Patent Publication Number: US-7581928-B1

Title: Serpentine microcircuits for hot gas migration

Description:
BACKGROUND 
     (1) Field of the Invention 
     The present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion. 
     (2) Prior Art 
     The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7. 
       FIG. 1  shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point  10  related to a new advanced serpentine microcircuit shown in  FIGS. 2   a - 2   c . This serpentine microcircuit includes a pressure side serpentine circuit  20  and a suction side serpentine circuit  22  embedded in the airfoil walls  24  and  26 . 
     The Table I below provides the operational parameters used to plot the design point in the durability map. 
     
       
         
           
               
             
               
                 TABLE I 
               
               
                   
               
               
                 Operational Parameters for 
               
               
                 serpentine microcircuit 
               
               
                   
               
             
            
               
                   
               
            
           
           
               
               
               
            
               
                   
                 beta 
                 2.898 
               
               
                   
                 Tg 
                 2581 [F] 
               
               
                   
                 Tc 
                 1365 [F] 
               
               
                   
                 Tm 
                 2050 [F] 
               
               
                   
                 Tm_bulk 
                 1709 [F] 
               
               
                   
                 Phi_loc 
                 0.437 
               
               
                   
                 Phi_bulk 
                 0.717 
               
               
                   
                 Tco 
                 1640 [F] 
               
               
                   
                 Tci 
                 1090 [F] 
               
               
                   
                 eta_c_loc 
                 0.573 
               
               
                   
                 eta_f 
                 0.296 
               
               
                   
                 Total Cooling Flow 
                 3.503% 
               
               
                   
                 WAE 
                 10.8 
               
               
                   
                   
               
               
                   
                 Legend for Table I 
               
               
                   
                 Beta =  heat load 
               
               
                   
                 Phi bc = local cooling effectiveness 
               
               
                   
                 Phi_bulk = bulk cooling effectiveness 
               
               
                   
                 Eta_c_bc = local cooling efficiency 
               
               
                   
                 Eta_f = film effectiveness 
               
               
                   
                 Tg = gas temperature 
               
               
                   
                 Tc = coolant temperature 
               
               
                   
                 Tm = metal temperature 
               
               
                   
                 Tm_bulk = bulk metal temperature 
               
               
                   
                 Tco = exit coolant temperature 
               
               
                   
                 Tci = inlet coolant temperature 
               
               
                   
                 WAE = compressor engine flow, pps 
               
            
           
         
       
     
     It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow.  FIG. 3  illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of  FIGS. 2   a - 2   c  embedded in the airfoils walls. 
     There are however field problems that can be addressed efficiently with peripheral microcircuit designs. One such field problem is illustrated in  FIGS. 4A and 4B . In  FIG. 4A , the streamlines of the gas path close to the external surface of the airfoil illustrate four different regions in which the gas flow changes direction or migration: a tip region, two mid-section regions, and a root region. In between the tip and the upper mid region, the flow transitions through a pseudo stagnation point(s). The momentum of the external gas seems to decelerate in such a way as to impose a local thermal load to the part. This manifests itself by regions where the propensity for erosion and oxidation increase in the airfoil surface. The superposition of  FIG. 4B  illustrates the local coincidence between the pseudo-stagnation region and the blade distress in the part surface. In the mid region, the upper and lower region also converge onto one another, but even though the space between streamlines decreases, the flow seems to accelerate and there is no pseudo-stagnation regions. A mild manifestation of the same tip-to-mid phenomena seems to initiate in the transition region between the mid-to-root regions. It is therefore necessary to tailor the peripheral microcircuit in such a manner as to address these local high thermal load regions. 
     SUMMARY OF THE INVENTION 
     In accordance with the present invention, a turbine engine component is provided with improved cooling. The turbine engine component broadly comprises an airfoil portion having a pressure side and a suction side. The turbine engine component further has a first cooling circuit within the pressure side for cooling the pressure side of the airfoil portion and a second cooling circuit within the suction side for cooling the suction side of the airfoil portion and for cooperating with means for creating a cooling film over the pressure side. 
     Other details of the serpentine microcircuits for hot gas migration of the present invention, as well as other objects and advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a graph showing cooling effectiveness versus film effectiveness for a turbine engine component; 
         FIG. 2A  shows an airfoil portion of a turbine engine component having a pressure side cooling microcircuit embedded in the pressure side wall and a suction side cooling microcircuit embedded in the suction side wall; 
         FIG. 2B  is a schematic representation of a pressure side cooling microcircuit used in the airfoil portion of  FIG. 2A ; 
         FIG. 2C  is a schematic representation of a suction side cooling microcircuit used in the airfoil portion of  FIG. 2A ; 
         FIG. 3  illustrates the cooling flow distribution for a turbine engine component with serpentine microcircuits embedded in the airfoil walls; 
         FIG. 4A  is a schematic representation illustrating the pressure side distress on an airfoil surface; 
         FIG. 4B  is a schematic representation of the local coincidence between the pseudo-stagnation region and the blade distress; 
         FIG. 5  is a schematic representation of a peripheral pressure side cooling circuit; 
         FIG. 6  is a schematic representation of a peripheral suction side cooling circuit; and 
         FIG. 7  is a schematic representation of main body internal cooling circuits. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S) 
     Referring now to  FIGS. 5 and 6 , there are depicted two peripheral cooling arrangements which may be used to address local increases in the airfoil thermal load of a turbine engine component  90  such as a turbine blade. The two peripheral cooling arrangements include a peripheral pressure side microcircuit  100  which may be incorporated or embedded within the wall forming the pressure side of an airfoil portion  104  and a suction side microcircuit  120  which may be incorporated or embedded within the wall forming the suction side of the airfoil portion  104 . 
     In  FIG. 5 , the pressure side peripheral microcircuit  100  is shown. In this circuit, the first leg  102  has an inlet  103  which receives cooling fluid from a source (not shown). The leg  102  provides a flow of cooling fluid which quenches the hot spot in the tip-to-mid region of the airfoil portion  104  shown in  FIG. 4B . The cooling fluid within the leg  102  proceeds around a 180 degree bend  106  which is supplemented with a plurality of film holes  108 , preferably three film holes. The film holes  108  ensure flow acceleration through the bend  106  to a second downstream leg  110  which ends below the platform  112  of the turbine engine component  90  in an exit  164 . Cooling fluid from the leg  110  is fed into an internal trailing edge circuit  114  to be discussed hereinafter via the exit  164  where it is used to further cool the airfoil portion  104 . 
     Referring now to  FIG. 6 , there is shown a peripheral suction side microcircuit  120 . The circuit  120  has a first leg  122  which communicates with a source (not shown) of cooling fluid. In the first leg  122 , the cooling flow convects heat away from the suction side. Since the circuit  120  has no film holes, effective cooling may not be done past the external gage point of the airfoil portion  104  where any film cooling would provide high aerodynamic penalties due to mixing. (PLEASE CHECK THIS TO SEE IF IT MAKES SENSE) Thus, the circuit  120  is used to feed cooling fluid to a leading edge microcircuit  124  which wraps around the leading edge  126  of the airfoil portion  104 . The circuit  120  feeds or supplies cooling fluid to the leading edge wrap around circuit  124  through a plurality of wall cross over holes  128 . As can be seen from  FIG. 6 , the circuit  120  has a bend  130  and a second leg  132 . The holes  128  are preferably located in the vicinity of the bend  130  and the second leg  132 . The second leg  132  may also communicate with the wrap around circuit  124  via a passageway  134 . As the microcircuit  124  wraps around the leading edge, several holes  136  are located in the leading edge and are used to cool the leading edge of the airfoil portion  104 . Further, the microcircuit  124  is provided with a plurality of film holes  138  for creating a film of cooling fluid over the pressure side of the airfoil portion. 
     Referring now to  FIG. 7 , there is shown the main body internal cooling circuits which include a leading edge internal cooling circuit  150  and the trailing edge internal cooling circuit  114 . The leading edge internal cooling circuit  150  communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet  151  and has one or more film cooling holes  152  adjacent the tip  154  of the airfoil portion  104  to provide tip cooling. The circuit  150  also has a plurality of cross-over holes  156  for supplying cooling fluid to the leading edge microcircuit  124 . 
     The trailing edge internal circuit  114  also communicates with a source (not shown) of cooling fluid, such as engine bleed air, via an inlet  157  and has one or more film cooling holes  158  adjacent the tip  154  to provide tip cooling. The circuit  114  also has a plurality of cross-over holes  160  for communicating with a trailing edge cooling circuit  162  for cooling the trailing edge of the airfoil portion  104 . As can be seen from  FIG. 7 , the tailing edge internal circuit  114  also receives cooling fluid from the peripheral pressure side microcircuit  100  via the exit  164 . 
     Each of the leading edge internal circuit  150  and the trailing edge internal circuit  114  may be provided with a plurality of film cooling holes  170  and  172  respectively to form cooling films over the pressure and suction sides of the airfoil portion  104 . 
     Using the pressure and suction side cooling circuits of the present invention, the airfoil portion of a turbine engine component may be very effectively convectively cooled. Using the pressure side circuit, the cooling flow is returned to the trailing edge internal circuit for further cooling of the airfoil. Using the suction side circuit, the leading edge of the airfoil is cooled first before discharging in pressure side film. This effective use of coolant allows for positive effects on cycle thermodynamic efficiency, turbine efficiency, rotor inlet temperature impacts, and specific fuel consumption. 
     It is apparent that there has been provided in accordance with the present invention serpentine microcircuits for hot gas migration which fully satisfy the objects, means, and advantages set forth hereinbefore. While the present invention has been described in the context of specific embodiments thereof, other unforeseeable alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations as fall within the broad scope of the appended claims.