Patent Publication Number: US-8985949-B2

Title: Cooling system including wavy cooling chamber in a trailing edge portion of an airfoil assembly

Description:
FIELD OF THE INVENTION 
     The present invention relates to a cooling system in a turbine engine, and more particularly, to a cooling system including a wavy cooling chamber for cooling a trailing edge portion of an airfoil assembly. 
     BACKGROUND OF THE INVENTION 
     In gas turbine engines, compressed air discharged from a compressor section and fuel introduced from a source of fuel are mixed together and burned in a combustion section, creating combustion products defining a high temperature working gas. The working gas is directed through a hot gas path in a turbine section of the engine, where the working gas expands to provide rotation of a turbine rotor. The turbine rotor may be linked to an electric generator, wherein the rotation of the turbine rotor can be used to produce electricity in the generator. 
     In view of high pressure ratios and high engine firing temperatures implemented in modern engines, certain components, such as airfoil assemblies, e.g., stationary vanes and rotating blades within the turbine section, must be cooled with cooling fluid, such as air discharged from a compressor in the compressor section, to prevent overheating of the components. 
     SUMMARY OF THE INVENTION 
     In accordance with a first aspect of the present invention, an airfoil is provided in a gas turbine engine. The airfoil comprises an outer wall, a cooling fluid cavity, and a cooling system. The outer wall comprises a leading edge portion including a leading edge, a trailing edge portion including a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer end. A chordal direction is defined between the leading and trailing edges and a radial direction is defined between the radially inner and outer ends. The cooling fluid cavity is defined in the outer wall and receives cooling fluid for cooling the outer wall. The cooling system receives cooling fluid from the cooling fluid cavity for cooling the trailing edge portion of the outer wall and comprises a plurality of radially spaced-apart first impingement channels extending generally in the chordal direction from the cooling fluid cavity to a first cooling fluid chamber for delivering cooling fluid from the cooling fluid cavity to the first cooling fluid chamber. The cooling system further comprises a plurality of radially spaced-apart second impingement channels extending generally in the chordal direction from the first cooling fluid chamber to a second cooling fluid chamber for delivering cooling fluid from the first cooling fluid chamber to the second cooling fluid chamber. The cooling system still further comprises a plurality of radially spaced-apart third impingement channels extending generally in the chordal direction from the second cooling fluid chamber to a third cooling fluid chamber for delivering cooling fluid from the second cooling fluid chamber to the third cooling fluid chamber. The third cooling fluid chamber is defined by opposing first and second sidewalls comprising respective alternating angled sections that provide the third cooling fluid chamber with a zigzag shape. 
     In accordance with a second aspect of the present invention, an airfoil assembly is provided in a gas turbine engine. The airfoil assembly comprises a platform assembly and an airfoil comprising an outer wall, a cooling fluid cavity, and a cooling system. The outer wall is coupled to the platform assembly and comprises a leading edge portion including a leading edge, a trailing edge portion including a trailing edge, a pressure side, a suction side, a radially inner end, and a radially outer end. A chordal direction is defined between the leading and trailing edges and a radial direction is defined between the radially inner and outer ends. The cooling fluid cavity is defined in the outer wall and receives cooling fluid from the platform assembly for cooling the outer wall. The cooling system receives cooling fluid from the cooling fluid cavity for cooling the trailing edge portion of the outer wall and comprises a plurality of radially spaced-apart first impingement channels extending generally in the chordal direction from the cooling fluid cavity to a first cooling fluid chamber for delivering cooling fluid from the cooling fluid cavity to the first cooling fluid chamber. The first cooling fluid chamber has a direction of elongation in the radial direction. The cooling system further comprises a plurality of radially spaced-apart second impingement channels extending generally in the chordal direction from the first cooling fluid chamber to a second cooling fluid chamber for delivering cooling fluid from the first cooling fluid chamber to the second cooling fluid chamber. The second cooling fluid chamber has a direction of elongation in the radial direction. The cooling system still further comprises a plurality of radially spaced-apart third impingement channels extending generally in the chordal direction from the second cooling fluid chamber to a third cooling fluid chamber for delivering cooling fluid from the second cooling fluid chamber to the third cooling fluid chamber. The third cooling fluid chamber has a direction of elongation in the radial direction and is defined by opposing first and second sidewalls that comprise respective alternating angled sections that provide the third cooling fluid chamber with a zigzag shape when viewed from a radially outer side of the cooling system. The cooling system additionally comprises a plurality of outlet passages extending from the third cooling fluid chamber to the trailing edge of the outer wall. The outlet passages receive cooling fluid from the third cooling fluid chamber and discharge the cooling fluid from the airfoil. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Although 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 side cross-sectional view of an airfoil assembly including a cooling system according to an embodiment of the invention, wherein a portion of a suction side of the airfoil assembly has been removed; 
         FIG. 2  is cross-sectional view taken along line  2 - 2  in  FIG. 1 ;  FIG. 3  is an enlarged cross-sectional view of section  3 - 3  from  FIG. 2 ; 
         FIG. 3A  is an enlarged portion of  FIG. 3  to show details of the cooling system; 
         FIG. 4A  is an enlarged cross-sectional view similar to  FIG. 3  and showing a portion of a cooling system for an airfoil assembly according to another embodiment of the invention. and 
         FIG. 4B  is an enlarged cross-sectional view of section  4 B- 4 B from  FIG. 4A . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     In the following detailed description of the preferred embodiments, 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, specific preferred embodiments 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 , an airfoil assembly  10  constructed in accordance with an embodiment of the present invention is illustrated. In the illustrated embodiment, the airfoil assembly  10  is a blade assembly comprising an airfoil, i.e., a rotatable blade  12 , although it is understood that the cooling concepts disclosed herein could be used in combination with a stationary vane. The airfoil assembly  10  is for use in a turbine section  14  of a gas turbine engine. 
     As will be apparent to those skilled in the art, the gas turbine engine includes a compressor section (not shown), a combustor section (not shown), and the turbine section  14 . The compressor section includes a compressor that compresses ambient air, at least a portion of which is conveyed to the combustor section. The combustor section includes one or more combustors that mix the compressed air from the compressor section with a fuel and ignite the mixture creating combustion products defining a high temperature working gas. The high temperature working gas travels to the turbine section  14  where the working gas passes through one or more turbine stages, each turbine stage comprising a row of stationary vanes and a row of rotating blades  12 . It is contemplated that the airfoil assembly  10  illustrated in  FIG. 1  may be included in a first row of rotating blade assemblies in the turbine section  14 . 
     The vane and blade assemblies in the turbine section  14  are exposed to the high temperature working gas as the working gas passes through the turbine section  14 . Cooling air, e.g., from the compressor section, may be provided to cool the vane and blade assemblies, as will be described herein. 
     As shown in  FIG. 1 , the airfoil assembly  10  comprises the blade  12  and a platform assembly  16  that is coupled to a turbine rotor (not shown) and to which the blade  12  is affixed. The blade  12  comprises an outer wall  18  (see also  FIG. 2 ) that is affixed at a radially inner end  18 A thereof to the platform assembly  16 . 
     Referring to  FIGS. 1 and 2 , the outer wall  18  comprises a leading edge portion  20 A including a leading edge  20 , a trailing edge portion  22 A including a trailing edge  22  spaced from the leading edge  20  in a chordal direction C, a concave-shaped pressure side  24 , a convex-shaped suction side  26 , the radially inner end  18 A, and a radially outer end  18 B, wherein a spanwise or radial direction R D  is defined between the inner and outer ends  18 A,  18 B. It is noted that a portion of the suction side  26  of the blade  12  illustrated in  FIG. 1  has been removed to show selected internal structures within the blade  12 , which will be described herein. 
     As shown in  FIG. 2 , an inner surface  18 C of the outer wall  18  defines a hollow interior portion  28  extending between the pressure and suction sides  24 ,  26  from the leading edge portion  20 A to the trailing edge portion  22 A and from the inner end  18 A to the outer end  18 B. A plurality of rigid spanning structures  30  extend within the hollow interior portion  28  from the pressure side  24  to the suction side  26  and from the inner end  18 A to the outer end  18 B to provide structural rigidity for the blade  12  and to divide the hollow interior portion  28  into a plurality of sections, which will be described below. The spanning structures  30  may be formed integrally with the outer wall  18 . A conventional thermal barrier coating (not shown) may be provided on an outer surface  18 D of the outer wall  18  to increase the heat resistance of the blade  12 , as will be apparent to those skilled in the art. 
     As shown in  FIGS. 1 and 2 , a cooling fluid cavity  34  is defined in the outer wall  18  between the pressure and suction sides  24 ,  26 . The cooling fluid cavity  34  is located in the hollow interior portion  28  of the outer wall  18  and extends generally radially between the inner and outer ends  18 A,  18 B of the outer wall  18 . The cooling fluid cavity  34  receives cooling fluid from the platform assembly  16  for cooling the trailing edge portion  22 A of the outer wall  18 , as will be described below. 
     In accordance with the present invention, the airfoil assembly  10  is provided with a cooling system  40  for effecting cooling of the trailing edge portion  22 A of the blade  12 . As noted above, while the description of the cooling system  40  herein pertains to a blade assembly, it is contemplated that the concepts of the cooling system  40  of the present invention could be incorporated into a vane assembly. 
     As shown in  FIGS. 1-3 , the cooling system  40  is located in the hollow interior portion  28  of the outer wall  18  near the trailing edge portion  22 A. The cooling system  40  comprises a plurality of radially spaced-apart first impingement channels  44  that extend generally in the chordal direction through a first one  30 A of the spanning structures. The first impingement channels  44  are in fluid communication with the cooling fluid cavity  34  and provide cooling fluid from the cooling fluid cavity  34  to a first cooling fluid chamber  46 . As shown in  FIG. 1 , the first cooling fluid chamber  46  has a direction of elongation in the radial direction R D  and extends from the radially inner end  18 A to the radially outer end  18 B of the outer wall  18  in the embodiment shown, although the first cooling fluid chamber  46  need not extend all the way to the inner and outer ends  18 A,  18 B of the outer wall  18 . 
     The cooling system  40  further comprises a plurality of radially spaced-apart second impingement channels  48  that extend generally in the chordal direction through a second one  30 B of the spanning structures. The second impingement channels  48  are in fluid communication with the first cooling fluid chamber  46  and provide cooling fluid from the first cooling fluid chamber  46  to a second cooling fluid chamber  50 . As shown in  FIG. 1 , the second cooling fluid chamber  50  has a direction of elongation in the radial direction R D  and extends from the radially inner end  18 A to the radially outer end  18 B of the outer wall  18  in the embodiment shown, although the second cooling fluid chamber  50  need not extend all the way to the inner and outer ends  18 A,  18 B of the outer wall  18 . 
     The cooling system  40  still further comprises a plurality of radially spaced-apart third impingement channels  52  that extend generally in the chordal direction through a third one  30 C of the spanning structures. The third impingement channels  52  are in fluid communication with the second cooling fluid chamber  50  and provide cooling fluid from the second cooling fluid chamber  50  to a third cooling fluid chamber  54 . Referring to  FIG. 1 , the third cooling fluid chamber  54  has a direction of elongation in the radial direction R D  and extends from the radially inner end  18 A to the radially outer end  18 B of the outer wall  18  in the embodiment shown, although the third cooling fluid chamber  54  need not extend all the way to the inner and outer ends  18 A,  18 B of the outer wall  18 . 
     As shown most clearly in  FIG. 3 , the third cooling fluid chamber  54  is defined by opposing first and second sidewalls  56 ,  58 , which sidewalls  56 ,  58  in the embodiment shown are portions of the outer wall  18  that have outer surfaces  56 A,  58 A that define respective sections of the pressure and suction sides  24 ,  26  of the outer wall  18 . 
     Referring to  FIG. 3A , the first and second sidewalls  56 ,  58  that define the third cooling fluid chamber  54  comprise respective alternating angled sections  60 A,  61 A,  62 A,  63 A,  64 A and  60 B,  61 B,  62 B,  63 B,  64 B that are angled toward the respective suction and pressure sides  24 ,  26  of the outer wall  18  and provide the third cooling fluid chamber  54  with a wavy or zigzag shape when viewed from a radially outer side of the cooling system  40 , i.e., as shown in  FIGS. 2 ,  3 , and  3 A. As discussed herein, the even numbered sections, i.e., sections  60 A,  60 B,  62 A,  62 B,  64 A,  64 B, are referred to as “first sections” that are angled toward the suction side  26  of the outer wall  18 , and the odd numbered sections, i.e., sections  61 A,  61 B,  63 A,  63 B, are referred to as “second sections” that are angled toward the pressure side  24  of the outer wall  18 . It is noted that while the first sections are angled toward the suction side  26  of the outer wall  18  and the second sections are angled toward the pressure side  24  of the outer wall  18  the first sections could be angled toward the pressure side  24  and the second sections could be angled toward the suction side  26 . 
     As shown in  FIG. 3A , angles θ of the respective first sections taken with respect to the chordal direction C, as measured from respective central inflection points CI p1  of the first sections, may be substantially equal and opposite to angles β of the respective second sections taken with respect to the chordal direction C, as measured from respective central inflection points CI p2  of the second sections. The angles θ of the first sections are preferably within a range of about (15) to about (60) degrees relative to the chordal direction C, and the angles β of the second sections are preferably with a range about (−15) to about (−60) degrees relative to the chordal direction C. Further, opposed angled sections  60 A and  60 B,  61 A and  61 B,  62 A and  62 B,  63 A and  63 B,  64 A and  64 B of the respective first and second sidewalls  56 ,  58  are generally parallel to one another and define outer inflection points OI p1 , OI p2  at apices thereof. 
     Although the turns between the adjacent first and second sections of each of the first and second sidewalls  56 ,  58  in the embodiment shown comprise continuously curved walls, the turns could be defined by relatively sharp intersecting angles or by generally linear wall portions with rounded corners at the turns. The continuously curved turns in the embodiment shown effect a turning of the cooling fluid passing through the third cooling fluid chamber  54  and also provide a boundary layer restart for the cooling fluid, resulting in more flow turbulence and higher heat transfer through the third cooling fluid chamber  54 . 
     Moreover, while the first and second sidewalls  56 ,  58  in the embodiment shown each comprise a total of five alternating angled sections  60 - 64 A,  60 - 64 B, additional or fewer alternating angled sections may be provided. However, the first and second sidewalls  56 ,  58  according to an aspect of the present invention comprise at least a first section angled toward one of the pressure side  24  and the suction side  26  of the outer wall  18  in a downstream direction of cooling fluid flow through the cooling system  40 , and at least a second section extending from the first section and angled toward the other of the pressure side  24  and the suction side  26  of the outer wall  18  in the downstream direction. 
     Referring back to  FIG. 1 , the first and second impingement channels  44 ,  48  are preferably radially offset with respect to one another, and the second and third impingement channels  48 ,  52  are also preferably radially offset with respect to one another. Hence, cooling fluid passing out of the first and second impingement channels  44 ,  48  strikes against respective radially facing surfaces of the second and third spanning structures  30 B,  30 C to provide impingement cooling thereto, as will be discussed further below. The first impingement channels  44  may be generally radially aligned with the third impingement channels  52  as shown in  FIG. 1 , or the first impingement channels  44  may be radially offset from the third impingement channels  52 . 
     The cooling system  40  further comprises a plurality of radially spaced-apart outlet passages  70  extending from the third cooling fluid chamber  54  to the trailing edge  22  of the outer wall  18 , see  FIGS. 1-3 . The outlet passages  70  receive cooling fluid from the third cooling fluid chamber  54  and discharge the cooling fluid from the blade  12 , i.e., the cooling fluid exits the blade  12  of the airfoil assembly  10  via the outlet passages  70 . The cooling fluid is then mixed with the hot working gas passing through the turbine section  14 . The outlet passages  70  may be located along substantially the entire radial length of the outer wall  18 , or may be selectively located along the trailing edge  22  to fine tune cooling provided to specific areas. 
     Referring now to  FIGS. 1 and 2 , the platform assembly  16  includes an opening  72  formed therein in communication with the cooling fluid cavity  34 . The opening  72  allows cooling fluid to pass from a cooling supply  74  (see  FIG. 1 ) provided in the platform assembly  16  into the cooling fluid cavity  34 . The cooling supply  74  may receive cooling fluid, such as compressor discharge air, as is conventionally known in the art. 
     A portion of the cooling fluid flowing through the cooling fluid cavity  34  flows toward the radially outer end  18 B of the outer wall  18  where it passes through an opening (not shown) and into a second cooling fluid cavity  82 , see  FIG. 2 . This portion of cooling fluid provides convective cooling to the blade  12  as it flows radially inwardly through the second cooling fluid cavity  82 . Upon reaching the radially inner end  18 A of the outer wall  18  within the second cooling fluid cavity  82 , this portion of cooling fluid flows through an additional opening  76  in the platform  16 , then makes a 180-degree turn and passes through another opening  78  in the platform  16 , see  FIG. 1 . This cooling fluid then flows radially outwardly through a third cooling fluid cavity  84  so as to provide additional convective cooling to the blade  12 , see  FIG. 2 . This portion of cooling fluid is then discharged from the blade  12  in any conventional manner, such as, for example, via an outlet (not shown) at the radially outer end  18 B of the outer wall  18 . 
     The platform assembly  16  may be provided with an additional opening  80  (see  FIGS. 1 and 2 ) that supplies cooling fluid to a leading edge cavity  86  (see  FIG. 2 ). Cooling fluid is provided from the cooling supply  74  in the platform assembly  16  into the leading edge cavity  86  to provide cooling to the leading edge portion  20 A of the blade  12 , as will be apparent to those skilled in the art. 
     During operation, cooling fluid is provided to the cooling supply  74  in the platform assembly  16  in any known manner, as will be apparent to those skilled in the art. The cooling fluid passes from the cooling supply  74  into the cooling fluid cavity  34  via the opening  72  and into the leading edge cavity  86  via the opening  80 , see  FIGS. 1 and 2 . 
     The cooling fluid passing into the cooling fluid cavity  34  flows radially outwardly through the cooling fluid cavity  34 . Portions of the cooling fluid flow into the first impingement channels  44  of the cooling system  40 , and an additional portion of the cooling fluid flows into the second cooling fluid cavity  82  as described above. 
     The first impingement channels  44  provide metering of the cooling fluid, wherein the cooling fluid provides convective cooling to the blade  12  while passing through the first impingement channels  44 . The cooling fluid is discharged from the first impingement channels  44  into the first cooling fluid chamber  46 , wherein the cooling fluid provides impingement cooling to the radially facing surface of the second spanning structure  30 B as mentioned above. The cooling fluid also provides convective cooling to the blade  12  while flowing within the first cooling fluid chamber  46 . 
     The cooling fluid then flows into the second impingement channels  48 , which provide additional metering of the cooling fluid, wherein the cooling fluid provides convective cooling to the blade  12  while passing through the second impingement channels  48 . The cooling fluid is discharged from the second impingement channels  48  into the second cooling fluid chamber  50 , wherein the cooling fluid provides impingement cooling to the radially facing surface of the third spanning structure  30 C as mentioned above. The cooling fluid also provides convective cooling to the blade  12  while flowing within the second cooling fluid chamber  50 . 
     The cooling fluid then flows into the third impingement channels  52 , which provide further metering of the cooling fluid, wherein the cooling fluid provides convective cooling to the blade  12  while passing through the third impingement channels  52 . The cooling fluid is discharged from the third impingement channels  52  into the third cooling fluid chamber  54 . 
     Due to the configuration of the third cooling fluid chamber  54 , i.e., due to the alternating angled sections  60 - 64 A,  60 - 64 B of the first and second sidewalls  56 ,  58 , the effective length of the third cooling fluid chamber  54  is increased, as opposed to a cooling fluid chamber defined by generally planar sidewalls. Hence, the effective surface area of the first and second sidewalls  56 ,  58  that define the third cooling fluid chamber  54  is increased, so as to increase cooling to the outer wall  18  provided by the cooling fluid passing through the third cooling fluid chamber  54 , again, as opposed to a cooling fluid chamber defined by generally planar sidewalls. The cooling fluid provides convective cooling for the outer wall  18  of the blade  12  at the trailing edge portion  22 A as it flows within the third cooling fluid chamber  54 , and also provides impingement cooling to the sidewalls  56 ,  58  as a result of striking against the alternating angled sections  60 - 64 A,  60 - 64 B after passing the turns between the first and second sections of each of the first and second sidewalls  56 ,  58 . 
     The cooling fluid then flows from the third cooling fluid chamber  54  into the outlet passages  70 , wherein the cooling fluid provides additional convective cooling for the outer wall  18  of the blade  12  at the trailing edge  22  as it flows out of the blade  12  through the outlet passages  70 . It is noted that the diameters of the outlet passages  70  may be sized so as to meter the cooling fluid passing out of the cooling system  40 . It is further noted that each outlet passage  70  may have the same diameter size, or outlet passages  70  located at select radial locations may have different sized diameters so as to fine tune cooling provided to the outer wall  18  at the corresponding radial locations. 
     According to one aspect of the invention, the cooling system  40  may be formed using a sacrificial ceramic core (not shown), which is dissolved or melted to form the voids that define the respective portions of the cooling system  40 . Alternatively, the cooling system  40  may be formed by other suitable methods. 
     Referring to  FIG. 4A , a portion of a blade  112  of an airfoil assembly  110  according to another aspect of the invention is shown, wherein structure similar to that described above with reference to  FIGS. 1-3A  includes the same reference number increased by  100 . According to this aspect of the invention, turbulating features  100  comprising grooves in the embodiment shown are formed in the first and second sidewalls  156 ,  158 . As clearly shown in  FIG. 4A  and as shown in more detail in  FIG. 4B , a chordal spacing S P  between adjacent turbulating features  100  is generally equal to or less than a chordal width C W  of the turbulating features  100 . The turbulating features  100  effect a turbulation of cooling fluid flowing through the third cooling fluid chamber  154  so as to increase cooling provided to the outer wall  118 . It is noted that other types of turbulating features than grooves could be used, such as elongate ribs or small bumps or dimples formed in the sidewalls  156 ,  158 .