Patent Publication Number: US-8123481-B1

Title: Turbine blade with dual serpentine cooling

Description:
GOVERNMENT LICENSE RIGHTS 
     None. 
     CROSS-REFERENCE TO RELATED APPLICATIONS 
     None. 
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates generally to a gas turbine engine, and more specifically to an air cooled blade in a gas turbine engine. 
     2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98 
     A gas turbine engine includes a turbine with multiple rows or stages of rotor blades that react with a high temperature gas flow to drive the engine or, in the case of an industrial gas turbine (IGT), drive an electric generator and produce electric power. It is well known that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage vanes and blades and the amount of cooling that can be achieved for these airfoils. 
     In latter stages of the turbine, the gas flow temperature is lower and thus the airfoils do not require as much cooling flow. In future engines, especially IGT engines, the turbine inlet temperature will increase and result in the latter stage airfoils to be exposed to higher temperatures. To improve efficiency of the engine, low cooling flow airfoils are being studied that will use less cooling air while maintaining the metal temperature of the airfoils within acceptable limits. Also, as the TBC (thermal barrier coating) gets thicker, less cooling air is required to provide the same metal temperature as would be for a thicker TBC. 
       FIG. 1  shows an external pressure profile for a turbine rotor blade. As indicated in the figure, the forward region of the pressure side surface experiences high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than the pressure side. The pressure side pressure profile in the line on the top while the suction side pressure profile is the line on the bottom in the  FIG. 1 . 
       FIG. 2  shows a prior art turbine rotor blade with a (1+5+1) forward flowing serpentine cooling circuit for a first stage rotor blade.  FIG. 3  shows a schematic view of the rotor blade of  FIG. 2  and  FIG. 4  shows a flow diagram of the flow path through the rotor blade, the prior art blade cooling circuit includes a leading edge cooling supply channel connected to a leading edge impingement cavity by a row of metering and impingement holes, and where the impingement cavity is connected to a showerhead arrangement of film cooling holes and gills holes on both sides to discharge a layer of film cooling air onto the leading edge surface of the airfoil. A forward flowing 5-pass serpentine cooling circuit is used in the airfoil mid-chord region with a first leg for supplying cooling air located adjacent to the trailing edge region of the airfoil. The second leg, third leg, fourth leg and fifth leg of the serpentine flow toward the leading edge in series with rows of film cooling holes connected to the 5 legs to discharge film cooling air onto one or both side of the airfoil. The cooling air flows from the trailing edge region toward the leading edge region and discharges into the hot gas side pressure section of the pressure side of the airfoil. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed for this particular design, and thus inducing a high leakage flow. In the prior art cooling arrangement of  FIG. 2 , the blade tip section is cooled with double tip turns in the serpentine circuit and with local film cooling. Cooling air bled off from the 5-pass serpentine flow circuit will thus reduce the cooling performance for the serpentine flow circuit. Independent cooling flow circuit is used to provide cooling circuits from the 5-pass serpentine flow circuit is used for cooling of the airfoil leading and trailing edges. 
     As the TBC technology improves and more industrial turbine blades are applied with thicker or low conductivity TBC, the amount of cooling flow required for the blade will be reduced. As a result, there is not sufficient cooling flow for the prior art design with the 1+5+1 forward flowing serpentine cooling circuits of  FIG. 2 . Cooling flow for the blade leading edge and trailing edge has to be combined with the mid-chord flow circuit to form a single 5-pass flow circuit. However, for a single forward flow 5-pass circuit with total blade cooling flow BFM (back flow margin) may become a design problem. 
     BRIEF SUMMARY OF THE INVENTION 
     It is an object of the present invention to provide for a turbine rotor blade with a thick TBC and low cooling flow for a low gas temperature condition. 
     The above objective and more are achieved with the cooling circuit for a rotor blade of the present invention which includes a maze flowing 5-pass serpentine flow circuit that includes an aft flowing 3-pass serpentine flow circuit in series with a forward flowing 2-pass serpentine flow circuit in which the 2-pass serpentine circuit provides cooling for the blade mid-chord region. The first leg of the aft flowing 3-pass serpentine circuit is connected to a row of showerhead film cooling holes to provide film cooling for the leading edge region of the blade. The last and third leg of the aft flowing 3-pass serpentine circuit is connected to a trailing edge cooling circuit that includes multiple impingement holes with a row of exit cooling holes to discharge the cooling air. The forward flowing 2-pass serpentine circuit discharges into a blade tip cooling channel to provide cooling for the tip region and to discharge cooling air though tip cooling holes. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         FIG. 1  shows a graph of a turbine rotor blade external pressure profile. 
         FIG. 2  shows a cross section top view of a prior art turbine rotor blade 1+5+1 forward flowing serpentine cooling circuit. 
         FIG. 3  shows a schematic view of the prior art turbine rotor blade. 
         FIG. 4  shows a flow diagram of the prior art 1+5+1 serpentine flow cooling circuit of  FIG. 2 . 
         FIG. 5  shows a cross section top view of the cooling circuit of the present invention. 
         FIG. 6  shows a flow diagram of the cooling circuit of the present invention. 
         FIG. 7  shows a cross section side view of the turbine rotor blade cooling circuit of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The serpentine flow cooling circuit for the rotor blade of the present invention is shown in  FIGS. 5 through 7  with the aft flowing 3-pass serpentine circuit having three channels or legs  11 - 13  and the forward flowing 2-pass serpentine circuit having two channels or legs  21  and  22  formed between the pressure side wall and the suction side wall of the blade. The first leg  11  of the aft flowing 3-pass serpentine circuit is located along the leading edge of the blade and is connected to an arrangement of showerhead film cooling holes  31 . The first leg  11  flows from the root to the blade tip and then into the second leg  12  which flows downward toward the root section of the blade. The second leg  12  is connected to the third leg  13  which is positioned adjacent to a trailing edge region of the blade. 
     The trailing edge region is cooled by a series of impingement holes that discharge into impingement cavities with the last impingement cavity connected to a row of exit cooling holes or slots formed on the pressure side wall of the trailing edge region to discharge the spent cooling air. In this particular embodiment, the trailing edge cooling circuit includes a row of first metering holes  32  that discharge into a first impingement channel  33 , and then through a second row of metering holes  34  and into a second impingement channel  35 , which is connected to the row of exit slots  36  to discharge the spent cooling air. 
     Positioned between the second leg  12  and the third leg  13  of the aft flowing 3-pass serpentine circuit is the forward flowing 2-pass serpentine circuit with the first leg  21  connected at the end of the third leg  13  of the 3-pass serpentine circuit, the second leg  22  is connected to the first leg  21  at a root turn as seen in  FIG. 6 . the second leg  22  flows up and into a blade tip cooling channel  41  located under the blade tip an aft of the tip turn of the first and second legs  11  and  12  of the 3-pass serpentine circuit, the tip cooling channel  41  is connected to tip cooling holes  42  that discharge most of the tip cooling air onto the external surface of the blade tip with the remaining tip channel cooling air being discharged through the tip cooling channel exit hole  43  located at the trailing edge. Two rows of film cooling holes  44  and  45  are connected to the second leg  22  of the 2-pass serpentine circuit on both the pressure side and the suction side walls to discharge film cooling air. 
       FIG. 7  shows a cross section side view of the cooling circuit for the blade. The first leg  11  of the 3-pass serpentine circuit is located adjacent to the leading edge and flows upward and into the tip turn that leads into the second leg  12 . The tip turn will provide cooling for the section of the blade tip not covered by the tip cooling channel  41 . The third leg  13  flows up and along the trailing edge region to supply cooling air to the metering holes and impingement channels formed in the trailing edge region. Any remaining cooling air from the third leg  13  will flow into the first leg  21  of the forward flowing 2-pass serpentine circuit and into the second leg  22 , where some of the cooling air flows through the pressure side and suction side film cooling holes  44  and  45 . the remaining cooling air in the second leg  22  will flow up and into the tip cooling channel  41  to provide cooling for the tip section and to discharge most of the tip channel cooling air through the tip cooling holes  42  with any remaining tip channel cooling air flowing out through the exit hole  43  connected to the tip cooling channel  41 . As seen in  FIG. 7 , all of the legs or channels of both serpentine flow circuits includes trip strips along the walls to promote heat transfer from the hot metal surfaces to the cooling air flow. 
     A cross-over hole  51  can be used to connect the tip turn of the aft flowing 3-pass serpentine circuit to the tip cooling channel and the second leg of the forward flowing 2-pass serpentine flow circuit as seen in  FIG. 7 . The rotor blade with the cooling circuit described above can be formed using the investment casting process in which the cooling channels, metering holes and impingement channels and other features are formed by core ties using the lost wax process. Film cooling holes can be drilled after the blade has been cast. 
     The cooling circuit of the present invention delivers fresh cooling air to the leading edge region first. Since the cooling air temperature is fresh (meaning it hasn&#39;t been heated up yet from the hot metal temperatures) and the blade leading edge section experiences the highest heat load on the entire airfoil external surface, the use of the cooling air potential to achieve a low local metal temperature is maximized, which yields a higher oxidation life for the blade. 
     The aft flowing 3-pass serpentine circuit used for the airfoil forward section will maximize the use of cooling air pressure potential. Since the cooling air is discharged on the airfoil leading edge region where the main stream hot gas side pressure is relatively high, the aft flowing 3-pass serpentine circuit will consume less pressure than will the forward flowing 5-pass serpentine flow circuit of the prior art. This result in a low cooling supply pressure needed to produce the required airfoil cooling. 
     At an end of the aft flowing 3-pass serpentine circuit is the forward flowing 2-pass serpentine flow circuit that provides cooling for the mid-portion of the blade mid-chord section. The forward flowing 2-pass serpentine circuit used for the airfoil mid section surface will maximize the use of cooling to main stream gas side heat load potential. The heat load at the blade mid-chord section is lower than at the leading edge and the trailing edge of the airfoil. The spent cooling air is then channeled through the blade tip section axial flow channel to provide blade section cooling and is then discharged along the blade pressure side peripheral at the aft section of the airfoil to provide cooling of the blade tip edge. This design yields a lower cooling air supply pressure requirement and a lower leakage flow.