to compressor  14  via a shaft  26 . Although only a single spool is depicted, embodiments of the present invention are equally applicable to multi-spool engines. In various embodiments, gas turbine engine  10  may include, in addition to engine core  12 , one or more fans, additional compressors and/or additional turbines. 
     During the operation of gas turbine engine  10 , air is supplied to the inlet of compressor  14 . Blades and vanes  16  compress air received at the inlet of compressor  14 , and after having been compressed, the air is discharged via OGV  18  into diffuser  20 . Diffuser  20  reduces the velocity of the pressurized air from compressor  14 , and directs the pressurized air to combustor  22 . Fuel is mixed with the air and combusted in combustor  22 , and the hot gases exiting combustor  22  are directed into turbine  24 . 
     Turbine  24  includes a plurality of blades and vanes  28 . Blades and vanes  28  extract energy from the hot gases to generate mechanical shaft power to drive compressor  14  via shaft  26 . In one form, the hot gases exiting turbine  24  are directed into an exhaust nozzle (not shown), which provides thrust output the gas turbine engine. In other embodiments, additional turbine stages in one or more additional rotors may be employed, e.g., in multi-spool gas turbine engines, which may include turbines upstream and/or downstream of turbine  24 . 
     Referring now to  FIGS. 2-4 , turbine  24  includes a vane assembly  30  and a static engine component  32 . In one form, vane assembly  30  is a second stage vane assembly. In other embodiments, vane assembly  30  may be a vane assembly of another turbine stage or of a compressor stage. In one form, second stage vane assembly  30  is formed of a plurality of vane packs  34 , described below, which are arranged circumferentially to form an annular gas path  36 . 
     Vane assembly  30  includes a plurality of circumferentially spaced composite airfoils  38 , an inner ring  40 , an outer ring  42 , a plurality of spokes  44  and a plurality of cooling air tubes  46 . Gas path  36  extends radially between inner ring  40  and outer ring  42 . Gas path  36  directs hot gases exiting the first stage turbine blades of blades and vanes  28  through composite airfoils  38  into the second stage turbine blades of blades and vanes  28 . In the depiction of  FIG. 3 , composite airfoils  38  are removed for purposes of illustration. 
     Composite airfoils  38  are turbine vane airfoils. Composite airfoils  38  are slidably disposed between inner ring  40  and outer ring  42 . In one form, each composite airfoil  38  includes an opening  48 . Opening  48  extends through composite airfoil  38  in the span-wise direction. In other embodiments, a greater or lesser number of openings  48  may be employed in vane assembly  30 . In still other embodiments, no openings  48  may be employed. Composite airfoils  38  are formed of a composite material. In one form, the composite material is a ceramic matrix composite. Other composite materials may be employed in other embodiments, e.g., including metal matrix composites, organic matrix composites and/or carbon-carbon composites. 
     Inner ring  40  defines the inner wall of gas path  36 . Inner ring  40  is formed of a circumferentially arranged plurality of arcuate inner shroud segments  50 . In one form, each inner shroud segment  50  includes an opening  52  adjacent and corresponding to the opening  48  of each composite airfoil  38  in vane pack  34 . In other embodiments, a greater or lesser number of openings  52  may be employed. In still other embodiments, no openings  52  may be employed. 
     Inner shroud segments  50  are metallic, e.g., nickel-base superalloys. An example of a material for inner shroud segments  50  is Inconel 718. In other embodiments, inner shroud segments  50  may be formed of a nonmetallic material, such as a composite material. 
     Inner ring  40  is coupled to static engine component  32  via piloting features  54 . In one form, piloting features  54  may be pilot diameter arrangements. In one form, static engine component  32  is a preswirler that provides a circumferential velocity component to cooling air being supplied to second stage turbine blades (not shown) from vane assembly  30 . In other embodiments, static engine component  32  may take other forms. In still other embodiments, static engine component  32  may house or support one or more rotating engine components. In some embodiments, static engine component  32  is coupled to and supported by inner ring  40 . In some embodiments, static engine component  32  is supported and positioned by only vane assembly  30 . In some embodiments, static engine component  32  provides hoop loading via piloting features  54  to retain inner shroud segments  50  in circumferential contact and/or close circumferential proximity to each other, to maintain the circularity of inner ring  40 , and/or to provide a structural ground for spokes  44  (via inner shroud segments  50 ) to allow spokes  44  to be in a state of tension. In yet other embodiments, inner ring  40  may not be coupled to any static engine component. In some forms, inner shroud segments  50  may include interlocking features (not shown) or interface features (not shown) that interface with interlocking features (not shown) in order to provide inner ring  40  with a hoop load carrying capacity. 
     Outer ring  42  defines the outer wall of gas path  36 . Outer ring  42  is formed of a circumferentially arranged plurality of arcuate outer shroud segments  56 . In one form, each outer shroud segment  56  includes an opening  58  adjacent and corresponding to the opening  48  of each composite airfoil  38  in vane pack  34 . In other embodiments, a greater or lesser number of openings  58  may be employed. In still other embodiments, no openings  58  may be employed. 
     Outer shroud segments  56  are metallic, e.g., nickel-base superalloys. An example of a material for outer shroud segments  56  is Inconel 718. In other embodiments, outer shroud segments  56  may be formed of a nonmetallic material, such as a composite material. 
     Outer ring  42  is coupled to a turbine structure  60  and a turbine structure  62 , which may be, for example, turbine case structures, such as vane case structures, or first and second stage turbine blade tracks, respectively. In other embodiments, outer ring  42  may be coupled to only a single component or may be coupled to more than two components. 
     In one form, each vane pack  34  includes a single inner shroud segment  50 , a single outer shroud segment  56  and two composite airfoils  38 . In some embodiments, each vane pack  34  includes a single inner shroud segment  50 , a single outer shroud segment  56  and two composite airfoils  38 . In other embodiments, a greater or lesser number of airfoils  38  may be associated with each vane pack  34 . A plurality of vane packs  34  are assembled together circumferentially to yield the annular vane assembly  30  illustrated in  FIG. 3 . 
     Spokes  44  are disposed within opening  48  each composite airfoil  38 . Spokes  44  extend between inner ring  40  and outer ring  42 , and are coupled to both inner ring  40  and outer ring  42 . In particular, spokes  44  are coupled to each inner shroud segment  50  and outer shroud segment  56  of inner ring  40  and outer ring  42 , respectively. Inner ring  40 , outer ring  42  and spokes  44  form a hub-and-spoke arrangement. Spokes  44  are operative to transfer loads between inner ring  40  and outer ring  42 , in conjunction with composite airfoils  38 . In one form, spokes  44  are pre-tensioned at assembly, i.e., pre-stressed with a tensile load. In one particular example, spokes  44  are preloaded to 80% of room temperature yield strength at assembly. In one form, there are two (2) spokes for each composite airfoil  38 . In other embodiments, a greater or lesser number of spokes  44  may be employed. In one form, spokes  44  are formed from wire. An example of a material for spokes  44  is Waspalloy, e.g., Waspalloy wire. 
     Cooling air tube  46  is disposed within opening  48 . In one form, cooling air tube extends through composite airfoil  38 . In another form, portions of cooling air tube  46  also extend into opening  52  of inner shroud segment  50  on one end, and extend into opening  58  of outer shroud segment  56  on the other end. Spokes  44  are disposed within and extend through cooling air tube  46 . Cooling air tube  46  is operative to transfer cooling air through second stage vane assembly  30  to static engine component  32 . In one form, cooling air tube  46  is metallic, e.g., a nickel-base superalloy. An example of a material for cooling air tube  46  is Inconel 718. In other embodiments, cooling air tube  46  may be formed of a nonmetallic material, such as a composite material. It will be understood that some embodiments may not employ a cooling air tube  46 . 
     Referring now to  FIG. 5 , vane assembly  30  includes a plurality of bushings  64  disposed around spokes  44  and inside cooling air tube  46 . Bushings  64  are operative to transfer loads from composite airfoil  38  to spokes  44 . In one form, two (2) bushings  64  are employed at each spoke  44 . In other embodiments, a greater or lesser number of bushings  64  may be employed. In still other embodiments, vane assembly  30  may not include any bushings  64 . 
     Cooling air tube  46  includes pads  66  adjacent the inner surface in opening  48  of each composite airfoil  38 . Pads  66  are operative to transfer loads between composite airfoil  38  and spokes  44  (in one form, via cooling air tube  46  and bushings  64 ). In one form, each cooling air tube  46  includes three (3) pads  66 . In other embodiments, a greater or lesser number of pads  66  may be employed. In still other embodiments, cooling air tubes  46  may not include any pads  66 . In yet other embodiments, opening  48  of each composite airfoil  38  may include raised pads to transfer loads from composite airfoil  38  to spokes  44 . In some embodiments, pads  66  are compliant, e.g., in order to distribute loading on the interior surface of opening  48  of composite airfoils  38 . 
     During the operation of gas turbine engine  10 , cooling air is supplied to static engine component  32  via openings  58  of outer ring  42 , cooling air tubes  46  disposed within openings  48  of composite airfoils  38 , and openings  52  of inner ring  40 . In some forms, vane assembly  30  includes seals  68  ( FIG. 2 ) to prevent leakage of cooling air from opening  48  and to prevent ingress of hot gases into openings  48 ,  52  and  58 . In one form, seals  68  are located in inner shroud segments  50  and outer shroud segments  56 . In other embodiments, seals  68  may be located in composite airfoils  38  or otherwise between surfaces of composite airfoils  38  and surfaces of inner shroud segments  50  and outer shroud segments  56 . In other embodiments, seals  68  may not be employed. In some embodiments, vane assembly  30  includes seals (not shown) operative to provide sealing between each adjacent pair of inner shroud segments  50 . In some embodiments, vane assembly  30  includes seals (not shown) operative to provide sealing between each adjacent pair of outer shroud segments  56 . 
     In some embodiments, cooling air tube  46  reduces undesirable heating of the cooling air supplied through vane assembly  30  by preventing or reducing contact of the cooling air with hot surfaces in opening  48  of composite airfoil  38 . In some embodiments, spokes  44  are kept relatively cool due to the passage of cooling air through cooling air tube  46 , which in some embodiments bathes the spokes  44  in the flow of cooling air and enhances the load-carrying capacity of spokes  44  by preventing or reducing degradation of the spoke  44  material properties that may otherwise result from operation at elevated temperatures. 
     Aerodynamic loading of composite airfoils  38 , e.g., in direction  70  ( FIG. 5 ), causes composite airfoils  38  to slide against inner shroud segments  50  and outer shroud segments  56  until pads  66  nest against composite airfoils  38 . Once nested, spokes  44  limit the movement of composite airfoils  38 , and aerodynamic loads on composite airfoils  38  are transferred to spokes  44  via pads  66 , cooling air tube  46  and bushings  64 . These loads are transferred to turbine structure  60 ,  62  via outer ring  42 . 
     Loads from static engine component  32  are transferred to inner ring  40 , e.g., via piloting features  54  that couple static engine component  32  and inner ring  40 . These loads are then transferred from inner ring  40  to outer ring  42  by composite airfoils  38  in conjunction with spokes  44 , and are transferred to turbine structure  60 ,  62  via outer ring  42 . 
     Embodiments of the present invention include vane segment for a gas turbine engine, comprising: an outer shroud; an inner shroud; a composite airfoil slidably disposed between the outer shroud and the inner shroud, the airfoil having a passage extending between the outer shroud and the inner shroud; and a spoke extending through the passage between the outer shroud and the inner shroud, wherein the spoke is operative to limit a movement of the airfoil. 
     In a refinement, the vane segment is structured to transfer aerodynamic loads from the airfoil to the spoke. 
     In another refinement, the spoke is pre-stressed at room temperature conditions with a tensile preload. 
     In yet another refinement, the spoke is a metal wire. 
     In still another refinement, the vane segment is structured to transfer loads from the inner shroud to the outer shroud via the spoke. 
     In yet still another refinement, the spoke is affixed to the inner shroud and to the outer shroud. 
     In a further refinement, the vane segment further comprises a tube disposed in the passage, wherein the spoke extends through the tube. 
     In yet a further refinement, the vane segment further comprises a compliant pad structured to transfer aerodynamic loads from the airfoil to the tube, and wherein the tube is structured to transfer the aerodynamic loads to the spoke. 
     In still a further refinement, the vane segment further comprises a second spoke extending through the passage and between the outer shroud and the inner shroud. 
     In yet still a further refinement, the vane segment further comprises a bushing disposed on the spoke and structured to transmit aerodynamic loads to the spoke. 
     Another embodiment includes a gas turbine engine, comprising: a compressor; a turbine; and a vane stage having an inner ring, an outer ring, a plurality of airfoils disposed between the inner ring and the outer ring, and a plurality of spokes extending between the inner ring and the outer ring through the plurality of airfoils, the spokes interconnecting the inner ring and the outer ring in a hub-and-spoke arrangement. 
     In a refinement, the gas turbine engine further comprises a vane case structure, wherein the outer ring is supported by the vane case structure. 
     In another refinement, the plurality of spokes are coupled to the inner ring and to the outer ring, and wherein the vane stage is structured to transfer loads from the inner ring to the vane case structure via the plurality of spokes and the outer ring. 
     In yet another refinement, the outer ring is formed of a plurality of vane segment outer shrouds. 
     In still another refinement, the inner ring is formed of a plurality of individual vane segment inner shrouds. 
     In yet still another refinement, the gas turbine engine further comprises an engine component supported by the inner ring, wherein the vane stage is structured to transfer loads from the engine component via the inner ring and the plurality of spokes to the outer ring. 
     In a further refinement, the engine component is a preswirler. 
     In a still further refinement, the plurality of airfoils are structured to float in a flowpath direction and transfer aerodynamic loads imposed on the plurality of airfoils to the plurality of spokes. 
     Embodiments also include gas turbine engine, comprising: at least one of a fan, a compressor and a turbine, wherein at least one of the fan, the compressor and the turbine includes a vane stage, the vane stage including: an inner ring; an outer ring; a plurality of airfoils slidably disposed between the inner ring and the outer ring; and means for transferring loads from the inner ring to the outer ring, wherein the means for transferring loads extends between the inner ring and the outer ring through the plurality of airfoils. 
     In a refinement, the means for transferring loads from the inner ring to the outer ring includes means for transferring aerodynamic loads from the plurality of airfoils to the outer ring. 
     In another refinement, the plurality of airfoils are composite airfoils. 
     While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.

Technology Classification (CPC): 5