Abstract:
A method for making a turbine airfoil includes providing a mold core and an outer shell which cooperatively define a cavity in the shape of a hollow airfoil having an outer wall, a root, and a tip. A tip portion of the core extends completely through the portion of the cavity defining the tip of the airfoil. The core is restrained to prevent movement between the core and outer shell. Molten metal is introduced into the cavity and solidified to form an airfoil having at least one outer wall which defines an open tip and a hollow interior. A metallic tip cap is formed on the outer wall which substantially closes off the open tip. The tip cap may be formed by packing the airfoil with metallic powder; and laser sintering the exposed powder so as to form a tip cap which is metallurgically bonded to the outer wall.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to the manufacture and repair of gas turbine engine components and more particularly to the production of tip cap structures for turbine airfoils. 
         [0002]    Cast turbine airfoils for advanced gas turbine engines have internal features that challenge the capability of current casting technologies. The castings require complex ceramic cores to form the internal features and these cores are fragile during the casting process. The result is that casting yields of 50 percent to 70 percent are not uncommon. The 30 percent to 50 percent casting scrap factors into the cost of the useable castings. 
         [0003]    The issue is compounded by exotic alloys such as single crystal materials that drive up the cost to cast a part and thus drive up the cost caused by scrapping hardware. If a mere 5 percent to 10 percent casting yield improvement can be achieved, the impact to each gas turbine engine is in the millions of dollars per year, based on volume. 
         [0004]    One basic casting limitation is that the ceramic core that forms the internal structure of the airfoil can only be secured by the lower (i.e. root) portion with the majority of the core “floating” within the casting wax form. The forces of the molten metal and thermally induced forces during the cooling and solidification cycle result in movement and/or breakage of the ceramic core (referred to as “core shift”). The motion can be such that the cast component no longer meets drawing requirements, for example by violating minimum casting wall thicknesses. If the core fractures during the process this will also cause the component to fail requirements. 
         [0005]    Cast turbine airfoils often include a tip cap and/or a “squealer tip”. A tip cap is a relatively small extension, having a cross-sectional shape conforming to that of the turbine blade, either integral with or mounted on the radially outer end of the turbine blade. The utilization of tip caps on turbine blades can effectively reduce the disadvantageous effects of rubbing between turbine blades and the shroud. However, prior art methods of casting turbine blades include the tip cap as an integral part, which exacerbates the core shift problems described above. It is also known to manufacture tip caps separately and then attach them to airfoils, but this requires welding processes, which require costly tooling and are relatively slow. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0006]    These and other shortcomings of the prior art are addressed by the present invention, which provides a method for supporting an airfoil core during casting, and for forming a tip structure on an airfoil while maintaining the metallurgical integrity of the finished component. 
         [0007]    According to an aspect of the invention, a method for making a turbine airfoil tip structure includes: (a) providing a turbine airfoil having an outer wall which defines a hollow interior, wherein a tip of the airfoil is open to the hollow interior; (b) packing the interior of the airfoil with metallic powder; and (c)using laser energy, sintering an exposed layer of the powder so as to form a tip cap which is metallurgically bonded to the outer wall. 
         [0008]    According to another aspect of the invention, a method for making a turbine airfoil includes: (a) providing a mold having: (i) a core; and (ii) an outer shell surrounding the core such that the core and the outer shell cooperatively define a cavity in the shape of a hollow airfoil having an outer wall, a root, and a tip; (iii) wherein at least a tip portion of the core extends completely through the portion of the cavity defining the tip of the airfoil; (b) restraining the tip portion of the core so as to prevent relative movement between the core and the outer shell; (c) introducing molten metal alloy into the cavity and surrounding the core; (d) solidifying the alloy to form an airfoil having at least one outer wall which defines an open tip and a hollow interior; (e) removing the mold and core so as to expose the airfoil; and (f) forming a metallic tip cap on the outer wall which substantially closes off the open tip. 
         [0009]    According to another aspect of the invention, a method is provided for repairing a turbine airfoil which includes an outer wall defining a hollow interior, wherein a tip of the airfoil is closed off by a tip cap. The method includes: (a) removing the tip cap from the airfoil; (b) packing the interior of the airfoil with metallic powder; and (c) using laser energy, sintering an exposed layer of the powder so as to form a new tip cap which is metallurgically bonded to the outer wall. 
         [0010]    According to another aspect of the invention, a method is provided for repairing a turbine airfoil which includes an outer wall defining a hollow interior, wherein a tip of the airfoil is closed off by a tip cap. The method includes: (a) removing the tip cap from the airfoil; (b) positioning a metallic plate-like member in contact with the outer wall at the tip; (c) using laser energy, heating the plate-like member and fusing it to the outer wall to form a tip cap; and (d) applying a metallic, radially-outwardly-extending tip wall to the tip cap. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0011]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0012]      FIG. 1  is a perspective view of an exemplary turbine blade constructed in accordance with an aspect of the present invention; 
           [0013]      FIG. 2  is a perspective view of a prior art mold core used in casting a blade; 
           [0014]      FIG. 3A  is a perspective view of a mold core constructed in accordance with an aspect of the present invention; 
           [0015]      FIG. 3B  is a perspective view of an alternative mold core; 
           [0016]      FIG. 4  is a schematic, partial cross-sectional view of an assembled mold constructed according to an aspect of the present invention; 
           [0017]      FIG. 5  is a schematic, cross-sectional view of the mold of  FIG. 4  with a portion of a blade casting therein; 
           [0018]      FIG. 6  is a schematic, partial cross-sectional view of an as-cast turbine blade, which includes an open tip; 
           [0019]      FIG. 7  is a schematic, partial cross-sectional view of the turbine blade of  FIG. 6 , with a tip plate applied thereto; 
           [0020]      FIG. 8  is another view of the as-cast turbine blade with a tip wall formed thereon; 
           [0021]      FIG. 9  is a schematic view of a laser welding apparatus; 
           [0022]      FIG. 10  is a schematic, partial cross-sectional view of an as-cast turbine blade, which includes an open tip; 
           [0023]      FIG. 11  is a schematic, partial cross-sectional view of the turbine blade of  FIG. 10 , with a tip shim applied thereto; 
           [0024]      FIG. 12  is a another view of the turbine blade of  FIG. 11  with a tip wall formed thereon; 
           [0025]      FIG. 13  is a schematic, partial cross-sectional view of an as-cast turbine blade, which includes an open tip, filled with metallic powder; 
           [0026]      FIG. 14  is a view of the turbine blade of  FIG. 13  after a laser sintering process; and 
           [0027]      FIG. 15  is a another view of the turbine blade of  FIG. 14  with a tip wall formed thereon. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0028]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates an exemplary turbine blade  10 . The turbine blade  10  includes a conventional dovetail  12 , which may have any suitable form including tangs that engage complementary tangs of a dovetail slot in a rotor disk (not shown) for radially retaining the blade  10  to the disk as it rotates during operation. A blade shank  14  extends radially upwardly from the dovetail  12  and terminates in a platform  16  that projects laterally outwardly from and surrounds the shank  14 . A hollow airfoil  18  extends radially outwardly from the platform  16 . The airfoil  18  has an outer wall  19  comprising concave pressure side outer wall  20  and a convex suction side outer wall  22  joined together at a leading edge  24  and at a trailing edge  26 . The trailing edge  26  may incorporate trailing edge cooling passages such as the illustrated holes  28 . The airfoil  18  has a root  30  and a tip  32 . The radially outermost portion of the airfoil  18  defines a peripheral tip wall  34 , sometimes referred to as a “squealer tip”. A tip cap  36  closes off the interior of the airfoil  18  and lies recessed a small distance radially inward from the tip  32 . The airfoil  18  may take any configuration suitable for extracting energy from the hot gas stream and causing rotation of the rotor disk. The blade  10  is preferably formed as a one-piece casting of a suitable “superalloy” of a known type, such as a nickel-based superalloy (e.g. Rene 80, Rene 142, Rene N4, Rene N5) which has acceptable strength at the elevated temperatures of operation in a gas turbine engine. 
         [0029]    The interior of the turbine blade  10  is mostly hollow and includes a number of internal cooling features of a known type, such as walls defining serpentine passages, ribs, turbulence promoters (“turbulators”), etc. While the turbine blade  10  is a high pressure turbine blade, the principles of the present invention are applicable to any type of airfoil having a hollow interior. 
         [0030]    Components such as the turbine blade  10  are manufactured using a known investment casting process. The method includes shaping the turbine blade in wax by enveloping a conventional alumina or silica based ceramic core which defines internal coolant passages. The wax assembly then undergoes a series of dips in liquid ceramic solution. The part is allowed to dry after each dip, forming a hard outer shell, typically a conventional zirconia based ceramic shell. After all dips are complete, and the wax assembly is encased by several layers of hardened ceramic shell, the assembly is placed in a furnace where the wax in the shell is melted out. 
         [0031]    After wax removal, the mold comprises the internal ceramic core surrounded by the outer ceramic shell. The cavity between the core and the outer shell defines the component and its interior features. The mold is again placed in the furnace, and liquid metal is poured into an opening at the top of the mold. The molten metal enters the space between the ceramic core and the ceramic shell, previously filled by the wax. After the metal is allowed to cool and solidify, the external shell is broken and removed, exposing the casting which has taken the shape of the cavity created by removal of the wax, and which encases the internal ceramic core. The casting is then placed in a leaching tank, where the core is dissolved. The component now has the shape of the wax form, and an internal cavity which was previously filled by the internal ceramic core. 
         [0032]    The relative thermal growths of the ceramic outer shell and the ceramic core material are different, so that after the metal has been poured and is allowed to cool, the relative shrinking of the shell and core components are different. This can cause varying wall thicknesses at areas of the metal nozzle part where one side of the wall is defined by the external shell, and the other side of the wall is engaged by the internal core. 
         [0033]      FIG. 2  illustrates a prior art core  38 . Relatively large cross-sectional areas of the root portion “R” of the core  38  define the shank  14  of the turbine blade  10 . This root portion R is configured so as to be rigidly supported during the casting process. However, a portion “A” which defines the airfoil  18  is normally unsupported, and is a substantial distance from the root portion R. Furthermore, a gap “G” is incorporated in the core  38  so as to define an integral tip cap  36  ( FIG. 1 ). This greatly reduces the ability to support the core  38 . 
         [0034]    In prior art processes, the core  38  is typically allowed to “float” and may thus shift its position relative to the outer shell during the casting process. This can cause the walls of a component such as an airfoil to be less than a required minimum thickness. The outer end of the core  38  can be supported by structures such as quartz rods. However, these are prone to breakage and can cause wall shift. 
         [0035]    To avoid core such shift, the turbine blade  10  is cast by a modification of the above process, which incorporates an “open-top” casting configuration, and an extended core.  FIG. 3A  illustrates an exemplary core  40  constructed according to an aspect of the present invention. It too includes a root portion “R′” which defines the shank  14  of the turbine blade  10 , and an airfoil portion “A′” which defines the airfoil  18 . In contrast to prior art configurations, the gap “G” is eliminated, and a substantial percentage of the cross-sectional area of the core  40  extends to the radially outer end of the airfoil  18 , or beyond. 
         [0036]      FIG. 3B  illustrates a slightly different core  40 ′ which has a gap “G′” that defines an airfoil tip cap, similar to that of the prior art core  38 . However, it also includes a bridge  41  of substantial cross-sectional area extending from the main body  43  of the core  40 ′ to a horizontal member  45  that would otherwise be unsupported. 
         [0037]      FIG. 4  is a schematic pre-casting view of a portion of the core  40  and a surrounding outer shell  42 . The space between the core  40  and the outer shell  42  defines the airfoil  18  and is filled with wax  44 . 
         [0038]    The extension of the core  40  permits the core  40  to be rigidly supported during the casting process, such that little or no relative motion will occur between the core  40  and the outer shell  42 . In the illustrated example, tie rods  46  (e.g. ceramic or metal) are provided which extend between the core  40  and the outer shell  42 . The core  40  could also be supported by attachment to external tooling which is shown schematically at  48 . 
         [0039]    An after-casting view shown in  FIG. 5 . After the casting process is complete, the core  40  is removed by caustic leaching using known techniques. However, by exposing up more area at the tip, the leaching process has a larger area of the core  40  to act on. The leaching process and core cleanout is therefore expected to be measured in hours instead of days for further cost reduction. In one example, it is estimated that the leaching process may be reduced from about seven to nine 10-hour-long cycles using the prior art configuration, to about two 10-hour-long cycles. 
         [0040]      FIG. 6  schematically illustrates the airfoil  18  after casting and removal of the outer shell  42  and the core  40 . At this point the radially outer end of the airfoil  18  is open and must be closed to provide a usable component. 
         [0041]      FIGS. 7 and 8  illustrate one method by which the outer end of the airfoil  18  may be closed. In a first step, a tip plate  50  is placed on the outer end of the airfoil  18 , in contact with the outer wall  19 . The tip plate  50  may be formed by casting, rolling, stamping, etc. In the illustrated example it is about 0.5 mm (0.02 in.) to about 1 mm (0.040 in.) thick. The tip plate  50  is constructed from a material that is metallurgically compatible with the airfoil body  50 . Representative examples of suitable materials include known “superalloys”, e.g. Rene 80, Rene 142, Rene N4, Rene N5. The tip plate  50  may be formed with a selected crystalline microstructure, such as directionally solidified (“DS”) or single-crystal (“SX”). 
         [0042]    The tip plate  50  is bonded to the outer wall  19  by laser welding. An example of a suitable apparatus for laser welding is disclosed in U.S. Pat. No. 5,622,638 to Schell et al., assigned to the assignee of this invention, and is schematically illustrated in  FIG. 9 . The apparatus includes a laser  52 , an enclosed beam delivery conduit  54 , laser focusing optics  56 , a part positioning system  58 , a vision system  60  for part location and laser path control, an optional preheating apparatus such as a preheat box or an induction coil (not shown), and a powder feed system  62  with a powder tube  64 . The working and coordination of the individual parts of the apparatus are controlled through a computerized system controller  66 . 
         [0043]    Using the apparatus of  FIG. 9 , laser energy is directed at the tip plate  50  from the end or the peripheral edges (see arrows “W” in  FIG. 7 ) so as to produce a through-weld and fuse the outer periphery of the tip plate  50  to the outer wall  19 . By using this process, the crystal microstructure of the airfoil  18  and the tip plate  50  (e.g. DS or SX) can be substantially maintained. 
         [0044]    Next, the tip wall  34  is formed through a freeform laser fabrication process. Using the apparatus shown in  FIG. 9 , molten alloy powder is deposited on the tip plate  50  in one or more passes. The powder alloy composition may be substantially the same as that of the basic turbine blade  10 . Alternatively, the powder alloy composition may be a material with better resistance to corrosion at high temperatures than the base alloy of the airfoil (e.g. nickel-based alloys such as RENE 142, RENE 195). This process, sometimes referred to as “reverse machining”, produces a structure that is metallurgically bonded to the tip plate  50 . With proper control of the process parameters, this process can produce the same microstructure in tip wall  34  (e.g. directionally solidified or single crystal) as that of the turbine blade  10 . For example, a continuous wave beam of about 300 W to about 1000 W power may be used, with a traverse rate of about 0.25 cm/s (0.1 in./s) to about 0.76 cm/s (0.3 in./s) and preferably about 0.44 cm/s (0.175 in./s) to about 0.51 cm/s (0.200 in./s). About 10 passes result in a tip wall  34  (see  FIG. 8 ) of a suitable height and a near-net shape. As used herein, the term “near-net” refers to a structure that does require substantial additional machining processes in order to arrive at a finished part. Once the freeform fabrication process is finished, the tip wall  34  may be further formed by known processes of machining, grinding, coating, etc. 
         [0045]      FIGS. 10-12  depict another method by which the outer end of an airfoil  18 ′ (similar to the airfoil  18 ) may be closed. In a first step, a shim  50 ′ is placed on the outer end of the airfoil  18 ′, in contact with the outer wall  19 ′. The shim  50 ′ is similar in construction and material to the tip plate  50 , but is substantially thinner, about as thick as a typical layer of laser-deposited material. In the illustrated example it is about 0.1 mm (0.004 in.) to about 0.3 mm (0.012 in.) thick, preferably about 0.2 mm (0.008 in.) to about 0.25 mm (0.010 in.) thick. 
         [0046]    The shim  50 ′ is bonded to the outer wall  19 ′ by laser welding as described above, so as to produce a through-weld and fuse the outer periphery of the shim  50 ′ to the outer wall  19 ′. 
         [0047]    Next, a freeform laser fabrication process is used to deposit material over the shim  50 ′, defining a tip cap  36 ′, and to build up a tip cap plenum wall  34 ′ of a suitable height and a near-net shape. Again, this may be done using the apparatus shown in  FIG. 9 . Once the freeform fabrication process is finished, the tip wall  34 ′ and tip cap  36 ′ may be further formed by known processes of machining, grinding, coating, etc. 
         [0048]      FIGS. 13-15  depict an airfoil  18 ″ formed by another alternative method.  FIG. 10  illustrates the airfoil  18 ″ in the as-cast condition with an outer wall  19 ″. The interior of the airfoil  18 ″ is filled with a suitable metallic alloy powder  68 , such as the powder described above, which is scraped flush or otherwise leveled with the outer end of the outer wall  19 ″. Then, using the apparatus illustrated in  FIG. 9 , the powder  68  is sintered together and bonded to the outer wall  19 ″ by directing laser energy at it, shown schematically at arrow “L”. The exact process parameters may vary to suit a specific application. In the illustrated example, a short pulsed infrared laser beam is used, with an average power of 1-100 W, pulse frequency of 1 Hz to 200 kHz. The translation speed or scanning speed, if the laser beam is used with a scanner, is approximately 5 mm/s (0.197 in./s) to about 500 mm/s (19.7 in./s).  FIG. 11  depicts the airfoil  18 ″ after the powder  68  has been sintered into a completed tip cap  36 ″ and the excess powder  68  removed (for example by blowing or vacuuming through passages in the base of the airfoil  18 ″). With proper control of the process parameters, this process can produce the same microstructure in the tip cap  36 ″ (e.g. directionally solidified or single crystal) as that of the airfoil  18 ″. 
         [0049]    Once the tip cap  36 ″ is formed, a tip wall  34 ″ is formed on top of the tip cap  36 ″, using the freeform laser fabrication process described above (see. Once the freeform fabrication process is finished, the tip cap  36 ″ and tip wall  34 ″ may be further formed by known processes of machining, grinding, coating, etc. 
         [0050]    The freeform laser fabrication method described above may be used to create a tip structure on any type of airfoil. For example, while it is described above in the context of new production, the method could also be used to create upgrade or repair an existing airfoil by first cutting off the existing tip structure to arrive at a partial airfoil similar to that shown in  FIG. 6 . 
         [0051]    The methods described above will significantly improve turbine airfoil casting yields by reducing or eliminating core shift. An additional benefit of this approach is the ability to precisely control the component wall thickness in various locations, thereby reducing scarfing (overdrilling) losses in downstream cooling hole drilling operations. Using this method the blades can be designed and cast with the wall thickness necessary only to provide required strength and without allowances for casting variation. This could potentially result in reduced component weight, and subsequently lower the overall weight of the engine and improve specific fuel consumption during operation of the engine. Furthermore, the freeform method of making the tip structure will lower production costs by eliminating complex and expensive tooling. 
         [0052]    The foregoing has described a method for making gas turbine engine airfoils and tip structures. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation.