Abstract:
An exemplary method of forming an endwall with a contour includes casting an endwall with at least one cooling channel having an opening from the endwall, and covering the opening with a cupped contour formed on the endwall. An exemplary gas turbine engine blade assembly includes an endwall with a plurality of cooling channels, an airfoil extending radially from the endwall to a tip, and a cupped contour formed on the endwall to provide a cooling chamber between the cupped contour and a radially facing surface of the endwall.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a contoured endwall of a gas turbine engine blade assembly. 
         [0002]    Gas turbine engines typically include a compressor section, a combustor section, and a turbine section. 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 are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
         [0003]    The compressor and turbine sections of a gas turbine engine typically include alternating rows of rotating blades and stationary vanes. The turbine blades rotate and extract energy from the hot combustion gases that are communicated through the gas turbine engine. The turbine vanes prepare the airflow for the next set of blades. 
         [0004]    The vanes and blades extend from endwalls that may be contoured to manipulate flow. The outer casing of an engine static structure may include one or more blade outer air seals (BOAS) providing endwalls that are be contoured to manipulate flow by reducing secondary flow losses and flow migration. 
       SUMMARY 
       [0005]    A method of forming an endwall with a contour according to an exemplary aspect of the present disclosure includes, among other things, casting an endwall with at least one cooling channel having an opening from the endwall, and covering the opening with a cupped contour that is formed on the endwall to provide a portion of a gas path surface. 
         [0006]    In a further non-limiting embodiment of the foregoing method, the method includes forming the cupped contour on the endwall using additive manufacturing. 
         [0007]    In a further non-limiting embodiment of any of the foregoing methods, the endwall has a first material composition, and the cupped contour has a second material composition different than the first material composition. 
         [0008]    In a further non-limiting embodiment of any of the foregoing methods, the method includes forming the cupped contour on a radially facing surface of the endwall. 
         [0009]    In a further non-limiting embodiment of any of the foregoing methods, the method includes covering an inlet to at least one cooling channel of the endwall with the cupped contour. 
         [0010]    In a further non-limiting embodiment of any of the foregoing methods, an outlet and the inlet open to a cooling chamber provided by the cupped contour. 
         [0011]    In a further non-limiting embodiment of any of the foregoing methods, the endwall is an endwall of a blade. 
         [0012]    In a further non-limiting embodiment of any of the foregoing methods, the method includes providing at least one cooling channel in the endwall using a core held within a mold during the casting. 
         [0013]    In a further non-limiting embodiment of any of the foregoing methods, a portion of the core providing the opening of the cooling channel provides an attachment point that secures the core within the mold. 
         [0014]    In a further non-limiting embodiment of any of the foregoing methods, the endwall is an endwall of a blade assembly and the attachment point is a first attachment point, a second attachment point that secures the core within the mold is adjacent a blade tip of the blade assembly, and a third attachment point that secures the core within the mold is adjacent a root of the blade assembly. 
         [0015]    A gas turbine engine assembly according to another exemplary aspect of the present disclosure includes, among other things, an endwall having a first material composition, and a cupped contour of a second material composition that is formed on the endwall. The first material composition is different than the second material composition. 
         [0016]    In a further non-limiting embodiment of the foregoing assembly, at least one cooling channel of the endwall opens to a cooling chamber provided by the cupped contour. 
         [0017]    In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is entirely radially misaligned from the cooling chamber. 
         [0018]    In a further non-limiting embodiment of any of the foregoing assemblies, the cupped contour covers one or more cooling channel openings. 
         [0019]    In a further non-limiting embodiment of any of the foregoing assemblies, the cupped contour covers one or more cooling channel inlets. 
         [0020]    In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is an endwall of a blade. 
         [0021]    In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is a cast endwall, and the cupped contour is an additively manufactured cupped contour. 
         [0022]    A gas turbine engine blade assembly according to an exemplary aspect of the present disclosure includes, among other things, an endwall with a plurality of cooling channels, an airfoil extending radially from the endwall to a tip, and a cupped contour formed on the endwall to provide a cooling chamber between the cupped contour and a radially facing surface of the endwall. 
         [0023]    In a further non-limiting embodiment of the foregoing assembly, at least one opening of the plurality of cooling channels opens to the cooling chamber, and at least one inlet of the plurality of cooling channels opens to the cooling chamber. The cupped contour covers the at least one opening and the at least one inlet. 
         [0024]    In a further non-limiting embodiment of any of the foregoing assemblies, the endwall is a cast endwall, and the cupped contour is and additively manufactured cupped contour. 
     
    
     
       DESCRIPTION OF THE FIGURES 
         [0025]    The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows: 
           [0026]      FIG. 1  illustrates an example gas turbine engine. 
           [0027]      FIG. 2  illustrates a gas turbine engine blade assembly from a turbine section of the gas turbine engine of  FIG. 1  having a cupped contour formed on an endwall. 
           [0028]      FIG. 3  illustrates a section view of the gas turbine engine blade assembly at the Line  3 - 3  in  FIG. 2 . 
           [0029]      FIG. 4  illustrates a close-up side view of the cupped contour from the gas turbine engine blade assembly of  FIG. 2 . 
           [0030]      FIG. 5  illustrates a bottom view of the cupped contour of  FIG. 3  showing a cooling chamber of the cupped contour. 
           [0031]      FIG. 6  illustrates a section view at a position of Line  3 - 3  of a mold and a core used to cast the gas turbine engine blade assembly of  FIG. 2 . 
           [0032]      FIG. 7  illustrates a section view of the gas turbine engine blade assembly of  FIG. 2  at Line  3 - 3  prior to forming a cupped contour formed on the endwall. 
           [0033]      FIG. 8  illustrates the section of the gas turbine engine blade assembly of  FIG. 5  at Line  3 - 3  when forming the cupped contour on the endwall. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]    This disclosure relates to a contoured endwall of an assembly of a gas turbine engine assembly. More particularly, this disclosure relates to a cupped contour that provides a cooling chamber. A cooling channel formed within an endwall of a gas turbine engine assembly opens to the cooling chamber within the cupped contour. 
         [0035]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26 , and a turbine section  28 . 
         [0036]    The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26 , and then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, the examples herein are not limited to use with two-spool turbofans and may be applied to other types of turbomachinery, including direct drive engine architectures, three-spool engine architectures, and ground-based turbines. 
         [0037]    The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
         [0038]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . 
         [0039]    The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports the bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with their longitudinal axes. 
         [0040]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes airfoils, which are in the core airflow path C. 
         [0041]    The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
         [0042]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines, including direct drive turbofans. 
         [0043]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]°  5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. 
         [0044]    Referring now to  FIG. 2  with continuing reference to  FIG. 1 , an example gas turbine engine blade assembly  60  from the turbine section  28  of the gas turbine engine of  FIG. 1  includes an endwall  64 , an airfoil  68 , and a root  72  or base. The airfoil  68  extends radially in a first direction from the endwall  64  to a tip region  76 . The root  72  extends radially in an opposite, second direction from the endwall  64 . 
         [0045]    The blade assembly  60  includes a cupped contour  80  formed on the endwall  64 . The cupped contour  80  helps manipulate and direct flow over and near the endwall  64  when the gas turbine engine blade assembly  60  is used in the gas turbine engine  20 . In another example, an endwall of a vane includes the cupped contour. 
         [0046]    Referring now to  FIGS. 3-5  with continuing reference to  FIG. 2 , the blade assembly  60  includes at least one cooling channel  84 . During operation, fluid, such as air from an air supply  88 , communicates through the cooling channel  84  to cool the gas turbine engine blade assembly  60 . The compressor section  24  of the gas turbine engine  20  ( FIG. 1 ) provides the air supply  88  in some examples. 
         [0047]    In this example, air enters the at least one cooling channel  84  near the root  72 . The air communicates through the cooling channel  84  and exits at an outlet near the tip region  76  of the airfoil  68 . The air communicates thermal energy from the blade assembly  60 . 
         [0048]    For simplicity, a single cooling channel  84  is shown in  FIG. 3 . The blade assembly  60  could include, however, a network of several cooling channels  84  that are each separate and distinct from one another. The cooling channels  84  could have a serpentine configuration. The cooling channels  84  could have an outlet in a leading edge of the airfoil  68 , the trailing edge of the airfoil  68 , or some other area of the blade assembly  60 . 
         [0049]    The example cooling channel  84  extends through a portion of the endwall  64 , such that the endwall  64  provides a portion of the at least one cooling channel  84 . The endwall  64  provides an opening  92  from the portion of the cooling channel  84  within the endwall  64 . Air communicates from the cooling channel  84  of the endwall  64  through the opening  92  into a cooling chamber  96  provided by the cupped contour  80 . Moving air into the cooling chamber  96  cools the cupped contour  80 , the endwall  64 , or both. The cooling chamber  96  is part of the cooling channel  84  in this example. 
         [0050]    Notably, in this example, the cupped contour  80  covers the opening  92  so that air does not exit directly from the cooling chamber  96 . Moving air through the cooling chamber  96  of the cooling channel  84  cools the cupped contour  80 . Prior art contours have lacked cooling chambers, and thus were prone to experience undesirable temperature extremes. 
         [0051]    The endwall  64  is, in this example, a cast metallic component. The endwall  64  is cast together with the airfoil  68  and the root  72 . The cupped contour  80  is then applied to the endwall  64  via an additive manufacturing process. 
         [0052]    Referring now to  FIG. 6 , with continuing reference to  FIG. 3 , a mold  100  and a core  104  provide an open area  108  that receives molten material when casting the endwall  64 , the airfoil  68 , and the root  72 . The open area  108  is filled with a wax that melts when molten material is introduced into the open area  108 . 
         [0053]    The core  104  is a silica-based ceramic core in this example, but other materials such as alumina based ceramics or refractory metals are possible. The core  104  blocks the molten material from entering areas that will form the cooling chamber  84  within the endwall  64 , the airfoil  68 , and the root  72 . 
         [0054]    The core  104  includes a first attachment point  112 , a second attachment point  116 , and a third attachment point  120 . The attachment points  112 ,  116 , and  120  provide an at least 3-point physical datum nest on the core  104  surface retaining the core  104  to design intent relative to the mold  100  when casting the endwall  64 , the airfoil  68 , and the root  72 . 
         [0055]    The first attachment point  112  extends through an area that will provide the opening  92  to the cooling chamber  96 . The second attachment point  116  is adjacent an area that will form the tip region  76 . The third attachment point  120  is adjacent an area that will provide the root  72 . 
         [0056]    Prior art cores used to create cooling channels within gas turbine engine blades have lacked at least an attachment point near the endwall, which can complicate the casting process as the core was more difficult to stabilize within a mold. With a lack of a third contact retention feature, outside of attachment points  116  and  120 , the core exhibits a rotational degree of freedom revolving around an axis created by the centroid points of  116  and  120 . This rotation, during casting, allows the core  104  to misalign and can negatively impact the desire casting wall thickness distribution. 
         [0057]    Referring now to  FIG. 7 , when the molten material has hardened within the open area  108 , the endwall  64 , the airfoil  68 , and the root  72  are removed from the mold  100 . The cooling chamber  84  extends through these portions of the gas turbine engine blade assembly  60 . 
         [0058]    The opening  92  within the endwall  64  extends through a surface  98  of the endwall  64  that will face radially outward when positioned within the gas turbine engine  20  of  FIG. 1 . The cupped contour  80  is radially outside the surface of the endwall  64  such that the endwall  64  is radially misaligned from the cooling chamber  96 . 
         [0059]    Referring now to  FIG. 8 , an additive manufacturing process is used, in this example, to form the cupped contour  80  on the endwall  64 , which will cover the opening  92 . Adding the cupped contour  80  after casting, rather than casting the cupped contour  80 , can reduce manufacturing complexity associated with casting components of complex geometries. 
         [0060]    The cupped contour  80  has a material composition that can be different than a material composition of the remaining portions of the blade assembly  60 . Typical materials are classed as superalloys of the nickel or cobalt base element variety with typical materials being PW1484, CMSX-4, Mar-M-247, Mar-M-509, Rene 80, Rene N5, Haynes Alloy X, Inconel 625, Inconel 723, and others whose material properties are suitable to high temperature applications. 
         [0061]    Additive manufacturing uses a wand  124  to stack successive layers L 1  and L 2  of material, and additional layers, until the desired cupped contour  80  is formed and covers the opening  92 . The endwall  64  and remaining portions of the blade assembly  60  can be preheated from 500 to 1900 degrees Fahrenheit to facilitate adhesion between the cupped contour  80  and the remaining portion of the blade assembly  60 , for example. The preheating temperature is typically 500 degrees Fahrenheit below the sintering temperature of the deposited material. Upon initial deposit, secondary heating operations with the laser can be used to further consolidate the material and provide heat treatment to further solution the alloy. 
         [0062]    Air can move from the cooling channel  84  of the endwall  64  through the opening  92  into the cooling chamber  96 , and from the cooling chamber  96  through the opening back to the cooling channel  84  of the endwall  64 . Thus, air can move to the cooling chamber  96  and from the cooling chamber  96  through the singular opening  92  of the cooling channel  84 . 
         [0063]    In another example, one or more cooling channels in the endwall  64  open to the cooling chamber  96  to permit air air into the cooling chamber  96 , and one or more other cooling channels open to the cooling chamber  96  to permit air to move from the cooling chamber  96  back to the network of cooling channels  84  in the endwall  64 . 
         [0064]    Air could move from to or from the cooling chamber  96  through an opening in the airfoil  68 , or another portion of the blade assembly  60  in other examples. 
         [0065]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.