Patent Publication Number: US-2018051566-A1

Title: Airfoil for a turbine engine with a porous tip

Description:
BACKGROUND OF THE INVENTION 
     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. 
     Turbine engines for aircraft, such as gas turbine engines, are often 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 a high temperature, it is cooler relative to the turbine air, and can be used to cool the turbine. 
     Contemporary turbine components, such as airfoil blades, can include one or more interior cooling circuits for routing the cooling air through the blades to cool different portions of the blades, and can include dedicated cooling circuits for cooling different portions of the blades, such as the leading edge, trailing edge, or tip. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, embodiments of the invention relate to an airfoil for a turbine engine. The airfoil includes a wall defining an interior and an exterior, including a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further includes a cooling passage disposed in the interior extending in the span-wise direction to the tip. A porous material is provided on at least a portion of the tip and fluidly coupled to the cooling passage. 
     In another aspect, embodiments of the invention relate to a component for a turbine engine. The component includes a wall extending radially between a root and a tip. The component further includes a cooling circuit located within the component having a cooling passage extending toward the tip. A porous material is provided on at least a portion of the tip and fluidly coupled to the cooling passage. 
     In yet another aspect, embodiments of the invention relate to a method of cooling a tip of an airfoil. The method includes passing a flow of cooling fluid through a cooling passage in a span-wise direction toward the tip and passing at least a portion of the flow of cooling fluid through a porous material based upon a porosity of the porous material. 
    
    
     
       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 of the turbine engine of  FIG. 1 . 
         FIG. 3  is a cross-sectional view of the airfoil of  FIG. 2  illustrating ribs defining passages within an interior of the airfoil. 
         FIG. 4  is a top-down view of the airfoil of  FIG. 3  illustrating a tip cap having a porous material forming a portion of the tip cap. 
         FIG. 5  is a cross-sectional view of the airfoil of  FIG. 4  including a partial-length rib, to define a tip turn, with the porous material disposed in the tip turn. 
         FIG. 6  is a top-down view of the airfoil of  FIG. 2  having porous material forming the tip cap at a trailing edge of the airfoil. 
         FIG. 7  is a cross-sectional view of the airfoil of  FIG. 6  with the porous material forming a tip flag. 
         FIG. 8  is a cross-sectional view of the airfoil of  FIG. 6  including a tip passage and having the porous material forming a portion of the tip passage at the trailing edge. 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     The described embodiments of the present invention are directed to a blade for a gas turbine engine. For purposes of illustration, the present invention will be described with respect to the blade for an aircraft gas turbine engine. It will be understood, however, that the invention is not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications. Additionally, the aspects will have applicability outside of a blade, and can extend to any engine component requiring cooling, such as a vane, shroud, or a combustion liner in non-limiting examples. 
     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. 
     Additionally, as used herein, the terms “radial” or “radially” 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 invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. 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  (also called a nozzle) 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 a disk  61 , which is mounted to the corresponding one of the HP and LP spools  48 ,  50 , with each stage having its own disk  61 . 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 , from a blade platform to a blade tip, 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 , with each stage having a dedicated disk  71 . 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 airflow  76  to the HP compressor  26 , which further pressurizes the air. The pressurized airflow  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 draw 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 , an engine component is shown in the form of an airfoil  90 , which can be one of the turbine blades  68  of the engine  10  of  FIG. 1 . Alternatively, the engine component can include a vane, a shroud, or a combustion liner in non-limiting examples, or any other engine component that can require or utilize cooling. The airfoil  90  includes a dovetail  92  and a platform  94 . The airfoil  90  extends radially between a root  96  and a tip  98  defining a span-wise direction. In the case of a shrouded airfoil, the tip  98  can be the furthest radial extend to the airfoil from the engine centerline  12  ( FIG. 1 ). The airfoil  90  extends axially between a leading edge  100  and a trailing edge  102  defining a chord-wise direction. The dovetail  92  can be integral with the platform  94 , which can couple to the airfoil  90  at the root  96 . The dovetail  92  can be configured to mount to a turbine rotor disk  71  on the engine  10  ( FIG. 1 ). The platform  94  helps to radially contain the turbine airflow. The dovetail  92  comprises at least one inlet passage  104 , shown as three inlet passages  104 , each extending through the dovetail  92  in fluid communication with the airfoil  90  at a passage outlet  106 . It should be appreciated that the dovetail  92  is shown in cross-section, such that the inlet passages  104  are housed within the dovetail  92 . 
     It should be appreciated that in additional components that are not airfoils  90 , the tip  98  can be defined at a terminal end of the component, such as at an end of a serpentine cooling circuit, where the circuit turns, in one non-limiting example. 
     Referring now to  FIG. 3 , a cross-sectional view of the airfoil  90  illustrates an outer wall  120  including a pressure side  122  and a suction side  124  extending between the leading edge  100  and the trailing edge  102 . The outer wall  120  separates the hot fluid flow H external of the airfoil  90  from the cooling fluid flow C within the airfoil  90 , having a hot surface  126  along the exterior of the airfoil  90  and a cooling surface  128  confronting the cooling fluid flow C. An interior  130  of the airfoil  90  is defined by the outer wall  120 . One or more internal ribs  132  separates the interior  130  into cooling passages  134  extending in the span-wise direction. The passages  134  can define one or more cooling circuits throughout the airfoil  90 . Additionally, the cooling circuits can be further includes micro-circuits, sub-circuits, near wall cooling circuits, leading edge passages, trailing edge passages, pin fins, pin banks, additional passages  134 , flow enhancers such as turbulators, or any other structures which can define the cooling circuits. 
     Referring now to  FIG. 4 , a top-down view of the airfoil  90  illustrates a tip cap  140  enclosing the airfoil  90  at the tip  98 . A porous material  142  forms at least a portion of the tip cap  140 . Tip edge  143  can be an extension of the outer wall  120  extending above the tip cap  140  forming the edges of the tip cap  140 . The porous material  142 , for example, can be disposed at a mid-chord  144  where a chord  146  is a straight line drawn from the leading edge  100  to the trailing edge  102 . Where the mid-chord  144  extends external of the airfoil  90 , the mid-chord position for placing the porous material  142  can be along a line  148  perpendicular to the chord-line  146  from the mid-point  144  of the chord, into the airfoil  90 . In another example, the airfoil  90  can form a camber line  150  as the line along the center position between the pressure side  122  and suction side  124  extending from the leading edge  100  to the trailing edge  102 . The porous material  142  can be disposed at a center  152  of the camber line  150 . It should be appreciated that the porous material  142  can be positioned anywhere along the tip cap  140 , such as any position along the chord-line  146  or the camber line  148 . 
     The porous material can  142  can form an area  154  of the tip cap  140 . The area  154  can extend fully between the pressure side  122  and the suction side  124 . Additionally, the area  154  can form multiple discrete areas  154  along the tip cap  140 . Alternatively, the area  154  can extend only partially between the pressure side  122  and the suction side  124 , being adjacent to either side  122 ,  124  or spaced therefrom. As such, it should be appreciated that the area  154  of the porous material  142  forming the tip cap  140  can form discrete shapes, or profiles, as cross-sectional areas  154 . Such shapes or profiles can be determined by the particular needs to the airfoil  90  or the tip  98 . Such shapes or profiles, in non-limiting examples, can include geometrical shapes, rectilinear shapes, curvilinear shapes, unique shapes, or any combination thereof. Furthermore, the multiple discrete areas  154  can be discretely shaped, or similarly shaped, in non-limiting examples. The porous material can be formed in the tip cap  140 , forming as least a portion of the boundary walls of the tip cap  140  or an adjacent passage. Formed in the in the tip cap  140  should not be construed as limiting to the interior of the tip cap  140  or adjacent passage. 
     The porous material  142  can be porous such that no single flow passage exists without interconnection to another flow passage, from ingress to egress of the fluid passing through the porous material  142 . A particular closed region within the porous material  142 , however, may not interconnect to another passage within the porous material  142 . Additionally, the porous material  142  can be the mean free path length of the flow is much greater than the geometric length for a discrete hole extending fully through the material. In a further non-limiting example, the porosity of the porous material  142  can include porous holes less than 0.005 inches for a random porous material and with deterministic sizes of 0.005 inches and less for a structured porous material. It should be appreciated that many different sizes can be included to meter the flow through the porous materials  142 . 
     The porous material  142  can be made by additive manufacturing, while it is contemplated that additive manufacturing can form the entire airfoil  90 . It should be appreciated that any portion of the airfoil  90  can be made by any known method including but not limited to, casting, machining, additive manufacturing, coating, or otherwise. The porous material  142  can define a porosity, being permeable by a volume of fluid, such as air. Additionally, the porous material  142  can be alternatively defined as a region having no single flow passage without interconnection among another flow passage. The porous material  142  can have a particular porosity to meter the flow of a fluid passing through the porous material  142  at a predetermined rate. It should be appreciated that additive manufacturing can be used to achieve a particular local porosity along the porous material  142 , as well as a consistent porosity across the entirety of the porous material  142 , as compared to traditional method of forming the porous material  142 . In alternative examples, the porous material  142  can be made of any of the materials described above, such that a porosity is defined. In one non-limiting example, the porous material  142  can be made of Ni, NiCrAlY, NiAl, or similar materials. The porous material  142  can further be made of a nickel foam, for example. 
     Additionally, the porous material  142  can be a structured porous material or a random porous material, or any combination thereof. A structured porous material includes a determinative porosity throughout the material, which can have particular local increases or decreases in porosity to meter a flow of fluid passing through the structured porous material. Such local porosities can be determined and controlled during manufacture. Additive manufacturing can be used to form a structured porous material, in one non-limiting example. Alternatively, the porous materials can have a random porosity. The random porosity can be adapted to have a porosity as the average porosity over an area or volume of the porous material  142 , having discrete variable porosities that are random. A random porous material can be made from a nickel foam, in one non-limiting example. 
     In one non-limiting example, the porosity can have an average pore size of less than 0.005 inches for a random porous material and a deterministic 0.005 inches or less for a structure porous material. 
     Referring to  FIG. 5 , an exemplary cross-sectional view of the airfoil of  FIG. 4  includes the internal ribs  132  as two full-length ribs  160  and one partial-length rib  162  disposed between the full-length ribs  160 . The passages  134  are defined between the ribs  132 . A cooling circuit  164  can be formed by the passages  134  for providing the cooling fluid flow C throughout the interior  130  of the airfoil  90 . The partial-length ribs  162  organized between the full-length ribs  160  define the cooling circuit  164 , having a substantially serpentine flow path as illustrated. It should be understood that the cooling circuit  164  as illustrated is exemplary, and can include additional structures to form the cooling circuit  164 , such as micro-circuits, sub-circuits, near wall cooling circuits, leading edge passages, trailing edge passages, pin fins, pin banks, additional passages  134 , or flow enhancers such as turbulators in non-limiting examples. A tip turn  166  is formed by the partial-length rib  162  adjacent the tip cap  140 , and forms a portion of the cooling circuit  164 . 
     The porous material  142  can be disposed in the tip cap  140  at the tip turn  166 . The cooling fluid flow C can be provided within the cooling circuit  164  through the passages  134  and to the tip turn  166 . At least a portion of the cooling fluid flow C can be exhausted through the porous material  142  in the tip cap  140  as a tip flow  168 . The porous material  142  can be disposed in the tip cap  140  for the entirety of the tip turn  166  or for a portion of the tip turn  166 . The tip flow  168  can form a cooling film along the tip  98  of the airfoil  90 , as well as maintain pressure within a tip cavity of the airfoil to maintain tip pressures. 
     In the example where the component is not an airfoil  90 , the tip  98  can be toward the end of a serpentine cooling circuit, such as the cooling circuit  164  having a substantially serpentine flow path. The serpentine cooling circuit can be located in the component to define a turn, such as the tip turn  166  at the end of one or more cooling passages  134 . The tip can be defined at the turn, such as the tip turn  166 . The porous material  142  can define a portion of the tip  98  and fluidly couple to the cooling passages  134 . In such a component, the porous material  142  can span the tip  98 , extending from both sides of the component. Additionally, the component can include a rib such as the partial-length rib  162  to at least partially define the serpentine cooling circuit as well as the turn  166 . The porous material  142  can be spaced from the partial-length rib  162 , such as aligned with the partial-length rib  162 , or disposed at the shortest distance from the partial-length rib  162  to the tip  98 , in non-limiting examples. 
     Additionally, the serpentine cooling circuit  164  can be disposed adjacent the aft edge  102  ( FIG. 4 ) of the component or the tip  98 . The porous material  142  can at least partially define the aft edge  102  and the tip  98 , or a junction between the two. 
     Referring now to  FIG. 6 , another exemplary tip  98  of the airfoil  90  can include the porous material  142  disposed along the tip cap  140  adjacent the trailing edge  102 . The porous material  142 , similar to that of  FIG. 4 , can extend between the pressure and suction sides  122 ,  124 , as well as form any shape or have any cross-sectional area. It should be appreciated that the porous material  142  can also be disposed adjacent the leading edge  100 , as well as at multiple positions on the tip cap  140 , such as at both the leading edge  100  and the trailing edge  102  in one non-limiting example, and should not be limited as shown in  FIGS. 4 and 6 . 
     Referring now to  FIG. 7 , a trailing edge cooling circuit  180  can be formed along the trailing edge  102 . The trailing edge cooling circuit  180  can include the passage  134 , at least partially defined by the rib  132 , extending in the span-wise direction. 
     A plurality of trailing edge ribs  182  can be disposed in span-wise arrangement along the trailing edge  102  defining a plurality of trailing edge ejection holes  184 . The trailing edge ejection holes  184  can also be elongated, shaped as slots. A tip flag  186  can be disposed along the tip cap  140  at the trailing edge  102 . The tip flag  186  can be at least partially formed of the porous material  142 . The trailing edge ejection holes  184  and the tip flag  186  can further define the trailing edge cooling circuit  180 . 
     The flow of cooling fluid C can be provided to the trailing edge cooling circuit  180  through the inlet passage  104  in the dovetail  92 . The flow of cooling fluid C can be exhausted through the trailing edge  102  at the trailing edge ejection holes  184  as a first exhaust flow  188 . Additionally, the cooling fluid flow C can be exhausted through the tip flag  186  as a second exhaust flow  190 , which can exhaust through the tip flag  186  at the tip  98 , the trailing edge  102 , or a combination thereof. 
     Referring now to  FIG. 8 , the airfoil  90  can include the cooling circuit  164  of  FIG. 5  and the trailing edge cooling circuit of  180 , having the cooling circuits  164 ,  180  terminate radially inward of the tip  98 . A full-length passage  200  can extend from the root  96  to the tip  98 , coupling to a tip passage  202  extending chord-wise along the tip  98  of the airfoil  90  to the trailing edge  102 . The full-length passage  200  and the tip passage  202  can form a tip cooling circuit  204 . The tip passage  202  can include the porous material  142  forming a portion of the tip passage  202 . The porous material  142  can extend through the entire span-wise length of the tip passage  202 , or can extend partially therethrough. The porous material  142  can be positioned adjacent the trailing edge  102 , in one example. In another example, the porous material  142  can be positioned along the tip passage  202  and spaced from the trailing edge  102 , such as the middle of the tip passage  202  or adjacent the full-length passage  200 . It should be appreciated that the cooling circuit pattern, or layout, as shown is exemplary, and numerous cooling circuits  164  varying from the exemplary embodiment are contemplated. 
     Alternatively, in the implementation in an engine component other than an airfoil, the tip passage  202  can extend along the tip  98  defined near the end of a serpentine cooling circuit  164  or the turn  166 , with the tip passage in fluid communication with the porous material  142  and forming at least a portion of the tip  98 . 
     In operation, the flow of cooling fluid C is provided to the full-length passage  200  from one of the inlet passage  104  in the dovetail  92 . The flow of cooling fluid C within the full-length passage  200  can move in a substantially span-wise or radial direction. The full-length passage  200  feeds the tip passage  202 , where the cooling fluid flow C turns from a radial direction to a substantially axial or chord-wise direction, moving toward the trailing edge  102 . The flow of cooling fluid C is provided to the porous material  142 , where it can be exhausted from the tip passage  202  as a third exhaust flow  206 , through the tip  98 , the trailing edge  102 , or a combination thereof. 
     Additionally, while it is shown that the tip passage  202  is fed from the full-length passage  200 , it is contemplated that the tip passage  202  can be fed from another cooling circuit  164 ,  180 , or that the cooling circuits can be integrated into a single cooling circuit defined throughout the airfoil  90 . 
     A method of cooling a tip of an airfoil can include (1) passing flow of cooling fluid through a cooling passage in a span-wise direction toward the tip and (2) passing at least a portion of the flow of cooling fluid through a porous material located in the tip. Passing the cooling fluid can include passing the flow of cooling fluid C in the passages  134 ,  200  shown in  FIG. 5, 7 , or  8 , moving toward the tip  98 . The porous material can be the porous material  142  as described and shown in  FIG. 5, 7 , or  8 . 
     Additionally, the method can include exhausting the flow of cooling fluid from the porous material. Such an exhausted cooling fluid flow can be the second or third exhaust flows  190 ,  206  of  FIG. 7 or 8 , or can be the tip flow  168  of  FIG. 5 . Additionally, the method can include exhausting the cooling fluid from at least one of the tip  98  or the trailing edge  102  of the airfoil  90 . Furthermore, the method can include metering the flow of cooling fluid C through the porous material  142  based upon the porosity of the porous material  142 . For example, the porous material  142  can be a structured porous material  142 , made by additive manufacturing, having a local determinative flow by having discrete porosities throughout the porous material  142 . As such, the cooling fluid flow C can be metered locally based upon the particular local porosity. Alternatively, the flow C can be metered by an average porosity throughout the porous material  142 , such as having a random porosity being made by a nickel foam. 
     It should be appreciated that the porous material  142 , as described herein provides for improved airfoil cooling at the tip  98  of the airfoil  90 , particularly, adjacent tip turns  166  or at the trailing edge  102 . The porous portions  142  permit a volume of cooling air to pass through the tip cap  140  to provide a particular cooling fluid C passing through the tip  98  from an interior airfoil cooling circuit. Additionally, the porous material  142  can be used to increase or maintain structural integrity of the airfoil  90 , while maintaining or decreasing system weight without sacrificing cooling efficiency, or even improving the cooling efficiency. The porous material  142  can be significantly lighter than the other portions or materials used in constructing the airfoil  90 . 
     It should be appreciated that while the description is directed toward a trailing edge of the airfoil, the concepts as described herein can have equal applicability in additional engine components, such as a blade, vane, or other airfoil-shaped elements, such as a strut or outlet guide vane, and can be any region of any engine component requiring cooling, such as regions typically requiring film cooling holes or multi-bore cooling. For example, the porous portions could be placed in the endwalls and platforms or other components. 
     It should be further appreciated that the porous material  142  can provide for improved film cooling, or tip pressure maintenance, such as providing improved directionality, metering, or local flow rates. 
     It should be appreciated that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turbo engines as well. 
     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.