Patent Publication Number: US-10787913-B2

Title: Airfoil cooling circuit

Description:
BACKGROUND 
     The present invention relates generally to cooling components of gas turbine engines and more particularly to cooling circuits for stationary vanes. 
     Hollow stationary vanes of a turbine section of a gas turbine engine can require internal structures to achieve a desired cooling air flow velocity and heat transfer coefficient with a minimum amount of cooling flow, while limiting deflections or bulging of the airfoil walls resulting from differences in internal and external pressures during operation. Improved cooling circuits are needed to address both heat transfer and bulge requirements while reducing cooling flow requirements. 
     SUMMARY 
     An airfoil for a gas turbine engine includes an axial flow cooling circuit defined within an airfoil body and a radial flow cooling circuit defined between the baffle and the trailing edge. The axial flow cooling circuit includes a baffle disposed in spaced relation to an inner surface of the airfoil with a plurality of impingement cooling holes configured to direct a cooling fluid at an inner surface of the airfoil body. The baffle has an axial extent from the leading edge defined by an aft wall with the axial extent being substantially constant between inner and outer end walls and defined by a plane perpendicular to an engine axis. The radial flow cooling circuit includes a first radially-extending rib and a second radially-extending rib. The first rib is angled with respect to the baffle aft wall to define a first passage between the first rib and the baffle that tapers in cross-sectional area between the inner end wall and the outer end wall becoming larger in cross-sectional area in a direction of cooling fluid flow through the first passage. 
     A method of cooling an airfoil for a gas turbine engine includes flowing cooling fluid through an axial flow cooling circuit and flowing the cooling fluid through the radial flow cooling circuit. The axial flow cooling circuit includes flowing the cooling fluid from a cavity of a baffle through a plurality of cooling holes and directing the flow of cooling fluid from the plurality of cooling holes in an axial direction to a radial cooling circuit defined between the baffle and a trailing edge of the airfoil. The cavity extends between an inner end wall and an outer end wall of the airfoil and has an axial extent from the leading edge defined by an aft wall, with the axial extent being substantially constant between the inner and outer end walls and defined by a plane perpendicular to an engine axis. Flowing the cooling fluid through the radial flow cooling circuit includes flowing the cooling fluid through a first radially-extending passage that tapers outward in cross-sectional area between the inner end wall and the outer end wall in a direction of cooling fluid flow through the first passage, and flowing the cooling fluid through a second radially-extending passage that tapers inward in cross-sectional area between the inner end wall and the outer end wall in a direction of cooling fluid flow through the second passage. 
     The present summary is provided only by way of example, and not limitation. Other aspects of the present disclosure will be appreciated in view of the entirety of the present disclosure, including the entire text, claims, and accompanying figures. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a quarter-sectional view of a gas turbine engine. 
         FIG. 2  is a schematized perspective view of a turbine section of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a schematized perspective view of one embodiment of a cooling circuit of a stator airfoil of  FIG. 2 . 
         FIG. 4  is a schematized perspective view of another embodiment of a cooling circuit of the stator airfoil of  FIG. 2 . 
         FIG. 5  is a schematized perspective view of yet another embodiment of a cooling circuit of a stator airfoil. 
     
    
    
     While the above-identified figures set forth one or more embodiments of the present disclosure, other embodiments are also contemplated, as noted in the discussion. In all cases, this disclosure presents the invention by way of representation and not limitation. It should be understood that numerous other modifications and embodiments can be devised by those skilled in the art, which fall within the scope and spirit of the principles of the invention. The figures may not be drawn to scale, and applications and embodiments of the present invention may include features and components not specifically shown in the drawings. 
     DETAILED DESCRIPTION 
       FIG. 1  is a quarter-sectional view of a gas turbine engine  20  that includes fan section  22 , compressor section  24 , combustor section  26  and turbine section  28 . Fan section  22  drives air along bypass flow path B while compressor section  24  draws air in along core flow path C where air is compressed and communicated to combustor section  26 . In combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through turbine section  28  where energy is extracted and utilized to drive fan section  22  and compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a low-bypass turbine engine, or a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
     The example engine  20  generally includes low speed spool  30  and 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. 
     Low speed spool  30  generally includes inner shaft  40  that connects fan  42  and low pressure (or first) compressor section  44  to low pressure (or first) turbine section  46 . Inner shaft  40  drives fan  42  through a speed change device, such as geared architecture  48 , to drive fan  42  at a lower speed than low speed spool  30 . High-speed spool  32  includes outer shaft  50  that interconnects high pressure (or second) compressor section  52  and high pressure (or second) turbine section  54 . Inner shaft  40  and outer shaft  50  are concentric and rotate via bearing systems  38  about engine central longitudinal axis A. 
     Combustor  56  is arranged between high pressure compressor  52  and high pressure turbine  54 . In one example, high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of low pressure turbine  46  as related to the pressure measured at the outlet of low pressure turbine  46  prior to an exhaust nozzle. 
     Mid-turbine frame  58  of engine static structure  36  is arranged generally between high pressure turbine  54  and low pressure turbine  46 . Mid-turbine frame  58  further supports bearing systems  38  in turbine section  28  as well as setting airflow entering low pressure turbine  46 . 
     The core airflow C is compressed by low pressure compressor  44  then by high pressure compressor  52  mixed with fuel and ignited in combustor  56  to produce high speed exhaust gases that are then expanded through high pressure turbine  54  and low pressure turbine  46 . Mid-turbine frame  57  includes airfoils/vanes  60 , which are in the core airflow path and function as an inlet guide vane for low pressure turbine  46 . Utilizing vanes  60  of mid-turbine frame  58  as inlet guide vanes for low pressure turbine  46  decreases the length of low pressure turbine  46  without increasing the axial length of mid-turbine frame  58 . Reducing or eliminating the number of vanes in low pressure turbine  46  shortens the axial length of turbine section  28 . Thus, the compactness of gas turbine engine  20  is increased and a higher power density may be achieved. 
     Each of the compressor section  24  and the turbine section  28  can include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. To improve efficiency, static outer shroud seals (not shown), such as a blade outer air seal (BOAS), can be located radially outward from rotor airfoils to reduce tip clearance and losses due to tip leakage. 
       FIG. 2 . is a schematized perspective view of high pressure turbine section  54 , which can include alternating rows of rotor assemblies  58  and stationary vane assemblies  61  (only one of which is shown). The illustrated stationary vane assembly  61  includes a plurality of vanes  62 . Each vane  62  includes radially inner and outer end walls  64 ,  66  joined by airfoil body  68  having leading edge  70  and trailing edge  72 . Airfoil body  68  includes internal cooling circuit  74 , through which cooling fluid F c  can flow (indicated with arrows). Cooling fluid F c  can be provided to vane  62  by any source of cooling fluid, such as bleed air, sourced from a location upstream of stationary vane assembly  61 . 
       FIG. 3  is a schematized perspective view of vane  62  with cooling circuit  74 . Cooling circuit  74  includes axial flow cooling circuit  76  and radial flow cooling circuit  78 . Axial flow cooling circuit  76  is defined within airfoil body  68  adjacent to leading edge  70  and is configured to cool leading edge  70  and up to 60 percent of chord length of airfoil body  68  from leading edge  70 . Radial flow cooling circuit  78  is defined within airfoil body  68  aft of axial flow cooling circuit and is configured to direct cooling fluid F c  through a series of predominantly radially-extending passages before cooling fluid F c  exits airfoil body  68  through trailing edge  72 . Axial flow cooling circuit  76  and radial flow cooling circuit  78  are characterized by carrying predominantly axial and radial cooling flow, respectively. 
     Axial flow cooling circuit  74  includes baffle  80  disposed in airfoil cavity  81  in spaced relation to inner surface  82  of airfoil body  68 . Baffle  80  can be formed from a metallic material, ceramic matrix composite (CMC) material, or other suitable material. Baffle  80  is a hollow structure having cavity  84  bounded by a U-shaped wall, which generally corresponds to a shape of inner surface  82 , and aft wall  86 , which can have a substantially flat surface. U-shaped wall includes a forward edge portion  88 , disposed adjacent to and in spaced relation to inner surface  82  along leading edge  70 , and opposing side walls  90 ,  92 , disposed adjacent to and in spaced relation to inner surface  82  along the pressure and suction sidewalls  93 ,  94  of the airfoil, respectively. Baffle  80  is configured to effectively reduce a cross-sectional area of airfoil cavity  81  to increase cooling along leading edge  70 . Baffle  80  can be a straight baffle with baffle aft wall  86  extending perpendicularly to inner end wall  64 , parallel to leading edge  70 , or in a plane perpendicular to engine axis A, such that baffle  80  has an axial extent from leading edge  70  that is substantially constant between inner end wall  64  and outer end wall  66 . In some embodiments, a cross-sectional area of baffle cavity  84  can remain substantially constant over the span of the airfoil body  68 . The use of a straight baffle allows for a reduction in cross-sectional area of airfoil body cavity  81  over a greater axial extent or airfoil chord length than a small end of a tapering baffle. Baffle  80  can generally extend from adjacent leading edge  70  to 30 percent to 60 percent of the chord length from leading edge  70 . Preferably, baffle  80  extends as far axially as possible to reduce the cross-sectional area of airfoil cavity  81 . The axial extent of baffle  80  is generally limited by the need for radial ribs to limit bulging or deflections of the airfoil walls. 
     Baffle  80  includes a plurality of impingement cooling holes  95  positioned along forward edge portion  88  to direct cooling fluid F c  along the inner surface of leading edge  70 . Impingement cooling holes  95  can be evenly sized and distributed along a radial length of forward edge portion  88  in one or more radially-extending rows. The size and distribution of impingement cooling holes  95  can be varied in alternative embodiments to tailor impingement cooling as may be necessary to target hot spots along leading edge  70 . For example, the density of impingement cooling holes  95  can be increased in regions corresponding to hot spots along leading edge  70 . Unlike conventional impingement baffles, aft wall  86  and side walls  90 ,  92  of baffle  80  are free of impingement cooling holes  95 . By limiting impingement cooling holes to the location of forward edge portion  88 , baffle  80  can increase heat transfer along leading edge  70  where heat load is highest by focusing all impingement cooling at the inner surface of leading edge  70 . 
     Cooling fluid F c  that impinges upon the inner surface of leading edge  70  is directed axially along inner surface  82  between inner surface  82  and baffle side walls  90 ,  92 . A plurality of axially-extending U-shaped ribs  96  can be disposed along inner surface  82  to channel or direct cooling fluid F c  that has exited impingement cooling holes  95  in an axial direction toward aft wall  86  and radial cooling circuit  78 . Ribs  96  can be distributed evenly as a function of span as shown in the embodiments represented in  FIGS. 2-4  or can be distributed non-uniformly as a function of span to achieve desired heat transfer at various radial locations along a span of airfoil body  68 . Heat transfer can be optimized by spacing ribs  96  to cover regions of interest such that hot regions are cooled and cold regions are not overcooled. Ribs  96  can extend from aft wall  86  along side wall  90 , around forward edge region  88 , and back to aft wall  86  along side wall  92 . Ribs  96  can extend substantially axially along side walls  90 ,  92 . Ribs  96  can be configured to contact forward edge portion  88  and walls  90 ,  92  of baffle  80  for locating baffle  80  during assembly and to limit radial flow of cooling fluid F c  through axial cooling circuit  76 . Ribs  96  can be formed integrally with airfoil body  68  via casting or additive manufacturing methods. In alternative embodiments ribs  96  can be formed on an outer wall of baffle  80 . 
     In some embodiments, a plurality of heat transfer features  98  (shown in phantom) can be disposed along inner surface  82  adjacent one or more side walls  90 ,  92  to increase heat transfer in the leading edge region of airfoil body  68 .  FIG. 3  shows these heat transfer features as pedestals, but the heat transfer features could also be trip strips, dimples, or other heat transfer features known in the art. Heat transfer features  98  can be used to move and redistribute cooling fluid F c  and can increase thermal heat transfer through the pressure and suction sidewalls  93 ,  94  of airfoil body  68 . Although illustrated only in a portion of axial cooling circuit  76 , heat transfer features  98  can be distributed along the full span of airfoil body  68  along baffle  80 . The distribution of heat transfer features  98  can be tailored to address regions of high heat load. For example, the concentration of heat transfer features can be increased in a region near leading edge  70  where heat load is highest and can be decreased over an axial extent toward baffle aft wall  86  as heat load decreases. 
     Cooling fluid F c  can enter baffle cavity  84  through inner end wall  64 , as shown in  FIG. 3  (indicated by arrow), or through outer end wall  66 . The construction of axial flow cooling circuit  76  and radial flow cooling circuit  78  can remain the same regardless of the direction in which cooling fluid F c  enters baffle cavity  84 . Cooling fluid F c  exits baffle cavity  84  through impingement cooling holes  95  and flows axially between adjacent ribs  96  toward baffle aft wall  86  and into first radially-extending passage  100  of radial flow cooling circuit  78 . The velocity of cooling fluid F c  between baffle  80  and airfoil body  68  in axial flow cooling circuit  76  can be tailored by modifying the spacing between baffle  80  and the inner surface of airfoil body  68  or by otherwise increasing or decreasing the cross-sectional area through which cooling fluid F c  flows. 
     Radial flow cooling circuit  78  can be designed to maintain a velocity of cooling fluid F c  exiting axial flow cooling circuit  76 . Radial flow cooling circuit  78  includes radially-extending ribs  102 ,  104 , which connect suction and pressure sidewalls of airfoil body  68  to define three cooling fluid passages  100 ,  106 ,  108 . Radially-extending rib  102  and baffle aft wall  86  define forward passage  100 ; radially-extending ribs  102  and  104  define central passage  106 ; and radially-extending rib  104  and trailing edge region  110  define aft passage  108 . To maintain cooling flow velocity F c , rib  102  is angled with respect to baffle aft wall  86 , such that forward passage  100  tapers in cross-sectional area between inner end wall  64  and outer end wall  66  becoming larger in cross-sectional area in the direction of cooling fluid flow through forward passage  100 . As illustrated in  FIG. 3 , cooling fluid F c  can flow from outer end wall  66  to inner end wall  64 . The cross-sectional area of forward passage  100  becomes larger as cooling fluid F c  is added from axial flow cooling circuit  76 . As illustrated in  FIG. 3 , axial flow cooling circuit  76  dumps cooling fluid F c  into forward passage  100  at locations along the airfoil span defined by axially-extending ribs  96  such that a volume of cooling fluid F c  increases in passage  100  from outer end wall  66  to inner end wall  64 . 
     A turn  112  (shown in  FIG. 2 ) connects forward passage  100  to central passage  106  at inner end wall  64  to channel cooling fluid F c  from forward passage  100  to central passage  106 . To maintain cooling fluid velocity, central passage  106  can have a substantially uniform cross-sectional shape over the span of the airfoil with rib  102  extending parallel to rib  104 . In alternative embodiments, a portion of cooling fluid F c  can be bled off through sidewalls of airfoil body  68  for film cooling of external surfaces of the airfoil. In these embodiments, central passage  106  can be tapered in cross-sectional area to maintain cooling fluid velocity as cooling fluid is bled from central passage  106 . As illustrated in  FIG. 3 , cooling fluid F c  flows through central passage  106  in a direction opposite to cooling fluid flow through forward passage  100 , (i.e., from inner end wall  64  to outer end wall  66 ). 
     A second turn  114  (shown in  FIG. 2 ) connects central passage  106  to aft passage  108  at outer end wall  66  to channel cooling fluid F c  from central passage  106  to aft passage  108 . Aft passage  108  connects radial flow cooling circuit  78  with trailing edge region  110 . Trailing edge region  110  includes a plurality of radially-spaced axially-extending ribs  116 , which channel cooling fluid F c  from radial flow cooling circuit  78  out of airfoil body  68  at trailing edge  72 . As shown in  FIG. 3 , cooling fluid F c  flows in a substantially radial direction through aft passage  108  from outer end wall  66  to inner end wall  64 . As cooling fluid F c  flows through aft passage  108 , a portion of cooling fluid F c  is exhausted through trailing edge slots (defined between adjacent ribs  116 ), flowing in an axial direction between adjacent ribs  116 . To maintain cooling fluid velocity through aft passage  108 , rib  104  can be angled with respect to trailing edge region  110  (or trailing edge  72 ) such that aft passage  108  tapers in cross-sectional area between inner end wall  64  and outer end wall  66  becoming smaller in cross-sectional area in the direction of cooling fluid flow through aft passage  108 . As illustrated in  FIG. 3 , cooling fluid F c  flows from outer end wall  66  to inner end wall  64 . The cross-sectional area of aft passage  108  becomes smaller as cooling fluid F c  is exhausted through trailing edge region  110 . As illustrated in  FIG. 3 , radial flow cooling circuit  78  exhausts cooling fluid F c  through trailing edge slots at locations along the airfoil span defined by axially-extending ribs  116  such that a volume of cooling fluid F c  decreases in passage  108  from outer end wall  66  to inner end wall  64 . In some embodiments, trailing edge region  110  can include axial ribs, oblong pedestals, round pedestals, and combinations thereof (not shown) to direct flow into trailing edge slots and prevent flow separation in trailing edge slots. 
     Radial flow cooling circuit  78  can include heat transfer features  118  to enhance heat transfer over the length of passages  100 ,  106 ,  108 .  FIG. 3  illustrates chevron-shaped trip strips  118  in each passage  100 ,  106 , and  108  pointing in a direction opposite the flow of cooling fluid F c  and located with non-uniform spacing. As will be understood by one of ordinary skill in the art, heat transfer features  118  can have different shapes, orientations, and spacing, or can otherwise be tailored to address different heat loads at different locations of airfoil body  68 . For example, trip strips can be concentrated or more closely spaced in areas of high heat load. 
       FIG. 4  is a schematized perspective view of vane  62  with alternative cooling circuit  74 ′. Cooling circuit  74 ′ is similar to cooling circuit  74  and, therefore, disclosure pertaining to cooling circuit  74  can be applied to cooling circuit  74 ′ with the modifications disclosed herein. Cooling circuit  74 ′ includes axial flow cooling circuit  76 ′ and radial flow cooling  78 ′. Like cooling circuit  74 , axial flow cooling circuit  76 ′ is defined within airfoil body  68  adjacent to leading edge  70  and is configured to cool leading edge  70  and up to 60 percent of an axial chord length of airfoil body  68  from leading edge  70 . Radial flow cooling circuit  78 ′ is defined within airfoil body  68  aft of axial flow cooling circuit and is configured to direct cooling fluid F c  through a series of radially-extending passages before cooling fluid F c  exits airfoil body  68  through trailing edge  72 . 
     Axial flow cooling circuit  76 ′ includes baffle  80  as described with respect to  FIG. 3 . Axial flow cooling circuit  76 ′ is configured similarly to axial flow cooling circuit  76 , but includes modified axially-extending U-shaped ribs  96 ′, which are angled with respect to inner end wall  64 , while maintaining a substantially axially-extending orientation. Modified ribs  96 ′ are angled to direct cooling fluid F c  toward a direction of cooling fluid flow through forward passage  100 ′ of radial flow cooling circuit  78 ′ to improve flow dynamics at the intersection of axial flow cooling circuit  76 ′ and radial flow cooling circuit  78 ′ 
     Cooling fluid F c  can enter baffle cavity  84  through outer end wall  66 , as shown in  FIG. 4  (indicated by arrow), or through inner end wall  64 . The construction of axial flow cooling circuit  76 ′ and radial flow cooling circuit  78 ′ can remain the same regardless of the direction in which cooling fluid F c  enters baffle cavity  84 . 
     Radial flow cooling circuit  78 ′ can be designed to maintain a velocity of cooling fluid F c  exiting axial flow cooling circuit  76 ′ as described with respect to radial flow cooling circuit  78  in  FIG. 3 . Radial flow cooling circuit  78 ′ includes radially-extending ribs  102 ′,  104 ′, which connect pressure and suction sidewalls  93 ,  94  of airfoil body  68  to define three cooling fluid passages  100 ′,  106 ′,  108 ′. Radially-extending rib  102 ′ and baffle aft wall  86  define forward passage  100 ′; radially-extending ribs  102 ′ and  104 ′ define central passage  106 ′; and radially-extending rib  104 ′ and trailing edge region  110  define aft passage  108 ′. To maintain cooling flow velocity F c , rib  102 ′ is angled with respect to baffle aft wall  86 , such that forward passage  100 ′ tapers in cross-sectional area between inner end wall  64  and outer end wall  66  becoming larger in cross-sectional area in the direction of cooling fluid flow through forward passage  100 ′. As illustrated in  FIG. 4 , cooling fluid F c  can flow through forward passage  100 ′ from inner end wall  64  to outer end wall  66 . To accommodate the addition of cooling fluid F c  into forward passage  100 ′, the cross-sectional area of forward passage  100 ′ tapers outward from inner end wall  64  to outer end wall  66 . 
     Modified turn  112 ′ (shown in phantom) connects forward passage  100 ′ to central passage  106 ′ at outer end wall  66  to channel cooling fluid F c  from forward passage  100 ′ to central passage  106 ′. As disclosed with respect to radial flow cooling circuit  78  of  FIG. 3 , central passage  106 ′ can be configured to maintain the cooling fluid velocity. As illustrated in  FIG. 4 , cooling fluid F c  flows through central passage  106 ′ in a direction opposite to flow through forward passage  100 ′, from outer end wall  66  to inner end wall  64 . Modified turn  114 ′ (shown in phantom) connects central passage  106 ′ to aft passage  108 ′ at inner end wall  64  to channel cooling fluid F c  from central passage  106 ′ to aft passage  108 ′. Aft passage  108 ′ connects radial flow cooling circuit  78 ′ with trailing edge region  110 , which exhausts air from radial flow cooling circuit  78 ′ as described with respect to radial flow cooling circuit  78 . As illustrated in  FIG. 4 , cooling fluid F c  flows through aft passage  108 ′ from inner end wall  64  to outer end wall  66 . To maintain cooling fluid velocity, the cross-sectional area of aft passage  108 ′ becomes smaller as cooling fluid F c  is exhausted through trailing edge region  110 . 
     Baffle placement is not limited to the leading edge cavity and baffle shape is not limited to the shape shown  FIGS. 2-4 . In some embodiments, the baffle can be located aft of and separate from an airfoil leading edge cooling circuit and can have a shape corresponding to the location of placement.  FIG. 5  is a schematized perspective view of another embodiment of a cooling circuit of a stator airfoil in which the baffle is spaced apart from a leading edge cooling circuit.  FIG. 5  shows vane  62 ″, which can replace vanes  62 ,  62 ′ of the disclosed gas turbine engine. Similar to stator vanes  62 ,  62 ′, vane  62 ″ has cooling circuit  74 ″, which includes axial flow cooling circuit  76 ″ and radial flow cooling circuit  78 ″. In addition, vane  62 ″ includes leading edge cooling circuit  120 . Axial and radial flow cooling circuits  76 ″,  78 ″ are similar in design to the axial and radial flow cooling circuits  76 ,  76 ′,  78 ,  78 ′ disclosed in  FIGS. 2-4 , with the exception of baffle  122 , which has a forward wall  124  corresponding to a shape of radially-extending rib  126  of leading edge cooling circuit  120 . Vane  62 ″ benefits from the advantages provided by a straight baffle coupled with a tapered radial flow cooling circuit, while providing a separate cooling circuit for leading edge  70 . 
     Leading edge cooling circuit  120  can include radial flow passage  128  and axial flow passage  130  separated by radially-extending rib  132 . Radial flow passage  128  is defined by opposing pressure and suction sidewalls  93 ,  94 , and by opposing radially-extending ribs  126  and  132 , which connect pressure and suction sidewalls  93 ,  94  of airfoil body  68  along the span. Radially-extending rib  132  can include a plurality of impingement cooling holes  134 , through which cooling air is directed from radial flow passage  128  to axial flow passage  130  to impinge upon the inner surface of leading edge  70  before exiting vane  62 ″ through leading edge cooling holes  136 . Leading edge cooling fluid F LE  can enter leading edge cooling circuit  120  from outer end wall  66  as shown in  FIG. 5  (indicated by arrow) or from inner end wall  64 . The use of leading edge cooling circuit  120  provides dedicated cooling to leading edge  70 , while axial flow cooling circuit  76 ″ provides cooling to pressure and suction sidewalls  93 ,  94 . 
     Axial flow cooling circuit  76 ″ includes baffle  122 , which can be a straight baffle with both baffle forward wall  124  and baffle aft wall  138  extending perpendicularly to inner end wall  64 , parallel to leading edge  70 , or in a plane perpendicular to engine axis A, such that baffle  122  has an axial extent from leading edge  70  that is substantially constant between inner end wall  64  and outer end wall  66 . In some embodiments, a cross-sectional area of baffle  122  can remain substantially constant over the span of the airfoil body  68 . The use of a straight baffle allows for a reduction in cross-sectional area of airfoil body cavity  81  over a greater axial chord length than a small end of a tapering baffle. Baffle  122  can be positioned in close proximity to or abutting radially-extending rib  126  of leading edge cooling circuit  120  with side walls  140 ,  142  in spaced relation to pressure and suction sidewalls  93 ,  94  of airfoil body  68 , respectively. Baffle  122  can generally extend from radially-extending rib  126  to up to 60 percent of the airfoil chord length from leading edge  70 . Preferably, baffle  122  extends as far axially as possible to reduce the cross-sectional area of airfoil cavity  81 . The axial extent of baffle  122  is generally limited by the need for radial ribs to limit bulging or deflections of the airfoil walls. 
     Baffle  122  includes a plurality of impingement cooling holes  144  positioned along opposing side walls  140 ,  142  to direct cooling air to pressure and suction sidewalls  93 ,  94 , respectively. Impingement cooling holes  144  can be evenly sized and distributed along a radial length of baffle  122  in one or more radially-extending rows. The size and distribution of impingement cooling holes  144  can be varied in alternative embodiments to tailor impingement cooling as may be necessary to target hot spots along the span of airfoil body  68  and pressure and suction sidewalls  93 ,  94 . Generally, the density of impingement cooling holes  144  can be concentrated along side walls  140 ,  142  toward baffle forward wall  124 , with few or no impingement cooling holes  144  in close proximity to baffle aft wall  138 . Baffle  122  can be free of impingement cooling holes on forward wall  124  and aft wall  138 , as radially-extending rib  126  adjacent to forward wall  124  is cooled by leading edge cooling fluid F LE  and baffle aft wall  138  is cooled by radial flow cooling circuit  78 ″ 
     Cooling fluid F c  that impinges upon the inner surface of pressure and suction sidewalls  93 ,  94  is directed axially along the inner surface of pressure and suction sidewalls  93 ,  94  and outer surface of baffle side walls  140 ,  142 . A plurality of axially-extending ribs  146  can be disposed along the inner surface of pressure and suction sidewalls  93 ,  94  to channel or direct cooling fluid F c  that has exited impingement cooling holes  144  in an axial direction toward aft wall  138  and radial cooling circuit  78 ″. Ribs  146  can be distributed evenly as a function of span as shown in the embodiment represented in  FIG. 5  or can be distributed non-uniformly as a function of span to achieve desired heat transfer at various radial locations along a span of airfoil body  68 . External heat transfer regions may not be uniform along the airfoil span. Heat transfer can be optimized by spacing ribs to cover a region of interest, such that hot regions are cooled and cold regions are not overcooled. Ribs  146  can extend along pressure and suction sidewalls  93 ,  94  from baffle forward wall  124  to baffle aft wall  138 . Ribs  146  can extend substantially axially along pressure and suction sidewalls  93 ,  94  or can be angled in a manner consistent with  FIG. 4  to direct cooling fluid F c  toward a direction of cooling fluid flow through forward passage  100 ″ of radial flow cooling circuit  78 ″. Ribs  146  can be configured to contact side walls  140 ,  142  of baffle  122  for locating baffle  122  during assembly and to limit radial flow of cooling fluid F c  through axial cooling circuit  76 ″. Ribs  146  can be formed integrally with airfoil body  68  via casting or additive manufacturing methods. In alternative embodiments ribs  144  can be formed on an outer wall of baffle  122 . 
     In some embodiments, a plurality of heat transfer features  148  can be disposed along the inner surface of pressure and suction sidewalls  93 ,  94  adjacent one or more baffle side walls  140 ,  142  to increase heat transfer as needed.  FIG. 5  shows these heat transfer features as chevron-shaped trip strips, but the heat transfer features could also be pedestals, dimples, trip strips of other shapes, or other heat transfer features known in the art. Heat transfer features  148  can be used to move and redistribute cooling fluid F c  and can increase thermal heat transfer through the pressure and suction sidewalls  93 ,  94  of airfoil body  68 . The distribution of heat transfer features  148  can be tailored to address regions of high heat load. 
     Cooling fluid F c  can enter baffle cavity  150  through outer end wall  66 , as shown in  FIG. 5  (indicated by arrow), or through inner end wall  64 . The construction of axial flow cooling circuit  76 ″ and radial flow cooling circuit  78 ″ can remain the same regardless of the direction in which cooling fluid F c  enters baffle cavity  150 . Cooling fluid F c  exits baffle cavity  150  through impingement cooling holes  144  and flows axially between adjacent ribs  146  toward baffle aft wall  138  and into first radially-extending passage  100 ″ of radial flow cooling circuit  78 ″. The velocity of cooling fluid F c  between baffle  122  and airfoil body  68  in axial flow cooling circuit  76 ″ can be tailored by modifying the spacing between baffle  122  and the inner surface of airfoil body  68  or by otherwise increasing or decreasing the cross-sectional area through which cooling fluid F c  flows. 
     Radial flow cooling circuit  78 ″ can be designed to maintain a velocity of cooling fluid F c  exiting axial flow cooling circuit  76 ″ as described with respect to radial flow cooling circuits  78  and  78 ′. Radial flow cooling circuit  78 ″ includes radially-extending ribs  102 ″,  104 ″, which connect pressure and suction sidewalls  93 ,  94  of airfoil body  68  to define three cooling fluid passages  100 ″,  106 ″,  108 ″. Radially-extending rib  102 ″ and baffle aft wall  138  define forward passage  100 ″; radially-extending ribs  102 ″ and  104 ″ define central passage  106 ″; and radially-extending rib  104 ″ and trailing edge region  110  define aft passage  108 ″. To maintain cooling flow velocity F c , rib  102 ″ is angled with respect to baffle aft wall  138 , such that forward passage  100 ″ tapers in cross-sectional area between inner end wall  64  and outer end wall  66  becoming larger in cross-sectional area in the direction of cooling fluid flow through forward passage  100 ″. As illustrated in  FIG. 5 , cooling fluid F c  can flow through forward passage  100 ″ from outer end wall  66  to inner end wall  64 . To accommodate the addition of cooling fluid F c  into forward passage  100 ″, the cross-sectional area of forward passage  100 ″ tapers outward from outer end wall  66  to inner end wall  64 . 
     Radial flow cooling circuit  78 ″ can have turns consistent with turns  112 ,  114 , as described with respect to  FIGS. 2 and 3  to form a serpentine cooling flow pathway. As disclosed with respect to radial flow cooling circuit  78  of  FIG. 3 , central passage  106 ″ can be configured to maintain the cooling fluid velocity. As illustrated in  FIG. 5 , cooling fluid F c  flows through central passage  106 ″ in a direction opposite to flow through forward passage  100 ″, from inner end wall  64  to outer end wall  66 . Aft passage  108 ″ connects radial flow cooling circuit  78 ″ with trailing edge region  110 , which exhausts air from radial flow cooling circuit  78 ″ as described with respect to radial flow cooling circuit  78 . As illustrated in  FIG. 5 , cooling fluid F c  flows through aft passage  108 ″ from outer end wall  66  to inner end wall  64 . To maintain cooling fluid velocity, the cross-sectional area of aft passage  108 ″ becomes smaller as cooling fluid F c  is exhausted through trailing edge region  110 . 
     The disclosed cooling circuit with straight baffle  80  and tapered radial flow passages addresses both heat transfer and bulge requirements while reducing cooling flow requirements. As disclosed herein, the cooling circuit is customizable and can be adapted to a variety of airfoil configurations. While the disclosed cooling circuit has been described with respect to a turbine vane, it should be understood that that it can be used for other types of vanes, as well as rotor blades. 
     Summation 
     Any relative terms or terms of degree used herein, such as “substantially”, “essentially”, “generally”, “approximately” and the like, should be interpreted in accordance with and subject to any applicable definitions or limits expressly stated herein. In all instances, any relative terms or terms of degree used herein should be interpreted to broadly encompass any relevant disclosed embodiments as well as such ranges or variations as would be understood by a person of ordinary skill in the art in view of the entirety of the present disclosure, such as to encompass ordinary manufacturing tolerance variations, incidental alignment variations, transient alignment or shape variations induced by thermal, rotational or vibrational operational conditions, and the like. Moreover, any relative terms or terms of degree used herein should be interpreted to encompass a range that expressly includes the designated quality, characteristic, parameter or value, without variation, as if no qualifying relative term or term of degree were utilized in the given disclosure or recitation. 
     Discussion of Possible Embodiments 
     The following are non-exclusive descriptions of possible embodiments of the present invention. 
     An airfoil for a gas turbine engine includes an airfoil body having a leading edge, a trailing edge, an inner end wall, and an outer end wall, an axial flow cooling circuit defined within the airfoil body, and a radial flow cooling circuit defined between the baffle and the trailing edge. The axial flow cooling circuit includes a baffle disposed in spaced relation to an inner surface of the airfoil. The baffle has an axial extent from the leading edge defined by an aft wall with the axial extent being substantially constant between the inner and outer end walls and defined by a plane perpendicular to an engine axis. The baffle also includes a plurality of impingement cooling holes configured to direct a cooling fluid at an inner surface of the airfoil body. The radial flow cooling circuit includes a first radially-extending rib and a second radially-extending rib. The first rib is angled with respect to the baffle aft wall to define a first passage between the first rib and the baffle that tapers in cross-sectional area between the inner end wall and the outer end wall becoming larger in cross-sectional area in a direction of cooling fluid flow through the first passage. 
     The airfoil of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components: 
     The airfoil of any of the preceding paragraphs, wherein the second rib can be positioned between the first rib and the trailing edge, and wherein the second rib can be angled with respect to the trailing edge to define a second passage between the second rib and the trailing edge that tapers in cross-sectional area between the inner end wall and the outer end wall becoming smaller in cross-sectional area in a direction of cooling fluid flow through the second passage. 
     The airfoil of any of the preceding paragraphs, wherein the baffle can further include a U-shaped wall together with the aft wall defining a central cavity, with the U-shaped wall having a forward edge portion proximate the leading edge of the airfoil and having the plurality of impingement cooling holes positioned to direct cooling fluid flow at an inner surface of the leading edge of the airfoil, a first side extending between the forward edge portion and the aft side, and a second side opposite the first side and extending between the forward edge portion and the aft side. The first side, the second side, and the aft wall can be free of impingement cooling holes. 
     The airfoil of any of the preceding paragraphs, can further include a forward wall free of impingement cooling holes, an aft wall opposite the forward wall with the aft wall being free of impingement cooling holes, and first and second opposing side walls separating the forward and aft walls. At least one of the first and second side walls includes the plurality of impingement cooling holes configured to direct cooling fluid flow at an inner surface of a pressure side or suction side of the airfoil. 
     The airfoil of any of the preceding paragraphs, wherein the inner surface of the airfoil can include a plurality of substantially axially-extending ribs configured to direct cooling fluid flow exiting the plurality of impingement cooling holes in an axial direction toward the first passage. 
     The airfoil of any of the preceding paragraphs, wherein the plurality of substantially axially-extending ribs can extend along the inner surface of the airfoil around a U-shaped wall of the baffle, extending from the aft wall of the baffle on a first side to the aft wall of the baffle on a second side opposite the first side. 
     The airfoil of any of the preceding paragraphs, wherein the plurality of substantially axially-extending ribs can be angled with respect to the inner end wall to direct cooling fluid flow toward a direction of cooling fluid flow in the first passage. 
     The airfoil of any of the preceding paragraphs, wherein the plurality of substantially axially-extending ribs can be non-uniformly distributed as a function of span between the inner and outer end walls. 
     The airfoil of any of the preceding paragraphs can further include a third passage defined between the first radially-extending rib and the second radially-extending rib, a first turn connecting the first passage and the third passage at one of the inner end wall and the outer end wall, and a second turn connecting the second passage and the third passage at the other of the inner end wall and outer end wall. 
     The airfoil of any of the preceding paragraphs, wherein the first passage can taper inward from the inner end wall to the outer end wall and the second passage can taper outward from the inner end wall to the outer end wall, and wherein the radial flow cooling circuit is configured to direct cooling fluid flow from the outer end wall to the inner end wall in the first and second passages. 
     The airfoil of any of the preceding paragraphs, wherein the first passage can taper outward from the inner end wall to the outer end wall and the second passage can taper inward from the inner end wall to the outer end wall, and wherein the radial flow cooling circuit is configured to direct cooling fluid flow from the inner end wall to the outer end wall in the first and second passages. 
     The airfoil of any of the preceding paragraphs can further include a plurality of heat transfer features selected from the group of heat transfer features comprising: first heat transfer features extending from the inner surface of the airfoil toward at least one of the first side of the baffle and the second side of the baffle, and second heat transfer features extending from the inner surface of the airfoil into the first, second, and third passages. 
     The airfoil of any of the preceding paragraphs, wherein a spacing between adjacent first or second heat transfer features can be non-uniform. 
     The airfoil of any of the preceding paragraphs, wherein the baffle can include a cavity inlet at the inner end wall or the outer end wall. 
     The airfoil of any of the preceding paragraphs, wherein the baffle aft wall can be disposed at 30 to 60 percent chord from the leading edge of the airfoil. 
     A method of cooling an airfoil for a gas turbine engine includes flowing cooling fluid through an axial flow cooling circuit and flowing the cooling fluid through the radial flow cooling circuit. The axial flow cooling circuit includes flowing the cooling fluid from a cavity of a baffle through a plurality of cooling holes and directing the flow of cooling fluid from the plurality of cooling holes in an axial direction to a radial cooling circuit defined between the baffle and a trailing edge of the airfoil. The cavity extends between an inner end wall and an outer end wall of the airfoil and has an axial extent from the leading edge defined by an aft wall, with the axial extent being substantially constant between the inner and outer end walls and defined by a plane perpendicular to an engine axis. Flowing the cooling fluid through the radial flow cooling circuit includes flowing the cooling fluid through a first radially-extending passage that tapers outward in cross-sectional area between the inner end wall and the outer end wall in a direction of cooling fluid flow through the first passage, and flowing the cooling fluid through a second radially-extending passage that tapers inward in cross-sectional area between the inner end wall and the outer end wall in a direction of cooling fluid flow through the second passage. 
     The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations, additional components, and/or additional steps: 
     The method of any of the preceding paragraphs, wherein the first passage can be defined between the baffle and a first rib angled with respect to the baffle and wherein the second passage can be defined between the trailing edge and a second rib angled with respect to the trailing edge. 
     The method of any of the preceding paragraphs, wherein the flow of cooling fluid can be directed in the axial direction by a plurality of ribs disposed along the inner surface of the airfoil adjacent to the baffle. 
     The method of any of the preceding paragraphs, wherein the plurality of cooling holes can be located to direct cooling fluid at an inner surface of a leading edge of the airfoil or at inner surfaces of pressure and suction sides of the airfoil. 
     The method of any of the preceding paragraphs can further include flowing the cooling fluid around a plurality of first heat transfer features disposed between the baffle and the inner surface of the airfoil, and flowing the cooling fluid across a plurality of second heat transfer features disposed in the first and second passages. 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.