Patent Publication Number: US-11377963-B2

Title: Component for a turbine engine with a conduit

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is a continuation of U.S. patent application Ser. No. 16/120,758 filed Sep. 4, 2018, which is incorporated herein in its entirety. 
    
    
     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. 
     Engine efficiency increases with temperature of combustion gases. However, the combustion gases heat the various components along their flow path, which in turn requires cooling thereof to achieve a long engine lifetime. Typically, the hot gas path components are cooled by bleeding air from the compressor. This cooling process reduces engine efficiency, as the bled air is not used in the combustion process. 
     Turbine engine cooling art is mature and is applied to various aspects of cooling circuits and features in the various hot gas path components. For example, the combustor includes radially outer and inner liners, which require cooling during operation. Turbine nozzles include hollow vanes supported between outer and inner bands, which also require cooling. Turbine rotor blades are hollow and typically include cooling circuits therein, with the blades being surrounded by turbine shrouds, which also require cooling. The hot combustion gases are discharged through an exhaust which may also be lined, and suitably cooled. 
     In all of these exemplary turbine engine components, thin metal walls of high strength superalloy metals are typically used for enhanced durability while minimizing the need for cooling thereof. Various cooling circuits and features are tailored for these individual components in their corresponding environments in the engine. These components typically include common rows of film cooling holes. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect the disclosure relates to an airfoil an airfoil for a turbine engine which generates a hot gas flow and provides a cooling fluid flow, the airfoil comprising a platform having an outer surface, at least a portion of which is exposed to the hot gas flow to define a hot surface; a cooling cavity located within the platform extending between a base wall and an outer wall to define a radial direction, the cooling cavity fluidly coupled to the cooling fluid flow; and a conduit defining an interior cooling passage extending into the cooling cavity between an inlet fluidly coupled to a clean portion of the cooling fluid flow proximate the base wall and an outlet fluidly coupled to the hot surface. 
     In another aspect the disclosure relates to a component for a turbine engine which generates a hot gas flow and provides a cooling fluid flow, the component comprising a body having an outer surface, at least a portion of which is exposed to the hot gas flow to define a hot surface; a cooling cavity located within the body extending between a base wall and an outer wall to define a radial direction, the cooling cavity fluidly coupled to the cooling fluid flow; and a conduit extending into the cooling cavity between at least one inlet fluidly coupled to a clean portion of the cooling fluid flow proximate the base wall and at least one outlet fluidly coupled to the hot surface. 
     In yet another aspect, the disclosure relates to a method for cooling a component with a cooling cavity, the method comprising flowing a cooling fluid flow through a conduit extending between an inlet and an outlet of a hollow pin located within the cooling cavity; ducting a clean portion of the cooling fluid flow proximate an interior surface of the cooling cavity; and emitting the cooling fluid flow through the outlet onto a heated surface. 
    
    
     
       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 an isometric view of an airfoil for the turbine engine of  FIG. 1  in the form of a blade and having a platform with cooling holes. 
         FIG. 3  is an enlarged cross-sectional perspective view of a portion of the platform with the cooling holes from  FIG. 1  showing hollow pins within a cooling cavity according to an aspect of the disclosure. 
         FIG. 4  is the enlarged cross-sectional perspective view of  FIG. 3  illustrating the path of cooling fluid through the hollow pins. 
         FIG. 5  is a variation of the hollow pins from  FIG. 3  according to another aspect of the disclosure herein. 
     
    
    
     DESCRIPTION OF EMBODIMENTS OF THE INVENTION 
     Aspects of the disclosure described herein are directed to the formation of a hole such as a cooling hole in an engine component such as an airfoil. For purposes of illustration, the aspects of the disclosure discussed herein will be described with respect to the platform portion of a blade. It will be understood, however, that the disclosure as discussed herein 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. 
     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 relative to the engine centerline. 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. Furthermore, as used herein, the term “set” or a “set” of elements can be any number of elements, including only one. 
     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, 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. Furthermore it should be understood that the term cross section or cross-sectional as used herein is referring to a section taken orthogonal to the centerline and to the general coolant flow direction in the hole. 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. 
     Referring to  FIG. 1 , an 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 mount to a disk  61 , which mounts 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 mount 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 mount to a disk  71 , which is mounts 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 mount 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  splits 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  mixes with fuel in the combustor  30  where the fuel combusts, thereby generating combustion gases. The HP turbine  34  extracts some work from these gases, which drives the HP compressor  26 . The HP turbine  34  discharges the combustion gases 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  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 . 
       FIG. 2  is a perspective view of an example of an engine component illustrated as an airfoil  90 , a platform  92 , and a dovetail  94 . The airfoil  90  is shown as one of the rotating blades  68 , but can alternatively be a stationary vane, such as the vane  72  of  FIG. 1 , while any suitable engine component is contemplated. The airfoil  90  includes a tip  96  and a root  98 , defining a span-wise direction there between. Additionally, the airfoil  90  includes a wall  100 . A pressure side  104  and a suction side  106  are defined by the airfoil shape of the wall  100 . 
     The airfoil  90  mounts to the platform  92  at the root  98 . The platform  92  is shown in section, but can be formed as an annular band for mounting a plurality of airfoils  90 . The airfoil  90  can fasten to the platform  92 , such as welding or mechanical fastening, or can be integral with the platform  92  in non-limiting examples. According to an aspect of the disclosure herein, at least one cooling hole  102  is formed in an outer wall  101  of the platform  92 . The at least one cooling hole  102  can be multiple cooling holes  102  as illustrated, and, by way of non-limiting example, can be located in the platform  92  on the pressure side  104  of the airfoil  90 . The airfoil  90  further includes a leading edge  108  and a trailing edge  110 , defining a chord-wise direction. 
     The dovetail  94  couples to the platform  92  opposite of the airfoil  90 , and can be configured to mount to the disk  71 , or rotor  51  of the engine  10  ( FIG. 1 ), for example. In one alternative example, the platform  92  can be formed as part of the dovetail  94 . The dovetail  94  can include one or more inlet passages  112 , illustrated as three inlet passages  112 . It is contemplated that the inlet passages  112  are fluidly coupled to the cooling holes  102  to provide a cooling fluid flow (C) for cooling the platform  92 . In another non-limiting example, the inlet passages  112  can provide the cooling fluid flow (C) to an interior of the airfoil  90  for cooling of the airfoil  90 . It should be appreciated that the dovetail  94  is shown in cross-section, such that the inlet passages  112  are housed within the body of the dovetail  94 . 
     The platform  92  can define a body  114  having an outer surface  116  of the outer wall  101  exposed to a hot gas flow (H) to define a hot surface. A cooling cavity  118  can be located within the body  114  and be fluidly coupled to the cooling fluid flow (C) via, by way of non-limiting example some internal cooling passage or other cooling cavity not shown, such that the cooling fluid flow (C) flows within the cooling cavity  118 . At least one hollow pin  120  can extend into the cooling cavity  118 . The at least one hollow pin  120  can extend in a radial direction with respect to the engine centerline  12 . The hollow pin  120  can be any conduit extending into the cooling cavity  118  and including a cooling passage. 
       FIG. 3  is an enlarged portion III of the platform  92  illustrating the cooling cavity  118  in more detail. It can more clearly be seen that the hollow pin  120  defines at least a portion of the cooling hole  102 , specifically an interior cooling passage  122 , illustrated in dashed line, extending between an inlet  124  and an outlet  126 . While illustrated as an oval shape, the outlet  126  can be any suitable shape, including but not limited to racetrack, circular, rounded rectangular, or rounded triangular. The hollow pin  120  can further define a pin wall thickness (T) between 0.1 mm and 3 mm (0.005 to 0.1 inches), and preferably between 0.2 mm and 2 mm (0.01 to 0.05 inches). The thickness (T) is tailored to reduce weight while still enabling producibility and mechanical support. Furthermore, the thickness (T) enables convection cooling. 
     The inlet  124  can be provided on one side of the hollow pin  120 , by way of non-limiting example on the end  127  of the hollow pin  120  as illustrated. The inlet  124  can be formed at any location of the hollow pin  120  proximate the cooling fluid flow (C) present in the cooling cavity  118 . Proximate the cooling fluid flow (C) refers to locating the inlet  124  anywhere along the length of the hollow pin  120  such that the inlet  124  can receive cooling fluid flow (C). An interior surface  128  of the cooling cavity  118  is in contact with the cooling fluid flow (C) to define a cooled surface. The cooling cavity  118  forms a large internal convection area with the at least one hollow pin  120  forming a conduction path from the hot surface to a cooled surface within the cooling cavity  118 . 
     At least a portion of the outer wall  101  at least partially defines the interior surface  128  such that the outer wall  101  extends between the interior surface  128  and the outer surface  116 . A base wall  130  can further define the interior surface  128  and be radially spaced from the outer wall  101  a radial dimension (D) to further define the cooling cavity  118 . The hollow pin  120  can be formed to extend from and be attached to both the base wall  130  and the outer wall  101 . During operation centrifugal loads on the engine component cause dust to move away from the base wall  130  forming a clean region  132  of the cooling fluid flow (C) located along the interior surface  128  at the base wall  130 . It is contemplated that the hollow pin  120  extends from the outer wall  101  towards the base wall  130  such that the inlet  124  is located proximate the clean region  132  of cooling fluid flow (C). The hollow pin  120  can extend radially into the cooling cavity  118  a length (L) less than the radial dimension (D). It should be understood that while illustrated as attached to the interior surface  128  in one of the hollow pins  120  illustrated, the hollow pin  120  can be a partial pin as illustrated in the other of the hollow pins  120  extending partially into the cooling cavity  118 . In this case, the length (L) is less than the radial dimension (D) and spaced (S) from the interior surface  128  with no connection to the interior surface  128 . When described as being proximate the cooling fluid flow (C), the inlet  124  can be touching the interior surface  128 , or spaced from the interior surface (S). Dust accumulating away from the base wall  130  can leave a majority of the cooling cavity  118  free of dust and defining the clean region  132 . 
     A bend  134  can be formed in the hollow pin  120  to enable a positioning of the inlet  124  toward the cooling fluid flow (C). While illustrated as one bend  134 , it is contemplated that a plurality of bends can be formed in the hollow pin  120  at multiple locations to help orient the inlet toward the clean region  132 . A vector (V) extending perpendicularly from a plane formed by the inlet  124  can align with the interior surface  128  to tailor inlet effects of the cooling fluid flow (C). It is also contemplated that the angle and orientation of the hollow pin  120  do not necessitate a bend  134  formed in the hollow pin  120 . 
     Turning to  FIG. 4 , a method is illustrated for cooling the engine component using the cooling cavity  118  and hollow pin  120 . The method includes flowing cooling fluid flow (C) through the cooling cavity  118  to supply the cooling fluid flow (C) to the interior cooling passage  122  that extends between the inlet  124  and the outlet  126 . The method further includes emitting the cooling fluid flow (C) through the outlet  126  onto the heated surface, or outer surface  116 , by way of non-limiting example, the outer surface  116  of the platform  92 . 
     The method can include flowing the cooling fluid flow (C) from the cooling cavity  118  into the interior cooling passage via the inlet  124 . The location of the inlet  124  can enable ducting a clean portion (C 132 ) of the cooling fluid flow (C) to the outer surface  116  from the clean region  132  proximate the interior surface  128  of the cooling cavity  118 . The clean region  132  is located along the interior surface  128  radially inboard with respect to the cooling cavity  118 . 
       FIG. 5  illustrates a hollow pin  220  that can be formed in the component as described herein. The hollow pin  220  is similar to the hollow pin  120  therefore, like parts will be described with like numerals increased by  100 , with it being understood that the description of the like parts of the hollow pin  120  applies to the hollow pin  220 , unless otherwise noted. 
     The hollow pin  220  can extend through a cooling cavity  218  as illustrated. The hollow pin can define a cooling hole  202  having an interior cooling passage  222  terminating in an outlet  226 . In an aspect of the disclosure herein an inlet  224 , hidden by a body  214  of the component and illustrated in dashed line, as described previously can be located outside of the cooling cavity  218  and fluidly coupled to another source, by way of non-limiting example a cooling cavity located elsewhere and having a cooling fluid flow (C). The hollow pin  220  can have a substantially curved S-shape  236 . An S-shape  236  can enable both an optimum inlet  224  location with respect to a clean region  232  of the cooling fluid flow (C), including when the clean region  232  is located outside of the cooling cavity  218 . 
     It is contemplated that a first cross-sectional area (CA 1 ) of the hollow pin  220  can decrease to a smaller second cross-sectional area (CA 2 ) along a length (L) extending towards the outlet  226 . The decrease in cross-sectional area can be a continuously decreasing cross-sectional area. It is also contemplated that the first cross-sectional area (CA 1 ) can define a constant cross-sectional area for a portion of the length (L) of the hollow pin  220  and the second cross-sectional area (CA 2 ) can define a constant cross-sectional area for another portion of the length (L) of the hollow pin  220 . A decrease of any kind in cross-sectional area of the hollow pin  220  can coordinate with a change in cross-sectional area of the interior cooling passage  222  such that the cooling fluid (C) is accelerated through a narrower passage before being emitted onto an exterior surface  216  of a platform  292 . The cross-sectional area can be any shape, including but not limited to circular or racetrack. 
     In one exemplary aspect of the disclosure herein, the internal cooling passage  222  can further include a metering section  240  having a circular cross section, though it could have any cross-sectional shape. The metering section  240  can be provided where the first cross-sectional area (CA 1 ) decreases to the second cross-sectional area (CA 2 ). The metering section can extend along the interior cooling passage and maintain a constant cross-sectional area. The metering section  240  defines the smallest, or minimum cross-sectional area of the interior cooling passage  222 . It is also contemplated that the metering section  240  can have no length and is located at any portion of the interior cooling passage  222  where the cross-sectional area is the smallest. It is further contemplated that the metering section  240  can define the inlet  224  without extending into the interior cooling passage  222  at all. The interior cooling passage  222  can include multiple metering sections and is not limited to one as illustrated. The metering section  240  is for metering of the mass flow rate of the cooling fluid flow (C). 
     In another aspect of the disclosure herein, the interior cooling passage can define an increasing cross-sectional area (CA 3 ) where at least a portion of the increasing cross-sectional area (CA 3 ) defines a diffusing section  242  having a maximum cross-sectional area of the passage and terminating in the outlet  226 . In some implementations the increasing cross-sectional area (CA 3 ) is continuously increasing as illustrated. The diffusing section  242  enables an expansion of the cooling fluid (C) to form a wider and slower cooling film on the exterior  216  along the heated surface. The diffusing section  242  can be in serial flow communication with the metering section  240 . It is alternatively contemplated that the cooling hole  202  have a minimal or no metering section  240 , or that the diffusing section  242  extends along the entirety of the cooling hole  202 . The S-shape  232  provides geometry necessary for a longer diffusing section  242  at the outlet  226 . 
     The hollow pins as described herein can be formed using additive or advanced casting manufacturing technologies. By way of non-limiting example these technologies can include fused deposition modeling (FDM), VAT Photopolymerisation, Powder-bed fusion (PBF), material jetting, binder jetting, sheet lamination, or directed energy deposition (DED). 
     Radially extending hollow pins with embedded apertures in them enable specific durability and performance benefits for the platform as described herein. Optimal diffuser lengths are possible by utilizing the hollow pin for elongation of the diffusing portion of the cooling hole to provide higher film effectiveness. Additionally the presence of a hollow pin increases internal convection. Furthermore, sourcing low-dirt-count air mass from the bottom of the platform increases cooling effectiveness which increases hot gas path durability which results in reduced services costs &amp; better SFC. 
     Turbine cooling is important in next generation architecture which includes ever increasing temperatures. Current cooling technology needs to expand to the continued increase in core temperature of the engine that comes with more efficient engine design. Optimizing cooling at the surface of engine components by designing more effective cooling hole geometry and placement enable more efficient engine designs. 
     It should be understood that while the description herein is related to an airfoil platform, it can have equal applicability in other engine components requiring cooling via cooling holes such as film cooling. One or more of the engine components of the engine  10  includes a film-cooled substrate, or wall, in which a film cooling hole, or hole, of the disclosure further herein may be provided. Some non-limiting examples of the engine component having a wall can include blades, vanes or nozzles, a combustor deflector, combustor liner, or a shroud assembly. Other non-limiting examples where film cooling is used include turbine transition ducts and exhaust nozzles. 
     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 illustrate the disclosure as discussed herein, including the best mode, and also to enable any person skilled in the art to practice the disclosure as discussed herein, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure as discussed herein 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.