Patent Document

CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of U.S. Provisional Application No. 61/682,870, filed Aug. 14, 2012, the contents of which are incorporated herein by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The present invention generally relates to processes for producing airfoil components of turbomachinery and airfoil components produced thereby. More particularly, this invention is directed to processes for producing ceramic-based airfoil components with tip caps, and airfoil components produced thereby. 
         [0003]    Components of turbomachinery, including blades (buckets) and vanes (nozzles) of gas turbines, are typically formed of nickel-, cobalt- or iron-base superalloys with desirable mechanical and environmental properties for turbine operating temperatures and conditions. Because the efficiency of a gas turbine is dependent on its operating temperatures, there is a demand for components that are capable of withstanding increasingly higher temperatures. As the maximum local temperature of a component approaches the melting temperature of its alloy, forced air cooling becomes necessary. For this reason, airfoils of gas turbines, and in particular their low pressure and high pressure turbine (LPT and HPT) blades, often require complex cooling schemes in which air is forced through internal cooling passages within the airfoil and then discharged through cooling holes at the airfoil surface. Airfoil components can be equipped with tip caps that regulate internal cavity pressure, allowing for proper air flow through the cooling passages and holes. Tip caps are typically cast, brazed or welded onto metallic air-cooled LPT and HPT blades. 
         [0004]    As higher operating temperatures for gas turbines are continuously sought in order to increase their efficiency, alternative materials have been investigated. Ceramic-based materials are a notable example because their high temperature capabilities significantly reduce cooling air requirements. As used herein, ceramic-based materials encompass homogeneous (monolithic) ceramic materials as well as ceramic matrix composite (CMC) materials. CMC materials generally comprise a ceramic fiber reinforcement material embedded in a ceramic matrix material. The reinforcement material may be discontinuous short fibers that are randomly dispersed in the matrix material or continuous fibers or fiber bundles oriented within the matrix material. The reinforcement material serves as the load-bearing constituent of the CMC in the event of a matrix crack. In turn, the ceramic matrix protects the reinforcement material, maintains the orientation of its fibers, and serves to dissipate loads to the reinforcement material. Silicon-based composites, such as silicon carbide (SiC) as the matrix and/or reinforcement material, have become of particular interest to high-temperature components of gas turbines, including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry. SiC fibers have also been used as a reinforcement material for a variety of other ceramic matrix materials, including TiC, Si 3 N 4 , and Al 2 O 3 . Continuous fiber reinforced ceramic composites (CFCC) are a particular type of CMC that offers light weight, high strength, and high stiffness for a variety of high temperature load-bearing applications, including shrouds, combustor liners, vanes (nozzles), blades (buckets), and other high-temperature components of gas turbines. A notable example of a CFCC material developed by the General Electric Company under the name HiPerComp® contains continuous silicon carbide fibers in a matrix of silicon carbide and elemental silicon or a silicon alloy. 
         [0005]    Various techniques may be employed in the fabrication of CMC components, including chemical vapor infiltration (CVI) and melt infiltration (MI). These fabrication techniques have been used in combination with tooling or dies to produce near-net-shape articles through processes that include the application of heat and chemical processes at various processing stages. Examples of such processes, particularly for SiC/Si—SiC (fiber/matrix) CFCC materials, are disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. One such process entails the fabrication of CMCs from prepregs, each in the form of a tape-like structure comprising the desired reinforcement material, a precursor of the CMC matrix material, and one or more binders. After partially drying and, if appropriate, partially curing the binders (B-staging), the resulting tape is laid-up with other tapes, debulked and, if appropriate, cured while subjected to elevated pressures and temperatures to produce a cured preform. The preform is then fired (pyrolized) in a vacuum or inert atmosphere to remove solvents, decompose the binders, and convert the precursor to the desired ceramic matrix material, yielding a porous preform that is ready for melt infiltration. During melt infiltration, molten silicon and/or a silicon alloy is typically infiltrated into the porosity of the preform, where it fills the porosity and may react with carbon to form additional silicon carbide. 
         [0006]    For purposes of discussion, a low pressure turbine (LPT) blade  10  of a gas turbine engine is represented in  FIG. 1 . The blade  10  is an example of a component that can be produced from ceramic-based materials, including CMC materials. The blade  10  is generally represented as being of a known type and adapted for mounting to a disk or rotor (not shown) within the turbine section of an aircraft gas turbine engine. For this reason, the blade  10  is represented as including a dovetail  12  for anchoring the blade  10  to a turbine disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented in  FIG. 1 , the interlocking features comprise one or more protrusions  14  that engage recesses defined by the dovetail slot. The blade  10  is further shown as having a platform  16  that separates an airfoil  18  from a shank  20  on which the dovetail  12  is defined. 
         [0007]    Current state-of-the-art approaches for fabricating ceramic-based turbine blades have involved integrating the dovetail  12 , platform  16 , and airfoil  18  as one piece during the manufacturing process, much like conventional investment casting techniques currently used to make metallic blades. Because of their relatively higher temperature capability, CMC airfoils such as the blade  10  have not been equipped with tip caps for the purpose described above for metallic airfoil components. Moreover, brazing and welding techniques used to attach tip caps to metallic air-cooled LPT and HPT blades processes are not generally practical for attaching tip caps to airfoil components formed of CMC materials. In addition, tip caps define a geometric feature that is oriented transverse to the span-wise direction of the blade  10 , such that the incorporation of a tip cap into a CMC blade would pose design and manufacturing challenges. Furthermore, the low strain-to-failure capabilities of typical CMC materials pose additional challenges to implementing tip caps in rotating CMC airfoil components such as turbine blades, where a tip cap would be subjected to high centrifugal forces. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0008]    The present invention provides a process for producing airfoil components containing ceramic-based materials, in which a tip cap formed of a ceramic-based material is incorporated to yield a component that may further incorporate air cooling cavities and cooling holes to provide an air cooling capability. 
         [0009]    According to a first aspect of the invention, a process is provided that entails forming an airfoil portion of an airfoil component from an airfoil portion material that contains a precursor of a ceramic-based material. The airfoil portion material defines concave and convex walls of the airfoil portion, and the concave and convex walls define a tip region of the airfoil portion and at least a first cavity within the airfoil portion. At least a first ply is formed that contains a precursor of a ceramic-based material, and the first ply at least partially closes the first cavity at the tip region of the airfoil portion. The airfoil portion material of the airfoil portion and the first ply are then cured so that the first ply forms a tip cap that closes the first cavity at the tip region and the precursors of the airfoil portion material and first ply are converted to the ceramic-based materials thereof. 
         [0010]    According to a preferred aspect of the invention, an airfoil component produced by the process described above may be, as a nonlimiting example, a turbine blade of a turbomachine. 
         [0011]    A technical effect of this invention is the ability to produce CMC airfoil components having tip caps suitable for use in combination with internal air cooling schemes, wherein the tip caps are capable of exhibiting strength and effective load transfer for inclusion on rotating airfoil components, including turbine blades. 
         [0012]    Other aspects and advantages of this invention will be better appreciated from the following detailed description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0013]      FIG. 1  is a perspective view schematically representing a turbine blade of a type that can be formed of a CMC material in accordance with embodiments of the present invention. 
           [0014]      FIGS. 2A and 2B  schematically represent, respectively, and end view and a span-wise cross-sectional view of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of a tip cap from prepreg plies in accordance with an embodiment of the present invention. 
           [0015]      FIGS. 3A and 3B  schematically represent, respectively, and end view and a span-wise cross-sectional view of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of a tip cap from prepreg plies in accordance with another embodiment of the present invention. 
           [0016]      FIG. 4  schematically represents a chord-wise cross-sectional view of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of a tip cap that closes multiple cavities within the blade in accordance with an embodiment of the invention. 
           [0017]      FIG. 5  schematically represents a chord-wise cross-sectional view of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of multiple tip caps each individually closing a cavity within the blade in accordance with an embodiment of the invention. 
           [0018]      FIG. 6  shows two perspective views of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of a reinforced tip cap in accordance with another embodiment of the invention. 
           [0019]      FIG. 7  is a perspective view of the tip region of a turbine blade (such as that of  FIG. 1 ), and represents the integration of holes in a tip cap constructed in accordance with embodiments of the invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0020]    The present invention will be described in terms of processes for producing components that contain ceramic-based materials, and particularly the incorporation of one or more tip caps that can be used to close one or more internal cavities of a component formed of a ceramic-based material, preferably a CMC material. While various applications are foreseeable and possible, applications of particular interest include high temperature applications, for example, turbine components of gas turbines, including land-based and aircraft gas turbine engines. The CMC turbine blade  10  of  FIG. 1  will serve as an example in the following discussion. While the invention is applicable to a wide variety of ceramic-based materials, ceramic-based materials of particular interest to the invention are believed to be CMC materials containing silicon, such as CMC&#39;s containing silicon carbide as the reinforcement and/or matrix material, for example, continuous silicon carbide fibers in a matrix of silicon carbide. However, other ceramic-based materials are also within the scope of the invention, nonlimiting examples of which include fibers and reinforcement materials formed of titanium carbide (TiC), silicon nitride (Si 3 N 4 ), and/or alumina (Al 2 O 3 ). 
         [0021]    As known in the art, the airfoil  18  of the blade  10  is an excellent candidate for being produced from a ceramic-based material, and especially a CMC material, because it is directly exposed to the hot combustion gases within the turbine section of a turbomachine, and has a generally linear geometry. On the other hand, the incorporation of an internal cooling cavity, cooling holes and a tip cap results in a more complex geometry, in the sense that the airfoil  18  has a generally linear geometry along its dominant span-wise axis, whereas a tip cap would be a geometric feature oriented transverse to the span-wise direction of the blade  10 . Furthermore, the off-axis geometry of a tip cap would be subjected to high mechanical loading during operation of the engine, and therefore require structural interface capabilities that pose substantial challenges to designing, manufacturing and integration with a blade formed of a CMC material. The present invention provides a process for taking advantage of the high-temperature capabilities of CMC materials, while addressing the difficulties of integrating a tip cap into an airfoil component formed of a CMC material. In particular, a preferred aspect of the present invention is the ability to produce a tip cap from plies, and to fully integrate the tip cap as part of an airfoil formed from plies utilizing a lay-up process. 
         [0022]      FIGS. 2A ,  2 B,  3 A,  3 B, and  4 - 7  schematically represent views of the tip region of the blade  10  of  FIG. 1 , and represent the integration of tip caps  22  from plies  24  in accordance with various non-limiting embodiments of the present invention. The airfoil  18  and tip cap  22  can be fabricated from ceramic-based materials produced using known processes, for example, with the use of prepregs. As a particular example, the airfoil  18  and its cap  22  can each be fabricated using a prepreg melt-infiltration (MI) process of a type previously described, wherein multiple prepregs are formed to contain one or more desired reinforcement materials and a precursor of the CMC matrix material, as well as one or more binders. The prepregs undergo lay-up, are debulked and cured while subjected to elevated pressures and temperatures, and may undergo various other processing steps to form a laminate preform. Thereafter, the laminate preform is heated (fired) in a vacuum or an inert atmosphere to decompose the binders and produce a porous preform, which then preferably undergoes melt infiltration. If the CMC material comprises a silicon carbide reinforcement material in a ceramic matrix of silicon carbide (a SiC/SiC CMC material), molten silicon or a silicon alloy is typically used to infiltrate and fill the porosity and, in preferred embodiments, react with a carbon constituent (carbon, carbon source, or carbon char) within the matrix to form silicon carbide. However, it will be apparent from the following discussion that the invention also applies to other types and combinations of ceramic and CMC materials. Furthermore, it is foreseeable that the unitary airfoil  18  and cap  22  could be fabricated with the use of materials other than prepregs, for example, cloth-reinforced CMCs, such as chemical vapor infiltrated (CVI) SiC reinforced with carbon fiber cloth (C/SiC), CVI/slurry cast/melt infiltrated SiC/SiC, and CVI SiC reinforced with SiC cloth. Polymer infiltration and pyrolysis (PIP) processes can also be used to deposit the matrix into a cloth reinforced preform, in which case a SiC or carbon cloth can be used. 
         [0023]    According to a preferred aspect of the invention, the fabrication of the tip cap  22  entails steps intended to fully integrate the tip cap  22  into the linear geometry of the airfoil  18 .  FIGS. 2A and 2B  represent an example of a blade tip region of the blade airfoil  18  during the fabrication of the tip cap  22 , which according to a preferred aspect of the invention can be entirely formed of a CMC material and produced by a CMC process as described above. As represented in  FIGS. 2A and 2B , the tip cap  22  is fabricated from multiple prepreg plies  24 .  FIGS. 2A and 2B  represent the plies  24  as disposed within a cavity  30  defined by and between the convex (suction) and concave (pressure) walls  26  and  28  of the airfoil  18 , which as represented in  FIG. 2B  are also fabricated from multiple plies  34 .  FIGS. 2A and 2B  further represent the plies  24  as extending in the chord-wise direction of the airfoil  18 . As previously noted, each of the plies  24  and  34  preferably contains a desired reinforcement material and a suitable precursor of a desired ceramic matrix material. The reinforcement material and ceramic matrix material of the tip cap plies  24  are preferably, though not necessarily, the same as those for the airfoil plies  34 . 
         [0024]    It should be appreciated that various numbers of plies  24  could be incorporated into the construction of the tip cap  22  of the blade  10 . To build up a suitable thickness for the tip cap  22  that completely fills the portion of the cavity  30  within the blade tip region of the airfoil  18 , most of the plies  24  are represented as having roughly equal chord-wise lengths ( FIG. 2A ). In addition, most of the plies  24  are represented as having roughly equal span-wise lengths ( FIG. 2B ), such that the tip cap  22  is substantially flush with the end of each wall  26  and  28  of the airfoil  18 . However, certain plies  24  are represented as being intentionally shorter than others in the chord-wise direction ( FIG. 2A ) to accommodate a varying width of the cavity  30 , and certain plies  24  are also represented as being intentionally shorter than others in the span-wise direction ( FIG. 2B ). It should be understood that the lengths and widths of the plies  24  can vary, for example, as a result of increasing or decreasing in length and/or width to yield what may be referred to as a stepped formation. Accordingly, shapes and sizes of the plies  24  other than the particular shapes and sizes represented in  FIG. 2  are foreseeable and within the scope of the invention. 
         [0025]    According to a preferred aspect of the invention, shorter plies  24  in the span-wise direction are utilized to create a wedge-shaped profile  32  at the radially-inward end of the tip cap  22 . As seen in  FIG. 2B , the wedge-shaped profile  32  of the cap  22  engages complementary notches  33  formed in the interior surfaces of the convex and concave walls  26  and  28  of the airfoil  18 . The wedge-shaped profile  32  of the tip cap  22  and the notch  33  within the airfoil cavity  30  cooperate to interlock the tip cap  22  within the cavity  30 , particularly after the plies  24  of the tip cap  22  are fired and melt infiltrated, enabling the tip cap  22  to withstand high centrifugal forces that exist during the operation of the blade  10 . 
         [0026]    To complete the manufacturing of the blade  10  and its tip cap  22 , the laid-up prepreg plies  24  and  34  are preferably debulked prior to undergoing curing, followed by firing during which binders are burned-off and a ceramic precursor is converted to the desired ceramic matrix material for the reinforcement material. Suitable debulking, curing and firing processes, as well as any additional processes necessary to achieve the final desired shape and properties of the blade  10 , are known in the art and therefore will not be described further. 
         [0027]    Whereas the plies  24  of the tip cap  22  are represented in  FIGS. 2A and 2B  as being oriented in the span-wise and chord-wise directions of the airfoil  18 ,  FIGS. 3A and 3B  represent another embodiment in which the plies  24  are oriented in the thickness-wise and chord-wise directions of the airfoil  18 . Aside from the difference in orientation of the plies  24 , the tip cap  22  can be fabricated and interlocked with the airfoil  18  in essentially the same manner as described for the embodiment of  FIGS. 2A and 2B . 
         [0028]      FIGS. 4 through 7  represent additional configurations for tip caps  22  that can be fabricated in accordance with various aspects of the invention. Whereas in  FIGS. 2A ,  2 B,  3 A and  3 B, a single tip cap  22  is represented as filling a single cavity  30  in an airfoil  18 ,  FIG. 4  represents a single tip cap  22  as closing multiple cavities (cooling passages)  30  within an airfoil  18 ,  FIG. 5  represents separate tip caps  22  as individually filling and closing each of multiple cavities (cooling passages)  30  within an airfoil  18 , and  FIG. 6  represents the incorporation of pins  38  to help secure a tip cap  22  used to close multiple cavities (cooling passages)  30  within an airfoil  18 . In the embodiment of  FIG. 4 , the tip cap  22  is fabricated on top of all blade cavities  30  to seal off cooling passage air flow at the blade tip. Plies  34  of the airfoil walls  26  and  28  are represented in  FIG. 4  as wrapped around the cooling cavities  30  and the plies  24  of the tip cap  22 , and the tip cap  22  is bonded to the interior surfaces of the airfoil walls  26  and  28  during curing of both the airfoil  18  and tip cap  22 . In  FIG. 5 , each cooling cavity  30  is individually sealed off by a separate tip cap  24 . 
         [0029]    The reinforced embodiment of  FIG. 6  is intended to increase the aerodynamic and centrifugal loading capability of the tip cap  22 . In  FIG. 6 , a single tip cap  22  (shown in the upper blade tip of  FIG. 6 , but omitted in the lower image to reveal the cavities  30  and pins  38 ) is represented as closing multiple cavities (cooling passages)  30  within an airfoil  18 , though it should be understood that separate tip caps  22  that individually fill and close multiple cavities  30  could also be reinforced in the same or similar manner. The embodiment represented in  FIG. 6  entails additional steps between the lamination and cure processes. In a particular example, holes (not shown) are drilled through the airfoil walls  26  and  28  and tip cap  22 , for example, using an ultrasonic needling process, and then the pins  38 , for example, formed of prepregs, are inserted into the holes to create an interlocking connection between the airfoil walls  26  and  28  and tip cap  22  following curing. 
         [0030]    Finally,  FIG. 7  represents the incorporation of holes  40  in the tip cap  22  that are fluidically connected to one or more cavities (not shown) within the airfoil  18 . As known in the art of blades formed of metallic materials, tip cap purge holes have been utilized to regulate internal cavity pressures within blades, which in turn determines the cooling air flow rates through the cooling passages and cooling holes of the blades. Holes  40  of the type represented in  FIG. 7  can be formed by drilling after melt infiltration, when the CMC plies  24  of the tip cap  22  have been fully processed. Drilling techniques that can be used include electrodischarge machining (EDM), ultrasonic machining, or another traditional machining technique. 
         [0031]    While the invention has been described in terms of specific embodiments, it is apparent that other forms could be adopted by one skilled in the art. For example, the number of tip cap plies  24  required to close a particular cavity  30  of a blade  10  can be modified, for example, by increasing the thickness of either or both airfoil walls  26  and  28 . Furthermore, the composition of the tip cap  22  can vary from that described above, for example, discontinuous (chopped) fiber reinforcement materials could be used in place of continuous fiber reinforcement materials, and in doing so could potentially eliminate the need for multiple laminated plies  24  to form the tip cap  22 . In addition, welding or fusing techniques could be adapted to bond the tip cap  22  to the airfoil  18  after melt infiltration, avoiding the process of forming the tip cap  22  as part of the initial composite laminate. Therefore, the scope of the invention is to be limited only by the following claims.

Technology Category: y