Abstract:
One embodiment of the present invention is a unique composite gas turbine engine component. In one form, the composite component is an airfoil. Another embodiment is a unique method for manufacturing a composite gas turbine engine component. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations composite gas turbine engine components. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims the benefit of U.S. Provisional Patent Application 61/290,698, filed Dec. 29, 2009, and is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to composite gas turbine engine components. 
       BACKGROUND 
       [0003]    Composite gas turbine engine components 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 composite gas turbine engine component. In one form, the composite component is an airfoil. Another embodiment is a unique method for manufacturing a composite gas turbine engine component. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations composite gas turbine engine components. 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 
         [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  illustrates a composite gas turbine engine component in accordance with an embodiment of the present invention. 
           [0007]      FIGS. 2A-2C  depict an example of openings in a composite component in accordance with an embodiment of the present invention. 
           [0008]      FIGS. 3A-3C  depict an example of openings in a composite component in accordance with an embodiment of the present invention. 
           [0009]      FIG. 4  schematically depicts a system for forming openings in a composite gas turbine engine component. 
       
    
    
     DETAILED DESCRIPTION 
       [0010]    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. 
         [0011]    Referring now to the drawings, and in particular  FIG. 1 , there is schematically illustrated a non-limiting example of a composite component  10  in accordance with an embodiment of the present invention. In one form, composite component  10  is a composite gas turbine engine component. As a composite component, gas turbine engine component  10  is formed of a composite material. In one form, the composite material is a ceramic matrix composite (CMC). In one form, the CMC material contains one or more of silicon carbide, transition metal carbide, carbon and/or oxide fibers with any combination of silicon carbide, transition metal carbide, carbon, boron carbide, aluminum oxide and/or any other ceramic, transition metal, and transition metal intermetallic. In other embodiments, other composite materials may be employed in addition to or in place of CMC, including metal matrix composites (MMC), organic matrix composites (OMC) and carbon-carbon composites. 
         [0012]    In one form, gas turbine engine component  10  is a hot gas flowpath component of a gas turbine engine, that is, a gas turbine engine component that is directly exposed to the primary hot gas flowpath of a gas turbine engine. In other embodiments, component  10  may be a gas turbine engine component other than a hot gas flowpath component and/or may be any flowpath component. In one form, component  10  is disposed at least partially in the hot gas flowpath. In a particular form, gas turbine engine component  10  is an airfoil, such as a blade or a vane airfoil. Gas turbine engine component  10  may also be a blade platform or a vane shroud or the like, which may or may not be integral with an airfoil. In some embodiments, component  10  may bound the hot gas flowpath, i.e., defines at least a portion of a boundary or wall of the hot gas flowpath. It is alternatively considered that component  10  may be any composite gas turbine engine component. 
         [0013]    In one form, composite gas turbine engine component  10  includes a composite structure  12  having a surface  14 , a cavity  16  and a plurality of openings  18 . Component  10  may include other features not shown in  FIG. 1  or described herein, or may include less features than shown in  FIG. 1  or described herein. In one form, surface  14  is a flowpath surface of component  10  and operable in the hot gas flowpath of the gas turbine engine. In a particular form, surface  14  is operable in the hot gas flowpath of a gas turbine engine during operation of the engine. As such, surface  14  is exposed to the hot gases passing through the flowpath. In other embodiments, surface  14  may be any surface of a composite gas turbine engine component that may or may not include a cavity such as cavity  16 . 
         [0014]    Cavity  16  is spaced apart from flowpath surface  14  by a wall thickness  20  of the composite material that forms component  10 . Thickness  20  may vary with location on component  10  or may be constant. In the depiction of  FIG. 1 , cavity  16  is illustrated as a single cavity inside component  10 . It will be understood that the present invention contemplates components having any number of cavities or no cavities. 
         [0015]    Openings  18  extend into surface  14  of composite component  10 . In one form, openings  18  are cooling air holes for supplying cooling air from cavity  16  to flowpath surface  14 , and extend through thickness  20  to cavity  16 . In the form of cooling holes, openings  18  are operable to discharge cooling air from cavity  16  to surface  14  and into the hot gas flowpath. In other embodiments, openings  18  may be other types of openings, e.g., recesses for receiving a mating part or for positioning component  10  relative to another part of the engine, and may or may not extend through thickness  20  to cavity  16  ( FIG. 1 ). The number, shape and location of openings  18  may vary with the needs of the particular application. 
         [0016]    Referring now to  FIGS. 2A-2C , a non-limiting example of an opening  18  in accordance with an embodiment of the present invention is depicted, and is identified as opening  18 A. Opening  18 A includes a geometric shape  22  extending from surface  14  through at least part of the composite material that defines component  10 , e.g., toward cavity  16 . In one form, a hole  24  extends between geometric shape  22  and cavity  16 . In one form, hole  24  is cylindrical in shape. In other embodiments, hole  24  may have any other shape suited to the particular application. In one form, geometric shape  22  extends to cavity  16 , e.g., wherein hole  24  may be considered a part of geometric shape  22 . In another form, hole  24  may be considered a part of cavity  16  or otherwise a separately formed feature, e.g., formed separately from geometric shape  22 . Geometric shape  22  may take a variety of forms. In one form, geometric shape  22  is noncylindrical. In one form, geometric shape  22  forms a diffuser, e.g., for diffusing cooling air received via cavity  16 . In a particular form, geometric shape  22  is fan shaped, e.g., as depicted in  FIGS. 2A-2C . In other embodiments, geometric shape  22  may have another shapes suited to the application. In one form, geometric shape  22  is formed ultrasonically, i.e., via an ultrasonic machining (USM) process. In another form, geometric shape  22  is formed by an electrical discharge machining (EDM) process. In other embodiments, other processes may be used to manufacture geometric shape  22 . 
         [0017]    Referring now to  FIGS. 3A-3C , another non-limiting example of an opening  18  in accordance with an embodiment of the present invention is depicted, and is identified as opening  18 B. Opening  18 B includes a geometric shape  26  extending from surface  14  through at least part of the composite material that defines component  10 , e.g., toward cavity  16 . In one form, geometric shape  26  includes a portion  26 A and a portion  26 B extending from portion  26 A. In one form, a hole  28  extends between geometric shape  26  and cavity  16 . In one form, hole  28  is cylindrical in shape. In other embodiments, hole  28  may have any other shape suited to the particular application. In one form, geometric shape  26  extends to cavity  16 , e.g., wherein hole  28  is considered a part of geometric shape  26 . In another form, hole  28  may be considered a part of cavity  16  or otherwise a separately formed feature, e.g., formed separately from geometric shape  26 . 
         [0018]    Geometric shape  26  may take a variety of forms. In one form, geometric shape  26  is noncylindrical. In one form, geometric shape  26  forms a diffuser, e.g., for diffusing cooling air received via cavity  16 . In a particular form, geometric shape  26  is laid-back fan shaped, e.g., as depicted in  FIGS. 3A-3C , wherein a fan shaped portion  26 B is laid back at an angle φ relative to the centerline of a fan shaped portion  26 A. In other embodiments, geometric shape  26  may have another shapes suited to the application. In one form, geometric shape  26  is formed ultrasonically, i.e., via a USM process. In another form, geometric shape  26  is formed by an EDM process. In other embodiments, other processes may be used to manufacture geometric shape  26 . 
         [0019]    In one form, component  10  is manufactured by forming a composite structure having surface  14 . The composite structure may also include cavity  16  spaced apart from the flowpath surface by thickness  20  of the composite material. The composite structure may be formed from one or more composite materials, e.g., those set forth herein, using conventional composite processing techniques. Once the composite structure is thus formed, openings  18  are formed by removing composite material from surface  14  to form one or more geometric shape, e.g., such as geometric shape  22  and/or  26  and/or other three-dimensional geometric shapes, which extend from the surface  14 , e.g., toward cavity  16  in embodiments where a cavity  16  is present. In other embodiments, composite material may be removed from cavity  16  to form a geometric shape extending toward surface  14 . 
         [0020]    Referring now to  FIG. 4 , a non-limiting example of a system  50  for forming openings  18  is described. System  50  includes a cutting tool  52 , such as a USM probe or an EDM probe. Cutting tool  52  includes one or more protrusions  54  that form one or more openings  18 . In particular, protrusions  54  have a shape correspond to the geometric shapes of openings  18 , such as geometric shapes  22 ,  26 . In one form, a plurality of protrusions  54  are used to simultaneously form a plurality of geometric shapes. In another form a single protrusion  54  is used to form one geometric shape at a time. 
         [0021]    In the form of an EDM system, system  50  electro-discharge machines the geometric shapes using cutting tool  52  with protrusions  54 . In the form of a USM system, system  50  ultrasonically machines the geometric shapes using cutting tool  52  with protrusions  54 . USM processing of openings  18  may be performed without masking surface  14 , which may be required for some other types of material removal processing. For example, some other processing techniques require the use of masking to protect surface  14  from the material removal processing and/or environment, e.g., where surface  14  has a coating, such as an environmental barrier coating, or is otherwise susceptible to chemical and/or physical damage. System  50  forms the geometric shapes with requiring the use of back-strike protection, which is required for some processing techniques, e.g., laser cutting or machining systems. 
         [0022]    Embodiments of the present invention include a composite gas turbine engine component, comprising: a composite structure having a flowpath surface operable in a hot gas flowpath of a gas turbine engine; a cavity spaced apart from the flowpath surface by a thickness of a composite material; and a cooling hole operative to discharge cooling air into the flowpath, wherein the cooling hole extends between the flowpath surface and the cavity, wherein the cooling hole includes an ultrasonically formed geometric shape extending from the flowpath surface through at least part of the composite material toward the cavity of the composite gas turbine engine component; and wherein the composite gas turbine engine component is disposed at least partially in the flowpath and/or bounds the flowpath. 
         [0023]    In a refinement, the ultrasonically formed geometric shape is noncylindrical. 
         [0024]    In another refinement, the ultrasonically formed geometric shape forms a diffuser for the cooling air. 
         [0025]    In yet another refinement, the ultrasonically formed geometric shape is fan shaped. 
         [0026]    In still another refinement, the ultrasonically formed geometric shape is laid-back fan shaped. 
         [0027]    In yet still another refinement, the composite gas turbine engine component is an airfoil. 
         [0028]    In a further refinement, the composite material is a ceramic matrix composite (CMC). 
         [0029]    Embodiments of the present invention include a method for manufacturing a composite gas turbine engine component, comprising: forming a composite structure that is operable in a gas turbine engine, the composite structure being defined by a composite material and having a surface; and machining a geometric shape into the surface and through at least part of the composite material using at least one of an ultrasonic machining process and an electrical discharge machining process. 
         [0030]    In a refinement, the machined geometric shape is fan shaped. 
         [0031]    In another refinement, the machined geometric shape is laid-back fan shaped. 
         [0032]    In yet another refinement, the composite gas turbine engine component is an airfoil. 
         [0033]    In still another refinement, the composite material is a ceramic matrix composite (CMC). 
         [0034]    Embodiments of the present invention include a method for manufacturing a composite airfoil, comprising: forming a composite airfoil structure having a flowpath surface and a cavity spaced apart from the flowpath surface by a thickness of a composite material; and a step for forming a geometric shape extending from the flowpath surface through at least part of the composite material toward the cavity of the composite airfoil. 
         [0035]    In a refinement, the step for forming the geometric shape includes ultrasonically machining the geometric shape in the composite airfoil. 
         [0036]    In another refinement, the step for forming includes using an ultrasonic probe that has a shape corresponding to the geometric shape. 
         [0037]    In yet another refinement, the step for forming the geometric shape includes electrical discharge machining the geometric shape in the composite airfoil. 
         [0038]    In still another refinement, the geometric shape forms at least part of a cooling hole for the composite airfoil. 
         [0039]    In yet still another refinement, the step for forming the geometric shape includes using a probe to simultaneously form a plurality of the geometric shapes in the flowpath surface, wherein the probe has a plurality of protrusions each having a shape corresponding to the geometric shape. 
         [0040]    In a further refinement, the flowpath surface has an environmental barrier coating; and wherein the step for forming the geometric shape is performed without using a masking material for protecting the environmental barrier coating. 
         [0041]    In a yet further refinement, the step for forming the geometric shape is performed without using a back-strike protection. 
         [0042]    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.