Patent Publication Number: US-11396819-B2

Title: Components for gas turbine engines

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims the benefit of an earlier filing date from U.S. Provisional Application Ser. No. 62/835,823, filed Apr. 18, 2019, the entire disclosure of which is incorporated herein by reference. 
    
    
     BACKGROUND 
     Illustrative embodiments pertain to the art of turbomachinery, and specifically to turbine rotor components. 
     Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine extracts energy from the expanding combustion gas, and drives the compressor via a common shaft. Energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both. 
     The compressor and turbine sections are typically subdivided into a number of stages, which are formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine. 
     Airfoils may incorporate various cooling cavities located adjacent external side walls. Such cooling cavities are subject to both hot material walls (exterior or external) and cold material walls (interior or internal). Although such cavities are designed for cooling portions of airfoil bodies, various cooling flow characteristics can cause hot sections where cooling may not be sufficient. Accordingly, improved means for providing cooling within an airfoil may be desirable. 
     BRIEF DESCRIPTION 
     According to some embodiments, components for gas turbine engines are provided. The components include an airfoil having a leading edge, a trailing edge, a pressure side, and a suction side, wherein the airfoil defines at least a leading edge cavity located proximate the leading edge and defined between the leading edge and a separator rib in an axial direction and between the pressure side and the suction side in a circumferential direction, the leading edge cavity comprising a baffle portion and a leading edge portion, with the baffle portion aft of the leading edge portion in the axial direction. A baffle is installed within the baffle portion of the leading edge cavity, the baffle having a first metering flow aperture. A first support element retention feature is located within the leading edge cavity and at least partially separating the baffle portion from the leading edge portion, the first support element retention feature on one of the pressure side and the suction side of the leading edge cavity. A first axial extending rib extends between an aft end proximate the separator rib of the leading edge cavity and a forward end proximate the first support element retention feature and formed on an interior surface of a same side as the first support element retention feature. A first axial extending flow channel is defined along the first axial extending rib between an exterior surface of the baffle and an interior surface of the airfoil and extending from the aft end to the forward end in an axial direction, and the first metering flow aperture is located proximate the aft end of the first axial extending flow channel such that air flowing through the first metering flow aperture into the first axial extending flow channel will flow forward toward the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include a plurality of additional axial extending ribs arranged on the same interior surface as the first axial extending rib, wherein a plurality of additional axial extending flow channels are defined between adjacent axial extending ribs. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include a second support element retention feature located within the leading edge cavity and at least partially separating the baffle portion from the leading edge portion, the second support element retention feature on the other of the pressure side and the suction side of the leading edge cavity from the first support element retention feature, a second axial extending rib extending between the aft end proximate the separator rib of the leading edge cavity and the forward end proximate the second support element retention feature and formed on an interior surface of a same side as the second support element retention feature, wherein a second axial extending flow channel is defined along the second axial extending rib between an exterior surface of the baffle and an interior surface of the airfoil and extending from the aft end to the forward end in an axial direction. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the baffle comprises at least one impingement aperture configured to fluidly connect to the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the first axial extending rib has a variable radial height in a direction from the aft end to the forward end. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the interior surface of the airfoil defining a portion of the first axial extending flow channel includes at least one heat transfer augmentation feature. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the at least one heat transfer augmentation feature comprises at least one of trip strips, pin fins, and pedestals. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include that the at least one heat transfer augmentation feature comprises a plurality of heat transfer augmentation features that extend along the interior surface of the airfoil from the separator rib into the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include at least one film cooling hole formed on the airfoil to fluidly connect the leading edge portion to an exterior of the airfoil. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include a second metering flow aperture defined at least partially by the first support element retention feature at the forward end of the first axial extending flow channel. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the components may include a second axial extending rib extending between the aft end and the forward end and the first axial extending rib and the second axial extending rib are not parallel to each other. 
     According to some embodiments, gas turbine engines are provided. The gas turbine engines include an airfoil having a leading edge, a trailing edge, a pressure side, and a suction side, wherein the airfoil defines at least a leading edge cavity located proximate the leading edge and defined between the leading edge and a separator rib in an axial direction and between the pressure side and the suction side in a circumferential direction, the leading edge cavity comprising a baffle portion and a leading edge portion, with the baffle portion aft of the leading edge portion in the axial direction; a baffle installed within the baffle portion of the leading edge cavity, the baffle having a first metering flow aperture; a first support element retention feature located within the leading edge cavity and at least partially separating the baffle portion from the leading edge portion, the first support element retention feature on one of the pressure side and the suction side of the leading edge cavity; a first axial extending rib extending between an aft end proximate the separator rib of the leading edge cavity and a forward end proximate the first support element retention feature and formed on an interior surface of a same side as the first support element retention feature, wherein a first axial extending flow channel is defined along the first axial extending rib between an exterior surface of the baffle and an interior surface of the airfoil and extending from the aft end to the forward end in an axial direction, and wherein the first metering flow aperture is located proximate the aft end of the first axial extending flow channel such that air flowing through the first metering flow aperture into the first axial extending flow channel will flow forward toward the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a plurality of additional axial extending ribs arranged on the same interior surface as the first axial extending rib, wherein a plurality of additional axial extending flow channels are defined between adjacent axial extending ribs. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a second support element retention feature located within the leading edge cavity and at least partially separating the baffle portion from the leading edge portion, the second support element retention feature on the other of the pressure side and the suction side of the leading edge cavity from the first support element retention feature; a second axial extending rib extending between the aft end proximate the separator rib and the forward end proximate the second support element retention feature and formed on an interior surface of a same side as the second support element retention feature, wherein a second axial extending flow channel is defined along the second axial extending rib between an exterior surface of the baffle and an interior surface of the airfoil and extending from the aft end to the forward end in an axial direction. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the baffle comprises at least one impingement aperture configured to fluidly connect to the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the first axial extending rib has a variable radial height in a direction from the aft end to the forward end. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the interior surface of the airfoil defining a portion of the first axial extending flow channel includes at least one heat transfer augmentation feature. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the at least one heat transfer augmentation feature comprises at least one of trip strips, pin fins, and pedestals. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include that the at least one heat transfer augmentation feature comprises a plurality of heat transfer augmentation features that extend along the interior surface of the airfoil from the separator rib into the leading edge portion of the leading edge cavity. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include at least one film cooling hole formed on the airfoil to fluidly connect the leading edge portion to an exterior of the airfoil. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a second metering flow aperture defined at least partially by the first support element retention feature at the forward end of the first axial extending flow channel. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the gas turbine engines may include a second axial extending rib extending between the aft end and the forward end and the first axial extending rib and the second axial extending rib are not parallel to each other. 
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: The subject matter is particularly pointed out and distinctly claimed at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which like elements may be numbered alike and: 
         FIG. 1  is a schematic cross-sectional illustration of a gas turbine engine; 
         FIG. 2  is a schematic illustration of a portion of a turbine section of the gas turbine engine of  FIG. 1 ; 
         FIG. 3A  is an elevation schematic illustration of an airfoil; 
         FIG. 3B  is a cross-sectional illustration of the airfoil of  FIG. 3A  as viewed along the line B-B; 
         FIG. 4A  illustrates an airfoil in accordance with an embodiment of the present disclosure in a top-down cross-sectional view showing an interior structure of the airfoil; 
         FIG. 4B  illustrates an elevational view of a portion of an interior surface of a leading edge cavity of the airfoil of  FIG. 4A , in accordance with an embodiment of the present disclosure; 
         FIG. 4C  illustrates the same top-down cross-sectional view of  FIG. 4A , with a “space-eater” baffle installed within the leading edge cavity of the airfoil; 
         FIG. 4D  illustrates the baffle of  FIG. 4C , in accordance with an embodiment of the present disclosure, in isolation and not installed within the airfoil; 
         FIG. 5  is a schematic partially exploded illustrative view of an airfoil and baffle in accordance with an embodiment of the present disclosure; 
         FIG. 6A  is an elevational schematic illustration of a portion of an airfoil in accordance with an embodiment of the present disclosure; 
         FIG. 6B  is a partial isometric illustration of a portion of the airfoil shown in  FIG. 6A ; 
         FIG. 7A  is a first elevational illustration of a portion of an airfoil in accordance with an embodiment of the present disclosure; 
         FIG. 7B  is a second elevational illustration of the portion of the airfoil shown in  FIG. 7A ; 
         FIG. 7C  is a schematic partial isometric illustration of the airfoil shown in  FIG. 7A ; 
         FIG. 8  is a schematic illustration of different types of axial extending ribs that may be employed in various embodiments of the present disclosure; 
         FIG. 9  is a schematic elevational illustration of a baffle installed within an airfoil in accordance with an embodiment of the present disclosure; 
         FIG. 10  is a schematic illustration of an axial extending rib configuration in accordance with an embodiment of the present disclosure; and 
         FIG. 11  is a schematic illustration of an airfoil having axial extending ribs in accordance with an embodiment of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     Detailed descriptions of one or more embodiments of the disclosed apparatus and/or methods are presented herein by way of exemplification and not limitation with reference to the Figures. 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  drives air along a bypass flow path B in a bypass duct, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . An engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The engine static structure  36  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one non-limiting example is a high-bypass geared aircraft engine. In a further non-limiting example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(514.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec). 
     Although the gas turbine engine  20  is depicted as a turbofan, it should be understood that the concepts described herein are not limited to use with the described configuration, as the teachings may be applied to other types of engines such as, but not limited to, turbojets, turboshafts, etc. 
     Referring now to  FIG. 2 , a cooling design in a turbine section  28  for a gas turbine engine  20  may utilize a vane  106  disposed between axially adjacent bladed full hoop disks  108 ,  108   a  having respective blades  109 ,  109   a . As shown, vane  106  is disposed radially between an inner air seal  112  and a full hoop case  114  on an outer side. Inner air seal  112  may be a full hoop structure supported by opposing vanes, including a plurality of vanes  106  that are separated in a circumferential direction. Vane  106  is supported by the full hoop case  114  through segmented vane hooks  117 ,  117   a . One or more full hoop cover plates  115 ,  115   a  may minimize leakage between the vane  106  and the blades  109 ,  109   a . The vane  106  is radially supported by the full hoop case  114  with segmented case hooks  116 ,  116   a  in mechanical connection with the segmented vane hooks  117 ,  117   a . The vane  106  may be circumferentially supported between circumferentially adjacent platforms  119 ,  119   a  which may include feather seals that can minimize leakage between the adjacent vanes  106  into the gas path. 
     Although  FIG. 2  depicts a second stage vane, as appreciated by those of skill in the art, embodiments provided herein can be applicable to first stage vanes as well. Such first stage vanes may have cooling flow supplied to the vane at both the inner and outer diameters, as opposed to the through-flow style cavity which goes from, for example, outer diameter to inner diameter. Thus, the present illustrations are not to be limiting but are rather provided for illustrative and explanatory purposes only. 
     In the present illustration, a turbine cooling air (TCA) conduit  125  provides cooling air into an outer diameter vane cavity  124  defined in part by an outer platform  119  and the full hoop case  114 . The vane  106  is hollow so that air can travel radially into and longitudinally downstream from the outer diameter vane cavity  124 , through the vane  106  via one or more vane cavities  122 , and into a vane inner diameter cavity  123 . The vane inner diameter cavity  123  is defined, in part, by an inner platform  119   a . Thereafter air may travel through an orifice  120  in the inner air seal  112  and into a rotor cavity  121 . Accordingly, cooling air for at least portions of the vane  106  will flow from a platform region, into the vane, and then out of the vane and into another platform region and/or into a hot gaspath/main gaspath. In some arrangements, the platforms  119 ,  119   a  can include ejection holes to enable some or all of the air to be injected into the main gaspath. 
     It is to be appreciated that the longitudinal orientation of vane  106  is illustrated in a radial direction, but other orientations for vane  106  are within the scope of the disclosure. In such alternate vane orientations, fluid such as cooling air can flow into the vane cavity  122  through an upstream opening illustrated herein as outer diameter cavity  124  and out through a downstream opening in vane cavity  122  illustrated herein as inner diameter cavity  123 . A longitudinal span of vane cavity  122  being between such openings. 
     The vane  106 , as shown, includes one or more baffles  126  located within the vane  106 . The baffles  126  are positioned within one or more respective baffle cavities  128 . The baffle cavities  128  are sub-portions or sub-cavities of the vane cavity  122 . In some embodiments, such as shown in  FIG. 2 , the baffle cavities  128  are internal cavities that are axially inward from the leading and trailing edges of the vane  106 , although such arrangement is not to be limiting. The TCA conduit  125  may provide cooling air that can flow into the baffles  126  and then impinge from the respective baffle  126  onto an interior surface of the vane  106 . 
     As shown and labeled in  FIG. 2 , a radial direction R is upward on the page (e.g., radial with respect to an engine axis) and an axial direction A is to the right on the page (e.g., along an engine axis). Thus, radial cooling flows will travel up or down on the page and axial flows will travel left-to-right (or vice versa). A circumferential direction C is a direction into and out of the page about the engine axis. 
     Turning now to  FIGS. 3A-3B , schematic illustrations of an airfoil  300  having a first baffle  302  and a second baffle  304  installed therein are shown. Each baffle  302 ,  304  has a baffle body that defines the structure and shape of the respective baffle  302 ,  304 . The airfoil  300  extends in an axial direction between a leading edge  306  and a trailing edge  308 . In a radial direction, the airfoil  300  extends between an inner platform  310  at an inner diameter  312  and an outer platform  314  at an outer diameter  316 . In this illustrative embodiment, the airfoil  300  has three internal cavities: a leading edge cavity  318 , a mid-cavity  320 , and a trailing edge cavity  322 . Although shown with a specific cavity configuration, those of skill in the art will appreciate that airfoils can have a variety of internal cavity configurations and implement embodiment of the present disclosure. Thus, the present illustration is merely for explanatory purposes and is not to be limiting.  FIG. 3A  is a side elevation illustration of the airfoil  300  illustrating an internal structure thereof.  FIG. 3B  is a cross-sectional illustration as viewed along the line B-B. 
     The cavities  318 ,  320 ,  322  may be separated by ribs  324   a ,  324   b , with fluid connections therebetween in some embodiments. The ribs  324   a ,  324   b  extend radially between the inner platform  310  at the inner diameter  312  to the outer platform  314  at the outer diameter  316 . A first rib  324   a  may separate the mid-cavity  320  from the leading edge cavity  318 , and may, in some embodiments, fluidly separate the two cavities  318 ,  320 . A second rib  324   b  may separate the mid-cavity  320  from the trailing edge cavity  322 , and may, in some embodiments, have through holes to fluidly connect the mid-cavity  320  to the trailing edge cavity  322 . 
     In this embodiment, the leading edge cavity  318  includes the second baffle  304  installed therein and the mid-cavity  320  includes the first baffle  302  therein. The first baffle  302  includes first baffle holes  326  (shown in  FIG. 3B ) to supply cooling air from within the first baffle  302  into the mid-cavity  320 . The cooling air within the mid-cavity  320  may flow into the trailing edge cavity  322  and subsequently exit the airfoil  300  as known in the art. The second baffle  304  includes second baffle holes  328  where cooling air within the second baffle  304  may impinge upon surfaces of the airfoil  300  of the leading edge cavity  318 . The cooling or impinged air may then exit the leading edge cavity  318  through film cooling holes  330 , as will be appreciated by those of skill in the art. 
     In some airfoils, the leading edge may not include a baffle, but rather may include a leading edge feed cavity and a leading edge impingement cavity, wherein flow from the leading edge feed cavity will flow through impingement apertures to impinge upon a leading edge hot wall, and then exit the leading edge impingement cavity through film cooling holes. Aft of the leading edge cavity arrangement may be one or more additional cavities, which typically includes a trailing edge cavity. In such airfoil arrangements, the leading edge is typically fed by a high pressure source for high compressor discharge air. The trailing edge, in contrast, may be fed from a mid-compressor bleed source, which is a lower pressure source. However, utilizing high pressure air may be undesirable from a thermodynamic cycle efficiency perspective. That being said, high pressure air is sometimes required to meet leading edge back flow margin requirements because a mid-compressor bleed feed source may not have high enough pressure to adequately ensure positive out-flow through leading edge film cooling holes. The leading edge film cooling may be required to effectively cool the leading region of the airfoil to prevent premature through-wall oxidation due to excessively high metal temperatures resulting from high external gas temperatures and heat flux. 
     Due to these considerations, the high pressure source that feeds the leading edge feed cavity can result in a significantly higher pressure ratio than required and/or desired. The leading edge pressure ratio is defined by the ratio of the supply feed pressure and the total free stream gas pressure at the leading edge of the airfoil, commonly referred to as the stagnation pressure. As a result of the high supply pressure and high leading edge pressure ratio, it may become desirable and necessary to “meter” the cooling air flow in order to meet allocated cooling flow requirements. 
     The high leading edge pressure ratio increases the cooling flow rate for a constant exit flow area (i.e., film cooling holes) resulting in high blowing ratios across the rows of leading edge film cooling holes. Although high blowing ratios are desirable to achieve high Reynolds numbers within the film cooling holes to increase convective cooling, such high blowing ratios may be undesirable from a film cooling perspective. Film cooling holes with excessively high blowing ratios have a tendency to have poor film cooling characteristics because the cooling flow emanating from the film holes may “blow off” and separate from the airfoil surface. This separation may prevent a desired “film” of cooling air along the exterior surface of the airfoil. 
     One solution is to meter the flow from the high pressure cooling supply source. The metering of the pressure may be achieved by introducing a relatively “small” feed orifice in order to reduce or “drop” the high supply pressure source by incurring additional pressure losses resulting from a small inlet feed area and sharped edge sudden contraction that results. However, reducing pressure in order to achieve a lower leading edge pressure ratio to mitigate cooling high flow rates is not desirable from a convective heat transfer perspective. This is because the lower pressure levels inherently result in a reduction in the absolute level of convective heat transfer that is otherwise achievable at a higher pressure level. 
     In some prior art embodiments, the incorporation of a small inlet feed aperture may be utilized to “meter” cooling flow rate in order to achieve desirably leading edge showerhead cooling flow levels and pressure ratios in order to improve local film cooling characteristics. The small inlet feed aperture serves as a flow restrictor in order to meter the cooling air flow rate by inducing significant pressure loss and, in this sense, the high supply pressure source is not utilized effective to provide necessary internal convective cooling. 
     In order to restrict the cooling flow rate, a single flow aperture having a relatively small cross-sectional area is required. However, a small leading edge feed orifice may be undesirable because it may be prone to plugging from debris within the engine, either from surrounding hardware such as brush seals, w-seals, and/or from dirt/sand particulate that is in the environment that the engine is subjected to in certain parts of the world. Further, in some solutions, due to other considerations, as discussed above, the sourced cooling air/pressure may be underutilized. 
     In an effort to utilize the high pressure supply source and, correspondingly, the high pressure ratio that exists across the leading edge film cooling holes, a more effective means of reducing or lowering the available pressure is to induce pressure losses through the incorporation of internal convective cooling features, such as baffle impingement apertures, turbulators, trip strips, pin fins and/or pedestal geometry features. In this sense, the high supply pressure source can be utilized more effectively by providing internal hot wall convective cooling in order to increase the local thermal cooling effectiveness, thereby reducing operating metal temperatures and improve overall part capability and durability. Those skilled in the art will appreciate that the increased frictional losses and pressure losses associated with the incorporation of internal cooling geometric features, which are used to promote local cooling flow vortices that induce turbulent mixing and in turn, enhance convective cooling characteristics immediate the hot internal wall surfaces. 
     Embodiments of the present disclosure are directed to incorporating a leading edge counter flow “space-eater” baffle concept. The “space-eater” baffle concept includes a plurality of predominantly axial rib offsets. Such axial rib offsets may include a second metering flow aperture defined at least partially by a first support element retention feature at the forward end of the first axial extending rib offset feature. The axial cooling channels are formed between the exterior surface of the “space-eater” baffle insert and the axial extending rib offset features, which serve to segregate the axial flow channels. The discrete axial channels may be smooth and/or rib roughened cooling channels to promote and enhance internal convective cooling. The channels may include various unique convective heat transfer cooling features proximate the internal surface of the leading edge of an airfoil. In various embodiments, an axial channel flow area formed between a baffle exterior and interior surfaces of the airfoil, in an axial stream wise direction, may be constant, converging, and/or diverging channel flow areas controlled by variable axial rib heights. As discussed herein, the term “axial” refers to a direction relative to an engine axis, when the airfoil is installed within such engine (e.g., as shown in  FIG. 2 ). The axial direction is a direction between a leading edge and a trailing edge of the airfoil, with a forward flow direction being a direction from the trailing edge toward the leading edge (an aft flow is a direction from the leading edge toward the trailing edge). 
     The “space-eater” baffle is a counter-flow (i.e., aft-to-forward flow) cooling concept in which a high pressure feed source can be leveraged by managing pressure losses within the cooling system in order to provide more efficient and effective use of cooling airflow for improved convective heat transfer and film cooling of the airfoil. Optimization or control of pressure loss within the cooling design, in accordance with embodiments of the present disclosure, may be achieved through various heat transfer features and orifices within the airfoil. For example, leading edge “space-eater” baffle feed and/or resupply flow apertures (e.g., size and shape thereof) may be independently tailored specifically for each pressure side and suction side axial flow channel to optimize both the radial and axial cooling flow distribution in each of the axial flow channels created between the exterior surface of the “space-eater” baffle and the interior surface of the airfoil external wall. Axial channel flow area, trip strip type, pitch, height, and spacing are other types of examples of creating the desired axial channel cooling flow Mach number, Reynolds number, convective heat transfer, pressure loss, and mass flow rate through axial channels of the present disclosure. 
     Turning to  FIGS. 4A-4D , schematic illustrations of an airfoil  400  in accordance with an embodiment of the present disclosure is shown.  FIG. 4A  illustrates the airfoil  400  in a top-down cross-sectional view showing an interior structure of the airfoil  400 .  FIG. 4B  illustrates an elevational view of a portion of an interior surface of a leading edge cavity of the airfoil  400 .  FIG. 4C  illustrates the same top-down cross-sectional view of  FIG. 4A , but with a “space-eater” baffle  401  installed within the cavity of the airfoil  400 .  FIG. 4D  illustrates the “space-eater” baffle  401 , in accordance with an embodiment of the present disclosure, in isolation, not installed within the airfoil  400 . As described herein, the described cavity (e.g., leading edge cavity) may be a compound cavity having multiple different portions/regions/aspects. That is, the term “compound” with respect to a cavity within an airfoil, as used herein, refers to a cavity having multiple functions, regions, sub-cavities, etc. that are distinct from each other—but a single cavity formed within the airfoil. For example, the leading edge cavity of the airfoil  400  shown in  FIGS. 4A-4D  may be “compound” and include a leading edge portion and a baffle portion, with the baffle portion aft of the leading edge portion. Generally, the cavities described herein are cavities having multiple substantially distinct portions or regions (e.g., sub-cavities) that can provide different functions within the airfoil cooling configuration, such as to receive a baffle. 
     As shown in  FIG. 4A , the airfoil  400  extends in a substantially axial direction between a leading edge  402  and a trailing edge  404 . In a circumferential direction, as described above, the airfoil  400  extends between a pressure side  406  and a suction side  408 . The airfoil  400  includes internal cavities that are configured to allow cooling air to enter and pass therethrough, for example, from a platform located at an inner or outer diameter location (e.g., as shown in  FIG. 2 ). In this illustrative embodiment, the airfoil  400  includes a leading edge cavity  410 , a midchord cavity  412 , and a trailing edge cavity  414 . One or more of the cavities  410 ,  412 ,  414  may be fluidly connected to allow cooling air to flow from one into another. However, in some embodiments, at least the leading edge cavity  410  may be fluidly separated from the other cooling cavities of the airfoil  400 . The separation of the leading edge cavity  410  from the midchord cavity  412  may be provided by a separator rib  416  that extends from between an inner diameter  418  to an outer diameter  420  of the airfoil  400  (as shown in  FIG. 4B ) and between the pressure side  406  and the suction side  408 . The separator rib  416  may provide structural support to the airfoil  400 , may fluidly separate the cavities thereof, and may provide a support surface for supporting a baffle within the leading edge cavity  410 . 
     The leading edge cavity  410  includes a baffle portion  422  and a leading edge portion  424 . The baffle portion  422  is partially separated from the leading edge portion  424  by one or more support element retention features  426 . The support element retention features  426  are located within the leading edge cavity and are arranged to partially separate the baffle portion from the leading edge portion. The support element retention features  426  extend between the pressure side and the suction side of the leading edge cavity. The support element retention features  426  are located within the leading edge cavity and are arranged to partially separate the baffle portion from the leading edge portion. The support element retention features  426  extend between the pressure side and the suction side of the leading edge cavity (i.e., circumferential direction). The baffle portion  422  is defined, in part, by a surface of the separator rib  416 , a pressure side surface  428 , and a suction side surface  430 . The separator rib  416  defines an aft most portion of the leading edge cavity  410  and the location of the support element retention features  426  defines the forward most extent of the baffle portion  422 . The support element retention features  426  may be discontinuous in the radial direction, allowing for fluid communication between the baffle portion  422  and the leading edge portion  424 . Further, in the circumferential direction, the space between opposing support element retention features  426  may be referred to as an impingement portion  432  of the leading edge cavity  410 . Forward of the impingement portion  432  is the leading edge portion  424  of the leading edge cavity  410 . Air within the leading edge portion  424  may exit the leading edge cavity  410  through one or more film cooling holes  434  (e.g., showerhead and gill row holes) located on the leading edge  402  of the airfoil  400 , as will be appreciated by those of skill in the art. 
     The baffle portion  422  of the leading edge cavity  410  is sized and shaped to receive a baffle, such as a space-eater baffle, therein. Further, the walls, and specifically the pressure side surface  428  and the suction side surface  430  that define the baffle portion  422  include one or more axial extending ribs  436 , as shown in  FIGS. 4A-4B . The axial extending ribs  436  extend between an aft end of the respective surface to a forward end of the respective surface. For example, as shown in  FIG. 4B , the axial extending ribs  436  extend between an aft end  438  to a forward end  440  of the suction side surface  430 . The separator rib  416  is located at and defines the aft end  438  and the support element retention features  426  are located at and define, in part, the forward end  440  of the axial extending ribs  436 . Between adjacent axial extending ribs  436  (in the radial direction between the inner diameter  418  and the outer diameter  420 ) are defined one or more axial extending flow channels  442 . The axial extending flow channels  442  are configured to allow a fluid flow, such as a cooling flow, tp pass therethrough. The surfaces of the axial extending flow channels  442  may be smooth or may include heat transfer augmentation features  444 , shown schematically in one of the axial extending flow channels  442 . The heat transfer augmentation features  444  may be, for example, discrete trip strips, pin fins, divots, pedestals, hemispherical protrusions, etc., as will be appreciated by those of skill in the art. It will be appreciated that the axial extending ribs  436  extend from the suction side surface  430  to define the axial extending flow channels  442 . That is, the axial extending ribs  436  may extend in a circumferential direction off of the suction side surface  430  and into the baffle portion  422  of the leading edge cavity  410 . 
     Turning now to  FIGS. 4C-4D , the airfoil  400  is shown with the “space-eater” baffle  401  installed within the baffle portion  422  of the leading edge cavity  410 . The “space-eater” baffle  401  is a space-eater baffle that is configured to provide a cooling flow of air into the airfoil  400 , and specifically to provide cooling along hot walls (i.e., pressure side  406  and suction side  408  of the airfoil  400 ) along the leading edge cavity  410 . Cooling flow enters in the middle of the “space-eater” baffle  401  (i.e., the interior portion of the “space-eater” baffle  401 ) from the vane inner diameter or outer diameter and exits out through one or more first metering flow apertures  446  into the axial extending flow channels  442  formed between the exteriors surface of the “space-eater” baffle insert and the axial extending rib offset features. The cooling flow will then pass into the leading edge portion  424  of the leading edge cavity  410  and is expelled out of leading edge portion  424  through the film cooling holes  434 . Additionally, a portion of the air within the “space-eater” baffle  401  may flow directly into the leading edge portion  424  through one or more impingement apertures  448  formed on/in a leading edge or forward wall  450  of the “space-eater” baffle  401 . When the “space-eater” baffle  401  is installed within the airfoil  400 , the axial extending flow channels  442  are defined between exterior surfaces of the “space-eater” baffle  401  and interior surfaces of the airfoil  400  (i.e., pressure side surface  428  and suction side surface  430 ). 
     As shown in  FIGS. 4C-4D , the “space-eater” baffle  401  includes the forward wall  450 , a pressure side wall  452 , a suction side wall  454 , and an aft wall  456 . When installed within the airfoil  400 , the pressure side wall  452  of the “space-eater” baffle  401  is arranged along and adjacent or proximate the pressure side surface  428  of the leading edge cavity  410  along the pressure side  406  of the airfoil  400 . Similarly, the suction side wall  454  of the “space-eater” baffle  401  is arranged along and adjacent or proximate the suction side surface  430  of the leading edge cavity  410  along the suction side  408  of the airfoil  400 . The aft wall  456  of the “space-eater” baffle  401  is arranged in contact with or proximate to and adjacent the separator rib  416  of the airfoil  400 . At the forward end the forward wall  450  of the “space-eater” baffle  401  contacts the support element retention features  426 . Accordingly, the “space-eater” baffle  401  is retained in a forward-aft direction between the support element retention features  426  and the separator rib  416 . 
     As shown in  FIG. 4D , the first metering flow apertures  446  are formed in the pressure side wall  452  and the suction side wall  454  of the “space-eater” baffle  401  proximate the aft wall  456 . As such, air exiting the interior of the “space-eater” baffle  401  will enter the axial extending flow channels  442  proximate the aft wall  454  and subsequently flow in a forward direction toward the leading edge  402  of the airfoil  400 . Furthermore, the impingement apertures  448  are formed in the forward wall  450  of the “space-eater” baffle  401 , and air can impinge through the impingement apertures  448  and enter the leading edge portion  424  of the airfoil  400 , and may impinge directly upon the leading edge hot wall of the airfoil  400 . The impingement apertures  448  are aligned with the impingement portion  432  defined between the support element retention features  426 . 
     The support element retention features  426  provide support and positioning for the “space-eater” baffle  401  as described above. Further, the support element retention features may control a flow entering the leading edge portion  424  of the leading edge cavity  410 . For example, the support element retention features  426  may define second metering flow apertures  458 , as shown in  FIG. 4B  (isometric illustration of a similar concept shown, for example, in  FIG. 5 , below). In operation, the first metering flow apertures  446  define an upstream or inlet end of a given axial extending flow channel  442  and the second metering flow apertures  458  define a downstream or outlet end of the given axial extending flow channel  442 . As such, as noted above, the axial extending flow channels  442  define forward flowing channels with an inlet of air at the aft end  438  of the respective axial extending flow channel  442  and an outlet of air at the forward end  440  of the respective axial extending flow channel  442 . 
     Although the support element retention features  426  shown in  FIG. 4B  are the same radial height as the axial extending ribs  436 , such configuration is not to be limiting. The support element retention features  426 , in some embodiments, may be integral with the axial extending ribs  436 . In some embodiments, the support element retention features  426  may be configured and defined in order to achieve the necessary flow distribution and pressure loss characteristics within each of the axial extending flow channels  442 . The flow area size, shape, fillet blends, surface contours, and spacing of the second metering flow apertures  458  may be independently configured and customized depending on local external heat load and cooling effectiveness requirements for a given airfoil or application. Further, the geometry of the support element retention features  426  (e.g., height, width, length) may be tailored to optimize local conduction and fin efficiency. The support element retention features  426  may also incorporate a variable taper depending on structural load and core die manufacturing requirements to mitigate die lock and die pull constraints. 
     As noted above, and shown in  FIG. 4B , the side surfaces  428 ,  430  can include heat transfer augmentation features  444  within the axial extending flow channels  442 . In some embodiments, the heat transfer augmentation features  444  may continue along the interior surfaces of the airfoil  400  into the leading edge portion  424 . In some embodiments, the interior surface of the leading edge portion  424  of the leading edge cavity  410  can include distinct or separate heat transfer augmentation features from those within the axial extending flow channels  442 . For example, in one non-limiting embodiment, the axial extending flow channels  442  may include chevron-type trip strips and the leading edge portion  424  can include divots or discrete hemispherical protrusion type heat transfer augmentation features. Various other configurations are possible without departing from the scope of the present disclosure. 
     Turning now to  FIG. 5 , a schematic illustration of a portion of an airfoil  500  and a “space-eater” baffle  501  in accordance with an embodiment of the present disclosure is shown.  FIG. 5  illustrates the “space-eater” baffle  501  separate from the airfoil  500 . 
     The airfoil  500  has a leading edge  502 , with pressure and suction sides extending aftward therefrom, as appreciated by those of skill in the art, and shown and described above. In this partial view, a portion of a suction side  508  proximate the leading edge  502  is shown. The airfoil  500  includes a leading edge cavity, as shown and described above, for receiving the “space-eater” baffle  501 . The airfoil  500  receives the “space-eater” baffle  501  between a separator rib  516  and a plurality of support element retention features  526 . Forward of the support element retention features  526 , and defined at or along the leading edge  502 , is a leading edge portion of the leading edge cavity, as shown and described above. It will be appreciated that the cavities of the airfoil  500  are not labeled for clarity of illustration, but are substantially similar to that shown and described above with respect to  FIGS. 4A-4D . 
     As shown, the airfoil  500  includes a plurality of axial extending ribs  536 . The axial extending ribs  536  extend between the separator rib  516  and the support element retention features  526 . Between radially adjacent, axial extending ribs  536  are defined axial extending flow channels  542 . The axial extending flow channels  542  may be channels extending from the separator rib  516  to the support element retention features  526  and may fluidly connect to the leading edge portion of the leading edge cavity through second metering flow apertures  558 , which define a downstream end of the axial extending flow channels  542 . Located at the upstream end of the axial extending flow channels  542  are first metering flow apertures  546 , which are illustratively shown relative to the axial extending flow channels  542  but are physically defined (and shown) by the “space-eater” baffle  501  in  FIG. 5 . The surfaces of the airfoil  500  that define the axial extending flow channels  542  between the axial extending ribs  536  may optionally, and as shown, include heat transfer augmentation features  544 . The illustrative heat transfer augmentation features  544  are shown as chevron-type trip strips, but other types of heat transfer augmentation features may be employed without departing from the scope of the present disclosure. 
     As depicted in  FIG. 5 , the axial extending ribs  536  may be defined to have a variable height along the axial extending flow channels  542 . The height of the axial extending ribs  536  is defined in a direction between pressure and suction side walls of an airfoil (e.g., circumferential, as defined in  FIG. 2 , when installed in a turbine engine). The height of the axial extending ribs  536 , in this configuration, increases in the streamwise direction (i.e., axial direction) from the separator rib  516 , proximate the first metering flow aperture  546 , to the support element retention features  526 , proximate the second metering flow apertures  558 , which defines the downstream end of the axial extending flow channels  542 . As shown, the axial extending ribs  536  have an aft-end rib height H a  and a forward-end rib height H f , with the aft-end rib height H a  being less than the forward-end rib height H f , resulting in an axially increasing rib height. The increase in the height of the axial extending ribs  536  increases the cross-sectional flow area of the axial extending cooling channels  542 . This increase in cross-sectional flow area causes a diffusion of the cooling flow. As such the Mach number in the streamwise direction along the axial extending flow channels  542  and the Reynolds number and convective heat transfer, in the streamwise flow direction, will be decreased. The variation in the height of the axial extending ribs  536  enables the local heat transfer, cooling air heat pickup, and pressure loss to be tailored to match local variations in the external airfoil heat flux. 
     Conversely, in an alternative embodiment, the height of the axial extending ribs  536  may decrease in the streamwise direction from the separator rib  516 , proximate the first metering flow aperture  546 , to the support element retention features  526 , proximate the second metering flow apertures  558  (e.g., as shown in  FIG. 7C ). The decrease in the rib height of the axial extending ribs  536  reduces the cross-sectional flow area of the axial extending cooling channels  542 . This decrease in cross-sectional area will cause an acceleration of the cooling flow, thereby increasing the Mach number and internal Reynolds number and convective heat transfer in the streamwise flow direction. 
     The “space-eater” baffle  501  includes a forward wall  550 , a pressure side wall  552 , a suction side wall  554 , and an aft wall  556 . The walls  550 ,  552 ,  554 ,  556  define an interior baffle cavity, as will be appreciated by those of skill in the art. The “space-eater” baffle  501  includes the first metering flow apertures  546  located proximate the aft wall  556  and within or on the pressure side wall  552  and the suction side wall  554 . The first metering flow apertures  546  are arranged to align with the axial extending flow channels  542  when the “space-eater” baffle  501  is installed within the airfoil  500 , and as illustratively shown in  FIG. 5 . The “space-eater” baffle  501  further includes one or more impingement apertures  548  formed on/in the forward wall  550  of the “space-eater” baffle  501 . 
     Although the first and second metering flow apertures and the impingement apertures are illustratively shown as rectangular and/or elongated, such illustrative geometry is not to be limiting, but rather for example purposes only. Any geometry, including, without limitation, circular, oval, square, and/or rectangular may be employed without departing from the scope of the present disclosure. 
     Turning now to  FIGS. 6A-6B , schematic illustrations of a portion of an airfoil  600  in accordance with an embodiment of the present disclosure is shown, with  FIG. 6A  being a partial elevation view and  FIG. 6B  being a partial isometric view. The airfoil  600  may be substantially similar to that shown and described above. For example, the airfoil  600  is arranged having a leading edge cavity configured to receive a baffle in a baffle portion thereof. The airfoil  600  includes a separator rib  616  at an aft end of the leading edge cavity. Along pressure and suction side walls of the airfoil  600  that define portions of the leading edge cavity are axial extending ribs  636 . The axial extending ribs  636  extend between an aft end of the respective surface to a forward end of the respective surface. For example, as shown in  FIGS. 6A-6B , the axial extending ribs  636  extend between an aft end  638  to a forward end  640  of a suction side surface  630 . The separator rib  616  is located at and defines the aft end  638  and one or more support element retention features  626  are located at the forward most end of the axial extending ribs  636  and define, in part, the forward end  640  of the axial extending ribs  636 . Between adjacent axial extending ribs  636  (in the radial direction between an inner diameter  618  and an outer diameter  620 ) are defined axial extending flow channels  642 . 
     The axial extending flow channels  642  are configured to allow fluid flow, such as a cooling flow, to pass therethrough. The surfaces of the axial extending flow channels  642  may be smooth or may include heat transfer augmentation features, as described above. It will be appreciated that the axial extending ribs  636  extend from the suction side surface  630  to define the axial extending flow channels  642 . That is, the axial extending ribs  636  may extend in a circumferential direction off of the suction side surface  630  and into the baffle portion of the leading edge cavity, as described above. 
     In this illustrative embodiment, the support element retention features  626  are arranged to provide metering of flow at the outlet or forward end  640  of the axial extending flow channels  642 . That is, second metering flow apertures  658  defined between radially adjacent support element retention features  626  are illustratively radially shorter or smaller than that shown in  FIG. 4B . This is achieved because the radial height H 1  of the support element retention features  626  is greater than the radial height H 2  of the axial extending ribs  636 . Stated another way, the radial height of the axial extending flow channels  642  is greater than the radial height of the second metering flow apertures  658 . This configuration can enable an impingement-type cooling flow through the second metering flow apertures  658 , which can impinge upon hot wall surfaces at the leading edge of the airfoil  600 . Such configurations can provide increased heat transfer augmentation on both the internal surface of the leading edge of the airfoil within the leading edge portion of the leading edge cavity. Additionally, such configurations may provide for enhanced heat transfer at the inlet of the leaded edge showerhead film cooling holes that emanate from the leading edge portion of the leading edge cavity due to the impinging flow effects created by the discrete flow jets provided by the discrete flow apertures  658 . 
     Cooling flow enters the axial extending flow channels  642  through one or more first metering flow apertures  646  of a “space-eater” baffle  601  into the axial extending flow channels  642 . Due to the increased height H 1  at the support element retention features  626 , the cooling flow will be funneled or otherwise converge upon the relatively narrow second metering flow apertures  658 , as shown in  FIG. 6 . 
     Turning now to  FIGS. 7A-7B , schematic illustrations of a portion of an airfoil  700  in accordance with an embodiment of the present disclosure are shown. The airfoil  700  may be substantially similar to that shown and described above. For example, the airfoil  700  is arranged having a leading edge cavity configured to receive a baffle in a baffle portion thereof.  FIG. 7A  is an elevation illustration of a portion of the airfoil  700  viewed in a direction from a leading edge toward a trailing edge of the airfoil  700 , viewing along a suction side of the airfoil  700 .  FIG. 7B  is an elevation illustration of the airfoil  700  as viewing the suction side interior surface proximate a leading edge of the airfoil  700 .  FIG. 7C  is a schematic partial isometric illustration of the airfoil shown in  FIG. 7A   
     The airfoil  700  includes a separator rib at an aft end of the leading edge cavity (e.g., similar to that shown in  FIGS. 5, 6A, 6B ). Along pressure and suction side walls of the airfoil  700  that define portions of the leading edge cavity are axial extending ribs  736 . The axial extending ribs  736  extend between an aft end of the respective surface to a forward end of the respective surface. For example, the axial extending ribs  736  extend between an aft end (not shown) to a forward end  740  of a suction side surface  730  of the airfoil  700 . The support element retention features  726  are located at and define, in part, the forward end  740  of the axial extending ribs  736 . The support element retention features  726  are located at the forward most end of the axial extending ribs  736 . Between adjacent axial extending ribs  736  (in the radial direction) are defined one or more axial extending flow channels  742 . 
     In this embodiment, the support element retention features  726  include metering elements  760  extending in a radial direction between radially adjacent support element retention features  726 . The support element retention features  726  and the metering elements  760  may be integral portions or part of the airfoil  700  and extend in a circumferential direction (i.e., away) from the suction side surface  730 . The metering elements  760  of the support element retention features  726  define, in part, second metering flow apertures  758  that restrict a flow cross-sectional area at the outlet or forward end of the axial extending flow channels  742 . For example, as shown in  FIGS. 7A-7C , a cross-sectional area  742   a  of the axial extending flow channels  742  is shown as larger than a cross-sectional area  758   a  of the second metering flow apertures  758 . As illustratively shown, the axial extending ribs  736  have a greater height relative to the suction side surface  730  than the height of the metering elements  760  and less than the height of the support element retention features  726 . 
     As depicted in  FIG. 7C , in this configuration, the axial extending ribs  736  may be defined to have a variable height along the axial extending flow channels  742 . The height of the axial extending ribs  736  is defined in a direction between pressure and suction side walls of an airfoil (e.g., circumferential, as defined in  FIG. 2 , when installed in a turbine engine). The height of the axial extending ribs  736 , in this configuration, decreases in the streamwise direction (i.e., axial direction) from the separator rib at the aft end to the support element retention features  726  at the forward end. As shown, the axial extending ribs  736  have an aft-end rib height H a  and a forward-end rib height H f , with the aft-end rib height H a  being greater than the forward-end rib height H f , resulting in an axially decreasing rib height. The decrease in the height of the axial extending ribs  736  decreases the cross-sectional flow area of the axial extending cooling channels  742 . 
     Turning now to  FIG. 8 , a schematic illustration of a suction side surface  830  of an airfoil  800  in accordance with an embodiment of the present disclosure is shown. In this illustrative configuration, the airfoil  800  includes three different configurations or styles of axial extending ribs in accordance with embodiments of the present disclosure. Incorporating variable rib widths may be implemented to address local airfoil stress, strain and/or panel bulge creep issues. 
     In some configurations, the tailoring of the internal axial flow area may be limited due to local thermal hot spots that can result from poor thermal fin efficiency related to unfavorable geometric aspect ratios of the axial extending ribs. Low H/W (height-to-width) ratios of the axial extending ribs can result in reduced local cooling effectiveness, resulting from lower convective heat transfer and increased conduction resistance due to the relatively large thermal mass associated with a poor fin efficiency design. A rib height (e.g., circumferential dimension, or distance extending from a hot wall) and/or a rib width (e.g., radial thickness) may be set to achieve a desired cooling. For example, as discussed above, the embodiment shown with respect to  FIG. 5  illustrates tapering of the height H of the axial extending rib along the streamwise axial direction.  FIG. 8  illustrates width configurations. It will be appreciated that both height and width may be defined for a specific purpose. 
     By linearly increasing or decreasing the height H of the axial extending rib, the flow area of the axial channels can be tailored to better manage the cooling air heat pickup, pressure loss, and internal convective heat transfer distribution in order to mitigate variations in external heat flux and gas temperature along the airfoil surface. In this sense, the local metal temperature, through-thickness, and in-plane thermal gradients in both the axial and radial directions along the airfoil surface can be minimizes to improve both oxidation and thermal mechanical fatigue failure modes. 
     With respect to a rib width (i.e., radial dimension), and turning to  FIG. 8 , a first axial extending rib  836   a  is arranged having a widening taper in rib width W extending from an aft end  838  to a forward end  840  and defining a wall or side of an axial extending flow channel  842 . The first axial extending rib  836   a  widens in the radial rib width Win a direction toward the leading edge of the airfoil  800  (i.e., in a direction from the aft end  838  to the forward end  840 ). A second axial extending rib  836   b  is shown having a constant radial width from the aft end  838  to the forward end  840 . A third axial extending rib  836   c  is shown having a narrowing radial width in the direction from the aft end  838  to the forward end  840 . Also, as shown, support element retention features  826  can take complimentary dimensions or may be different in dimension from the forward end of a respective axial extending rib. Although shown as a mix of different axial extending rib styles, a given airfoil may be arranged having all of a single or complimentary type, to enable a desired cooling flow through axial extending flow channels, with either narrowing, widening, or constant radial width channels in a direction from aft end to forward end. 
       FIG. 9  illustrates a partial elevation view of a portion of an airfoil  900  having a baffle  901  installed therein, in accordance with an embodiment of the present disclosure. The arrangement of the airfoil  900  and the baffle  901  may be substantially similar to that shown and described above. The baffle  901  is arranged within the airfoil  900  and supported at a forward end by support element retention features  926  that extend from a pressure side  906  and a suction side  908  of the airfoil  900 . As described above, the support element retention features  926  extend into a leading edge cavity defined at the leading edge of the airfoil  900 . The baffle  901  is configured to form axial extending flow channels  942  between axial extending ribs (not visible in this view), a suction side surface  930 , a pressure side surface  928 , a pressure side wall  952  of the baffle  901 , and a suction side wall  954  of the baffle  901 . 
     As shown, the pressure side wall  952  and the suction side wall  954  of the baffle  901  are arranged to contact or engage with the support element retention features  926  at the forward end of the baffle  901 . A forward wall  950  extends in a direction from the pressure side to the suction side (or circumferentially; left-right in  FIG. 9 ) between the pressure side wall  952  and the suction side wall  954  of the baffle  901 . The forward wall  950  defines a wall of the baffle  901  and, when installed within the airfoil  900 , defines the forward most wall of the baffle  901  (i.e., closest to a leading edge of the airfoil  901 ). The forward wall  950  can define, in part, a portion of a leading edge portion of the leading edge cavity. As shown, a series of impingement apertures  948  are formed on/in the forward wall  950  of the baffle  901 . The impingement apertures  948  enable air from within the baffle  901  to impinge into the leading edge portion of the leading edge cavity. 
     As shown, the support element retention features  926  are separated by a circumferential gap  962 . As such, the support element retention features  926  do not span the full extent between the suction side surface  930  and the pressure side surface  928 . Such circumferential gap  962  may reduce the weight of the airfoil  900 , while providing for support and positioning of the baffle  901  within the airfoil  900 . 
     Turning now to  FIG. 10 , a schematic illustration of a portion of an airfoil  1000  in accordance with an embodiment of the present disclosure is shown. The airfoil  1000  may be substantially similar to that shown and described above. For example, the airfoil  1000  is arranged having a leading edge cavity configured to receive a baffle in a baffle portion thereof. The airfoil  1000  includes, along pressure and suction side walls of the airfoil  1000 , axial extending ribs  1036 . The axial extending ribs  1036  extend between an aft end of the respective surface to a forward end of the respective surface, as shown and described above. In this illustrative embodiment, the axial extending ribs  1036  are not evenly or equally distributed in a radial direction between an inner diameter  1018  and an outer diameter  1020 . As shown, some of the axial extending ribs  1036  may be separated by a first radial separation distance S 1  and other axial extending ribs  1036  may be separated by a second radial separation distance S 2 . As shown, the first radial separation distance S 1  is greater than the second radial separation distance S 2 . The disclosed configuration of this embodiment may be combined with other features, such as variable heights, variable width, etc. as shown and described herein. 
     Although shown illustratively as having the axial extending ribs oriented in substantially parallel arrangements, such configurations are not to be limiting, but are rather provided for illustrative and explanatory purposes. In some embodiments of the present disclosure, the ribs may not be purely axial and may vary spatially relative to any rib in order to create a passage width that is converging, diverging, and/or both converging and diverging. It will be appreciated that the ribs of such configurations will have a substantially axial extend or direction, but the structure an orientation is not limited to only axial in extent. That is, the illustrative embodiments are merely provided for explaining the functionality of the ribs and are not intended to be limiting on the structure, orientation, relative configurations, geometries, shapes, sizes, etc., as will be appreciated by those of skill in the art in view of the teachings provided herein. 
     For example, turning now to  FIG. 11 , a schematic illustration of an airfoil  1100  in accordance with an embodiment of the present disclosure is shown. The airfoil  1100  extends in a radial direction R between an inner diameter  1102  and an outer diameter  1104  and between a leading edge  1106  and a trailing edge  1108  in an axial direction A. As shown, the outer diameter  1104  of the airfoil  1100  may not be parallel to the inner diameter  1102 , as may be used for some airfoils within gas turbine engines. The airfoil  1100  includes internal cavities, such as shown and described above. The interior surfaces of at least one of the internal cavities includes axial extending ribs  1110  and associated support element retention features  1112 , as shown and described above. 
     In this embodiment, the axial extending ribs  1110  are not purely axial along the axial direction A, but rather may be angled relative to the axial direction A, but have a general axial extent. The axial extending ribs  1110  may be evenly or unevenly distributed in the radial direction and may be separated by different radial separation distances S 1 , S 2 , etc. (e.g., constant separation distance along axial length) or may have varying radial separation distances D 1 , D 2  (along axial length) between two radially adjacent axially extending ribs  1110 , as shown. As such, the axial extending ribs  1110  define different configurations of axial extending flow channels  1114  therebetween. The axial extending flow channels  1114  define flow paths for cooling air from first metering flow apertures  1116  (in a “space-eater baffle”) at an aft end and second metering flow apertures  1118  at a forward end. As shown in this embodiment, the axial extending flow channels  1114  may be configured with multiple associated first metering flow apertures  1116  that supply cooling air into the axial extending flow channels  1114 . 
     The configuration shown in  FIG. 11  illustrates that the axial extending ribs do not need to be purely parallel to an engine axis, or even parallel to each other. That is, the beginning and ends of the axial extending ribs may not be at a single radial position (e.g., relative to the inner diameter  1102  of the airfoil  1100 ), and may not span a single constant radius position. Further, in some embodiments, the axial extending ribs may be fanned, such that no two axial extending ribs are parallel (or some subset comprises a non-parallel set). Accordingly, configurations implementing the features illustrated in  FIG. 11  can enable control of streamwise channel flow area (e.g., defined in part by separation distances D 1 , D 2 ). The control of the streamwise channel flow area may also be controlled by varying the rib height (e.g., as shown in  FIG. 5  and/or  FIG. 7C ). Furthermore, the relative pitch or separation distances S 1 , S 2  of the axially extending ribs  1110  may be unique and/or varying depending on local external heatload and backside heat transfer, cooling flow, cooling air temperature heat pickup, and pressure loss requirement. Furthermore, the support element retention features  1112  do not have to be aligned linearly (e.g., along a single radii through the airfoil  1100  in a direction between the inner diameter  1102  and the outer diameter  1104 ) and may be offset in a monotonically or curvilinear with no inflections from any adjacent “space-eater” baffle retention feature  1112 , to ensure “space-eater” baffle insertion and contact along each of the support element retention features  1112 . 
     Advantageously, embodiments described herein provide for improved cooling schemes for airfoils. In accordance with embodiments of the present disclosure, airfoils, such as vanes for gas turbine engines, may be formed to receive a baffle and be arranged to have forward flowing cooling flow proximate the leading edge of the airfoil. In some embodiments, airfoils incorporate a leading edge “space-eater” baffle arranged adjacent segregated axial extending ribs to form axial (and forward) flowing cooling channels. Advantageously, in accordance with various embodiments of the present disclosure, an axial channel flow area in the axial streamwise direction may be constant, converging, diverging, or combinations thereof, with such flow area controlled by variable rib heights or widths. 
     In accordance with some embodiments of the present disclosure, a “space-eater” baffle is provided to form a counter-flow cooling concept in which a high pressure feed source can be optimally leveraged by managing pressure losses within a cooling system in order to provide more efficient and effective use of cooling airflow for improved convective and film cooling of a vane airfoil. Advantageously, in accordance with some design concepts of the present disclosure, a larger inlet feed may be incorporated along the outer diameter of a leading edge rail to mitigate plugging caused by internal sources (e.g., compressor rub strip material, blade outer air seal coating, w-seal/brush seal material, etc.) and external environmental sources (e.g., dirt, sand, debris, etc.). Further, advantageously, optimization of pressure loss may be achieved through various heat transfer augmentation features and orifices within the system. 
     Features that may be incorporated into embodiments of the present disclosure may include, but are not limited to, leading edge “space-eater” baffle feed/resupply flow apertures sizes and shapes that may be tailored specifically for each pressure side and suction side axial flow channel to optimize both the radial and axial cooling flow distribution in each of the axial flow channels. Further, the axial channel flow area, Mach number, trip strip or heat transfer augmentation type (e.g., pitch, height, spacing, geometry, etc.) may be varied and included or omitted as desired for a specific airfoil application. Metering apertures and/or baffle retention features (support element retention features) located immediately upstream of the leading edge portion of the cavity may be customized for a specific application in terms of size, shape, blocking characteristics, etc. Such support element retention features may be spaced radially along the internal pressure side and/or suction side of the interior airfoil surfaces. It is noted that although shown and described above as having the support element retention features on the suction side with a mirror image implied upon the pressure side, in some embodiments, the support element retention features may be arranged on only one of the pressure or suction sides. 
     Orifices or apertures as described herein may be integral with axial ribs and/or may tailored radially in both flow area size and spacing depending on external heat load and cooling effectiveness requirements. Further, because the metering apertures are located adjacent to the leading edge pressure side and suction side internal surfaces, trip strips may be incorporated in the leading edge portion of the cavity to augment the local convective heat transfer and thermal cooling effectiveness at the leading edge of the airfoil. Moreover, the geometry of the support element retention features (e.g., height, width, length) may be tailored to optimize local conduction and fin efficiency. Further, advantageously, in some embodiments, the support element retention features may also incorporate a variable taper depending on structural load and core die manufacturing requirements to mitigate die lock and die pull constraints. 
     Although the various above embodiments are shown as separate illustrations, those of skill in the art will appreciate that the various features can be combined, mix, and matched to form an airfoil having a desired cooling scheme that is enabled by one or more features described herein. Thus, the above described embodiments are not intended to be distinct arrangements and structures of airfoils, but rather are provided as separate embodiments for clarity and ease of explanation. For example, different axial extending rib orientations, geometries, dimensions, etc. and features thereof may be selected for a desired cooling scheme of an airfoil, and each individual disclosed and described embodiment is not intended to be limiting, but rather provided for explanatory and illustrative purposes only. 
     As used herein, the terms “about” and “substantially” are intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” may include a range of ±8%, or 5%, or 2% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein. Further, for example, the term “substantially” allows for deviations with the skill of those in the art. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a,” “an,” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” “radial,” “axial,” “circumferential,” and the like are with reference to normal operational attitude and should not be considered otherwise limiting. 
     While the present disclosure has been described with reference to an illustrative embodiment or embodiments, 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 present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.