Abstract:
A component for a gas turbine engine includes a wall that adjoins an interior cooling passage and provides an exterior surface. A film cooling hole fluidly connects the interior cooling passage and the exterior surface. The film cooling passage includes inlet and outlet passages that fluidly interconnect and adjoin one another in a misaligned non-line of sight relationship.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a gas turbine engine. In particular, the disclosure relates to cooling features and an example core manufacturing process that produces a core providing such features. 
         [0002]    Components, such as airfoils, particularly those used in a hot section of a gas turbine engine, incorporate internal cooling features. Current airfoil manufacturing techniques limit possible cooling configurations. Typically, the airfoil is cast within a mold having at least first and second portions secured to one another to define an exterior airfoil surface. The core structure used to form the impingement holes and cooling passages must be retained between the mold portions, which limit the location and configuration of the core, which is quite fragile. The core is typically assembled from multiple elements constructed from different material. The elements are glued to one another through a painstaking assembly process, which may result in scrapped cores. 
         [0003]    Film cooling holes are provided at various locations on the airfoil to supply cooling fluid to the exterior airfoil surface. One type of film cooling hole geometry extends from an inlet passage that communicates with a cooling passage to an outlet that terminates at the exterior airfoil surface. The outlet may be diffuser shaped, if desired. Using traditional film cooling hole forming methods, such as laser machining, electro discharge machining, or drilling, the inlet and outlet passages must be generally coaxially aligned with one another. 
       SUMMARY 
       [0004]    In one exemplary embodiment, a component for a gas turbine engine includes a wall that adjoins an interior cooling passage and provides an exterior surface. A film cooling hole fluidly connects the interior cooling passage and the exterior surface. The film cooling passage includes inlet and outlet passages that fluidly interconnect and adjoin one another in a misaligned non-line of sight relationship. 
         [0005]    In a further embodiment of the above, the inlet and outlet passages are generally linear. 
         [0006]    In a further embodiment of any of the above, the inlet and outlet passages are arranged at an angle relative to one another. 
         [0007]    In a further embodiment of any of the above, the angle is acute. 
         [0008]    In a further embodiment of any of the above, the outlet passage provides a diffuser shape. 
         [0009]    In a further embodiment of any of the above, the inlet passage provides a metering section that has a cross-sectional area that is less than a cross-sectional area of the outlet portion. 
         [0010]    In a further embodiment of any of the above, the inlet passage includes first and second metering portions. The second metering portion adjoins the outlet passage and includes a length L and a diameter D having an L/D ratio of greater than 1. 
         [0011]    In a further embodiment of any of the above, the L/D ratio is greater than 3. 
         [0012]    In a further embodiment of any of the above, the film cooling hole is additively manufactured. 
         [0013]    In a further embodiment of any of the above, the component is one of an airfoil combustor BOAS and platform. 
         [0014]    In another exemplary embodiment, a method of manufacturing airfoil component for a gas turbine engine includes depositing multiple layers of a powdered metal onto one another. The layers are joined to one another with reference to CAD data relating to a particular cross-section. A film cooling hole geometry is produced corresponding to inlet and outlet passages fluidly interconnecting and adjoining one another in a misaligned, non-line of sight relationship. 
         [0015]    In a further embodiment of any of the above, the producing step includes manufacturing a core that provides the film cooling hole geometry. 
         [0016]    In a further embodiment of any of the above, the core includes molybdenum. 
         [0017]    In a further embodiment of any of the above, the producing step includes forming the film cooling hole in response to the joining step producing a wall. 
         [0018]    In a further embodiment of any of the above, the inlet and outlet passages are generally linear. 
         [0019]    In a further embodiment of any of the above, the inlet and outlet passages are arranged at an angle relative to one another. 
         [0020]    In a further embodiment of any of the above, the angle is acute. 
         [0021]    In a further embodiment of any of the above, the outlet passage provides a diffuser shape. 
         [0022]    In a further embodiment of any of the above, the inlet passage is provided during the depositing step. The outlet passage is provided by machining. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0023]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0024]      FIG. 1  is a schematic view of a gas turbine engine incorporating the disclosed airfoil. 
           [0025]      FIG. 2A  is a perspective view of the airfoil having the disclosed cooling passage. 
           [0026]      FIG. 2B  is a plan view of the airfoil illustrating directional references. 
           [0027]      FIG. 3A  is a partial cross-sectional view of an example core. 
           [0028]      FIG. 3B  is a perspective view of a portion of an airfoil with the core of  FIG. 3A . 
           [0029]      FIG. 3C  is a perspective view of the airfoil of  FIG. 3B  without the core. 
           [0030]      FIG. 3D  is a cross-sectional view of the airfoil of  FIG. 3C . 
           [0031]      FIG. 4  is a perspective view of another example core. 
           [0032]      FIG. 5A  is a cross-sectional view of a portion of another example core and airfoil. 
           [0033]      FIG. 5B  is a cross-sectional view of the airfoil shown in  FIG. 5A  without the core. 
           [0034]      FIG. 6A  is a cross-sectional view of another example core and airfoil. 
           [0035]      FIG. 6B  is a perspective view of a portion of the core shown in  FIG. 6A . 
           [0036]      FIG. 7  is a cross-sectional view of another example core and airfoil. 
           [0037]      FIG. 8  is a cross-sectional view of another example core and airfoil. 
           [0038]      FIG. 9  is a flow chart depicting an example refractory metal core manufacturing process. 
           [0039]      FIG. 10  is a flow chart depicting an example airfoil manufacturing process. 
           [0040]      FIG. 11  illustrates a film cooling hole and passage of  FIG. 7  in more detail. 
           [0041]      FIG. 12  is an elevational view of the film cooling hole shown in  FIG. 11 . 
           [0042]      FIG. 13  illustrates another film cooling hole and passage. 
           [0043]      FIG. 14  illustrates still another film cooling hole and passage. 
       
    
    
     DETAILED DESCRIPTION 
       [0044]      FIG. 1  schematically illustrates a gas turbine engine  10  that includes a fan  14 , a compressor section  16 , a combustion section  18  and a turbine section  11 , which are disposed about a central axis A. As known in the art, air compressed in the compressor section  16  is mixed with fuel that is burned in combustion section  18  and expanded in the turbine section  11 . The turbine section  11  includes, for example, rotors  13  and  15  that, in response to expansion of the burned fuel, rotate, which drives the compressor section  16  and fan  14 . 
         [0045]    The turbine section  11  includes alternating rows of blades  20  and static airfoils or vanes  19 . It should be understood that  FIG. 1  is for illustrative purposes only and is in no way intended as a limitation on this disclosure or its application. 
         [0046]    An example blade  20  is shown in  FIG. 2A . The blade  20  includes a platform  24  supported by a root  22 , which is secured to a rotor, for example. An airfoil  26  extends radially outwardly from the platform  24  opposite the root  22  to a tip  28 . While the airfoil  26  is disclosed as being part of a turbine blade  20 , it should be understood that the disclosed airfoil can also be used as a vane. 
         [0047]    Referring to  FIG. 2B , the airfoil  26  includes an exterior airfoil surface  38  extending in a chord-wise direction C from a leading edge  30  to a trailing edge  32 . The airfoil  26  extends between pressure and suction sides  34 ,  36  in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple airfoils  26  are arranged circumferentially in a circumferential direction H. The airfoil  26  extends from the platform  24  in a radial direction R to the tip  28 . The exterior airfoil surface  38  may include multiple film cooling holes. The disclosed film cooling holes may also be provided in combustion liners, blade outer air seals and platforms, for example. 
         [0048]    An example core  40  and resultant airfoil  26  is shown in  FIGS. 3A-3D . As described in relation to  FIG. 10 , the airfoils disclosed may also be manufactured directly without the need of a core. In direct airfoil manufacturing, the airfoil features described as being provided by a core can be provided in the airfoil during the airfoil forming process. 
         [0049]    The core  40  is provided by a refractory metal structure, constructed from molybdenum, for example, having a variable thickness. The core  40  is defined by an exterior surface  60  providing a contour. The exterior surface  60  including a uniform surface finish from the core manufacturing process, described in connection with  FIG. 9  below, which results in a surface roughness to which suitable coatings will adhere. The exterior surface  60  is without machining, that is, milling, chemical etch, filing, or sanding. However, the exterior surface  60  may be finished in a slurry. As a result, coatings will adhere to the edges of the core  40 . 
         [0050]    In the example, first and second thicknesses  42 ,  44  are different than one another. In one example, the first thickness  42  is less than 0.060 inch (1.52 mm). The core  40  may include an aperture  46  with a radiused edge  48  providing the second thickness  44  of less 0.020 inch (0.51 mm), for example. The cast airfoil  26  provides a cooling passage  53  and standoff  50  corresponding to the aperture  46 . The standoff  50  illustrated in  FIG. 3D  is shown in the area indicated in  FIG. 6A . 
         [0051]    Referring to  FIG. 4 , the core  140  includes an exterior surface  160  having a perimeter  58 . A recess  54  may be arranged inboard of the perimeter  58 , for example. In another example, a protrusion  56  is arranged inboard of the perimeter  58 . The recess  54  and protrusion  56  are not machined. The thickness of the core  140  surrounding the recess  54  and protrusion  56  is less than 0.060 inch (1.52 mm) in one example, and less than 0.020 inch (0.51 mm) in another example. 
         [0052]    In another example shown in  FIGS. 5A-5B , the core  240  may be used to produce an airfoil  226  including a cooling passage  253  having a tapered wall. The core  240  is tapered between first and second portions  62 ,  64  with the second portion  64  having a second thickness  244  less than 0.020 inch (0.51 mm). The first portion  62  has a first thickness  242  that is greater than 0.020 inch (0.51 mm). 
         [0053]    Referring to  FIGS. 6A-6B , which depicts a core within an airfoil, the exterior airfoil surface  360  is defined by a perimeter wall  76 . First, second and third cooling passages  66 ,  68 ,  70  are provided within the airfoil  326 , for example. An interior wall  78  is arranged interiorly and adjacent to the perimeter wall  76  to provide the second cooling passage  67 , for example. A cooling passage, for example, first cooling passage  66  is tapered and respectively has different thickness, for example, as described above with respect to  FIGS. 5A and 5B . At least one of the passages, for example, second passage  67 , may include a thickness less than 0.060 inch (1.52 mm). The cooling passages are formed by correspondingly shaped core structure. The core  340  is provided by a unitary structure having uniform material properties, for example. That is, multiple core elements constructed from different core materials glued to one another need not be used. 
         [0054]    The core  340  may include first and second portions  77 ,  79  overlapping one another. The first and second portions  77 ,  79  are less than 0.060 inch (1.52 mm) thick, in one example, and of varying thickness. At least one of the first and second portions  77 ,  79  may provide a film cooling hole  74  in the exterior airfoil surface  360 . The first and second portions  77 ,  79  may be joined to one another by a standoff  72  that produces a hole interconnecting the resultant overlapping cooling passages. Standoffs  72  can be used to integrally connect and join all passages  66 ,  68 ,  70  to eliminate the need for core assembly and better stabilize the core during casting. However, directly manufacturing the airfoil, as shown in  FIG. 10 , would not require these features. 
         [0055]    Similarly, a unitary body having uniform material properties throughout the structure provides the cores  440 ,  540  shown in  FIGS. 7 and 8 . The core  440 ,  540  have at least one portion with a thickness of less than 0.060 inch (1.52 mm). Referring to the airfoil  426  of  FIG. 7 , the perimeter wall  476  defines first and second cooling passages  477 ,  479  with the interior wall  478 . The first and second cooling passages  477 ,  479  are arranged in a switch back configuration, and one of the passages may provide a film cooling hole  474 . The standoff  472  interconnects a central wall  473 , which splits the channel to increase hot wall contact and maintain flow speed and pressure. 
         [0056]    Referring to the airfoil  526  of  FIG. 8 , the first cooling passage  566  is arranged adjacent to the perimeter wall  576  at the leading edge  530 . The third cooling passage  570  is arranged between the perimeter wall  576  and the interior wall  578  to provide a microcircuit of less than 0.060 inches (1.52 mm) thickness, interconnecting the first and second cooling passages  566 ,  568  to one another. 
         [0057]    The core geometries and associated airfoil cooling passages disclosed in  FIGS. 3A-8  may be difficult to form using conventional casting technologies. Thus, an additive manufacturing process  80  may be used, as schematically illustrated in  FIG. 9 . Powdered metal  82  suitable for refractory metal core applications, such as molybdenum, is fed to a machine  84 , which may provide a vacuum, for example. The machine  84  deposits multiple layers of powdered metal onto one another. The layers are joined to one another with reference to CAD data  86 , which relates to a particular cross-section of the core  40 . In one example, the powdered metal  82  may be melted using a direct metal laser sintering process or an electron-beam melting process. With the layers built upon one another and joined to one another cross-section by cross-section, a core with the above-described geometries may be produced, as indicated at  88 . A single piece core can be produced that requires no assembly and can be directly placed into a mold after being coated with wax. 
         [0058]    The coating  90  may be applied to the exterior surface of the core  40 , which enables the core  40  to be more easily removed subsequently. The core  40  is arranged in a multi-piece mold and held in a desired orientation by features on the mold, as indicated at  92 . The core  40  is more robust and can better withstand handling as it is positioned within the mold. The airfoil  26  is cast about the core  40 , as indicated at  94 . The core  40  is then removed from the airfoil  26 , as indicated at  96 , to provide desired cooling passage features. 
         [0059]    An additive manufacturing process  180  may be used to produce an airfoil, as schematically illustrated in  FIG. 10 . Powdered metal  182  suitable for aerospace airfoil applications is fed to a machine  184 , which may provide a vacuum, for example. The machine  184  deposits multiple layers of powdered metal onto one another. The layers are joined to one another with reference to CAD data  186 , which relates to a particular cross-section of the airfoil  20 . In one example, the powdered metal  182  may be melted using a direct metal laser sintering process or an electron-beam melting process. With the layers built upon one another and joined to one another cross-section by cross-section, an airfoil with the above-described geometries may be produced, as indicated at  194 . The airfoil may be post-processed  196  to provide desired structural characteristics. For example, the airfoil may be heated to reconfigure the joined layers into a single crystalline structure. 
         [0060]    The film cooling hole  474  of  FIG. 7  is shown in more detail in  FIGS. 11 ,  12 ,  13  and  14 . The first and second cooling passages  477 ,  479 , or outlet and inlet passages, in the wall  476  of airfoil  426  respectively correspond to outlet and inlet passages. The outlet passage  477  provides a fluid exit in the exterior airfoil surface  460 . The outlet passage  477  provides a diffuser portion that has a diffuser shape such that its cross-sectional area increases for at least a portion of the length of the outlet passage  477 , decreasing the velocity of the fluid. The inlet passage  479  provides a metering portion that has a cross-sectional area that is less than the cross-sectional area of the outlet passage  477  and which is designed to meter the flow of the fluid from the interior cooling passage  466  through the film cooling hole  474 . 
         [0061]    Using the additive manufacturing techniques described in this disclosure, the outlet and inlet passages  477 ,  479  can be substantially misaligned with one another, i.e., non-line of sight with respect to one another. In the example shown, the outlet and inlet passages  477 ,  479  are each generally linear and arranged at an acute angle with respect to one another in the wall thickness direction T ( FIG. 11 ). The outlet and inlet passages  477 ,  479  may also be angled relative to one another in another direction, such as the radial direction R ( FIG. 12 ). Such a switch back configuration can be packaged more easily in some wall geometries and may reduce cooling flow blow off. 
         [0062]    The film cooling hole  474  may be formed during an additive manufacturing process of the airfoil  426 . Alternatively, a core  440  ( FIG. 11 ) may be constructed using additive manufacturing and which is used to provide the correspondingly shaped film cooling hole  474  when casting the airfoil  426 . Thus, the film cooling hole  474  is not machined. Cores  540 ,  640  may also be used to form the cooling holes  574 ,  674  described in  FIGS. 13 and 14  below. 
         [0063]    A hybrid manufacturing method may also be used. That is, the diffusion portion and an aligned metering section ( FIGS. 13 and 14  below), if any, can be electro discharge machined, while the non-aligned metering section can be additively manufactured by either additively making the core, or by additively making the airfoil directly. 
         [0064]    Another film cooling hole  574  is shown in  FIG. 13 . The first and second cooling passages in the wall  576  of airfoil  526  respectively correspond to the outlet passage  577  and inlet passage having metering portions  579   a,    579   b.  The outlet passage  577  provides a fluid exit in the exterior airfoil surface  560 . The outlet passage  577  has a diffuser shape such that its cross-sectional area increases for at least a portion of the length of the outlet passage  577 , decreasing the velocity of the fluid. 
         [0065]    The inlet passage is provided by multiple metering portions  579   a,    579   b,  which have a cross-sectional area that is less than the cross-sectional area of the outlet passage  577  and which is designed to meter the flow of the fluid from the interior cooling passage  566  through the film cooling hole  574 . In the example shown in  FIG. 13 , the metering portion  579   a  is straight. The metering portion  579   b  has a diameter D (hydraulic diameter D if the cross-sectional area of the metering section  579   b  is not circular) and a length L that provides an L/D ratio of greater than  1 , and in one example, greater than  3 . 
         [0066]    Another example film cooling hole  674  is shown in  FIG. 14 . The first and second cooling passages in the wall  676  of airfoil  626  respectively correspond to the outlet passage  677  and inlet passage having metering portions  679   a,    679   b.  The outlet passage  677  provides a fluid exit in the exterior airfoil surface  660 . The outlet passage  677  has a diffuser shape such that its cross-sectional area increases for at least a portion of the length of the outlet passage  677 , decreasing the velocity of the fluid. 
         [0067]    The inlet passage is provided by multiple metering portions  679   a,    679   b,  which have a cross-sectional area that is less than the cross-sectional area of the outlet passage  677  and which is designed to meter the flow of the fluid from the interior cooling passage  666  through the film cooling hole  674 . In the example shown in  FIG. 14 , the metering portion  679   a  is non-linear. The metering portion  679   b  has a diameter D (hydraulic diameter D if the cross-sectional area of the meetering section  579   b  is not circular) and a length L that provides an L/D ratio of greater than  1 , and in one example, greater than  3 . 
         [0068]    The outlet passage  477 ,  577 ,  677  provides the diffuser, which could be compound diffusion shaped, multi-lobe shaped, chevron-shaped, for example. The cross-section of the inlet passage  479 ,  579   a/   579   b,    679   a/   679   b,  which provides the metering portion, could be circular, oblong-shaped, crescent, or cusp-shaped, for example. 
         [0069]    Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For example, the disclosed cores, cooling passages and cooling holes may be used for applications other than airfoils, such as combustor liners, blade outer air seals (BOAS) and platforms. For that reason, the following claims should be studied to determine their true scope and content.