Abstract:
One embodiment of the present invention is a unique method for producing a turbomachine airfoil. Other embodiments include unique methods for manufacturing an airfoil for a gas turbine engine. Still other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for airfoils for gas turbine engines and other turbomachinery. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/428,710, filed Dec. 30, 2010, entitled AIRFOIL FOR GAS TURBINE ENGINE, which is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to airfoils, and more particularly, airfoils for gas turbine engines and other turbomachinery. 
       BACKGROUND 
       [0003]    Airfoils for gas turbine engines and other turbomachinery 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 
       [0004]    One embodiment of the present invention is a unique method for producing a turbomachine airfoil. Other embodiments include unique methods for manufacturing an airfoil for a gas turbine engine. Still other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for airfoils for gas turbine engines and other turbomachinery. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  depicts some aspects of a non-limiting example of an airfoil in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  is a cross section of the airfoil of  FIG. 2   
           [0009]      FIG. 4  is a cross section illustrating some aspects of a non-limiting example of an airfoil in accordance with an embodiment of the present invention that includes a hollow substrate. 
           [0010]      FIG. 5  is a cross section illustrating some aspects of a non-limiting example of an airfoil in accordance with an embodiment of the present invention having a substrate removed. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    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. 
         [0012]    Referring to the drawings, and in particular  FIG. 1 , there are illustrated some aspects of a non-limiting example of a gas turbine engine  20  in accordance with an embodiment of the present invention. In one form, engine  20  is a propulsion engine, e.g., an aircraft propulsion engine. In other embodiments, engine  20  may be any other type of gas turbine engine, e.g., a marine gas turbine engine, an industrial gas turbine engine, or any aero, aero-derivative or non-aero gas turbine engine. In one form, engine  20  is a two spool engine having a high pressure (HP) spool  24  and a low pressure (LP) spool  26 . In other embodiments, engine  20  may include only a single spool, or may include three or more spools, e.g., may include an intermediate pressure (IP) spool and/or other spools. In one form, engine  20  is a turbofan engine, wherein LP spool  26  is operative to drive a propulsor  28  in the form of a turbofan (fan) system, which may be referred to as a turbofan, a fan or a fan system. In other embodiments, engine  20  may be a turboprop engine, wherein LP spool  26  powers a propulsor  28  in the form of a propeller system (not shown), e.g., via a reduction gearbox (not shown). In yet other embodiments, LP spool  26  powers a propulsor  28  in the form of a propfan. In still other embodiments, propulsor  28  may take other forms, such as one or more helicopter rotors or tilt-wing aircraft rotors. 
         [0013]    In one form, engine  20  includes, in addition to fan  28 , a bypass duct  30 , a compressor  32 , a diffuser  34 , a combustor  36 , a high pressure (HP) turbine  38 , a low pressure (LP) turbine  40 , a nozzle  42 A, a nozzle  42 B, and a tailcone  46 , which are generally disposed about and/or rotate about an engine centerline  49 . In other embodiments, there may be, for example, an intermediate pressure spool having an intermediate pressure turbine. In one form, engine centerline  49  is the axis of rotation of fan  28 , compressor  32 , turbine  38  and turbine  40 . In other embodiments, one or more of fan  28 , compressor  32 , turbine  38  and turbine  40  may rotate about a different axis of rotation. 
         [0014]    In the depicted embodiment, engine  20  core flow is discharged through nozzle  42 A, and the bypass flow is discharged through nozzle  42 B. In other embodiments, other nozzle arrangements may be employed, e.g., a common nozzle for core and bypass flow; a nozzle for core flow, but no nozzle for bypass flow; or another nozzle arrangement. Bypass duct  30  and compressor  32  are in fluid communication with fan  28 . Nozzle  42 B is in fluid communication with bypass duct  30 . Diffuser  34  is in fluid communication with compressor  32 . Combustor  36  is fluidly disposed between compressor  32  and turbine  38 . Turbine  40  is fluidly disposed between turbine  38  and nozzle  42 A. In one form, combustor  36  includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustor  36  may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, a pulse detonation combustion system or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
         [0015]    Fan system  28  includes a fan rotor system  48  driven by LP spool  26 . In various embodiments, fan rotor system  48  may include one or more rotors (not shown) that are powered by turbine  40 . In various embodiments, fan  28  may include one or more fan vane stages (not shown in  FIG. 1 ) that cooperate with fan blades (not shown) of fan rotor system  48  to compress air and to generate a thrust-producing flow. Bypass duct  30  is operative to transmit a bypass flow generated by fan  28  around the core of engine  20 . Compressor  32  includes a compressor rotor system  50 . In various embodiments, compressor rotor system  50  includes one or more rotors (not shown) that are powered by turbine  38 . Compressor  32  also includes a plurality of compressor vane stages (not shown in  FIG. 1 ) that cooperate with compressor blades (not shown) of compressor rotor system  50  to compress air. In various embodiments, the compressor vane stages may include a compressor discharge vane stage and/or a diffuser vane stage. 
         [0016]    Turbine  38  includes a turbine rotor system  52 . In various embodiments, turbine rotor system  52  includes one or more rotors (not shown) operative to drive compressor rotor system  50 . Turbine  38  also includes a plurality of turbine vane stages (not shown in  FIG. 1 ) that cooperate with turbine blades (not shown) of turbine rotor system  52  to extract power from the hot gases discharged by combustor  36 . Turbine rotor system  52  is drivingly coupled to compressor rotor system  50  via a shafting system  54 . Turbine  40  includes a turbine rotor system  56 . In various embodiments, turbine rotor system  56  includes one or more rotors (not shown) operative to drive fan rotor system  48 . Turbine  40  also includes a plurality of turbine vane stages (not shown in  FIG. 1 ) that cooperate with turbine blades (not shown) of turbine rotor system  56  to extract power from the hot gases discharged by turbine  38 . Turbine rotor system  56  is drivingly coupled to fan rotor system  48  via a shafting system  58 . In various embodiments, shafting systems  54  and  58  include a plurality of shafts that may rotate at the same or different speeds and directions for driving fan rotor system  48  rotor(s) and compressor rotor system  50  rotor(s). In some embodiments, only a single shaft may be employed in one or both of shafting systems  54  and  58 . Turbine  40  is operative to discharge the engine  20  core flow to nozzle  42 A. 
         [0017]    During normal operation of gas turbine engine  20 , air is drawn into the inlet of fan  28  and pressurized by fan rotor  48 . Some of the air pressurized by fan rotor  48  is directed into compressor  32  as core flow, and some of the pressurized air is directed into bypass duct  30  as bypass flow. Compressor  32  further pressurizes the portion of the air received therein from fan  28 , which is then discharged into diffuser  34 . Diffuser  34  reduces the velocity of the pressurized air, and directs the diffused core airflow into combustor  36 . Fuel is mixed with the pressurized air in combustor  36 , which is then combusted. The hot gases exiting combustor  36  are directed into turbines  38  and  40 , which extract energy in the form of mechanical shaft power to drive compressor  32  and fan  28  via respective shafting systems  54  and  58 . The hot gases exiting turbine  40  are discharged through nozzle system  42 A, and provide a component of the thrust output by engine  20 . 
         [0018]    Gas turbine engine  20  employs many airfoils in the form of blades and vanes in order to pressurize, expand and/or direct the flow of air and/or combustion products in and through engine  20 . The airfoils are used in fan  28 , compressor  32  and turbines  38  and  40 . It is desirable that the airfoils be light in weight, e.g., in order to reduce the weight of gas turbine engine  20  and increase the damage tolerance of engine  20 . Accordingly, embodiments of the present invention envision, among other things, airfoils having a skin formed from one or more nano-metals. Nano-metals may have superior properties relative to conventional metals, e.g., including strength and crack resistance, due to the very small grain size of nano-metals. Although embodiments are described herein as with respect to airfoils for gas turbine engines, the present application also envisions embodiments pertaining to airfoils for other types of turbomachinery. 
         [0019]    Referring to  FIGS. 2-5 , some aspects of a non-limiting example of an airfoil  60  in accordance with an embodiment of the present invention is depicted. Airfoil  60  includes a nano-metal skin  62  formed over a composite substrate  64 . A portion of nano-metal skin  62  is removed in the illustration of  FIG. 3  in order to illustrate some aspects of substrate  64 . Substrate  64  is formed of a composite material having electrically conductive elements  66  dispersed therein. In one form, electrically conductive elements  66  are configured to provide electrical conductivity to substrate  64 , e.g., sufficient for use in depositing nano-metal material onto substrate  64  via an electrodeposition process. 
         [0020]    Substrate  64  is formed into an airfoil shape. In one form, substrate  64  is formed into the airfoil shape by an injection molding process. In other embodiments, other manufacturing processes may be used in addition to or in place of injection molding to form substrate  64  into an airfoil shape. In one form, composite substrate  64  includes a resin having electrically conductive elements  66  disposed therein. In one form, electrically conductive elements  66  are fibers. In other embodiments, electrically conductive elements  66  may take other forms, e.g., one or more conductive powders dispersed throughout the resin in addition to or in place of fibers. In a particular form, composite substrate  64  is a carbon-fiber composite, wherein electrically conductive elements  66  are carbon fibers, e.g., in a carbon fiber fabric. 
         [0021]    In one form, substrate  64  is solid, e.g., as depicted in  FIG. 3 . In another form, the substrate may be hollow, depicted in the cross-section of  FIG. 4  as substrate  64 A. In other embodiments, substrate  64  may take other geometric forms that provide an airfoil shape on the external surface of substrate  64 . For example and without limitation, in some embodiments, substrate  64  may be a hollow substrate having reinforcing ribs or struts extending through the hollow. 
         [0022]    After forming substrate  64 , a nano-metal is deposited onto substrate  64  to form nano-metal skin  62 . In one form, the nano-metal is deposited onto the surface of substrate  64  to form skin  62  via an electrodeposition process. In other embodiments, other processes may be employed to deposit the nano-metal onto substrate  64 . In one form, the nano-metal is a nickel-based alloy. In other embodiments, other metals and/or alloys may be employed in addition to or in place of a nickel-based alloy. In some embodiments, coatings or other treatments may be applied to the surface of nano-metal skin  62  and/or may be applied to substrate  64  prior to the deposition of the nano-metal material. The nano-metal skin is formed to a desired thickness  68 . Thickness  68  may vary with the needs of the application. In addition, the thickness of skin  62  may vary with location about skin  62 . For example, in one form, the nano-metal layer(s) forming skin  62  is configured to withstand thermal, mechanical and aerodynamic loading associated with its location and function during service in engine  20 , and hence may have different thickness values at different locations about skin  62 . In one form, the nano-metal used to form skin  62  has a grain size in the range of 15 nanometers to 100 nanometers. In other embodiments, larger and/or smaller grain sizes may be employed in addition to or in place of values within the range of 15 nanometers to 100 nanometers. 
         [0023]    In one form, skin  62  is sintered after electrodeposition onto substrate  64 . In other embodiments, skin  62  may not be sintered, e.g., depending upon the type of process used to deposit the nano-metal onto substrate  64 . In various embodiments, other treatments may be applied to skin  62 , for example and without limitation, a hot isostatic press (HIP). In some embodiments, as depicted in  FIG. 5 , substrate  64  may be removed after skin  62  is formed, e.g., via electrical, chemical and/or mechanical processing. In embodiments where skin  62  is sintered, substrate  64  may be removed after sintering, or may be removed prior to sintering or during sintering. In some embodiments, substrate  64  may be retained as part of airfoil  60 , e.g., to provide enhanced damage tolerance and/or airfoil damping. Although depicted as an airfoil alone, in various embodiments, airfoil  60  may be formed to include attachment features for affixing airfoil  60  to engine  20  and/or other components, or may include other features. For example and without limitation, when implemented as a fan, compressor or turbine blade, airfoil  60  may include a rotor attachment feature and/or may include a mid-span snubber and/or tip shroud; when implemented as a vane, airfoil  60  may include an attachment feature for securing airfoil  60  to a vane ring or a vane segment. In some embodiments, attachment features and/or other features may be formed separately and affixed to airfoil  60 . 
         [0024]    Embodiments of the present invention include a method for producing a turbomachine airfoil, comprising: providing a composite material interspersed with electrically conductive elements; forming the composite material into a substrate having an airfoil shape; and depositing a nano-metal onto the substrate. 
         [0025]    In a refinement, the nano-metal is a nickel based alloy. 
         [0026]    In another refinement, the nano-metal has a grain size in the range of 15 nanometers to 100 nanometers. 
         [0027]    In yet another refinement, the electrically conductive elements are fibers. 
         [0028]    In still another refinement, the electrically conductive elements are carbon fibers. 
         [0029]    In yet still another refinement, the composite material is a carbon fiber composite. 
         [0030]    In an additional refinement, the composite material includes a resin. 
         [0031]    In a further refinement, the nano-metal is deposited into the airfoil shape using electrodeposition. 
         [0032]    In a yet further refinement, the method further comprises sintering the nano-metal subsequent to depositing a nano-metal onto the airfoil shape. 
         [0033]    In a still further refinement, the method further comprises removing the substrate from the airfoil. 
         [0034]    In a yet still further refinement, the substrate is hollow. 
         [0035]    Embodiments of the present invention include a method for manufacturing an airfoil for a gas turbine engine, comprising: interspersing a composite material with electrically conductive elements; forming a substrate having an airfoil shape from the composite material with electrically conductive elements; and forming a nano-metal layer on the substrate. 
         [0036]    In a refinement, the method further comprises sintering the nano-metal layer. 
         [0037]    In another refinement, the nano-metal layer has a thickness configured to withstand thermal and mechanical loads of the airfoil in service in a gas turbine engine. 
         [0038]    In yet another refinement, the composite material is a carbon-fiber composite. 
         [0039]    In still another refinement, the nano-metal used to form the nano-metal layer is a nickel based alloy having a grain size in the range of 15 nanometers to 100 nanometers. 
         [0040]    In yet still another refinement, the nano-metal layer is formed onto the substrate using electrodeposition. 
         [0041]    In an additional refinement, the method further comprises removing the substrate from the airfoil. 
         [0042]    In a further refinement, the airfoil shape is formed by injection molding the composite material with electrically conductive elements. 
         [0043]    Embodiments of the present invention include a method for manufacturing an airfoil for a gas turbine engine, comprising: a step for forming a composite material interspersed with electrically conductive elements into an airfoil shape having an electrically conductive surface; a step for depositing a nano-metal layer at one or more desired thicknesses onto the airfoil shape; and a step for solidifying the nano-metal layer. 
         [0044]    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.