Patent Publication Number: US-2018051568-A1

Title: Engine component with porous holes

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, which generates a hot fluid flow and provides a cooling fluid flow. The component includes a wall separating the hot fluid flow from the cooling fluid flow and having a hot surface along with the hot fluid flow and a cooling surface facing the cooling fluid flow. The component further includes a cooling region defined in the hot surface. A plurality of holes extend between the cooling surface and the hot surface with at least some of the plurality of holes located within the cooling region. A first porous material fills at least some of the plurality of holes. 
     In another aspect, embodiments of the invention relate to an airfoil for a turbine engine including a perimeter wall bounding an interior and defining a pressure side and a suction side extending axially between a leading edge and a trailing edge, and extending between a root and a tip. The airfoil further includes a radially extending leading edge region disposed along the leading edge and at least partially extending between the root and the tip. A plurality of film holes are disposed in the leading edge region. A first porous material fills at least some of the film holes. 
     In yet another aspect, embodiments of the invention relate to a method of providing a cooling film along a leading edge region of an airfoil for a turbine engine. The method includes: (1) supplying cooling air to the interior of the airfoil; (2) exhausting at least a portion of the supplied cooling air through at least one film hole disposed in the leading edge region; and (3) exhausting the cooling air through the at least one film hole by passing the cooling air through a first porous material in the film hole. 
    
    
     
       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 engine component of the gas turbine engine of  FIG. 1  illustrated as an airfoil. 
         FIG. 3  is a cross-sectional view of the rotating blade of  FIG. 2  including a leading edge region. 
         FIG. 4  is perspective view of a portion of the leading edge region of  FIG. 3  including a plurality of film holes filled with porous material. 
         FIG. 5  is a view of the leading edge region of  FIG. 4  taken along section  5 - 5  illustrating an angled disposition of the film holes. 
         FIG. 6  is a cross-sectional view of an alternative rotating blade of  FIG. 2  having a porous leading edge region. 
         FIG. 7  is a perspective view of a portion of the porous leading edge region of  FIG. 6 . 
         FIG. 8  is a flow chart illustrating a method of providing a cooling film along the leading edge region of the airfoil. 
     
    
    
     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 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 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 . 
     A cooling region, shown as a leading edge region  108 , defines a portion of the engine component, which requires cooling. The leading edge region  108  can be defined extending along the leading edge  100 , extending at least partially between the root  96  and the tip  98 . A plurality of holes, such as film holes  110 , can be provided in the leading edge cooling region  108 . During operation of the gas turbine engine, a hot fluid flow H drives the blades to drive the compressor section of the engine. The combined core flow and exhaust momentum generate thrust. The hot fluid flow H is often of an excessive temperature to maximize engine thrust. A cooling fluid flow C is provided to the airfoil  90  for cooling. The cooling fluid flow C can be exhausted through the film holes  110  in the leading edge region  108  to cool the leading edge of the airfoil  90 . 
     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  121  along the exterior of the airfoil  90  and a cooling surface  123  confronting the cooling fluid flow C. An interior  126  of the airfoil  90  is defined by the outer wall  120 . One or more internal ribs  128  separates the interior  126  into passages  130  extending in the span-wise direction. The passages  130  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  130 , flow enhancers such as turbulators, or any other structures which can define the cooling circuits. 
     The cooling region or leading edge region  108  can be disposed at least partially within the pressure side  122  and the suction side  124 , and can be symmetric about the leading edge  100 . Alternatively, the cooling region or leading edge region  108  can be asymmetric about the leading edge  100 , having a larger portion on either the pressure or suction side  122 ,  124 , or having a unique shape. Additionally, it is contemplated that the cooling region or leading edge region  108  can be disposed entirely on the pressure side  122  or the suction side  124  terminating at or near the leading edge  100  that requires cooling such as a film cooling during engine operation. 
     The cooling region or leading edge region  108  can be a portion of the outer wall  120  requiring cooling at, adjacent to, or near the leading edge  100 . The cooling region can be any shape or size, having any geometry. The cooling region can extend at least partially in the span-wise direction between the root and the tip, and can extend fully between the root and the tip. The cooling region can extend along the outer wall  120  in the axial, or chord-wise, directions for any length such that cooling is needed such as film cooling in one example. 
     A porous material  132  can be provided in the film holes  110 . The porous material  132  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  132  can define a porosity, being permeable by a volume of fluid, such as air. The porous material  132  can have a particular porosity to meter the flow of a fluid passing through the porous material  132  at a predetermined rate. It should be appreciated that additive manufacturing can be used to achieve a particular local porosity along the porous material  132 , as well as a consistent porosity across the entirety of the porous material  132 , as compared to traditional method of forming the porous material  132 . In alternative examples, the porous material  132  can be made of any of the methods described above, such that a porosity is defined. In one non-limiting example, the porous material  132  can be made of Ni, NiCrAlY, NiAl, or similar materials. The porous material  132  can further be made of a nickel foam, for example. 
     Referring now to  FIG. 4 , a perspective view of a portion  140  of the leading edge region  108  includes a plurality of film holes  110  having the porous material  132  filling the film holes. The film holes  110  can be organized within the leading edge region  108 . Such an organization, for example, can be rows of film holes  110  extending in the span-wise direction. In other examples, the film holes  110  can be organized into patterns, groups, rows, columns, clusters, or can be based upon the particular needs of the airfoil  90 , such as areas requiring more or less cooling or are more or less susceptible to thermal aggregation. 
     Referring to  FIG. 5 , taken across section  5 - 5  of  FIG. 4 , the film holes  110  are disposed at an angle  142  relative to the surface of the outer wall  120 . The angle  142  is measured radially with respect to the engine centerline  12 , or in the span-wise direction relative to the airfoil  90 . The angle  142 , for example, can be between 15-degrees and 30-degrees, and can be 20-degrees in one non-limiting example. Alternatively, it is contemplated that the angle  142  can be between 1-degree and 45-degrees. It should be appreciated that the smaller value for the angle  142  can provide for improved surface cooling along the leading edge region  108 . Additionally, it should be understood that while the angles  142  are shown in the span-wise direction, they can be formed in any direction, such as span-wise, chord-wise, radial, axial, or any combination thereof in three-dimensional space. 
     Referring now to  FIG. 6 , a cross-section of an alternative airfoil  150  is shown having a cooling region  108  illustrated as a leading edge region  152  with a plurality of film holes  154 . A first porous material  156  fills the film holes  154 . A second porous material  158  forms the leading edge region  152 . 
     It should be understood that the first porous material  156  can fill some or all of the film holes  154  within the leading edge region  152 . Additionally, the second porous material  158  can form a portion of the leading edge region  152 , or the entirety of the leading edge region  152 . The first porous material  156  can have a greater porosity than the second porous material  158 . In one non-limiting example, the first porous material  156  can have a porosity up to one-hundred times the porosity of the second porous material  158 . The first and second porous materials  156 ,  158  can be formed similar to the porous material  132  as discussed regarding  FIG. 3 , such as by additive manufacturing, while it is further contemplated that additive manufacturing forms the entire airfoil  150  or engine component. Alternatively, it is contemplated that one of the first or second porous material  156 ,  158  is formed by additive manufacturing while the other is formed by other manufacturing methods, such as with a nickel foam. 
     Referring now to  FIG. 7 , a portion  160  of the leading edge region  152  is illustrated including a pattern of the film holes  154 . The pattern can be any organization of the film holes  154 , such as parallel rows or columns, groups, sets, or clusters in non-limiting examples. The film holes  154  can be disposed at the angle  142  to provide a cooling fluid at an angle along the leading edge region  152 . 
     It should be understood that the porous materials described herein, such as the porous material  132  of  FIGS. 3-5  or the porous materials  156 ,  158  of  FIGS. 6-7  can be a structured porous material or a random porous material, or any combination thereof. A structured porous material includes a structured, 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. In another example, a structured porous material can include a porous material having a non-random arrangement. 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 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. 
     Referring now to  FIG. 8 , a method  200  of providing a cooling film along a leading edge region of an airfoil for a turbine engine can include: at 202, (1) providing a flow of cooling fluid to the interior of the airfoil; and, at 204, (2) exhausting at least a portion of the cooling fluid through a first porous material in at least one film hole disposed in a cooling region. Alternatively, the method  200  can include, at  206 , exhausting the flow of cooling air through a leading edge region at the film hole. Additionally, the method can include, at  208 , exhausting a portion of the cooling fluid flow through the leading edge region having a second porous material. 
     At step  202 , a flow of cooling fluid C can be provided to the interior such as shown in  FIG. 2 , having the cooling flow provided through inlet passages in the dovetail. At step  204 , the cooling fluid is exhausted through a first porous material in at least one film hole in the cooling region. such as the leading edge regions  108 ,  152  of  FIGS. 2 and 6 , for example. As such, the component can be an airfoil, such as the airfoil described herein, with the cooling region being the leading edge region near or at the leading edge of the airfoil. At step  206 , the cooling air can be exhausted through the leading edge region at the film holes, through the first porous material in the film holes. As the flow of cooling fluid C passes through the porous material, the porosity or local porosities can particularly meter or direct the flow of cooling fluid C through the porous material. As such, the required flow of cooling fluid can be reduced to improve efficiency. Additionally, at step  208 , the leading edge region  152  can include the porous material, such as the second porous material  158 . A portion of the cooling fluid flow exhausts through the leading edge region  152 , such as through the second porous material  158 , as well as the first porous material in the film holes. The porosity of the leading edge region  152  can be less than that of the first porous material  132 ,  156  disposed in the film holes, permitting greater flow rates of the cooling fluid passing through the film holes. 
     It is contemplated that the porous material, the airfoils, or the other components described herein can be made with additive manufacturing. Additive manufacturing, such as 3D printing, can be used to form complex cooling circuit designs, having shaping or metering sections, complex circuits, holes, conduits, channels, or similar geometry, which is otherwise difficult to achieve with other manufacturing methods like drilling or casting. Additionally, the porous material can be formed with additive manufacturing. Typical methods for forming porous metals can result in uneven porosity among areas of the porous metals. Utilizing additive manufacturing can enable a manufacturer to achieve a uniform porosity along the entire porous structure. Alternatively, the manufacturer can achieve variable local porosities throughout the porous material as is desirable. Furthermore, such manufacturing can provide a more precisely made product, having a higher yield as compared to other manufacturing strategies. 
     It should be appreciated that the airfoil or engine component, utilizing porous material provides for even cooling distribution for a flow of cooling fluid. An additive manufacturing build of the regions could provide a precise distribution, particularly permitting an even porosity for the porous material(s). Additionally, the use of additive manufacturing can permit particular shaping or tailoring of the porous material or the airflow to control the flows throughout the airfoil. Utilizing such a porous material permits the flow of a fluid through the engine component, while retaining less heat to remain cooler. As such, the cooling, such as surface film cooling, provided through the walls of such engine components is enhanced. The enhanced cooling reduces the required flow of cooling fluid, such as up to 30-50% in one example. Such a reduction can increase engine efficiency. Furthermore, reduced blowing ratios can obtain better surface film cooling to increase component lifetime or reduce required maintenance. 
     It should be appreciated that while the description is directed toward a leading edge of the airfoil, 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 the leading edge region 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 holes as shown are non-limiting, and can be any shape, size, orientation, or include any geometry. 
     It should be further appreciated that the region and film holes having the porous material can provide for improved film cooling, such as providing improved directionality, metering, or local flow rates. Additionally, the porous material include in the region and the film holes can further improve the film cooling to an entire region beyond just the areas local to the film holes. 
     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.