Patent Document

CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    Not applicable. 
       STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0002]    Not applicable. 
       BACKGROUND OF THE INVENTION 
       [0003]    This invention is generally in the field of gas turbine engines. More particularly, the present invention is directed to a composite blade for a turbine rotor. 
         [0004]    Turbine blades are typically manufactured from a casting process in which a molten alloy is poured into a ceramic mold, heated, and then cooled. When the mold is broken off, the blade is then machined to its final shape. This results in a turbine blade having a substantially uniform composition from the root of the blade to the tip. Thus, the alloy chosen for the turbine blade must have suitable performance properties for the thermal and mechanical stresses encountered at various locations on the blade. Such a manufacturing process may not generally allow for a designer to independently select an optimal alloy for different portions of the turbine blade. 
         [0005]    In general, the turbine blade is cast from a creep resistant superalloy. In an exemplary turbine blade casting process, the superalloy is directionally solidified from root tip. During operation, turbine blades tips are exposed to extreme temperatures and stresses which cause them to oxidize and crack. A turbine blade may crack along grain boundaries at or near the tip of the airfoil and the crack will propagate along the length of the airfoil. Eventually a blade may suffer enough damage to compromise the turbine&#39;s efficiency. A blade is typically replaced before it reaches this level of damage. 
       BRIEF SUMMARY OF THE INVENTION 
       [0006]    In one aspect, the present invention comprises a composite turbine blade. The composite turbine blade comprises a turbine blade portion comprising a first material and a first tip plate comprising a second material. The turbine blade portion comprises an exterior wall and an interior wall surrounding a hollow interior cavity. The turbine blade portion further comprises a top surface extending from the exterior wall to the interior wall, and the top surface bounds an orifice that is fluidly connected to the hollow interior cavity. The composite turbine blade further comprises a first tip plate comprising a second material attached to the turbine blade along the top surface and extending from proximate the exterior wall of the turbine blade across the orifice to cover the orifice. 
         [0007]    In another aspect, the present invention comprises a composite turbine blade having a reinforced platform. The composite turbine blade comprises a turbine portion comprising a first material and an insert portion comprising a second material. The turbine blade portion comprises an airfoil portion having a tip and a root and a platform portion attached to the airfoil portion at the root. The platform portion comprises an orifice passing therethrough and an insert attached within the orifice of the platform portion. 
         [0008]    In another aspect, the present invention comprises a method of manufacturing a composite turbine blade. The method comprises (1) providing a turbine blade portion having an exterior wall and an interior wall surrounding a hollow interior cavity, and a top surface extending from the exterior wall to the interior wall bounding an orifice that is fluidly connected to the hollow interior cavity; and (2) attaching a first tip plate to the turbine blade along the top surface so that the first tip plate extends from proximate the exterior wall of the turbine blade across the orifice to cover the orifice. The turbine blade portion comprises a first material and the first tip plate comprises a second material. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  is a sectioned perspective view, illustrating a prior art turbine blade tip. 
           [0010]      FIG. 2  is a sectioned perspective view, illustrating a part of a process for repairing a turbine blade tip in accordance with an embodiment of this invention. 
           [0011]      FIG. 3  is a sectioned perspective view, illustrating a part of a process for repairing a turbine blade tip in accordance with an embodiment of this invention. 
           [0012]      FIG. 4  is a sectioned perspective view, illustrating a composite turbine blade tip in accordance with an embodiment of this invention. 
           [0013]      FIG. 5  is a perspective view, illustrating a composite turbine blade platform in accordance with an embodiment of this invention. 
           [0014]      FIGS. 6-9  are section views, illustrating profiles for the insert of a composite turbine blade platform in accordance with an embodiment of this invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0015]      FIG. 1  is a sectioned perspective view, illustrating the blade tip  12  of a conventional turbine blade  10 . The turbine blade  10  has a hollow interior  15  which is bounded at one end by a tip plate  14 . The tip plate  14  mates with a flange  16  which extends inward towards the center of the turbine blade  10 . The tip plate  14  is offset from the end of blade tip  12  to form a tip cavity  18  bounded by a wall  20 . The tip cavity  18  allows for cooling air to escape the airfoil between the blade tip  12  and the shroud of the casing during operation. 
         [0016]    As mentioned previously, the blade tip  12  is exposed to extreme temperatures and stress during operation. This can cause the blade tip  12  to deteriorate over time. Other components of the turbine blade are also subject to extreme stresses which can possibly lead to deterioration of the turbine blade. In one aspect, the present invention comprises a method of repairing a turbine blade to improve the performance or longevity of the blade. In another aspect, the present invention comprises a composite turbine blade. In another aspect, the present invention comprises a method of manufacturing a composite blade. 
         [0017]    As illustrated in  FIG. 2 , a method of manufacturing a composite blade in accordance with an embodiment of this embodiment begins with the step of preparing a turbine blade  100  having a modified top surface  22 . The turbine blade  100  of  FIG. 2  may be prepared by removing the tip plate  14  and the material of the wall  20  above the flange  16  of the turbine blade  10  of  FIG. 1 . Alternatively, the turbine blade  100  may be originally cast to have the profile shown in  FIG. 2 . The modified top surface  22  forms a plane across the flange  19  from the innermost point  21  of the flange  19  to the outer surface of the turbine blade  100 . 
         [0018]    As illustrated in  FIG. 3 , the plates  24  and  26  are attached to the modified top surface  22  of  FIG. 2 . The plate  24  may be attached the plate  26  and the flange  19  by various processes including, but not limited to, brazing, welding, or diffusion bonding. Alternatively, the plates  24  and  26  may consist of weld-deposited materials. The plates  24  and  26  may comprise the same or different materials. In one embodiment, the plate  24  comprises a material that provides excellent mechanical tolerance at high temperatures. In particular, plate  24  may comprise a material which is more resistant to creep than the material of the airfoil of turbine blade  100 . Although the preferred material for the plate  24  may vary, René 142™, René 80™, René N4™, René N5™, GTD 111™, and GTD 222™ alloys (General Electric Company) are exemplary materials for the plate  24  because of their resistance to stress rupture at high temperature. In certain embodiments, the plate  26  comprises a material that may withstand even higher temperatures without oxidizing. René 142™ and HAYNES 214™ (Haynes International) alloys are exemplary materials for the plate  26  because of their resistance to oxidation, however many other materials may be used for the plate  26  including, but not limited to, René 195™ (General Electric Company) and HAYNES 230™ (Haynes International) alloys. 
         [0019]    Although the present embodiment illustrates the use of two plates (plates  24  and  26 ), it should be noted that any number of plates may be used. For example, in some embodiments a single plate being both resistant to low cycle fatigue and oxidation may be used. Alternatively, a plurality of plates may be stacked to produce a gradient effect with each plate possessing the optimal properties for the thermodynamic and mechanical stresses at the particular location on the airfoil. For example, an intermediate plate comprising a material having an intermediate level of creep resistance and oxidation resistance relative to the plates  24  and  26  may be added between the plates  24  and  26 . 
         [0020]    The expression “different material” and variations thereof as used herein encompasses the use of different alloys among different components. The term also encompasses the use of the same alloy in different orientations among different components where the difference in orientation appreciably affects the manner in which the component responds to thermodynamic and mechanical stresses at the particular location where the component is placed on the turbine blade. 
         [0021]    Unlike conventional blade tip designs (e.g., the design of  FIG. 1 ), the quality of the bond between the plates  24  and  26  and the turbine blade  100  may be easily inspected without destroying the attached components or the bond. For example, the bond quality may be visually inspected or may be inspected using ultrasonic imaging techniques. As such, a bond quality assessment may be made before proceeding to the next step in the manufacturing process. 
         [0022]    As illustrated in  FIG. 4 , a blade tip  27  is then formed by machining the plates  24  and  26  to produce a cavity  28  bounded by a wall  29  and a tip plate  25 . The cavity  28  is preferably formed by milling away material from the plates  24  and  26  using a CNC milling machine; however, other machining methods may also be used. Further, the exterior walls of plates  24  and  26  may be machined to match the contours of the turbine blade. As such, the interior and exterior profile of wall  29  of the composite blade tip  27  may be made to mimic the wall profiles of the conventional blade tip  12  of  FIG. 1  or a new design may be employed. It should be noted that the unique manufacturing process for producing composite blade tip  27  allows for the manufacture of profile designs which would normally be disallowed by the constraints of the casting processes. Although not illustrated herein, in some embodiments the wall  29  may not entirely surround machined cavity  28 . For example, the wall  29  may comprise one or more gaps to allow cooling air to escape from the machine cavity  28 . As such, the term “substantially surrounding” and variations thereof when referring to the wall  29  of blade tip  27  herein is intended to encompass embodiments where the wall  29  completely surrounds the machined cavity  28  and embodiments where gaps are provided in the wall  29 . 
         [0023]    The foregoing process may be either used for manufacturing a new turbine blade or retrofitting a composite blade tip  27  to a used turbine blade (for repairing the used turbine blade or improving the performance of the used turbine blade). As mentioned previously, the principle variation in the process relates to the method of producing the modified top surface  22  of  FIG. 2 . In repairing or retrofitting applications, material must generally be removed from the used turbine blade before the composite blade tip  27  may be added. In new manufacturing applications, the turbine blade component may be manufactured to be shorter in length, and the composite blade tip  27  is then added to the end of the manufactured airfoil component. 
         [0024]    In another aspect, the present invention comprises a composite turbine blade  100  having a blade tip  27  produced by the foregoing method. One additional benefit of the blade tip configuration of the present invention is that the tip plate  25  is attached to the turbine blade  100  over a larger contact area than the tip plate  14  of the conventional blade tip design of  FIG. 1 . This reduces the risk of the tip plate  25  becoming disconnected from the turbine blade  10  during operation. Furthermore, the configuration of the present invention avoids the complexity associated with providing sufficient weld penetration in the conventional blade tip design of  FIG. 1 . 
         [0025]    The turbine blade  100  having a composite blade tip  27  benefits from variation in metallurgical properties at the tip of the blade. As described previously, the material of the plate  24  and the plate  26  may be generally selected to possess the optimal properties for the thermodynamic and mechanical stresses encountered at the particular location on the airfoil. In some embodiments, the blade tip  27  may be designed to simply prevent cracks which initiate in the airfoil from propagating to the tip of the airfoil. In embodiments where this is the principle design criteria, it may not be necessary to use an entirely different alloy for the blade tip  27 . For example, the blade tip  27  may comprise the same alloy as the cast portion of the airfoil where the grain orientation of the alloy of blade tip  27  is generally perpendicular to the grain orientation of the cast portion of the airfoil. Such a variation in grain orientation may be considered a “different material” from the material of the cast portion of the airfoil since the orientation appreciably affects the manner in which the component responds to thermodynamic and mechanical stresses at the particular location where the component is placed on the turbine blade (i.e. the orientation of the grain arrests the propagation of the crack). 
         [0026]    As illustrated in  FIG. 5 , the turbine blade  100  may be further reinforced by modifying the platform  30  to which the root of the airfoil is attached. An insert  32 , which comprises a different material from the material of the platform  30 , is provided within an orifice formed in platform  30 . The orifice may be formed during the casting process used to produce the turbine blade  100 . Alternatively, the orifice may be formed after the turbine blade is cast by milling away a portion of the material of the platform  30 . The insert  32  adds strength beyond that which is normally provided by the material of the platform  30 . As such, the insert  32  makes the platform  30  more strain tolerant. Although various materials may be used for the insert  32 , René 80™, René 142™, René 195™ alloys are exemplary materials for the insert  32  because of their excellent resistance to low cycle fatigue. 
         [0027]    As with the composite blade tip  27 , the insert  32  may be added during the manufacture of a new turbine blade or may be employed as retrofit strengthening or repair solution for a used turbine blade. Similar to the composite blade tip  27 , the utilization of an insert  32  allows the metallurgical properties of the platform  30  to be optimized for the thermodynamic and mechanical stresses encountered at each location of the platform  30 . 
         [0028]    Many different profiles may used for the insert. As illustrated in  FIG. 6 , the insert  32  may have vertically-straight sidewalls which mate with the vertically-straight sidewalls of the orifice in the platform  30 . Alternatively, as illustrated in  FIG. 7 , the insert  34  may have a “stepped” sidewall which mates with a vertically-straight sidewall of the orifice in the platform  30 . In this embodiment, the insert  34  comprises a flange which overlaps a portion of the platform  30 . In another embodiment, as illustrated in  FIG. 8 , the insert  36  may have a stepped sidewall which mates with a orifice having a stepped sidewall in the platform  30 . In this embodiment, the platform  30  has a counterbore which mates with the flange of the insert  36 . In another embodiment, as illustrated in  FIG. 9 , the insert  38  has a tapered sidewall which mates with a orifice having a tapered sidewall in the platform  30 . In each of the foregoing examples, the insert may be attached to the platform  30  by various processes including, but not limited to, brazing, welding, or diffusion bonding. 
       Example 
       [0029]    In one non-limiting example, a turbine blade of a conventional design is uniformly cast using a René 41 superalloy. The turbine blade tip is then modified as shown in  FIG. 2  using a CNC machine to form a modified top surface  22  having a flat plane across the flange  19  from the innermost point of the flange  19  to the outer surface of the turbine blade  100 . A plate of HAYNES 230™ alloy, corresponding to the plate  24 , is then welded to the modified surface  22  as illustrated in  FIG. 3 . A plate of HAYNES 214™ alloy, corresponding to the plate  26 , is then welded to the plate of HAYNES 230™ alloy as illustrated in  FIG. 3 . The plates of HAYNES 230™ alloy and HAYNES 214™ alloy are then milled using the CNC machine to form the shape of the profile illustrated in  FIG. 4 . 
         [0030]    The platform  30  of the cast turbine blade is then milled using a CNC machine to remove the cast superalloy material in the region of the platform  30  occupied by the insert  32  of  FIG. 5  (i.e., the region of the platform  30  partially encircled by the curved face of the turbine blade). An insert  32  is then cut from a plate of HAYNES 214™ alloy to match the shape of the resulting void. The insert  32  is then joined by welding or brazing to the platform  30  as illustrated in  FIG. 5 . 
         [0031]    This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Technology Category: 2