Abstract:
The flow of cooling fluid through a core tie hole formed between a pair of internal cooling passageways of an airfoil component is reduced by providing a restriction that meters coolant flow through the inlet passage of one of the passageways so that the pressures in the two passageways are equalized, thereby minimizing the flow of cooling fluid through the hole. The restriction can be a metering plate disposed at the entrance of the inlet passage or a restriction integrally formed in the inlet passage.

Description:
BACKGROUND OF THE INVENTION 
     The present invention relates generally to gas turbine engines, and more particularly to internally cooled airfoils used in such engines. 
     Gas turbine engines, such as aircraft jet engines, include many components (e.g., turbines, compressors, fans and the like) that utilize airfoils. Turbine airfoils, such as turbine blades and nozzle vanes, which are exposed to the highest operating temperatures, typically employ internal cooling to keep the airfoil temperatures within certain design limits. A turbine rotor blade, for example, has a shank portion that is attached to a rotating turbine rotor disk and an airfoil blade portion which is employed to extract useful work from the hot gases exiting the engine&#39;s combustor. The airfoil is attached to the shank and includes a blade tip that is the free end of the airfoil blade. Typically, the airfoil of the turbine rotor blade is cooled by air (normally bled from the engine&#39;s compressor) passing through an internal circuit, with the air entering the airfoil through the shank and exiting through airfoil tip holes, airfoil film cooling holes and blade trailing edge slots or holes. Known turbine blade cooling circuits include a plurality of radially-oriented passageways that are series-connected to produce a serpentine flow path, thereby increasing cooling effectiveness by extending the length of the coolant flow path. It is also known to provide additional, unconnected passageways adjacent to the serpentine cooling circuit. 
     Turbine rotor blades with internal cooling circuits are typically manufactured using an investment casting process commonly referred to as the lost wax process. This process comprises enveloping a ceramic core defining the internal cooling circuit in wax shaped to the desired configuration of the turbine blade. The wax assembly is then repeatedly dipped into a liquid ceramic solution such that a hard ceramic shell is formed thereon. Next, the wax is melted out of the shell so that the remaining mold consists of the internal ceramic core, the external ceramic shell and the space therebetween, previously filled with wax. The empty space is then filled with molten metal. After the metal cools and solidifies, the external shell is broken and removed, exposing the metal that has taken the shape of the void created by the removal of the wax. The internal ceramic core is dissolved via a leaching process. The metal component now has the desired shape of the turbine blade with the internal cooling circuit. 
     In casting turbine blades with serpentine cooling circuits, the internal ceramic core is formed as a serpentine element having a number of long, thin branches. This presents the challenge of making the core sturdy enough to survive the pouring of the metal while maintaining the stringent requirements for positioning the core. Furthermore, the thin branches of the serpentine core can experience relative movement if not stabilized in some manner. Thus, core ties (i.e., small ceramic connectors between various branches) are used to strengthen the core. This prevents relative movement of the core branches such that the airfoil external wall thicknesses are controlled better. After casting, when they have been removed along with the core, the core ties leave holes in the ribs or walls separating adjacent passageways. These core tie holes provide unwanted flow communication between adjacent passageways if a pressure differential exists between the two passageways. That is, cooling fluid in the higher pressure passageway will flow into the lower pressure passageway through the core tie hole. This will result in an undesirable cooling flow distribution compared to the original design intent. 
     Accordingly, there is a need for an airfoil component in which cooling fluid flow through core tie holes is minimized. 
     SUMMARY OF THE INVENTION 
     The above-mentioned needs are met by the present invention which provides an airfoil component comprising at least two internal cooling passageways separated by a rib having a core tie hole formed therein. A means for metering flow through the inlet passage of one of the passageways is provided so that the pressures in the two passageways are substantially equal. This reduces the flow of cooling fluid through the core tie hole. 
     Other objects and advantages of the present invention will become apparent upon reading the following detailed description and the appended claims with reference to the accompanying drawings. 
    
    
     DESCRIPTION OF THE DRAWINGS 
     The subject matter which is regarded as the invention is particularly pointed out and distinctly claimed in the concluding part of the specification. The invention, however, may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
     FIG. 1 is a longitudinal cross-sectional view of a prior art turbine blade. 
     FIG. 2 is a longitudinal cross-sectional view of a turbine blade in accordance with a first embodiment of the present invention. 
     FIG. 3 is a longitudinal cross-sectional view of a turbine blade in accordance with a second embodiment of the present invention. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, FIG. 1 shows a prior art gas turbine engine rotor blade  10  having a hollow airfoil  12  and an integral shank  14  for mounting the airfoil  12  to a rotor disk (not shown) in a conventionally known manner. The airfoil  12  extends longitudinally or radially upwardly from a blade platform  16  disposed at the top of the shank  14  to a blade tip  18 . The airfoil  12  includes an internal serpentine cooling circuit having five series-connected, generally radially extending cooling passageways  20 - 24 . 
     The first passageway  20  receives a cooling fluid (usually a portion of relatively cool compressed air bled from the compressor (not shown) of the gas turbine engine) through a first inlet passage  46  in the shank  14 . The cooling fluid travels radially outwardly through the first passageway  20 , passes into the second passageway  21  and then flows radially inwardly through the second passageway  21 . From there, the cooling fluid similarly passes in series through the other passageways  22 - 24 , thereby cooling the airfoil  12  from the heating effect of the combustion gases flowing over the outer surfaces thereof. As is known, the cooling fluid exits the airfoil  12  through film cooling holes (not shown) and an opening  26  in the blade tip  18 . 
     The airfoil  12  includes a leading edge cooling passageway  28  in addition to the serpentine cooling circuit. The leading edge passageway  28  extends radially between the airfoil leading edge  30  and the first passageway  20  and is not connected to the serpentine cooling circuit. A separate flow of cooling fluid is introduced through a second inlet passage  48  in the shank  14 . The cooling fluid flows radially through the leading edge passageway  28  and is discharged from the airfoil  12  through conventional film cooling holes and/or a tip hole (not shown) formed through the exterior wall of the airfoil  12 . Similarly, a radially extending trailing edge cooling passageway  32  is disposed between the airfoil trailing edge  34  and the fifth passageway  24  of the serpentine cooling circuit. The trailing edge passageway  32  is also not connected to the serpentine cooling circuit and receives another separate flow of cooling fluid through a third inlet passage  50  in the shank  14 . This cooling fluid flows radially through the trailing edge passageway  32  and is discharged from the airfoil  12  through a conventional row of trailing edge film holes or slots and/or a tip hole (not shown). The arrows in FIG. 1 indicate the various paths of cooling fluid flow. 
     As seen in FIG. 1, each one of the passageways  20 - 24 ,  28 ,  32  is separated from adjacent passageways by six radially extending ribs  36 - 41 . That is, the leading edge passageway  28  and the first passageway  20  of the serpentine cooling circuit are separated by a first rib  36 , the first passageway  20  and the second passageway  21  are separated by a second rib  37 , and so on. At least some of the ribs  36 - 41  have a core tie hole  42  formed therein due to the use of core ties in the casting process. Specifically, the prior art blade  10  of FIG. 1 has core tie holes  42  formed in the first rib  36 , the third rib  38 , the fifth rib  40  and the sixth rib  41 , although other configurations are possible depending on how the core ties are deployed during the casting process. Core tie holes, which are often elliptical in cross-section, typically have an equivalent diameter of about 0.03-0.1 inches. 
     The cooling fluid, which is typically air bled from the compressor, is supplied to each of the three inlet passages  46 , 48 , 50  at the same pressure. However, the cooling fluid pressure in the passageways  20 - 24  tends to decrease along the serpentine flow path due to friction and turning losses in the five pass serpentine circuit. The first passageway  20 , the leading edge passageway  28  and the trailing edge passageway  32 , which are all directly connected to a corresponding one of the inlet passages  46 , 48 , 50 , all have substantially the same pressure, but the pressure in the fifth passageway  24 , the last pass of the serpentine circuit, will be substantially less. Accordingly, there is a pressure differential between the fifth passageway  24  and the adjacent trailing edge passageway  32 , which is a single pass circuit not subject to the same pressure loss as the five pass serpentine circuit. Because of this pressure differential, cooling fluid will pass from the trailing edge passageway  32  to the fifth passageway  24  through the core tie hole  42  in the sixth rib  41 , starving the tip region of the trailing edge passageway  32  of cooling fluid. 
     Referring now to FIG. 2, a turbine blade  110  is shown in which cooling fluid flow through core tie holes is minimized. For purposes of illustration only, the blade  110  has the same cooling circuit configuration as the blade  10  of FIG.  1 . However, it should be noted that the present invention is applicable to turbine blades having other cooling circuit configurations. Furthermore, the present invention is not limited to turbine blades and could be used with other types of airfoil components such as turbine nozzles. As will become apparent from the following description, the present invention is applicable to any airfoil component having individually fed cooling passageways that are short-circuited by core tie holes. 
     The blade  110  has a hollow airfoil  112  and an integral shank  114 . The airfoil  112  includes a serpentine cooling circuit having five series-connected, generally radially extending cooling passageways  120 - 124 , a leading edge cooling passageway  128  extending radially between airfoil leading edge  130  and the first passageway  120 , and a radially extending trailing edge cooling passageway  132  disposed between airfoil trailing edge  134  and the fifth passageway  124 . The first passageway  120  is supplied with cooling fluid through a first inlet passage  146  in the shank  114 , the leading edge passageway  128  is supplied with cooling fluid through a second inlet passage  148  in the shank  114 , and the trailing edge passageway  132  is supplied with cooling fluid through a third inlet passage  150 . Each one of the passageways  120 - 124 ,  128 , 132  is separated from adjacent passageways by six radially extending ribs  136 - 141 . A core tie hole  142  is formed in the first rib  136 , the third rib  138 , the fifth rib  140  and the sixth rib  141 , although other configurations are possible depending on how the core ties are deployed during the casting process. 
     The blade  110  includes a root metering plate  152  disposed on the radially inner surface of the shank  114  so as to completely cover the third inlet passage  150 . The metering plate  152  is a thin plate of any suitable material attached to the shank  114  by an appropriate means such as brazing. A metering hole  154  is formed in the metering plate  152  to allow a metered flow of cooling fluid to pass into the third inlet passage  150 . The cross-sectional area of the metering hole  154  is smaller than the cross-sectional area of the third inlet passage  150 . Thus, the metering hole  154  presents a restriction at the entrance of the third inlet passage  150  that causes a pressure drop such that the pressure in the trailing edge passageway  132  is less than what it would be without the metering plate  152 . 
     The size of the metering hole  154  is selected to meter the cooling fluid flow through the third inlet passage  150  such that the pressure in the trailing edge passageway  132  is substantially equal to the pressure in the fifth passageway  124 , thereby minimizing the pressure differential across the core tie hole  142  in the sixth rib  141 . The specific size of the metering hole l 54  to achieve this result will be dependent on the overall cooling fluid flow level and the pressure differential that would exist between the trailing edge passageway  132  and the fifth passageway  124  without the metering plate  152 . By minimizing the pressure differential across the core tie hole  142  in the sixth rib  141 , the present invention lessens the adverse impact of the core tie hole  142  on the effectiveness of the airfoil cooling scheme. 
     Turning to FIG. 3, an alternative embodiment of the present invention is shown in the form of a turbine blade  210 . For purposes of illustration only, the blade  210  is similar to the blade  110  of FIG. 2, although, as before, it should be noted that this alternative embodiment of the present invention is applicable to turbine blades having other cooling circuit configurations as well as other types of airfoil components. 
     The blade  210  is similar to the blade  110  of FIG. 2 in that it has a hollow airfoil  212  and an integral shank  214 . The airfoil  212  includes a serpentine cooling circuit having five series-connected, generally radially extending cooling passageways  220 - 224 , a leading edge cooling passageway  228  extending radially between airfoil leading edge  230  and the first passageway  220 , and a radially extending trailing edge cooling passageway  232  disposed between airfoil trailing edge  234  and the fifth passageway  224 . The first passageway  220  is supplied with cooling fluid through a first inlet passage  246  in the shank  214 , the leading edge passageway  228  is supplied with cooling fluid through a second inlet passage  248  in the shank  214 , and the trailing edge passageway  232  is supplied with cooling fluid through a third inlet passage  250 . Each one of the passageways  220 - 224 ,  228 ,  232  is separated from adjacent passageways by six radially extending ribs  236 - 241 . A core tie hole  242  is formed in the first rib  236 , the third rib  238 , the fifth rib  240  and the sixth rib  241 , although other configurations are possible depending on how the core ties are deployed during the casting process. 
     The blade  210  differs from the blade  110  of FIG. 2 in that it has no metering plate. Instead, a restriction  256  is formed in the third inlet passage  250 . Preferably, the restriction  256  is cast as an integral part of the blade  210 . The restriction  256  presents a reduced cross-sectional area so as to cause a pressure drop such that the pressure in the trailing edge passageway  232  is less than what it would be if the restriction  256  was omitted. 
     Like the metering hole  154  of FIG. 2, the size of the restriction  256  is selected to meter the cooling fluid flow through the third inlet passage  250  such that the pressure in the trailing edge passageway  232  is substantially equal to the pressure in the fifth passageway  224 , thereby minimizing the pressure differential across the core tie hole  242  in the sixth rib  241 . The specific size of the restriction  256  to achieve this result will be dependent on the overall cooling fluid flow level and the pressure differential that would exist between the trailing edge passageway  232  and the fifth passageway  224  without the restriction  256 . By minimizing the pressure differential across the core tie hole  242  in the sixth rib  241 , the present invention lessens the adverse impact of the core tie hole  242  on the effectiveness of the airfoil cooling scheme. 
     The foregoing has described a turbine airfoil component in which cooling fluid flow through a core tie hole is minimized. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention as defined in the appended claims.