Patent Publication Number: US-2018051571-A1

Title: Airfoil for a turbine engine with porous rib

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 blades, can include one or more interior cooling circuits for routing the cooling air through the component to cool different portions of the component, and can include dedicated cooling circuits for cooling different portions of the component, such as the leading edge, trailing edge, or tip of the blade. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, embodiments of the invention relate to a component for a turbine engine. The component includes a wall bounding an interior. A cooling circuit is located in the interior having at least one rib that at least partially defines a flow channel. A porous material is provided in at least one rib to define a flow path through the at least one rib. 
     In another aspect, embodiments of the invention relate 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 extending axially between a leading edge and a trailing edge to define a chord-wise direction and extending radially between a root and a tip to define a span-wise direction. A cooling circuit is located within the interior and has at least one rib that at least partially defines a flow channel. A porous material is provided in at least one rib to define a flow path through the at least one rib. 
     In yet another aspect, embodiments of the invention relate to a method of reducing flow separation at a turn in a cooling circuit formed at least in part by a partial-length rib within an interior of an airfoil for a turbine engine. The method includes flowing cooling fluid through a porous material at an end of the partial length rib. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings: 
         FIG. 1  is a schematic cross-sectional diagram of a gas turbine engine for an aircraft. 
         FIG. 2  is a perspective view of an airfoil of the gas 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 section view of the airfoil of  FIG. 3  illustrating a cooling circuit within the interior defined by the ribs, with a partial-length rib having a porous portion. 
         FIG. 5  is a cross-sectional view of a turn in the cooling circuit of  FIG. 4  defined by the partial-length rib, with the porous portion spaced from the turn. 
         FIG. 6  is a cross-sectional view of the partial-length rib of  FIG. 5  having a solid structure within the porous portion. 
         FIG. 7  is a cross-sectional view of an alternative partial-length rib having the porous portion connecting the partial-length rib to a tip, while the porous portion can define the turn. 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     The described embodiments of the present invention are directed to an airfoil for a turbine engine. For purposes of illustration, the present invention will be described with respect to the airfoil for an aircraft 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 an airfoil, and can extend to any engine component requiring cooling, such as a vane, blade, 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. 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 on the engine  10 . The platform  94  helps to radially contain the turbine airflow. The dovetail  92  comprises at least one inlet passage, 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 . 
     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 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 to  FIG. 4 , a section view of the airfoil  90  illustrates an exemplary system of ribs  132  defining a cooling circuit  150  extending in the span-wise direction within the interior  126 . The ribs  132  are separated into first ribs and second ribs, illustrated as full-length ribs  140  and partial length ribs  142 , respectively. The full-length ribs  140  extend fully in the span-wise direction between the root  96  and the tip  98 . The partial-length ribs  142  extend only partially between the root  96  and the tip  98 , terminating at a rib end  144 . The partial-length ribs  142  organized between the full-length ribs  140  define a cooling circuit  150 , having a substantially serpentine flow path as illustrated. It should be understood that the cooling circuit  150  as illustrated is exemplary, and can include additional structures to form the cooling circuit  150 , 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. 
     The partial-length ribs  142  can include a porous portion  146  made of porous material. The porous portions  146  can extend from the rib end  144  radially along at least a portion of the partial-length ribs  142 . The porous portions  146  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 portions  146  can define a porosity, being permeable by a volume of fluid, such as air. The porous portions  146  can have a particular porosity to meter the flow of a fluid passing through the porous material at a predetermined rate. It should be appreciated that additive manufacturing can be used to achieve a particular local porosity along the porous portions  146 , as well as a consistent porosity across the entirety of the porous portions  146 , as compared to traditional method of forming the porous portions  146 . In alternative examples, the porous portions  146  can be made of any of the materials described above, such that a porosity is defined. In one non-limiting example, the porous portions  146  can be made of Ni, NiCrAlY, NiAl, or similar materials. The porous portions  146  can further be made of a nickel foam, for example. 
     Additionally, the porous material in the porous portions  146 , can be a structured porous material or a random porous material, or a 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, such as a non-structured porous material. The random porosity can be adapted to have a porosity as the average porosity over an area of the porous material, having discrete variable porosities that are random. A random porous material can be made from a nickel foam, in one non-limiting example. 
     A plurality of flow channels  148  can be defined between adjacent ribs  132  to further define the cooling circuit  150 . The partial-length ribs  142  at the rib end  144  forms a turn  152  within the cooling circuit, such as a tip turn or a root turn. The turns  152  include about a 180-degree change in direction from moving radially inward to radially outward relative to the engine centerline  12  ( FIG. 2 ). 
     The flow of cooling fluid C can be provided to the cooling circuit  150  from the inlet passage  104  in the dovetail  92 . The flow of cooling fluid C can pass through the serpentine path of the cooling circuit  150 . The flow cooling fluid C turns within the turns  152 . Additionally, a portion  154  of the flow of cooling fluid C can pass through the porous portions  146 , bypassing the turns  152 . The porosity of the porous portions  146  can be adapted to determine the flow rate of the portion of cooling fluid  154  through the porous portions  146 . 
     Referring now to  FIG. 5 , illustrating one exemplary position for the porous portion  146 , as positioned along the partial-length rib  142 , being spaced from the rib end  144 . The porous portion  146 , in one example, can be spaced from the rib end  144  by a distance less than or equal to a length L of the porous portion  146 . Alternatively, the porous portion  146  can be space from the rib end  144  by a distance of less than three times a width W of the porous portion  146 . In another example, the porous portion  146  need not extend full through the rib  142  between the pressure side  122  and the suction side  124 , but can extend only partially through the rib  142  with the porous portion  146  adjacent the pressure side  122 , the suction side  124 , or disposed in the middle of the rib  132 . Furthermore, it is contemplated that the porous portion  146  can be positioned anywhere along the partial-length rib  142 , however it is advantageous to place the porous portion  146  near to the turn  152  to prevent any cycling of the cooling fluid flow C through the cooling circuit  150 . 
     Referring now to  FIG. 6 , the porous portion  146  can include a framework  160 , which can be made of a plurality of solid elements. The framework  160  can be a single integral unit, or can be multiple discrete elements. In the case of multiple discrete elements, some or none of the framework  160  can couple to one another. The framework  160  can be linear, curved, or any combination thereof, having any cross-sectional shape or profile, such that any geometry is contemplated. As such, a myriad of framework  160  disposed within the porous portion are contemplated. 
     A plurality of interstitial spaces  162  are defined between the framework  160 . The porous material of the porous portion  146  can fill the interstitial spaces  162 . Discrete orifices  164  can be formed in the framework  160  to provide a flow path for the portion of cooling fluid  154  to pass through the framework  160  within the porous portion  146 . 
     As such, the framework  160  can be used to provide directionality to the portion of cooling fluid  154  passing through the porous portion  146 . Additionally, the framework  160  can meter the portion of cooling fluid  154  passing through the porous portion  146 , as well as increase structural integrity where desirable. The framework  160  can be made of any material, such as a similar material to that of the rib or the porous material. 
     Referring now to  FIG. 7 , another example airfoil  190  is illustrated having a partial-length rib  242  connected to a tip  198  with a porous material  246 . It should be appreciated that the airfoil  190  of  FIG. 7  can be substantially similar to the airfoil  90  of  FIGS. 4-6 , and that similar elements will be identified with similar numerals increased by a value of one hundred. 
     The partial-length rib  242  terminates at a rib end  244  spaced from the tip  198  of the airfoil  190 . The porous material  246  extends from the rib end  244  to a cooling surface  226  of the tip  198 . A turn  252  is formed through the porous material  246 . A portion of the cooling fluid  254  can pass through the porous material  246  in the turn  252  to pass from one flow channel  248  to the next. 
     It should be appreciated that the example illustrated in  FIG. 7  can provide for increased structural integrity of the airfoil  190  while permitting the cooling fluid C to pass within a cooling circuit  250  within the airfoil  190 . Additionally, it should be appreciated that the partial-length rib  242  having the porous material  246  connected to the tip  198  is effectively a full-length rib. As such, a porous material  246  formed in a full-length rib at the tip  198  can define the turn  252  for forming the cooling circuit  250 . 
     It should be appreciated that the porous portions  146 ,  246  described in  FIGS. 4-7  provide for reduced flow separation within cooling circuits, particularly in portions of the cooling circuit requiring drastic changes in flow direction such as a turn. The porous portions  146 ,  246  permit a volume of cooling air to pass through the partial-length ribs  142 ,  242  to reduce flow separation of the cooling fluid C passing through the turns within the cooling circuit. Additionally, the porous portions  146 ,  246  can be used to increase or maintain structural integrity of the airfoil  90 , without increasing system weight or sacrificing cooling efficiency. The porous material  146 ,  246  can be significantly lighter than the other portions or materials used in constructing the airfoil  90 . 
     A method of reducing flow separation within a cooling circuit within an airfoil for a turbine engine can include forming a portion of a partial-length rib with a porous material to permit a portion of a flow in the cooling circuit to pass through the partial-length rib. The cooling circuit can be the cooling circuit  150  formed within the airfoil  90 . The partial-length rib  142 ,  242  includes the porous portion  146 ,  246  to permit a portion of the cooling fluid flow  154  to pass through the partial-length rib  142 ,  242 . 
     In one example, the method can further include forming the end of the partial-length rib  142 , such as shown in  FIG. 4 , with the porous portion  146 . In another example, the porous portion  146  can be spaced from the end of the partial-length rib  142 , such as that shown in  FIGS. 5-6 . Additionally, the method can include metering the portion of cooling fluid  154 ,  254  passing through the porous portions  146 ,  246 . In non-limiting example, the metering can be accomplished by utilizing a structured porous material in the porous portions  146  or using framework  160 , such as shown in  FIG. 6 . 
     It should be appreciated that such a method can reduce flow separation within the cooling circuit  150 . Such flow separation is common at cooling circuit geometry such as turns, requiring a cooling fluid C to make a drastic turn, such as 180-degrees. Utilizing the porous material can permit a portion of the cooling fluid C to pass through the partial-length ribs  142 , minimizing the amount of fluid required to make the turn, and reducing the flow separation at the turn. The reduced flow separation can improve cooling circuit efficiency that requires less cooling flow, which can improve overall engine efficiency. 
     It should be appreciated that while embodiments are shown for blade internal ribs, such designs could also apply to endwall and shroud cooling circuits, or other component containing internal flow passages or turns, appreciating that the concepts as described herein can have equal applicability in additional engine components, such as a vane, shroud, or combustion liner in non-limiting examples, and can be any region of any engine component requiring cooling, such as regions typically requiring film cooling holes or multi-bore cooling. 
     It should be further appreciated that the region having the porous portion can provide for improved cooling, such as providing improved directionality, metering, or local flow rates. Additionally, the porous material include in the region can further improve the cooling to an entire region beyond just the areas local to the porous material. 
     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.