Patent Publication Number: US-7901182-B2

Title: Near wall cooling for a highly tapered turbine blade

Description:
This invention was made with U.S. Government support under Contract Number DE-FC26-05NT42644 awarded by the U.S. Department of Energy. The U.S. Government has certain rights to this invention. 
    
    
     FIELD OF THE INVENTION 
     This invention is directed generally to an airfoil for a gas turbine engine and, more particularly, to a turbine blade airfoil having cooling cavities for conducting a cooling fluid to provide near wall cooling in a highly tapered turbine blade. 
     BACKGROUND OF THE INVENTION 
     A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a shaft and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the shaft and disc assembly, to rotate. 
     Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures. 
     Typically, turbine blades comprise a root, a platform and an airfoil that extends outwardly from the platform. The airfoil is ordinarily composed of a tip, a leading edge and a trailing edge. Most blades typically contain internal cooling channels forming a cooling system. The cooling channels in the blades may receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade. 
     A conventional cooling system in a turbine blade assembly may include an intricate maze of cooling flow paths through various portions of the turbine blade. 
     Further, third row turbine blades, which may comprise a highly tapered large airfoil, present additional cooling problems associated with the geometry of the airfoil. Specifically, the lower or radially inward portion of the airfoil comprises a large cross section area, and the cross section tapers to a smaller thickness toward the tip of the blade. Accordingly, the configuration of the cooling circuit requires particular consideration to providing an airflow in contact with the lower portions of the airfoil to maintain the heat transfer coefficient in the wider cross section portion of the blade. 
     While many of the known cooling systems for turbine blades have operated successfully, a need still exists to provide increased cooling capability, particularly in turbine blades having highly tapered large airfoils. 
     SUMMARY OF THE INVENTION 
     In accordance with one aspect of the invention, a turbine blade is provided comprising an outer wall extending from a blade root to a blade tip. The outer wall comprises a pressure sidewall and a suction sidewall, and the pressure and suction sidewalls are connected at chordally spaced leading and trailing edges. A cooling cavity is defined between the pressure and suction sidewalls. A pressure side inner wall extends radially within the cooling cavity from a location adjacent the blade root toward the blade tip and defines a pressure side near wall chamber. A suction side inner wall extends radially within the cooling cavity from a location adjacent the blade root toward the blade tip and defines a suction side near wall chamber. The suction side inner wall intersects the pressure side inner wall at an intermediate location between the blade root and the blade tip. A plurality of pressure side channels extend radially from the intermediate location to a tip passage at the blade tip for connecting the pressure side near wall chamber in fluid communication with the tip passage, and a plurality of suction side channels extend radially from the intermediate location to the blade tip for connecting the suction side near wall chamber in fluid communication with the tip passage. 
     In accordance with another aspect of the invention, a turbine blade is provided comprising an outer wall extending from a blade root to a blade tip. The outer wall comprises a pressure sidewall and a suction sidewall, and the pressure and suction sidewalls are connected at chordally spaced leading and trailing edges. A cooling cavity is defined between the pressure and suction sidewalls. A pressure side inner wall extends radially within the cooling cavity from a location adjacent the blade root toward the blade tip and defines a pressure side near wall chamber. A suction side inner wall extends radially within the cooling cavity from a location adjacent the blade root toward the blade tip and defines a suction side near wall chamber. A leading edge flow channel extends radially adjacent to the leading edge, and a trailing edge flow channel extends radially adjacent to the trailing edge. A cooling fluid supply provides cooling fluid to at least the leading edge flow channel, and the cooling fluid flows in at least a triple-pass serpentine path through the leading edge flow channel, the pressure side near wall chamber, the suction side near wall chamber and the trailing edge flow channel. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein: 
         FIG. 1  is a perspective view of a turbine blade incorporating the present invention; 
         FIG. 2A  is a cross-sectional view of the turbine blade shown in  FIG. 1  taken along line  2 - 2 ; 
         FIG. 2B  is a sectional view of the turbine blade shown in  FIG. 1  with the pressure sidewall cut away; 
         FIG. 3  is a cross-sectional view of the turbine blade shown in  FIG. 1  taken along line  3 - 3 ; 
         FIG. 4  is an enlarged view of the area of  FIG. 2A  identified by bracket A 4 ; 
         FIG. 5  is a cross-sectional view of the turbine blade shown in  FIG. 4  taken along line  5 - 5 ; 
         FIG. 6  is a cross-sectional view of the turbine blade shown in  FIG. 4  taken along line  6 - 6 ; 
         FIG. 7  is a cross-sectional view of the turbine blade shown in  FIG. 4  taken along line  7 - 7 ; 
         FIG. 8  is a cross-sectional view of the turbine blade airfoil taken at the location indicated by line  8 - 8  in  FIG. 2A ; 
         FIG. 9  is a cross-sectional view of an alternative configuration of the turbine blade taken at the location indicated by line  2 - 2  in  FIG. 1 ; 
         FIG. 10  is a cross-sectional view of the turbine blade shown in  FIG. 9  taken along line  10 - 10 ; 
         FIG. 11  is a cross-sectional view of the turbine blade shown in  FIG. 9  taken along line  11 - 11 ; and 
         FIG. 12  is a cross-sectional view of the turbine blade airfoil taken at the location indicated by line  12 - 12  in  FIG. 9 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention. 
     Referring now to  FIG. 1 , a turbine blade  10  constructed in accordance with the present invention is illustrated. The blade  10  is adapted to be used in a gas turbine (not shown) of a gas turbine engine (not shown). The gas turbine engine includes a compressor (not shown), a combustor (not shown), and a turbine (not shown). The compressor compresses ambient air. The combustor combines compressed air with a fuel and ignites the mixture creating combustion products defining a high temperature working gas. The high temperature working gas travels to the turbine. Within the turbine are a series of rows of stationary vanes and rotating blades. Each pair of rows of vanes and blades is called a stage. Typically, there are four stages in a turbine. It is contemplated that the blade  10  described herein may define blade configuration for a third stage of blades in the gas turbine. 
     The stationary vanes and rotating blades are exposed to the high temperature working gas. To cool the vanes and blades, cooling air from the compressor is provided to the vanes and the blades. 
     The blade  10  includes an airfoil  12  and a blade root  14  which is used to conventionally secure the blade  10  to a rotor disk of the engine for supporting the blade  10  in the working medium flow path of the turbine where working medium gases exert motive forces on the surfaces thereof. The airfoil  12  has an outer wall  16  comprising a generally concave pressure sidewall  18  and a generally convex suction sidewall  20 . The pressure and suction sidewalls  18 ,  20  are joined together along an upstream leading edge  22  and a downstream trailing edge  24 . The leading and trailing edges  22 ,  24  are spaced axially or chordally from each other. The airfoil  12  extends radially along a longitudinal or radial direction of the blade  10 , defined by a span of the airfoil  12 , from a radially inner airfoil platform  26  to a radially outer blade tip  28 . 
     Referring to  FIGS. 2A and 2B , the airfoil  12  defines a radially extending cooling cavity  30  located between the pressure sidewall  18  and the suction sidewall  20  and extending between the blade root  14  and the blade tip  28 . A leading edge partition  32  extends radially through the cooling cavity  30  adjacent to the leading edge  22 . The leading edge partition  32  extends between the pressure and suction sidewalls  18 ,  20  to define a leading edge flow channel  34 . A trailing edge partition  36  extends radially through the cooling cavity  30  adjacent to the trailing edge  24 . The trailing edge partition  36  extends between the pressure and suction sidewalls  18 ,  20  to define a trailing edge flow channel  38 . 
     Referring to  FIG. 3 , a pressure side inner wall  40  extends radially within the cooling cavity  30  from a location adjacent the blade root  14  toward the blade tip  28 . The pressure side inner wall  40  extends from the leading edge partition  32  to the trailing edge partition  36  and is located in spaced relation to the interior surface  19  of the pressure sidewall  18  to define a pressure side near wall chamber  42 . A suction side inner wall  44  extends radially within the cooling cavity  30  from a location adjacent the blade root  14  toward the blade tip  28 . The suction side inner wall  44  extends from the leading edge partition  32  to the trailing edge partition  36  and is located in spaced relation to the interior surface  21  of the suction sidewall  20  to define a suction side near wall chamber  46 . The pressure side inner wall  40  and suction side inner wall  44  extend in converging relation toward each other and intersect at an intermediate location  48  intermediate the blade root  14  and the blade tip  28 . 
     The pressure side near wall chamber  42  may include a plurality of pin fins  50  to provide extended convection cooling surfaces and to increase the stiffness of the pressure sidewall  18 . Similarly, the suction side near wall chamber  46  may include a plurality of pin fins  52  for extending the convection cooling surfaces in the pressure side near wall chamber  46  and to increase the stiffness of the suction sidewall  20 . In addition, the pin fins  50 ,  52  increase the conduction of heat from the pressure and suction sidewalls  18 ,  20  to the respective pressure and suction side inner walls  40 ,  44 . 
     Referring to  FIGS. 2A ,  2 B,  4  and  8 , the upper or radially outer portion of the airfoil  12 , between the intermediate location  48  and the blade tip  28 , includes a plurality of chordally spaced mid-chord partitions  54   a ,  54   b ,  54   c ,  54   d ,  54   e ,  54   f  ( FIGS. 4 and 8 ) extending between the pressure sidewall  18  and the suction sidewall  20 . The mid-chord partitions  54   a ,  54   b ,  54   c ,  54   d ,  54   e ,  54   f  define a plurality of radially extending chordally spaced mid-chord channels  56   a ,  56   b ,  56   c ,  56   d ,  56   e ,  56   f ,  56   g  generally positioned along a chordal centerline  55  ( FIG. 8 ) of the airfoil  12 , i.e., located generally centrally between the pressure sidewall  18  and the suction sidewall  20 . The mid-chord channels  56   b  and  56   e  extend radially from the pressure side near wall chamber  42  to define pressure side channels, and the mid-chord channels  56   c  and  56   f  extend radially from the suction side chamber  46  to define suction side channels. In addition, the mid-chord channels  56   a ,  56   d  and  56   g  define return channels that extend radially past the intermediate location  48  to connect a tip passage  72 , extending chordally between the blade tip  28  and the upper edges of the mid-chord partitions  54   a ,  54   b ,  54   c ,  54   d ,  54   e ,  54   f  and the leading edge partition  32 , to a collection cavity  60  ( FIG. 3 ) located between the pressure side inner wall  40  and the suction side inner wall  44 . 
     Referring to  FIG. 5 , in which the pressure side channel  56   b  is typical of both pressure side channels  56   b ,  56   e , it can be seen that the pressure side inner wall  40  extends to the suction sidewall  20  such that the inner wall  40  is continuous with the inner surface  21  of the suction sidewall  20 . Referring to  FIG. 6 , in which the suction side channel  56   c  is typical of both suction side channels  56   c ,  56   f , it can be seen that the suction side inner wall  44  extends to the pressure sidewall  18  such that the inner wall  44  is continuous with the inner surface  19  of the pressure sidewall  18 . In addition, as seen in  FIG. 7 , in which the return channel  56   a  is typical of the return channels  56   a ,  56   d  and  56   g , a return passage  68  is defined between the return channel  56   a  and the collection cavity  60  to permit flow of spent cooling fluid from the return passage  56   a  into the collection cavity  60 . It should also be noted that the interior surfaces  19 ,  21  of pressure sidewall  18  and the suction sidewall  20  as well as surfaces of the inner walls  40 ,  44  may be provided with trip strips  70  to facilitate heat transfer at the boundary layer between the cooling fluid and the interior surfaces of the airfoil  12 . Trip strips  70  may additionally be provided along the interior surfaces of the leading edge flow channel  34  and trailing edge flow channel  38  ( FIGS. 2A and 2B ). 
     Referring to  FIGS. 2A ,  2 B and  3 , the pressure side near wall chamber  42  and suction side near wall chamber  46  are connected to cooling fluid supply openings  66   a ,  66   b ,  66   c  in the blade root  14  via respective conduits  62 ,  64 . Cooling fluid, such as cooling air supplied from the compressor for the gas turbine engine, flows from the conduits  62 ,  64  into the respective near wall chambers  42 ,  46  where heat is transferred to the cooling fluid from the lower half of the pressure sidewall  18  and suction sidewall  20  of the airfoil  12 . The cooling fluid further flows radially outwardly through the pressure side channels  56   b ,  56   e  and through the suction side channels  56   c ,  56   f  to the tip passage  72  extending chordally adjacent to the blade tip  28 . In addition, cooling fluid flows through the leading edge flow channel  34  to the tip passage  72 . 
     The spent cooling fluid flows radially inwardly through the return channels  56   a ,  56   d  and  56   g  and is collected in the collection cavity  60 . The spent cooling fluid further flows out of the collection cavity  60  though a trailing edge passage  74  ( FIG. 2A ) and into the trailing edge flow channel  38  where the fluid flows radially outwardly. The trailing edge  24  includes exit holes  76  and trailing edge slots  78  for providing a film of cooling fluid to the trailing edge  24 . Further, the leading edge  22  may be providing with exit holes  80  ( FIG. 8 ) extending from the leading edge flow channel  34  to provide a film of cooling fluid at the leading edge  22 . 
     The above described structure effectively provides a triple-pass serpentine path for the cooling fluid where the cooling fluid initially flows radially outwardly to the tip passage  72 , flows radially inwardly into the collection cavity  60  and then flows radially outwardly through the trailing edge flow channel  38 . Flow of the cooling fluid into the collection cavity  60  places the cooling fluid in contact with the interior surfaces of the inner walls  40 ,  44 , permitting the spent or warmed cooling fluid to transfer heat to the inner walls  40 ,  44  and thereby reduce the temperature differential between the inner walls  40 ,  44  and the pressure and suction sidewalls  18 ,  20 . In addition, the pin fins  50 ,  52  may conduct heat inwardly to the inner walls  40 ,  44  to reduce the thermal gradient. 
     The size and distribution or spacing of the pin fins  50 ,  52  may be selected based on the airfoil external heat load. Also, the heat transfer performance of the near wall chambers  42 ,  46  may be controlled by forming the near wall chambers  42 ,  46  as tapered convective channels to control the flow velocity in relation to the desired heat transfer. Further, it should be noted that the pressure and suction side channels  56   b ,  56   c ,  56   e ,  56   f  provide a reduced flow area, operating to accelerate the flow velocity of the cooling fluid it leaves the near wall chambers  42 ,  46  and thereby generates an increased heat transfer coefficient to maintain the cooling efficiency as the cooling fluid flows through the radially outer portion of the airfoil  12 . 
     Referring to  FIGS. 9-12  and alternative configuration for the turbine blade of the present invention is disclosed where elements corresponding to elements of the first described configuration are labeled with the same reference numeral increased by 100. 
     As in the first described configuration, the turbine blade  110  includes an airfoil with a pressure side inner wall  140  located adjacent a pressure sidewall  118  and a suction side inner wall  144  located adjacent a suction sidewall  120  to define near wall chambers  142  and  146 , respectively, at the radially inner portion of the airfoil  112 . In addition, flow channels  156   a - 156   g  are provided extending to a tip passage  172  from an intermediate location  148  at the outer end of the inner walls  140 ,  144 . 
     In the present configuration, each of the flow channels  156   a - 156   g  are in direct fluid communication with either the pressure side near wall chamber  142  or the suction side near wall chamber  146 . Specifically, each of the flow channels  156   b ,  156   d  and  156   f  extend from the pressure side near wall chamber  142  and comprise a structure, as illustrated for the flow channel  156   b  in  FIG. 10 ; and each of the flow channels  156   a ,  156   c ,  156   e  and  156   g  extend from the suction side near wall chamber  146 , as illustrated by the flow channel  156   g  in  FIG. 11 . The pressure side near wall chamber  142  is connected to a cooling fluid supply opening  166   a  in the blade root  114  by one or more conduits  162 . In addition, the blade root  114  includes an opening  166   b  that is covered by a cover plate  182 , and the suction side near wall chamber  146  is connected to the opening  166   b  by one or more conduits  164 . The opening  166   b  is further open to the trailing edge flow channel  138  ( FIG. 9 ). 
     A cooling fluid, such as cooling air supplied from the compressor, enters the blade  110  through the supply opening  166   a , flowing radially outwardly through the leading edge flow channel  134  and through the pressure side near wall chamber  142  and associated flow channels  156   b ,  156   d ,  156   f  to the tip passage  172 . From the tip passage  172 , the cooling fluid flows radially inwardly through the flow channels  156   a ,  156   c ,  156   e  and  156   g  and through the suction side near wall chamber  146 . The cooling fluid then passes through the conduits  164  to the opening  166   b , and subsequently flows radially outwardly through the trailing edge flow channel  138  and exits the airfoil  112  through exit holes  176  to trailing edge slots  178 . Accordingly, the cooling fluid circuit of the configuration described with reference to  FIGS. 9-12  provides a triple-pass serpentine path for the cooling fluid. 
     In both of the above described configurations, the pressure side near wall chamber  42 ,  142  is not in flow communication with the suction side near wall chamber  46 ,  146 , thus permitting the individual flow chambers to be individually designed based on the external heat load on the pressure sidewall  18 ,  118  and the suction sidewall  20 ,  120  of the airfoil  12 ,  112 . In addition, the individual flow channels may be designed with reference to the heat load at particular locations on the airfoil  12 ,  112 . Further, in both configurations of the cooling circuit, the triple-pass configuration comprises an aft flowing fluid path directing the cooling fluid to flow radially through separated mid-chord section near wall cooling paths, defined by the near wall chambers  42 ,  46  and  142 ,  146  and associated flow channels  56   a - g  and  156   a - g , as it flows to the trailing edge flow channel  38 ,  138 . 
     While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.