Patent Publication Number: US-7717675-B1

Title: Turbine airfoil with a near wall mini serpentine cooling circuit

Description:
BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with a cooling circuit. 
   2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98 
   In a gas turbine engine, especially in an industrial gas turbine engine, a turbine section includes multiple stages of stator or guide vanes and rotor blades to extract mechanical energy from a hot gas flow passing through the turbine. Increasing the turbine inlet temperature can increase the turbine efficiency, and therefore the engine efficiency. However, the maximum turbine inlet temperature is limited to the material characteristics of the turbine airfoils, especially the first stage guide vanes and rotor blades, since these airfoils are exposed to the highest temperature. 
   In order to allow for a higher gas flow temperature, the turbine airfoils include complex internal cooling circuits to provide the maximum amount of cooling for the airfoil while making use of the minimum amount of cooling air in order to maximize the efficiency of the turbine and therefore the engine. In a prior art turbine blade with near wall cooling, the airfoil main body includes radial flow channel plus re-supply holes in conjunction with film discharge cooling holes from the near wall channel. In this prior art airfoil, spanwise (the direction from root to tip) and chord wise (the direction from leading edge to trailing edge) cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, a single radial channel flow is not the best method of utilizing cooling air since it results in low convective cooling. 
   It is therefore an object of the present invention to provide for a turbine airfoil with a cooling circuit that will reduce the main body metal temperature and therefore reduce the cooling flow requirement and improve the turbine efficiency. 
   BRIEF SUMMARY OF THE INVENTION 
   A turbine blade with a near wall mini serpentine flow cooling circuit for the airfoil main body is used to reduce the airfoil main body metal temperature. The mini serpentine cooling circuit is constructed of a plurality of small module formations of serpentine cooling passages arranged along the pressure and suction side walls in an array from the leading edge to the trailing edge. Each module can have a triple 3-pass near wall serpentine flow circuit with a feed hole on the forward end and a collection cavity cooling air return hole on the aft end of the circuit. In an alternate embodiment, a row of multi-film cooling holes can be used in the passage connecting adjacent serpentine passages within each module. Each individual module can be designed based on the airfoil gas side pressure distribution in both the chord wise and the spanwise directions. Also, each individual module can be designed based on the airfoil local external heat load to achieve a desired local metal temperature. 

   
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       FIG. 1  shows a schematic view of a turbine blade with the near wall mini serpentine cooling modules of the present invention. 
       FIG. 2  shows a cross section top view of the near wall mini serpentine cooling circuit of the  FIG. 1  turbine blade. 
       FIG. 3   a  shows a detailed view of a triple 3-pass near wall serpentine cooling circuit of the present invention. 
       FIG. 3   b  shows a detailed view of a second embodiment of the present invention with rows of film cooling holes. 
       FIG. 4  shows a third embodiment of the near wall mini serpentine flow cooling channel of the present invention. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   The present invention is a turbine blade used in an industrial gas turbine engine with a near wall mini serpentine flow cooling circuit arranged in modules along the airfoil walls to reduce the main body metal temperature.  FIG. 1  shows the turbine blade of the present invention. However, the cooling circuits of the present invention can also be used in an aero gas turbine engine, or in stator vanes of both an industrial and an aero gas turbine engine.  FIG. 1  shows the turbine blade with a pressure side airfoil wall with a plurality of the near wall mini serpentine cooling modules arranged extending from the blade platform to the tip, and from the leading edge region to the trailing edge region. 
     FIG. 2  shows a cross section view of the turbine blade of  FIG. 1  with a leading edge having a showerhead arrangement of film cooling holes  11  connected to a leading edge cooling supply cavity  12 . Located aft of the cooling supply cavity  12  is a number of cooling air discharge cavities  13  each separated by a rib. In the embodiment shown in  FIG. 2 , three cooling supply cavities  12  each with two cooling discharge cavities are arranged in the chord wise direction and extend from the leading edge to the trailing edge region of the airfoil. A near wall mini serpentine flow cooling channel  15  is located on both sides of the airfoil and between the supply cavity  12  and the aft most discharge cavity  13  as seen in  FIG. 2 . Cooling holes connect the mini serpentine channels  15  to the supply cavity  12  and each of the discharge cavities  13 . The aft most discharge cavity is connected to a film cooling hole on one or both sides of the airfoil to discharge cooling air to the airfoil external surface. Suction side film cooling holes  25  and pressure side film cooling holes  26  are shown in  FIG. 2 . The aft most cooling discharge cavity  13  is connected to a trailing edge cooling slot  18  to discharge cooling air out the trailing edge of the airfoil. 
     FIG. 3   a  shows a detailed view of a first embodiment of the near wall mini serpentine flow cooling channel used in the blade of  FIG. 2 .  FIG. 3  shows four of the mini serpentine flow channels  15  each having a cooling air feed hole  21  that is connected to a cooling supply cavity  12  and a cooling air return hole  22  that is connected to a cooling air discharge cavity  13 . As seen in  FIG. 3   a , the mini serpentine flow channel includes a triple 3-pass near wall mini serpentine flow channel with a first 3-pass serpentine flow channel  31  having three legs extending in the airfoil chord wise direction, a second 3-pass serpentine flow channel  32  and a third 3-pass serpentine flow channel  33  each connected by a airfoil spanwise channel  41 . The last spanwise channel  41  connects the third 3-pass serpentine flow channel  33  to the cooling air return hole  22 . 
   The second embodiment is shown in  FIG. 3   b  and is similar to the first embodiment of  FIG. 3   a  in which three 3-pass serpentine flow channels  31  through  33  are arranged along the airfoil wall in the chord wise direction with a cooling air feed hole  21  connected to a cooling supply cavity  12 . The  FIG. 3   b  embodiment eliminates the cooling air return holes  22  and includes a row of film cooling holes  17  in each of the spanwise channels  41 . On the suction side wall, the film cooling holes would be suction side film cooling holes  16  to discharge onto the suction side wall. 
   In the two embodiments of  FIGS. 3   a  and  3   b , cooling air is supplied through the cooling supply cavity  12 , metered through the cooling feed hole  21  and into the axial mini serpentine flow module  15 . Cooling air is then passed through the chord wise serpentine flow channel and then discharged through the return hole  22  into the spent cooling air collector cavity  13  within the airfoil mid-chord section or out the row of film cooling holes  16  or  17  on the pressure side or the suction side walls if used. Multiple film cooling holes can be used to discharge cooling air from the collector cavity  13  or from the mini serpentine cooling passage to provide film cooling for the airfoil external surface. 
   A third embodiment of the present invention is shown in  FIG. 4  in which the cooling circuit include a two 3-pass serpentine flow channels  31  and  32  instead of three 3-pass channels as shown in  FIGS. 3   a  and  3   b . Each mini serpentine flow channel includes a first 3-pass channel  31  and a second 3-pass channel  32  connected by a spanwise channel  41 . A feed hole  21  supplies cooling air to the first 3-pass channel and a discharge hole  22  discharges cooling air from the second 3-pass channel  32 . As in the  FIG. 3   b  embodiment, the discharge holes  22  can be replaced with a row of film cooling holes to discharge the cooling air onto the external surface of the airfoil. 
   In each of the near wall mini serpentine flow channels of the above embodiments, the cooling air flow through the individual module can be regulated according to the airfoil gas side pressure distribution in both the chord wise and the span wise directions to control the airfoil main body metal temperature. Also, each individual module can be designed based on the airfoil local external heat load to achieve a desired local metal temperature. Varying the size of the supply hole  21  or the discharge hole  22  can accomplish this adjustment. The mini serpentine module can be designed as a 5-pass counter and parallel flow serpentine network or a triple-pass counter and parallel flow serpentine network. Also, the individual small modules can be constructed in a multiple array along the airfoil main body wall in an inline or staggered array. For example, it can be a triple 3-pass mini serpentine flow circuit as seen in  FIGS. 3   a  and  3   b , or a double 3-pass mini serpentine flow circuit, or a single 3-pass mini serpentine flow circuit depending on the airfoil local heat load or required design metal temperatures. Also, the mini serpentine passages can be any arrangement of 2, 3, 4, or 5 pass chordwise channels in series such. For example, the chordwise extending mini serpentine circuits can be 2 by 5-pass channels, 3 by 3-pass channels, 4 by 3-pass channels, or any other combination. With the near wall mini serpentine cooling modules of the present invention, a maximum usage of cooling air for a given airfoil inlet gas temperature and pressure profile can be achieved. Also, the multi-pass of cooling air in the serpentine channels yields a higher internal convection cooling effectiveness than does the single pas radial flow channel used in the prior art near wall cooling circuit.