Patent Abstract:
A method for casting an airfoil for a turbine engine is provided. The method includes forming a casting core to define a hollow portion in the airfoil and forming a print out region at one end of the casting core. The method also includes coupling the casting core to the print out region with at least one frusto-conical member to facilitate structurally supporting the casting core.

Full Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to turbine engines, and more specifically to turbine blades used with turbine engines. 
   At least some known turbine engines include a turbine that includes a plurality of rotor blades that extract rotational energy from fluid flow entering the turbine. Because the turbine is subjected to high temperatures, turbine components are cooled to reduce thermal stresses that may be induced by the high temperatures. Accordingly, at least some known rotating blades include hollow airfoils that are supplied cooling air through cooling circuits defined within the airfoil. More specifically, the airfoils include a cooling cavity bounded by sidewalls that define the cooling cavity. 
   To fabricate the cooling passages, at least some known turbine blades are cast using an internal core that forms the internal cooling passageways within the blades. Because of the relative large size of blades and/or vanes that may be used within industrial turbine engines, at least some known cores are reinforced to enable the core to withstand the injection pressures of the wax and the subsequent casting process. More specifically, a tip of at least some known casting cores is supported during the casting process by at least one rod that has a substantially constant diameter along its length. 
   When the casting process is complete, a print out coupled between the rod and the core is removed. An opening created by the rod may provide a channel for cooling the tip cap portion of the blade. In some known blade designs, the opening is sealed to facilitate cooling other portions of the blade. In such cases, the openings are sealed using known sealing techniques, such as welding or brazing. To facilitate forming a smaller diameter opening, some known castings use rods that have a diameter less than approximately 0.035 inches. However, as an overall size and/or weight of the casting is increased, a smaller diameter rod may not provide enough structural support to the core. 
   BRIEF SUMMARY OF THE INVENTION 
   In one aspect of the invention, a method for casting an airfoil for a turbine engine is provided. The method includes forming a casting core to define a hollow portion in the airfoil and forming a print out region at one end of the casting core. The method also includes coupling the casting core to the print out region with at least one frusto-conical member to facilitate structurally supporting the casting core. 
   In another aspect, an airfoil casting core for a turbine blade is provided. The casting core includes at least one of a leading edge path region, a center path region, and a trailing edge path region. The casting core also includes a core print region coupled to at least one of a leading edge path region, a center path region, and a trailing edge path region by at least one frusto-conical member. 
   In a further aspect of the invention, an airfoil core for use in casting an airfoil is provided. The airfoil core includes at least one of a leading edge path region, a center path region, and a trailing edge path region, extending between a core tip and a core root. The airfoil core also includes a print out region coupled to at least one of the core tip and the core root by at least one frusto-conical rod. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a perspective partial cut away view of an exemplary turbine; 
       FIG. 2  is a partial perspective view of an exemplary rotor assembly that may be used with the turbine shown in  FIG. 1 ; 
       FIG. 3  is a perspective view of an exemplary airfoil core that may be used to fabricate an airfoil used with the rotor assembly shown in  FIG. 2 ; and 
       FIG. 4  is an enlarged schematic view of a portion of the airfoil core shown in FIG.  3  and taken along area  4 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a schematic illustration of a gas turbine engine  10  including a generator  12 , a compressor  14 , a combustor  16  and a turbine  18 . Engine  10  has an inlet or upstream side  20 , an exhaust or downstream side  22 , and a gas fuel inlet  24 . The gas fuel passes through a gas control module  26  containing an isolation valve  27 , known as the stop-ratio valve (SRV) and a gas control valve (GCV)  28 . In one embodiment, engine  10  is a turbine engine commercially available from General Electric Power Systems, Schenectady, N.Y. 
   In operation, highly compressed air is delivered from compressor  14  to combustor  16 . Gas fuel is delivered to the combustor  16  through a plurality of fuel nozzles (not shown in  FIG. 1 ) and hot exhaust gas from combustor  16  is discharged through a turbine nozzle assembly (not shown in  FIG. 1 ) and is used to drive turbine  18 . Turbine  18 , in turn, drives compressor  14  and generator  12 . 
     FIG. 2  is a perspective view of a rotor assembly  40  that may be used with a turbine, such as turbine engine  10  (shown in FIG.  1 ). Assembly  40  includes a plurality of rotor buckets or blades  42  mounted to rotor disk  44 . In one embodiment, blades  42  form a high-pressure turbine rotor blade stage (not shown) of turbine engine  10 . 
   Rotor blades  42  extend radially outward from rotor disk  44 , and each blade  42  includes an airfoil  50 , a platform  52 , a shank  54 , and a dovetail  56 . Each airfoil  50  includes first sidewall  60  and a second sidewall  62 . First sidewall  60  is convex and defines a suction side of airfoil  50 , and second sidewall  62  is concave and defines a pressure side of airfoil  50 . Sidewalls  60  and  62  are joined at a leading edge  64  and at an axially-spaced trailing edge  65  of airfoil  50 . More specifically, airfoil trailing edge  65  is spaced chord-wise and downstream from airfoil leading edge  64 . A plurality of trailing edge slots  67  are formed in airfoil  50  to discharge cooling air over trailing edge  65 . The cooling air facilitates reducing the temperatures, thermal stresses, and strains experienced by trailing edge  65 . 
   First and second sidewalls  60  and  62 , respectively, extend longitudinally or radially outward in span from a blade root  68  positioned adjacent platform  52 , to an airfoil tip cap  70 . Airfoil tip cap  70  defines a radially outer boundary of an internal cooling chamber (not shown in FIG.  2 ). The cooling chamber is bounded within airfoil  50  between sidewalls  60  and  62 , and extends through platform  52  and through shank  54  and into dovetail  56 . More specifically, airfoil  50  includes an inner surface (not shown in  FIG. 2 ) and an outer surface  74 , and the cooling chamber is defined by the airfoil inner surface. 
   Platform  52  extends between airfoil  50  and shank  54  such that each airfoil  50  extends radially outward from each respective platform  52 . Shank  54  extends radially inwardly from platform  52  to dovetail  56 . Dovetail  56  extends radially inwardly from shank  54  and facilitates securing rotor blade  42  to rotor disk  44 . More specifically, each dovetail  56  includes at least one tang  80  that extends radially outwardly from dovetail  56  and facilitates mounting each dovetail  56  in a respective dovetail slot  82 . In the exemplary embodiment, dovetail  56  includes an upper pair of blade tangs  84 , and a lower pair of blade tangs  86 . 
     FIG. 3  shows an exemplary airfoil core  100  used in fabricating turbine blades  42  (shown in FIG.  2 ).  FIG. 4  is an enlarged schematic view of a portion of airfoil core  100  taken along area  4  (shown in FIG.  3 ). In one embodiment, core  100  is used to fabricate Stage  2  Bucket castings. Airfoil core  100  includes a leading edge path  102 , a center path  104 , a trailing edge path  106 , and a root cooling path  108 . Trailing edge path  106  has a plurality of fingers  110  extending from trailing edge path  106 . 
   During casting, leading edge path  102  and center path  104  form a first cooling passage (not shown), and a second cooling passage (not shown), respectively, in the resulting airfoil. Trailing edge path  106  forms a third cooling passage (not shown), and fingers  108  extending from trailing edge path  106 , form a plurality of trailing edge slots, such as slots  67  (shown in FIG.  2 ). In one embodiment, at least one of leading edge path  102 , center path  104 , and trailing edge path  106  includes an extension that forms a recess in the resulting airfoil cooling chamber. Thus, after a cooling passage is formed, the recess facilitates controlling airflow within the cooling cavity by forming an air flow restriction in the cooling chamber. 
   Airfoil core  100  also includes at least one “print out” region that facilitates handling of core  100 . More specifically, in the exemplary embodiment, airfoil core  100  includes a core tip print out region  112 . Core tip print out region  112  is coupled to at least one of leading edge path  102 , center path  104 , and trailing edge path  106  by at least one member  116 . First member  116  includes a first end  118  and a second end  120 . Specifically, first end  118  is coupled to at least one of leading edge path  102 , center path  104 , and trailing edge path  106  and second end  120  is coupled to core tip print out region  112 . Alternatively, core tip print out region  112  is coupled to root cooling path  108  by at least one member  116 . 
   Member  116  is frusto-conical and has a first end  118  that has a smaller diameter d 1  than a diameter d 2  at a second end  120 . Frusto-conical rod  116  reduces the area of weak mechanical strength in the regions of airfoil core  100  which exhibit break potential and subsequent loss of the casting. In another embodiment, member  116  can have any cross-sectional shape, such as a substantially square or triangular shape, with first end  118  having a smaller cross-sectional dimension than second end  120 . 
   Airfoil core  100  is fabricated by injecting a liquid ceramic and graphite slurry into core die (not shown). The slurry is heated to form a solid ceramic airfoil core  100 . The airfoil core  100  is suspended by core print out  112  in an airfoil die (not shown) and hot wax is injected into the airfoil die to surround the ceramic airfoil core. The hot wax solidifies and forms an airfoil (not shown in  FIG. 1 ) with the ceramic core suspended in the airfoil. 
   The wax airfoil with the ceramic core is then coated with multiple layers of ceramic and heated to remove the wax, thus forming a cavity shell having the shape of the airfoil. The shell is then cured in a heated furnace. Molten metal is then poured into the shell and thus forming a metal airfoil with the ceramic core remaining in place. The airfoil is then cooled, and the ceramic core is removed from the solidified casting by leaching or other means, leaving a casting having a hollow interior corresponding to the configuration of the airfoil core  100 . 
   The above-described airfoil core is cost-effective and highly reliable. The airfoil core includes at least one conical rod for attaching a core print out to the airfoil core. An area/diameter of the rods increases from the first end to the second end adding mechanical strength in regions of the airfoil core which exhibit break potential and subsequent loss of the casting. Additionally, the increased strength of the conical rod enables the conical rod to suspend a larger airfoil core. As a result, the geometry design of the conical rod, allows for the expansion of as cast feature geometry into the original casting design with an acceptable approach for manufacturing introduction, the conical rod facilitates maintaining material fatigue life and extending a useful life of the airfoil core during the casting process in a cost-effective and reliable manner. 
   Exemplary embodiments of airfoil casting cores are described above in detail. The systems are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each airfoil casting core component can also be used in combination with other airfoil casting cores and turbine components. 
   While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Technology Classification (CPC): 1