Patent Publication Number: US-7722327-B1

Title: Multiple vortex cooling circuit for a thin airfoil

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
   This application is related to co-pending U.S. patent application Ser. No. 11/642,258 filed on Dec. 20, 2006 by George Liang and entitled THIN TURBINE ROTOR BLADE WITH SINUSOIDAL FLOW COOLING CHANNELS and to co-pending U.S. patent application Ser. No. 11/642,255 filed on Dec. 20, 2006 by George Liang and entitled LARGE TAPERED ROTOR BLADE WITH NEAR WALL COOLING. 
   BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates generally to fluid reaction surfaces, and more specifically to thin walled turbine airfoils with cooling circuits. 
   2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98 
   In a gas turbine engine such as an industrial gas turbine engine, a turbine section includes a plurality of rotor blades that react with the hot gas flow passing through the turbine to produce mechanical work by rotating the rotor shaft. In an industrial gas turbine, four stages of rotor blades and stator vanes are used to extract the energy from the flow. As the inlet temperature to the turbine increases, the size of the fourth stage rotor blade also increases because the flow into the fourth stage has higher energy than previous lower temperature engines. These fourth stage rotor blades can be over 30 inches from platform to blade tip, and also have very large taper and twist in order to react with the flow. 
   With the higher gas flow temperature exposed to the fourth stage blade, internal air cooling is required in order to increase the life of the rotor blade. However, prior art methods of casting turbine blades having internal cooling circuits are not practical with these larger blades. Radial holes cannot be drilled into the blade because of the large amount of twist from the root to the tip. A straight hole cannot be placed within the blade. These large twist blades have large cross sectional areas in the lower span but have thin cross sectional areas in the upper span. Thus, the rotor blade in the upper span is very thin and thus not acceptable to casting processes of the prior art. Also, ceramic cores used for investment casting of these blades cannot be used in these long and highly twisted blades because the ceramic core would also have a long length with high twist. This produces a very brittle core which would un-twist when hanging within the mold used to cast the blade with the internal cooling passages. Core ties would break and result in improper positioning of the core within the mold. Defective blades would be cast that would also increase the overall cost of manufacturing the usable rotor blades. Therefore, there is a need in the art for producing a long rotor blade with thin airfoil walls with a cooling circuit to provide cooling for the blade. 
   It is an object of the present invention to provide a thin walled turbine airfoil with an internal cooling air circuit to provide cooling for the airfoil. 
   BRIEF SUMMARY OF THE INVENTION 
   A turbine airfoil with a thin wall cross sectional area, the airfoil having a cooling air supply channel positioned along the leading edge of the airfoil, and a plurality of chordwise extending cooling channels extending from the leading edge to the trailing edge, where each channel includes a plurality of vortex chambers connected in series by inlet metering holes. Cooling air from the leading edge supply channel flows through a metering hole and into a first vortex chamber, then through a second metering hole and into a second vortex chamber, and continues in this process until exiting through a trailing edge exit hole. The vortex chambers are circular in shape and include trip strips or a roughened surface on the inner surfaces to promote heat transfer to the cooling air flow. 

   
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       FIG. 1  shows a cross section side view of a turbine blade of the present invention. 
       FIG. 2  shows a detailed view of the vortex chambers used in the cooling circuit of the present invention. 
       FIG. 3  shows a detailed view of one of the vortex chambers from  FIG. 2 . 
       FIG. 4  shows a cross section top view of one of the cooling passages from  FIG. 1 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   The present invention is a turbine airfoil having thin wall cross section with an internal cooling air circuit to provide cooling for the airfoil. The airfoil can be a stator vane or a rotor blade. In the preferred embodiment, the airfoil is a rotor blade used in the fourth or last stage of a turbine in an industrial gas turbine engine. The fourth stage rotor blade includes an upper span portion with thin airfoil walls. However, the airfoil can include the cooling circuit of the present invention extending from the platform  14  to the blade tip as shown in  FIG. 1 . The blade includes a leading edge cooling supply channel  12  supplied with cooling air from the root channel  11 . A showerhead arrangement of leading edge film cooling holes can be used (not shown in  FIG. 1 ) connected to the leading edge cooling supply channel  12  to provide film cooling. Connected to the leading edge cooling supply channel  12  are a plurality of multiple vortex channels  13  extending along the chordwise length of the blade and ending at the trailing edge along exit holes. Adjacent multiple vortex channels  13  are offset (180 degrees out of phase) as shown in  FIGS. 1 and 2  in order to maximize the space these channels occupy. 
   A more detailed view of the multiple vortex channels  13  is shown in  FIG. 2  in which the vortex channel  13  includes an inlet metering hole  21  connected to the supply channel  12 , a first vortex chamber  22  immediately downstream from the inlet metering hole  21 , a second vortex chamber connected to the first vortex chamber through a metering hole, and additional vortex chambers connected in series through metering holes connecting adjacent vortex chambers. The last vortex chamber  22  is connected to an exit hole  24  that discharges the cooling air out through the trailing edge region of the blade. The exit holes  24  can be holes opening onto the trailing edge of the airfoil, or they can be slots opening onto the pressure side wall of the trailing edge region, or any other prior art trailing edge region discharging and cooling holes. 
   Each vortex chambers  22  has a circular cross sectional shape as shown in the figures, and is offset from the vortex chamber above or below in order to maximize the space for the cooling circuit by compacting as many of the vortex chambers into the space provided along the airfoil. The vortex chambers  22  can be any shape that will provide for a vortex flow within the chamber for the cooling air. Each vortex chamber  22  also includes trip strips  25  or a roughened surface  26  to promote the heat transfer from the metal to the cooling air flow. The space between the vortex channels  13  is solid material of the airfoil. 
     FIG. 3  shows a detailed view of one of the vortex chambers  22  used in the present invention. The inlet metering hole  21  delivers cooling air into the vortex chamber  22  which is formed by an upper wall  27  and a lower wall  28 . Trip strips  25  extend along the inner surface of the vortex chamber  22  to promote heat transfer to the cooling air flow. A cooling air exit hole  23  allows for the cooling air to flow out form the vortex chamber and into the next metering hole and vortex chamber within the channel  13 . As the cooling air flows through the inlet metering hole  21  and into the vortex chamber  22 , the cooling air will flow in the direction of the two arrows shown in  FIG. 2 . The trip strips  25  will force the cooling air to flow against the inner surface of the chamber  22  repeatedly. Then, the cooling air will flow toward the exit hole  23  and into the next chamber to repeat this process again. 
   The upper walls  27  and the lower walls  28  and the metering holes  21  extend from the pressure side wall to the suction side wall of the airfoil (as seen in  FIG. 4 ) and form the holes and chambers of the vortex cooling channel  13 . These  21  holes and chambers  22  are cast into the airfoil during the casting process. Ceramic core ties are used to form the channels  13  within the airfoil. 
     FIG. 4  shows a top view of one of the vortex channels  13  from the  FIG. 1  airfoil. The leading edge supply channel  12  is shown in the leading edge region of the blade. The first metering hole  21  connects the supply channel  12  to the vortex channel  13  that extends along the airfoil chordwise direction. The exit hole  24  connects the vortex channel  13  to the trailing edge of the blade to discharge the cooling air from the channel  13 . 
   The multiple vortex chambers can be designed based on airfoil hot gas side pressure distribution in both chordwise and spanwise directions. This is done by varying the metering holes at the inlet of each individual channel  13  as well as varying the metering flow orifice within each vortex channel. Also, each individual vortex chamber can be designed based on the airfoil local external heat load to achieve a desired local metal temperature level. This is achieved by varying the tangential velocity and pressure level within the vortex chamber with different pressure ratio across the cooling metering flow orifice. Trip strips in the vortex flow direction or two dimensional bumps built into the inner walls of the vortex chambers will further enhance the internal heat transfer performance. 
   In operation, the cooling air flow initiated from the airfoil leading edge radial cooling flow channel is bled off through a row of metering holes for the proper distribution of cooling air into each individual vortex flow channel. The cooling flow can be distributed based on the airfoil spanwise metal temperature requirement. The inter-linked vortex chambers provide a long flow path for the coolant parallel to the chordwise direction of the gas path pressure and temperature profile. The cooling flow can be distributed based on the airfoil chordwise metal temperature requirement by varying the inter-linked metering orifice. The vortex chambers create a high overall coolant velocity and high heat transfer while the long flow path yields high overall cooling effectiveness. The injection process for the cooling air repeats throughout the entire inter-linked vortex chambers and then discharges the coolant from the airfoil trailing edge through multiple cooling holes or slots.