Abstract:
A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a first sidewall and a second sidewall joined together at a leading edge and a trailing edge and extending from a base to a tip. A plenum is defined inside the airfoil. A first cooling cavity merges into the plenum and a second cooling cavity merges into the plenum. A rib extends from at least one of the first sidewall and the second sidewall at least partially into the plenum to separate the first cooling cavity from the second cooling cavity.

Description:
[0001]    This invention was made with government support under Contract No. N00019-12-D-0002 awarded by the United States Navy. The Government therefore has certain rights in this invention. 
     
    
     BACKGROUND 
       [0002]    This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component and a core assembly for defining internal cooling features within a completed component. 
         [0003]    Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
         [0004]    Due to exposure to hot combustion gases, some gas turbine engine components include an internal cooling scheme that routes cooling air through the part. For example, the internal cooling scheme may define multiple hollow passages through which the cooling air may be circulated. Thermal energy is transferred from the component to the airflow as the cooling air passes through the cooling scheme to cool the component. 
         [0005]    Some components, such as airfoils, are typically molded parts. The internal cooling passages required to communicate cooling air through the part are typically formed using core assemblies that are over-molded during a casting or other molding process to define the hollow passages inside the component. 
       SUMMARY 
       [0006]    A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a first sidewall and a second sidewall joined together at a leading edge and a trailing edge and extending from a base to a tip. A plenum is defined inside the airfoil. A first cooling cavity merges into the plenum and a second cooling cavity merges into the plenum. A rib extends from at least one of the first sidewall and the second sidewall at least partially into the plenum to separate the first cooling cavity from the second cooling cavity. 
         [0007]    In a further non-limiting embodiment of the foregoing component, the component is a blade. 
         [0008]    In a further non-limiting embodiment of either of the foregoing components, the component is a vane. 
         [0009]    In a further non-limiting embodiment of any of the foregoing components, the rib extends across the plenum. 
         [0010]    In a further non-limiting embodiment of any of the foregoing components, the rib extends between opposing sides of the plenum. 
         [0011]    In a further non-limiting embodiment of any of the foregoing components, the rib extends transversely between the opposing sides. 
         [0012]    In a further non-limiting embodiment of any of the foregoing components, the rib defines a surface that constricts a flow of cooling fluid that exits at least one of the first cooling cavity and the second cooling cavity into the plenum. 
         [0013]    In a further non-limiting embodiment of any of the foregoing components, a portion of the rib extends from at least one of the first sidewall and second sidewall and terminates prior to the other of the first sidewall and the second sidewall. 
         [0014]    In a further non-limiting embodiment of any of the foregoing components, a third cooling cavity merges into the plenum. 
         [0015]    In a further non-limiting embodiment of any of the foregoing components, the third cooling cavity exits adjacent to another rib. 
         [0016]    In a further non-limiting embodiment of any of the foregoing components, the plenum is positioned in the tip. 
         [0017]    In a further non-limiting embodiment of any of the foregoing components, the rib includes a face that is offset from an outlet of at least one of the first cooling cavity and the second cooling cavity. 
         [0018]    A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a component disposed in at least one of a compressor section and a turbine section. The component includes a body that includes a first sidewall and a second sidewall joined together at a leading edge and a trailing edge and extending from a base to a tip. The component has an internal cooling scheme comprising a cooling cavity that merges into a plenum and a rib that extends from at least one of the first sidewall and the second sidewall into the plenum to control expansion of a cooling fluid from the cooling cavity into the plenum. 
         [0019]    A core assembly for fabricating a component according to an exemplary aspect of the present disclosure includes, among other things, a first core that defines a first cooling cavity of a component and a second core that defines a second cooling cavity of the component. A partial rib extends from the first core and the second core and defines a plenum of the component. At least one cut-out in the partial rib defines a rib in the component. 
         [0020]    In a further non-limiting embodiment of the foregoing core assembly, the component is a completed component. 
         [0021]    In a further non-limiting embodiment of either of the foregoing core assemblies, a third core defines a third cooling cavity of the component. 
         [0022]    In a further non-limiting embodiment of any of the foregoing core assemblies, the partial rib extends from the third core. 
         [0023]    In a further non-limiting embodiment of any of the foregoing core assemblies, at least one cut-out includes a plurality of cut-outs disposed on opposing faces of the partial rib. 
         [0024]    The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0025]      FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
           [0026]      FIG. 2  illustrates a gas turbine engine component. 
           [0027]      FIG. 3  illustrates a gas turbine engine component prior to removal of a core assembly. 
           [0028]      FIG. 4  illustrates a partial cut-away view of a gas turbine engine component. 
           [0029]      FIG. 5  illustrates a side view of a gas turbine engine component. 
           [0030]      FIG. 6  illustrates a top view of a gas turbine engine component. 
           [0031]      FIG. 7  illustrates an exemplary core assembly. 
       
    
    
     DETAILED DESCRIPTION 
       [0032]    This disclosure is directed to a gas turbine engine component and a core assembly for defining internal cooling features within a completed component. Among other features, an exemplary cooling scheme may include partial ribs that both structurally transition between a cavity and a plenum and improve effective heat transfer between a cooling fluid and the component. For example, the ribs described herein increase the amount of surface area available for exchanging heat and conduct heat away from the walls of the component. Flow of the cooling fluid is constrained in certain directions by reducing area, thereby maintaining heat transfer coefficients relatively high as the cooling fluid enters the plenum of the component. 
         [0033]      FIG. 1  schematically illustrates a gas turbine engine  20 . The exemplary gas turbine engine  20  is a two-spool turbofan engine that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems for features. The fan section  22  drives air along a bypass flow path B, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 . The hot combustion gases generated in the combustor section  26  are expanded through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines, including but not limited to, three-spool engine architectures. 
         [0034]    The gas turbine engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine centerline longitudinal axis A. The low speed spool  30  and the high speed spool  32  may be mounted relative to an engine static structure  33  via several bearing systems  31 . It should be understood that other bearing systems  31  may alternatively or additionally be provided. 
         [0035]    The low speed spool  30  generally includes an inner shaft  34  that interconnects a fan  36 , a low pressure compressor  38  and a low pressure turbine  39 . The inner shaft  34  can be connected to the fan  36  through a geared architecture  45  to drive the fan  36  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  35  that interconnects a high pressure compressor  37  and a high pressure turbine  40 . In this embodiment, the inner shaft  34  and the outer shaft  35  are supported at various axial locations by bearing systems  31  positioned within the engine static structure  33 . 
         [0036]    A combustor  42  is arranged between the high pressure compressor  37  and the high pressure turbine  40 . A mid-turbine frame  44  may be arranged generally between the high pressure turbine  40  and the low pressure turbine  39 . The mid-turbine frame  44  can support one or more bearing systems  31  of the turbine section  28 . The mid-turbine frame  44  may include one or more airfoils  46  that extend within the core flow path C. 
         [0037]    The inner shaft  34  and the outer shaft  35  are concentric and rotate via the bearing systems  31  about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor  38  and the high pressure compressor  37 , is mixed with fuel and burned in the combustor  42 , and is then expanded over the high pressure turbine  40  and the low pressure turbine  39 . The high pressure turbine  40  and the low pressure turbine  39  rotationally drive the respective high speed spool  32  and the low speed spool  30  in response to the expansion. 
         [0038]    Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades  25 , while each vane assembly can carry a plurality of vanes  27  that extend into the core flow path C. The blades  25  create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine  20  along the core flow path C. The vanes  27  direct the core airflow to the blades  25  to either add or extract energy. 
         [0039]    Various components of the gas turbine engine  20 , including but not limited to the airfoils of the blades  25  and the vanes  27  of the compressor section  24  and the turbine section  28 , may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section  28  is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling schemes for cooling the parts during engine operation. 
         [0040]      FIG. 2  illustrates a component  50  that may be employed by a gas turbine engine, such as the gas turbine engine  20  of  FIG. 1 . The component  50  can be manufactured in a casting or molding process. Exemplary casting processes include investment casting, die casting, other molding processes or additive manufacturing processes. In one embodiment, the component  50  is a turbine blade. However, the features of this disclosure are applicable to any cast part of a gas turbine engine, or any other part, including compressor parts. Rotating structures, vanes or other components may all benefit from the teachings of this disclosure. 
         [0041]    The component  50  includes an airfoil  52  (or other body portion) that axially extends between a leading edge  54  and a trailing edge  56 . The airfoil  52  may additionally include a pressure sidewall  58  (i.e., a first sidewall) and a suction sidewall  60  (i.e., a second sidewall) that are spaced apart from one another and that join together at each of the leading edge  54  and the trailing edge  56 . The component  50  in one embodiment additionally includes a platform  51  and a root  53 . The airfoil  52  extends outwardly from the platform  51  and the root  53  extends outwardly in an opposed direction from the platform  51 . The airfoil  52  extends from a base  62  adjacent to the platform  51  to a tip  64 . 
         [0042]    The component  50  may include an internal cooling scheme  55  for cooling the component  50 . The internal cooling scheme  55  includes one or more cooling cavities  72  (in this embodiment, two cooling cavities are shown in phantom). It should be appreciated that the component  50  could include additional cooling cavities or only a single cooling cavity. The cooling cavities  72  may be in fluid communication with one another, such as along a serpentine path, or could alternatively be fluidly isolated from one another. 
         [0043]    The cooling cavities  72  extend radially, axially and/or circumferentially inside of the airfoil  52  or other sections of the component  50  and establish hollow passages for receiving and circulating a cooling fluid  68 , such as relatively cool air from the compressor section  24 , to cool the component  50 . Although not shown by  FIG. 2 , the internal cooling scheme  55  may feed various cooling holes disposed through the component  50 , including through the airfoil  52 , the platform  51 , and/or the root  53 , to provide a layer of film cooling air at surfaces of the component  50 . 
         [0044]      FIG. 3  illustrates the component  50  (shown in phantom) of  FIG. 2  prior to removal of a core assembly  80  (shown in solid) that can be used during a casting or other molding process to define at least a portion of the internal cooling scheme  55  (best illustrated in  FIGS. 2 and 4 ). For example, the component  50  may be cast to include hollow portions that make up the internal cooling scheme  55  that extends inside of the component  50 . 
         [0045]    The component  50  may be manufactured in a casting process. One exemplary casting process includes the initial step of fabricating the core assembly  80  to include features that define the internal cooling cavities  72  (see  FIGS. 2 and 4 ) and various other passages and features of the internal cooling scheme  55 . The core assembly  80  is inserted into a mold or other molding fixture and surrounded by a molten material. The molten material cures and hardens about the core assembly  80  to define the internal and external surfaces of the component  50 . Once the molten metal has cured, the core assembly  80  may be removed through known methods, such as leeching, to form a completed component. 
         [0046]    In one embodiment, the core assembly  80  is a ceramic core. In another embodiment, the core assembly  80  is a refractory metal core (RMC). In another embodiment, the core assembly  80  is a hybrid core (for example, a hybrid of a ceramic core and a RMC core). Other materials are also within the scope of this disclosure.  FIG. 3  illustrates but one embodiment of a core assembly  80  that can be used to cast internal features into the component  50 . Additional features of the core assembly  80  are described below and are illustrated by  FIG. 7 . Although not shown, various other core assemblies may be utilized to create additional internal features of the component  50 . In addition, the core assembly  80  may be made using any known process, including but not limited to an additive manufacturing process. 
         [0047]      FIG. 4  illustrates portions of a completed component  50 A. That is, the completed component  50 A is illustrated post-cast and post-removal of any core assemblies used to create the internal cooling scheme  55  of the completed component  50 A. In one embodiment, the internal cooling scheme  55  of the completed component  50 A includes a first cooling cavity  72 A, a second cooling cavity  72 B and a plenum  82  that extend inside of a body  75  of the completed component  50 A. The body  75  may be representative of an airfoil, platform, root or any other section of the completed component  50 A. The first cooling cavity  72 A and the second cooling cavity  72 B extend to the plenum  82  to feed the plenum with a cooling fluid  68 . 
         [0048]    In one embodiment, the plenum  82  is positioned in the tip  64  of the body  75 ; however, the plenum  82  could be located elsewhere. In addition, the cooling cavities  72 A,  72 B and the plenum  82  are not limited to the configuration shown in which the cooling cavities  72 A,  72 B radially feed the axially disposed plenum  82 . For example, the cooling cavities  72 A,  72 B could axially feed a radially disposed plenum  82  within the scope of this disclosure. 
         [0049]    A rib  84  separates the first cooling cavity  72 A from the second cooling cavity  72 B. In one embodiment, one or more ribs  84  extend into the plenum  82 . In this way, the rib(s)  84  structurally supports the transition area between the cooling cavities  72 A,  72 B and the plenum  82 . 
         [0050]    In one embodiment, the rib  84  extends from the suction sidewall  60  into the plenum  82 . However, the rib  84  could also extend from the pressure sidewall  58  within the scope of this disclosure. The rib  84  may be as thick as the sidewall  58 ,  60 , or could extend further into part of the plenum  82 . In one embodiment, the rib  84  terminates prior to the opposite sidewall  58 ,  60  (see rib  184  of  FIG. 6 , for example). 
         [0051]    As best illustrated by  FIG. 5 , the rib  84  of a completed component  50 A extends between opposing sides  85 ,  87  of the plenum  82  and holds the plenum  82  together near a transition area TA between a cooling cavity  72  and the plenum  82 . In one embodiment, the rib  84  extends transversely between the opposing sides  85 ,  87  of the plenum  82 . The rib  84  may reduce deflections or bulging of the plenum  82  that can occur during engine operation and may reduce the likelihood of the plenum  82  liberating from the component  50 . 
         [0052]      FIG. 6  illustrate an internal cooling scheme  155  of a completed component  50 A. The internal cooling scheme  155  includes a first cooling cavity  172 A, a second cooling cavity  172 B and a third cooling cavity  172 C that merge into a plenum  182 . Ribs  184  are positioned between the cooling cavities  172 A,  172 B and  172 C and at least partially extend into the plenum  182 . The number and configuration of cooling cavities and ribs is exemplary only and is not intended to limit this disclosure. 
         [0053]    The ribs  184  can control the expansion of cooling fluid  68  into the plenum  182 . For example, cooling fluid  68  that exits the cooling cavities  172 A,  172 B and  172 C is forced to travel along a surface  99  of the ribs  184  prior to circulating through the plenum  182 . Put another way, the ribs  184  increase the amount of area available to perform heat transfer and conduct heat away from the suction sidewall  160  (or pressure sidewall  158 ) of the completed component  50 A by constricting the flow of the cooling fluid  68  in a specific direction (here, a direction toward a center  101  of the completed component  50 A). 
         [0054]    The ribs  184  may also act as augmentation features that convect heat away from hot surfaces of the suction sidewall  160  and/or the pressure sidewall  158 . In one embodiment, the ribs  184  include a face  103  that extends past, or is circumferentially offset from, an outlet  105  of each cooling cavity  172 A,  172 B and  172 C. The surfaces  99  of the rib  184  force the cooling fluid  68  to flow across the face  103 , thereby increasing heat transfer. 
         [0055]      FIG. 7 , with continued reference to  FIGS. 1-6 , illustrates an exemplary core assembly  80  that can be used to define one or more portions of an internal cooling scheme of a completed component (see, for example, the internal cooing schemes  55 ,  155  of  FIGS. 2, 4 and 6 ). The exemplary core assembly  80  includes a first passage core  90  that defines a first cooling cavity of a completed airfoil, a second passage core  92  that defines the second cooling cavity of the completed airfoil, a third passage core  93  that can define a third cooling cavity of the completed airfoil, and a partial rib  94  that extends from the passage cores  90 ,  92  and  93 . Although three cores are illustrated, the core assembly  80  could include additional or fewer cores depending on the number of desired cooling cavities. 
         [0056]    The partial rib  94  defines a plenum of the completed component. One or more cutouts  96  may be disposed in a face  95  of the partial rib  94 . The cutouts  96  define the ribs of a completed component. In one embodiment, the cutouts  96  are disposed on opposing faces  95  of the partial rib  94  in order to cast ribs on both a first sidewall and a second sidewall of the plenum. The core assembly  80  could include additional features. 
         [0057]    Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
         [0058]    It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
         [0059]    The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.