Abstract:
A turbine airfoil or a substrate exposed to a high temperature environment having a plurality of modular formed cooling circuits with diffusion chambers and cooling holes for each module. Each module includes diffusion chambers and transpiration cooling holes and is placed on the airfoil substrate and a refractory material is formed over the modules. The modules are then leached away leaving the diffusion chambers and cooling holes formed between the substrate and the refractory coating.

Description:
BACKGROUND OF THE INVENTION 
     Field of the Invention 
     The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine airfoil with film cooling holes. 
     Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98 
     A gas turbine engine includes a turbine section that has a plurality of stages of stator vanes and rotor blades reacting to a high temperature gas flow passing through the turbine to convert the chemical energy from combustion into mechanical energy by rotating the turbine shaft. The efficiency of the turbine, and therefore of the engine, can be increased by increasing the hot gas flow that enters the turbine. 
     To allow for higher turbine entrance temperatures, the upper stage vanes and blades are made from exotic nickel alloys that can withstand very high temperatures and have complex internal cooling air passages to provide cooling to these airfoils. A thermal barrier coating (TBC) is also applied to the airfoil surfaces exposed to the hot gas flow in order to provide further protection from the heat. A TBC is typically made from a ceramic material. Also, the TBC is typically applied after the film cooling holes have been drilled into the airfoil surface to provide for the film cooling. These film cooling holes are limited to the diameter because of the drilling process. Thicker TBC layers have been proposed to provide more protection to the airfoil substrate from the high temperature gas flow. As the TBC gets thicker, the thermal stresses developed in the TBC will tend to cause spalling. 
     In some prior art applications, a thin refractory coating is used in the turbine airfoil cooling design to provide a protective coating for the turbine airfoil and thus reduce the cooling flow consumption and improve turbine efficiency. The refractory coating is made of a material that is very expensive. The refractory coating is made so thin that cooling holes are not used in the coating because the hole length to diameter ratio cannot be larger than 2, which is required for cooling holes. Because the thin refractory coating is so thin—in the order of 2 to 4 mils (one mil is 0.001 inch)—the cooling hole would have to be at least 4 to 8 mils in diameter to maintain the hole ratio of 2 to 1. 
     As the turbine inlet temperature increases, the cooling flow demand for cooling the airfoil increases as well, and as a result the turbine efficiency is reduced. One alternative way for reducing the cooling air consumption while increasing the turbine inlet temperature for higher turbine efficiency is to use transpiration film cooling on the cooled thicker layer of the protective coating in order to reduce the heat load on the airfoil. 
     It is therefore an object of the present invention to provide for an improved high temperature resistant coating applied to a turbine airfoil. 
     It is another object of the present invention to provide for a high temperature resistant coating with smaller diameter film cooling holes. 
     It is another object of the present invention to provide for a refractory material coating on a turbine airfoil with smaller diameter cooling holes. 
     It is another object of the present invention to provide for a process of forming small diameter cooling holes in a refractory material using modules that form the holes. 
     BRIEF SUMMARY OF THE INVENTION 
     The present invention is a turbine airfoil with a refractory coating applied to the surface in which the coating includes small diameter cooling holes formed therein. The cooling holes are formed by placing a module of a leachable ceramic material into trenches already formed within the surface of the airfoil substrate. The module includes an array of trusses extending chordwise and spanwise, each truss having a plurality of hole forming extensions to form the cooling holes. The module is placed within the trenches formed on the blade substrate, a refractory coating is applied over the module, and the module is leached away leaving the cooling holes and the diffusion openings formed within the refractory coating. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         FIG. 1  shows a top view of a cross section of a turbine blade with the cooling holes of the present invention. 
         FIG. 2  shows a close-up view of the cooling holes of  FIG. 1 . 
         FIG. 3  shows a side view of one of the modules used to form the cooling holes of the present invention. 
         FIG. 4  shows a top view of the module of  FIG. 3 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The present invention is a turbine airfoil, such as a rotor blade or a stator vane, used in a gas turbine engine, in which the turbine airfoil includes a thick refractory coating to provide protection form a higher external gas flow temperature than would a typical ceramic TBC used on the airfoil. The airfoil  10  in the present invention is shown in  FIG. 1  and has a leading edge and a trailing edge, and a pressure side and a suction side. Internal cooling air supply channels  11  are formed within the airfoil walls and are separated by ribs  12  that also reinforce the airfoil walls. Exit cooling holes  16  are located in the trailing edge of the blade  10  and discharge cooling air from the downstream channel of the blade. Cooling holes  13  are formed in the main wall or substrate  14  of the blade and connect the internal cooling air supply channels to the cooling holes of the present invention best described in  FIG. 2 . 
       FIG. 2  shows the details of the small cooling holes formed in the coating applied to the outer surface of the airfoil on the substrate  14 . Cooling supply holes  13  are formed in the substrate by any of the well known processes such as drilling. The cooling holes  13  function as metering holes for the individual cooling holes  22  that are formed within the coating  21 . Each cooling supply hole  13  ends into a diffusion chamber  13  that is also formed within the substrate  14 . The cooling holes  22  connect the diffusion chamber  23  to the exterior surface of the coating  21 . 
     The cooling holes  22  are formed into the coating  21  by a process that uses a plurality of modules or mini cores  31  shown in  FIGS. 3 and 4  that form a number of the cooling holes  22  in the coating  21 . The module or mini core  31  is rectangular in shape and includes core trusses that extend in the vertical and horizontal directions as seen in  FIG. 4 . Two horizontal trusses  33  and three vertical trusses  32  form a rectangular shaped module with two openings  34  inside. Cooling hole shaped pins  22  extend from the flat surface of the trusses the length equal to about that of the thickness of the coating to be applied. One metering hole  13  would supply cooling air to the diffusion chamber formed by one of the vertical trusses  32  of the module  31  shown in  FIG. 4 . Thus, the module  31  shown in  FIG. 4  would be associated with three metering holes  13  with one metering hole for each of the three vertical trusses  32 . The substrate  14  has an arrangement of trenches machined or cast into the blade wall and having a spherical cross sectional shape as seen in  FIG. 2 . The size and shape of the trenches formed in the substrate  14  will be the same as the module or min core  31 , since the module will be placed into the trenches before the coating is applied. The module or mini core  31  is made of a leachable ceramic material of the kind used to form hollow turbine airfoils with internal cooling passages using the lost wax process. 
     To produce the turbine blade (or stator vane), the blade is cast and the trenches that will form the diffusion chamber  23  will be machined into the blade substrate or cast with the blade. The blade substrate thus has an array of trenches formed in the shape of the module  31  shown in  FIG. 4  in which three vertical or primary trenches extend between two horizontal or secondary trenches with three metering holes  13  drilled in the substrate at about the midpoint of each of the three vertical or primary trenches. The primary trenches include a metering hole connected to the trench. The secondary trenches connect two adjacent primary trenches. The metering holes  13  for each of the trenches that form the diffusion chamber  23  are drilled into the blade to connect the trench to the cooling supply channel  11 . Primary diffusion chambers are formed from the vertical or primary trenches, and secondary diffusion chambers are formed from the horizontal or secondary trenches. The modules  31  are placed within the trenches such that the outer substrate surface and the top surface of the modules are flush. The cooling hole forming pins  35  extend outward in the size and length of the cooling holes that will be formed later. The coating  21  is applied to the substrate with all of the modules  31  in place. When the coating is dried, the ceramic material that forms the modules is leached out. With the ceramic material leached out, the diffusion chamber  23  and the cooling hole  22  remains and forms the cooling air passage from the metering hole  13  to the opening on the surface of the coating  21 . In the present embodiment, the coating is a refractory material such as Iridium or Rhodium that can withstand higher gas flow temperatures than the typical ceramic thermal barrier coatings. Thus, a turbine airfoil with the refractory coating and the small diameter cooling holes can produce transpiration cooling of the airfoil that will allow for exposure to the higher gas flow temperatures. This will allow for a gas turbine engine with a higher turbine inlet temperature, which will provide for higher engine efficiency. Also, because of the small cooling holes that will allow transpiration cooling for the refractory coating, the refractory coating can be thicker than a non-cooled refractory coating. The thicker refractory coating will also provide for additional protection to the blade substrate from the extreme gas flow temperature. In the present invention, the refractory coating has a thickness of about 0.005 inches to 0.008 inches. With a thickness in the smaller range of 0.005 inches, to keep a cooling hole length to diameter ratio of 2, the diameter of the cooling hole would have to be 0.0025 inches. The process of forming cooling holes of the present invention is capable of forming cooling holes of this small diameter. 
       FIG. 4  shows the grid of trench forming trusses extending in a vertical and horizontal direction with openings  34  formed between the trusses that are in the shame shape and size as the trenches on the blade substrate. The present invention shows three vertical trenches and two horizontal trenches. However, this could be rotated 90 degrees without departing from the spirit and scope of the present invention. Also, instead of the trusses forming a rectangular array or grid, a triangular array or grid can be used. Three trenches in which the two side trenches could extend at about 30 degrees from the normal while the base trench would connect the two. The metering holes would be associated with the longer side trenches, with the base trench acting as the secondary diffuser connecting the two primary diffusers together. 
       FIG. 1  shows a portion of the airfoil wall to include the cooling holes with diffusion chambers as described in the present invention above for the purpose of clarity. However, the entire airfoil wall from the leading edge to the trailing edge along the pressure side and the suction side includes the cooling holes.