Patent Publication Number: US-2016222795-A1

Title: Turbine Airfoil Cooling Core Exit

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application claims priority to U.S. Provisional Application No. 61/894496, filed Oct. 23, 2013. 
    
    
     BACKGROUND OF THE INVENTION 
     This application relates to cooling a tip of a gas turbine engine airfoil. 
     Gas turbine engines are known and, typically, include a fan delivering air into a compressor section. The air is compressed and delivered into a combustor section. In the combustor section, fuel is mixed with the compressed air and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
     The turbine rotors include rotating blades and static vanes, all of which are exposed to the hot products of combustion. As such, it is known to provide cooling air passages within the airfoils. One known cooling scheme directs cooling air through a plurality of channels on a suction side of the airfoil, and extending toward a radially tip. The plurality of channels merge into a plenum at the tip. Air from this plenum then passes toward a pressure side of the airfoil and leaves through cooling holes at the tip and adjacent the pressure side. The air in the plenum loses velocity and, thus, the cooling is not as efficient as may be desired. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine component has an airfoil extending from a platform to a tip at an end of the airfoil spaced from the platform. The airfoil has a suction wall and a pressure wall, with at least one channel extending toward the tip from the platform. A plenum communicates with the at least one channel. The plenum flows from the suction wall toward the pressure wall at the tip to communicate with cooling holes near the pressure wall. The plenum has a reduced cross-sectional area between the suction wall and the pressure wall, and an increase in cross-sectional area downstream of the reduced cross-sectional area. 
     In another embodiment according to the previous document, the plenum has a first enlarged cross-sectional area portion, the reduced cross-sectional area portion, and then a second enlarged cross-sectional area flow portion. 
     In another embodiment according to any of the previous embodiments, a plurality of cavities form the reduced cross-sectional area. 
     In another embodiment according to any of the previous embodiments, the first enlarged cross-sectional area portion is formed by a single plenum that communicates cooling air downstream to the plurality of cavities. The second enlarged cross-sectional area portion is formed by a single plenum that receives cooling air from the plurality of cavities. 
     In another embodiment according to any of the previous embodiments, a plurality of cavities form the reduced cross-sectional area. 
     In another embodiment according to any of the previous embodiments, the first enlarged cross-sectional area portion is formed by a single plenum that communicates cooling air downstream to the plurality of cavities. The second enlarged cross-sectional area portion is formed by a single plenum receiving cooling air from the plurality of cavities. 
     In another embodiment according to any of the previous embodiments, the at least one channel is a plurality of suction wall channels. 
     In another embodiment according to any of the previous embodiments, the at least one channel is a serpentine channel that extends between the suction and pressure walls, and communicates with the plenum adjacent the suction wall. 
     In another embodiment according to any of the previous embodiments, the cooling holes are formed in an outer tip face of the airfoil. 
     In another embodiment according to any of the previous embodiments, the component is a turbine blade. 
     In another featured embodiment, a mold core for use in forming cooling passages within a gas turbine component has at least one finger merging into a single solid portion. A plurality of ribs connect the first single solid portion to a second single solid portion. 
     In another embodiment according to the previous embodiment, the at least one finger is a plurality of fingers spaced from each other. 
     In another featured embodiment, a gas turbine engine has a compressor section and a turbine section that include rotating blades and static vanes, with at least one of the rotating blades and static vanes including an airfoil extending from a platform to a tip defined at an end of the airfoil spaced from the platform. The airfoil has a suction wall and a pressure wall, with at least one channel extending toward the tip from the platform. A plenum communicates with the at least one channel. The plenum flows from the suction wall toward the pressure wall at the tip to communicate with cooling holes near the pressure wall. The plenum has a reduced cross-sectional area between the suction wall and the pressure wall, and an increase in cross-sectional area downstream of the reduced cross-sectional area. 
     In another embodiment according to the previous embodiment, the plenum has a first enlarged cross-sectional area portion, the reduced cross-sectional area portion, and then a second enlarged cross-sectional area flow portion. 
     In another embodiment according to any of the previous embodiments, a plurality of cavities form the reduced cross-sectional area. 
     In another embodiment according to any of the previous embodiments, the first enlarged cross-sectional area portion is formed by a single plenum that communicates cooling air downstream to the plurality of cavities. The second enlarged cross-sectional area portion is formed by a single plenum receiving cooling air from the plurality of cavities. 
     In another embodiment according to any of the previous embodiments, the component is a turbine blade. 
     In another embodiment according to any of the previous embodiments, a plurality of cavities form the reduced cross-sectional area. 
     In another embodiment according to any of the previous embodiments, the cooling holes are formed in an outer tip face of the airfoil. 
     In another embodiment according to any of the previous embodiments, the component is a turbine blade. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2A  shows an example turbine component. 
         FIG. 2B  shows a top view of the  FIG. 2A  component. 
         FIG. 2C  is a cross-sectional view along line C-C of  FIG. 2A . 
         FIG. 2D  shows a prior art cooling scheme. 
         FIG. 2E  shows a mold core for forming the  FIG. 2D  cooling scheme. 
         FIG. 3  shows a novel cooling scheme. 
         FIG. 4  shows a core component for forming the  FIG. 3  cooling scheme. 
         FIG. 5  schematically shows a mold. 
         FIG. 6A  shows a second embodiment. 
         FIG. 6B  is a view along line B-B of  FIG. 6A . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
       FIG. 2A  shows a sample turbine component, which is illustrated as a blade  80 . While a blade  80  is illustrated, it should be understood the teachings of this disclosure extend to other components having an airfoil, such as vanes. An airfoil  82  extends radially outwardly from a platform  84 . An outer tip face  86  is spaced from the platform at a radial end. If used as a blade, this will be the radially outer end. If used as a vane, this will be the radially inner end. 
       FIG. 2B  shows the airfoil  82  extending from a trailing edge  90  to a leading edge  88 . There is also a suction wall  94  and a pressure wall  92 . 
       FIG. 2C  is a cross-sectional view along line C-C as shown in  FIG. 2A . As shown, there are main body cooling air channels  96  which receive cooling air, such as from a source radially beyond the platform  84  and pass cooling air toward the outer tip  86 . There are also suction sides skin channels  98  formed adjacent the suction wall  94 . 
     As shown in  FIG. 2D , the skin channels  98  extend radially from the platform  84  toward the tip  86 , then merge into a single plenum  100 . Air from the plenum  100  then leads to a plurality of cooling holes  102  in the outer tip face  86  and adjacent the pressure wall  92 . 
     As known, the cooling channels within an airfoil may be formed by a mold core in a lost core molding process. A mold core  110  for forming the cooling scheme as shown in  FIG. 2D  is illustrated in  FIG. 2E . A plurality of mold core fingers  197  extend to a plenum core portion  200 . As known, the mold core  110  is placed in a mold and surrounded by molten metal. That metal then solidifies. The mold core  110  may then be leached leaving the cavity such as shown in  FIG. 2D  within the airfoil  82 . 
     As mentioned, the cooling air in the enlarged plenum  100  may lose velocity and does not cool as efficiently as may be desired.  FIG. 3  shows a novel airfoil  198 . Airfoil  198  has a suction wall  202  with a plurality of suction wall channels  206 . A pressure wall  204  has cooling holes  216  on an outer tip face  215 . The suction wall channels  206  merge into a first plenum  207 . There is an end wall  208  of the first plenum portion  207  and then necked or reduced cross-sectional area portion  210 . Downstream of the reduced area portion  210 , a second plenum portion  212  expands outwardly to an end wall  214 . While the reduced area portion is formed of a plurality of areas or cavities  210 , it could be a single channel. 
     Now, the cooling air will flow radially outwardly through the suction wall channels  206 . The cooling air from the plurality of suction wall channels  206  will merge together in the first plenum portion  207 , then be reduced down to the necked portion  210  and expand back into the second plenum portion  212 . The plenum has a reduced cross-sectional area between the suction wall and the pressure wall, and an increase in cross-sectional area downstream of the reduced cross-sectional area. At that point, the cooling air may hit the wall  214 . This then provides additional cooling at or adjacent to the pressure wall  204  before the air leaves through the cooling holes  216 . 
     Further, the area reduction ensures that the air will continue to have adequate and high speeds to provide increased convection cooling adjacent the tip. 
     In one embodiment, a ratio of a cooling air flow area in the first plenum portion  207 , at a location immediately before the necked or reduced area portion  210 , compared to an area of the reduced area portion  210  immediately before the end of the reduced area portion is between ⅞ and ¾. 
     Of course, other ratios would come within the scope of this disclosure. 
       FIG. 4  shows a mold core  300  for forming the cooling scheme of  FIG. 3 . As shown, a plurality of fingers  302  will form the suction wall channels  206 . Those fingers all merge into an enlarged plenum portion  304 . The plenum portions then merge into a plurality of ribs  306 . Intermediate the ribs  306  are spaces  307 . A worker of ordinary skill in this art would recognize that when molding of an airfoil occurs around the mold core  300 , the ribs  306  will result in cavities such as the reduced flow cross-sectional areas  210 , while the spaces  307  will result in solid material intermediate the areas  210 . That solid material will form ribs. The ribs provide additional rigidity at the tip of the airfoil  198 . 
     A second plenum portion  308  is connected to each of the ribs  306  and will form the second plenum portion  212 . It should be understood that the second plenum portion  308  may actually be a plurality of portions. 
       FIG. 5  schematically shows a mold  400 . As known, a space  402  surrounds the mold core  300 . The space  402  receives molten metal and the metal is allowed to solidify. After the metal has solidified, the mold core  300  is leached away leaving the cavity such as shown in  FIG. 3 . 
     A mold core such as shown in  FIG. 4  would preferably be formed by an additive or layer manufacturing process. One desired process would be direct metal laser centering. 
     Another embodiment  500  is illustrated in  FIG. 6A and 6B . Rather than having the suction wall channels supplying the plenum, a serpentine feed channel  510  extends between the suction wall S and pressure wall P. The airfoil  506  is in large part cooled by this serpentine feed passage. As shown, the serpentine feed channel  510  communicates at  514  to the plenum  512 , and adjacent the suction wall S. There is again a reduced flow cross-sectional area  516 , which may be a plurality of areas formed by the ribs. A second plenum portion  518  communicates with outlet holes  520 . 
     As shown in  FIG. 6B , the serpentine feed channel  510  may include a plurality of legs  550 ,  552 , and  554  moving cooling fluid in a serpentine path through the airfoil  506 . 
     The embodiment of  FIG. 6A and 6B  may be formed by an appropriate mold core, and molded by a method similar to that described with regard to  FIG. 5 . 
     To further enhance heat transfer, augmentation features such as tip-strips, pedestals, dimples and radial fins may be provided. In particular, such structure may be included in the second plenum portion  212 , as shown schematically at  600  in  FIG. 3 . 
     Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.