Patent Publication Number: US-2022228493-A1

Title: Airfoil tip rail and method of cooling

Description:
CROSS-REFERENCE TO RELATED APPLICATION(S) 
     This application is a continuation of U.S. patent application Ser. No. 16/223,308, filed Dec. 18, 2018, now allowed, which is incorporated herein by reference in its entirety. 
    
    
     BACKGROUND 
     Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of combusted gases passing through the engine onto a multitude of rotating turbine blades. 
     Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency, so cooling of certain engine components, such as the high pressure turbine and the low pressure turbine, can be beneficial. Typically, cooling is accomplished by ducting cooler air from the high and/or low pressure compressors to the engine components that require cooling. Temperatures in the high pressure turbine are around 1000° C. to 2000° C. and the cooling air from the compressor is around 500° C. to 700° C. While the compressor air is at a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine. 
     Contemporary turbine engine airfoils generally include one or more interior cooling circuits for routing the cooling air through the airfoil to cool different portions of the airfoil, and can include dedicated cooling circuits for cooling different portions of the airfoil. 
     BRIEF DESCRIPTION 
     In one aspect, the disclosure relates to an airfoil for a turbine engine. The airfoil includes an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending axially between a leading edge and a trailing edge to define a chord-wise direction, and also extending radially between a root and a tip to define a span-wise direction, at least one cooling conduit formed in the interior of the airfoil, a tip rail projecting from the tip in the span-wise direction, the tip rail including an exterior surface spaced from an interior surface with a tip surface connecting the exterior and interior surfaces, and a three-dimensional plexus of fluidly interconnected cooling passages provided within the tip rail between the exterior and interior surfaces and fluidly coupled to the at least one cooling conduit. 
     In another aspect, the disclosure relates to a turbine engine. The turbine engine can include a compressor, a combustor, and a turbine in axial flow arrangement, and an airfoil including an outer wall bounding an interior and defining a pressure side and a suction side, the outer wall extending axially between a leading edge and a trailing edge to define a chord-wise direction, and also extending radially between a root and a tip to define a span-wise direction, at least one cooling conduit formed in the interior of the airfoil, a tip rail projecting from the tip in the span-wise direction, the tip rail including an exterior surface spaced from an interior surface with a tip surface connecting the exterior and interior surfaces, and a three-dimensional plexus of fluidly interconnected cooling passages provided at least partially within the tip rail between the exterior and interior surfaces and fluidly coupled to the at least one cooling conduit. 
     In yet another aspect, the disclosure relates to a method of cooling an airfoil having a tip rail in a turbine engine. The method includes flowing cooling air through at least one cooling conduit within an interior of the airfoil, flowing the cooling air from the at least one cooling conduit through a three-dimensional plexus of fluidly interconnected cooling passages located at least partially within the tip rail, and ejecting the cooling air from the three-dimensional plexus through a set of outlets located on the tip rail. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG. 1  is a schematic cross-sectional diagram of a turbine engine for an aircraft. 
         FIG. 2  is a perspective view of an airfoil assembly in the turbine engine of  FIG. 1  including an airfoil according to various aspects described herein. 
         FIG. 3  is a perspective view of the airfoil of  FIG. 2  including a tip rail and a network of cooling passages. 
         FIG. 4  is a side sectional view of the airfoil of  FIG. 2 . 
         FIG. 5  is a schematic view of the network of cooling passages of  FIG. 3 . 
     
    
    
     DESCRIPTION OF EMBODIMENTS 
     The described embodiments of the present disclosure are directed to a cooled airfoil and tip rail for a turbine engine. For purposes of illustration, the present disclosure will be described with respect to an aircraft turbine engine. It will be understood, however, that the disclosure is not so limited and may have general applicability within an engine, including a compressor section or turbine section of a turbine engine, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. 
     As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component. 
     As used herein, “a set” can include any number of the respectively described elements, including only one element. Additionally, the terms “radial” or “radially” as used herein refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference. 
     All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader&#39;s understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. 
       FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine  10  for an aircraft. The engine  10  has a generally longitudinally extending axis or centerline  12  extending forward  14  to aft  16 . The engine  10  includes, in downstream serial flow relationship, a fan section  18  including a fan  20 , a compressor section  22  including a booster or low pressure (LP) compressor  24  and a high pressure (HP) compressor  26 , a combustion section  28  including a combustor  30 , a turbine section  32  including a HP turbine  34 , and a LP turbine  36 , and an exhaust section  38 . 
     The fan section  18  includes a fan casing  40  surrounding the fan  20 . The fan  20  includes a plurality of fan blades  42  disposed radially about the centerline  12 . The HP compressor  26 , the combustor  30 , and the HP turbine  34  form a core  44  of the engine  10 , which generates combustion gases. The core  44  is surrounded by core casing  46 , which can be coupled with the fan casing  40 . 
     A HP shaft or spool  48  disposed coaxially about the centerline  12  of the engine  10  drivingly connects the HP turbine  34  to the HP compressor  26 . A LP shaft or spool  50 , which is disposed coaxially about the centerline  12  of the engine  10  within the larger diameter annular HP spool  48 , drivingly connects the LP turbine  36  to the LP compressor  24  and fan  20 . The spools  48 ,  50  are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor  51 . 
     The LP compressor  24  and the HP compressor  26  respectively include a plurality of compressor stages  52 ,  54 , in which a set of compressor blades  56 ,  58  rotate relative to a corresponding set of static compressor vanes  60 ,  62  to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage  52 ,  54 , multiple compressor blades  56 ,  58  can be provided in a ring and can extend radially outwardly relative to the centerline  12 , from a blade platform to a blade tip, while the corresponding static compressor vanes  60 ,  62  are positioned upstream of and adjacent to the rotating blades  56 ,  58 . It is noted that the number of blades, vanes, and compressor stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  56 ,  58  for a stage of the compressor can be mounted to (or integral to) a disk  61 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 . The vanes  60 ,  62  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     The HP turbine  34  and the LP turbine  36  respectively include a plurality of turbine stages  64 ,  66 , in which a set of turbine blades  68 ,  70  are rotated relative to a corresponding set of static turbine vanes  72 ,  74  (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage  64 ,  66 , multiple turbine blades  68 ,  70  can be provided in a ring and can extend radially outwardly relative to the centerline  12  while the corresponding static turbine vanes  72 ,  74  are positioned upstream of and adjacent to the rotating blades  68 ,  70 . It is noted that the number of blades, vanes, and turbine stages shown in  FIG. 1  were selected for illustrative purposes only, and that other numbers are possible. 
     The blades  68 ,  70  for a stage of the turbine can be mounted to a disk  71 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 . The vanes  72 ,  74  for a stage of the compressor can be mounted to the core casing  46  in a circumferential arrangement. 
     Complementary to the rotor portion, the stationary portions of the engine  10 , such as the static vanes  60 ,  62 ,  72 ,  74  among the compressor and turbine section  22 ,  32  are also referred to individually or collectively as a stator  63 . As such, the stator  63  can refer to the combination of non-rotating elements throughout the engine  10 . 
     In operation, the airflow exiting the fan section  18  is split such that a portion of the airflow is channeled into the LP compressor  24 , which then supplies pressurized air  76  to the HP compressor  26 , which further pressurizes the air. The pressurized air  76  from the HP compressor  26  is mixed with fuel in the combustor  30  and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine  34 , which drives the HP compressor  26 . The combustion gases are discharged into the LP turbine  36 , which extracts additional work to drive the LP compressor  24 , and the exhaust gas is ultimately discharged from the engine  10  via the exhaust section  38 . The driving of the LP turbine  36  drives the LP spool  50  to rotate the fan  20  and the LP compressor  24 . 
     A portion of the pressurized airflow  76  can be drawn from the compressor section  22  as bleed air  77 . The bleed air  77  can be drawn from the pressurized airflow  76  and provided to engine components requiring cooling. The temperature of pressurized airflow  76  entering the combustor  30  is significantly increased. As such, cooling provided by the bleed air  77  is necessary for operating of such engine components in the heightened temperature environments. 
     A remaining portion of the airflow  78  bypasses the LP compressor  24  and engine core  44  and exits the engine assembly  10  through a stationary vane row, and more particularly an outlet guide vane assembly  80 , comprising a plurality of airfoil guide vanes  82 , at the fan exhaust side  84 . More specifically, a circumferential row of radially extending airfoil guide vanes  82  are utilized adjacent the fan section  18  to exert some directional control of the airflow  78 . 
     Some of the air supplied by the fan  20  can bypass the engine core  44  and be used for cooling of portions, especially hot portions, of the engine  10 , and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor  30 , especially the turbine section  32 , with the HP turbine  34  being the hottest portion as it is directly downstream of the combustion section  28 . Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor  24  or the HP compressor  26 . 
     Referring now to  FIG. 2 , a perspective view of an airfoil assembly  95  is shown that can be utilized in the turbine engine  10  ( FIG. 1 ). The airfoil assembly  95  can be utilized anywhere in the turbine engine  10 , including the compressor section  22  or turbine section  32  in non-limiting examples. 
     The airfoil assembly  95  includes an airfoil  100  with an outer wall  102  bounding an interior  104  and defining a pressure side  106  and suction side  108 . The outer wall  102  extends axially between a leading edge  110  and a trailing edge  112  to define a chord-wise direction  121 . The outer wall  102  also extends radially between a root  114  and a tip  116  to define a span-wise direction  122 . 
     The airfoil  100  can also be mounted to a dovetail  130  at the root  114 . The dovetail  130  can include a set of dovetail cooling passages  132 . A set of cooling conduits  125  can be formed in the interior  104  of the airfoil  100 . The set of dovetail cooling passages  132  can fluidly couple with the set of cooling conduits  125  to supply cooling air to the interior  104  of the airfoil  100 . 
     A tip rail  140  can project from the tip  116  of the airfoil  100  in the span-wise direction  122 . The tip rail  140  can include a suction-side portion  142  located on the suction side  108 , as well as a pressure-side portion  144  located on the pressure side  106  of the airfoil  100 . A set of suction-side outlets  146  can be located on the suction-side portion  142 . The suction-side outlets  146  can be fluidly coupled to the set of cooling conduits  125  or the set of dovetail cooling passages  132 . Optionally, a set of pressure-side outlets  148  can be provided on the pressure-side portion  144  of the tip rail  140  and fluidly couple to the set of cooling conduits  125 . 
     The tip rail  140  can also include an exterior surface  150  spaced from an interior surface  152 , where a tip surface  154  can connect the exterior and interior surfaces  150 ,  152 . It is contemplated that the suction-side outlets  146  can be located on at least one of the exterior surface  150 , tip surface  154 , or interior surface  152  of the suction-side portion  142  of the tip rail  140  as desired. While shown on the tip surface  154 , it will be understood that the suction-side outlets  146  and pressure-side outlets  148  can also be located on the exterior surface  150  or interior surface  152  of the tip rail  140 . In addition, it is contemplated that the suction-side outlets  146  and pressure-side outlets  148  can include any type of outlets, such as a film hole, ejection hole, in-line diffuser, diffusing slot, or channel, in non-limiting examples. 
     Turning to  FIG. 3 , the airfoil  100  and tip rail  140  are shown in dotted-line for clarity, with the suction side  108  more clearly seen. It is further contemplated that a three-dimensional plexus  160  of fluidly interconnected cooling passages  162 , shown in solid line, can be provided within the airfoil  100 . The three-dimensional plexus  160  is illustrated schematically with “flat” passages and regions. It should be understood that the three-dimensional plexus  160  shown in solid line represents a three-dimensional void within the airfoil  100 . 
     The three-dimensional plexus  160  can be formed using a variety of methods, including additive manufacturing, casting, electroforming, or direct metal laser melting, in non-limiting examples. It is contemplated that the airfoil  100  having the plexus  160  can be an additively manufactured component. As used herein, an “additively manufactured” component will refer to a component formed by an additive manufacturing (AM) process, wherein the component is built layer-by-layer by successive deposition of material. AM is an appropriate name to describe the technologies that build 3D objects by adding layer-upon-layer of material, whether the material is plastic, metal, composite, or other. AM technologies can utilize a computer, 3D modeling software (Computer Aided Design or CAD), machine equipment, and layering material. Once a CAD sketch is produced, the AM equipment can read in data from the CAD file and lay down or add successive layers of liquid, powder, sheet material or other material, in a layer-upon-layer fashion to fabricate a 3D object. It should be understood that the term “additive manufacturing” encompasses many technologies including subsets like 3D Printing, Rapid Prototyping (RP), Direct Digital Manufacturing (DDM), layered manufacturing and additive fabrication. Non-limiting examples of additive manufacturing that can be utilized to form an additively-manufactured component include powder bed fusion, vat photopolymerization, binder jetting, material extrusion, directed energy deposition, material jetting, or sheet lamination. 
     The three-dimensional plexus  160  can be provided at least partially within the tip rail  140  between the exterior and interior surfaces  150 ,  152 . In the illustrated example, the set of cooling conduits  125  can further include a first cooling conduit  126  and a leading-edge cooling conduit  128  each fluidly coupled to the three-dimensional plexus  160 . It should be understood that any number of cooling conduits can be included within the interior  104  of the airfoil  100 . In an alternate example (not shown), the three-dimensional plexus can have a first plexus portion fluidly coupled to a first cooling conduit and a second plexus portion, locally separated from the first portion, fluidly coupled to a second cooling conduit. 
     The three-dimensional plexus  160  can include a lattice portion  164  having at least one impingement zone  166  fluidly coupled to either or both of the first or leading-edge cooling conduit  126 ,  128 . In the example shown, three impingement zones  166  are fluidly coupled to the first cooling conduit  126  and one impingement zone  166  is fluidly coupled to the leading-edge cooling conduit  128 . Any number of impingement zones can be provided. In addition, an impingement zone  166  can optionally have a narrower cross-sectional area than its fluidly-coupled cooling conduit (e.g. the cooling conduits  126 ,  128 ) such that air flowing through the impingement zone  166  is accelerated into the lattice portion  164 . 
     The lattice portion  164  in one example can form a lattice pin bank, and can also extend at least in the span-wise direction  122  and the chord-wise direction  121 . It should be understood that the cooling passages  162  forming the lattice portion  164  can extend in multiple directions, including in a cross-wise direction  123  between the pressure side  106  and suction side  108 , such that the lattice portion  164  can provide for airflow in three dimensions within the airfoil  100 . The cooling passages  162  can have a constant or variable cross-sectional width, and can also have any desired cross-sectional geometry including rounded or squared corners in non-limiting examples. 
     A set of elongated ejection holes  167  can be included in the three-dimensional plexus  160  and extend from the lattice portion  164 . The ejection holes  167  can extend at least axially (e.g. in the chord-wise direction  121 ) or radially (e.g. in the span-wise direction  122 ) from the lattice portion  164 . The set of elongated ejection holes  167  can be fluidly coupled to a set of outlets, such as the set of suction-side outlets  146  on the tip rail  140 . For example, a first suction-side outlet  171  is shown on the exterior surface  150  and fluidly coupled to a first ejection hole  181 , a second suction-side outlet  172  is shown on the interior surface  152  and fluidly coupled to a second ejection hole  182 , and a third suction-side outlet  173  is shown on the tip surface  154  and fluidly coupled to a third ejection hole  183 . It will be understood that any number and arrangement of suction-side outlets  146  and ejection holes  167  can be utilized. In this manner, the set of suction-side outlets  146  can be fluidly coupled to the three-dimensional plexus  160 . 
       FIG. 4  schematically illustrates an axial view of the airfoil  100  (shown in solid line), where arrows  195  illustrate the flow of cooling air through the cooling conduits  126 ,  128  (shown in dotted line, and overlapping in this view) and three-dimensional plexus  160  (shown in solid line). The first, second, and third ejection holes  181 ,  182 ,  183  are shown coupled to the respective first, second, and third suction-side outlets  171 ,  172 ,  173 . 
     It is further contemplated that the airfoil  100  can include at least one winglet  190 . In the illustrated example, the winglet  190  is included in the suction-side portion  142  of the tip rail  140  and defines a portion of the airfoil  100  or tip rail  140  that extends in the cross-wise direction  123  orthogonal to the span-wise direction  122 . For example, winglets can be included in airfoils to improve aerodynamic performance of the airfoil or tailor airflows surrounding the airfoil. In addition, the three-dimensional plexus  160  can be located at least partially within the winglet  190 , as shown by the lattice portion  164  and set of elongated ejection holes  167  curving in the cross-wise direction  123  along with the winglet  190 . 
     In the illustrated example, the pressure-side portion  144  of the tip rail  140  can also include at least one pressure-side cooling passage  145  between the exterior and interior surfaces  150 ,  152 . The pressure-side cooling passage  145  is shown as being fluidly coupled to a second cooling conduit  127  within the interior  104  of the airfoil  100 . A set of pressure-side outlets  148  can be located on the pressure-side portion  144  of the tip rail  140  and fluidly coupled to the pressure-side cooling passage  145 . As described above, while shown on the tip surface  154  of the pressure-side portion  144 , the pressure-side outlets  148  can also be located on the exterior surface  150  or interior surface  152 . It is contemplated that the second cooling conduit  127  can be fluidly separated from the first cooling conduit  126  or the leading-edge cooling conduit  128 . Alternately, any or all of the first, second, or leading-edge cooling conduits  126 ,  127 ,  128  can be fluidly coupled, such as via a common dovetail cooling passage  132  ( FIG. 2 ). In this manner, a first cooling conduit can be fluidly coupled to the three-dimensional plexus  160  and a second cooling conduit can be fluidly coupled to at least one cooling passage provided within the pressure-side portion  144  of the tip rail  140 . 
     Turning to  FIG. 5 , a perspective view of the tip rail  140  is shown in dotted line proximate the leading edge  110  of the airfoil  100 . A portion of the three-dimensional plexus  160  is schematically illustrated in solid line. It is contemplated that the three-dimensional plexus  160  can extend or curve three-dimensionally in the span-wise direction  122 , the chord-wise direction  121 , or the cross-wise direction  123 . The three-dimensional plexus  160  can curve within the tip rail  140 , such as within the winglet  190 , and the elongated ejection holes  167  can extend at least axially in the chord-wise direction  121 , at least radially in the span-wise direction  122 , or at least in the cross-wise direction  123 . In addition, the elongated ejection holes  167  can be parallel with one another, or adjacent ejection holes  167  can be unaligned or intertwined with one another. In the illustrated example, the first, second, and third ejection holes  181 ,  182 ,  183  can curve in three dimensions within the winglet  190  and fluidly couple to their respective suction-side outlets  171 ,  172 ,  173  on the respective exterior surface  150 , interior surface  152 , and tip surface  154  as described above. 
     During operation, cooling air supplied from the set of dovetail cooling passages  132  can flow from the set of cooling conduits  125  through the impingement zone  166  to impinge at least one passage wall within the lattice portion  164  and provide initial cooling to the lattice portion  164 . Cooling air can be directed through the lattice portion  164  in a variety of directions, including at least two dimensions or in three dimensions, and can then flow through the elongated ejection holes  167 . Cooling air can exit the tip rail  140  via the set of suction-side outlets  146  or the pressure-side outlets  148 . Optionally, the winglet  190  can provide for improved aerodynamic flow around the airfoil  100  as well as additional directional flow for the cooling air exiting the suction-side portion  142  of the tip rail  140 . 
     It should be understood that the airfoil and plexus described above can include any suitable arrangement or position of ejection holes, outlets, lattice portion, and the like. For example, cooling air can move radially outward, radially inward, axially, or the like, or any combination thereof within the airfoil interior while moving through the plexus and toward the outlets. 
     In one non-limiting example (not shown), the lattice portion of the plexus can extend along at least a portion of both the trailing edge and the tip of the airfoil, and the ejection holes can extend from the lattice portion to outlets located on both the tip rail and trailing edge. In addition, a single ejection hole can branch or divide to be fluidly coupled to multiple outlets, such as to a first outlet on the tip rail and to a second outlet on the trailing edge in one example. 
     In another non-limiting example (not shown), at least a portion of the lattice portion can be directly fluidly coupled to tip rail outlets without the elongated ejection holes. In such a case, the lattice portion can extend fully through the tip rail and fluidly couple to the outlets, including outlets located on the exterior surface, interior surface, or tip surface of the tip rail. The lattice portion can also be fluidly coupled to other outlets located on the pressure side or suction side of the airfoil, including by way of the elongated ejection holes or by directly fluidly coupling to the outlets without such ejection holes. 
     In yet another non-limiting example (not shown), the plexus can further include multiple discrete groups of cooling passages each fluidly supplied by a separate cooling conduit. Each of the multiple discrete groups can include any or all of the impingement zone, lattice portion, or elongated ejection holes. The multiple discrete groups can be fluidly coupled, for example by a single connecting fluid passage, or they can be separated within the airfoil interior. In addition, the multiple discrete groups can form multiple impingement zones arranged radially within the airfoil, such that cooling air supplied from the cooling conduit can impinge a first zone, impinge a second zone, impinge a third zone, and so on, until exiting via a cooling hole outlet. 
     In still another non-limiting example (not shown), the cooling conduit can “blend” or transition into the lattice portion without an intervening impingement zone. In such a case, cooling air supplied by the cooling conduit can flow directly into the lattice portion and be smoothly directed into passages therein without impingement. 
     In still another non-limiting example (not shown), a radially-oriented elongated ejection hole can be fluidly coupled to a second radially-oriented passage with an outlet, such as a film hole on the pressure side of the airfoil. In such a case, cooling air flowing through the plexus can flow radially outward through the ejection hole, radially inward through the passage, and then out through the film hole. 
     Aspects provide for a method of cooling an airfoil having a tip rail in a turbine engine. The method includes flowing cooling air through at least one cooling conduit, such as the set of cooling conduits  125 , within the interior  104  of the airfoil  100 . The method also includes flowing the cooling air from the set of cooling conduits  125  the three-dimensional plexus  160  of fluidly interconnected cooling passages  162  located at least partially within the tip rail  140 . Optionally, the method can further include impinging the cooling air within the impingement zone  166  of the three-dimensional plexus  160 . The method also includes ejecting the cooling air from the three-dimensional plexus  160  through a set of outlets, such as the suction-side outlets  146 , located on the tip rail  140 . 
     Aspects of the disclosure provide for a variety of benefits. The use of the three-dimensional plexus, impingement zones, and elongated ejection holes can increase the heat transfer to the cooling air from the airfoil and direct the now-heated cooling air to a location of the airfoil that can handle higher coolant temperatures. It can be appreciated that the elongated ejection holes can improve bore cooling, while the lattice with impingement zones can provide impingement cooling while directing the cooling air to various portions of the airfoil, including those portions where film cooling performance is traditionally limited. This enables higher convection of cooling air within the airfoil using the same or less supplied airflow compared with traditional cooling designs. The lattice portion can provide for increased working or mixing of the cooling air within the airfoil, which can improve cooling performance. In addition, the elongated ejection holes can be utilized to eject or sink the cooling air to a region with a lower air or sink pressure. Such lower sink pressures can maintain the flow of cooling air moving through and out of the three-dimensional plexus with reduced back flow, which further improves cooling performance. 
     It should be understood that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turboshaft engines as well. 
     To the extent not already described, the different features and structures of the various embodiments can be used in combination, or in substitution with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be so illustrated, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.