Abstract:
A method of manufacturing a gas turbine engine component includes providing a core having a brittle feature, supporting the feature with a first meltable material, arranging the core with the first meltable material in a first mold, and surrounding the core and the first meltable material with a second meltable material to provide a component shape. The method also includes coating the second meltable material with a refractory material to produce a second mold, removing the first and second meltable material, and casting a component in the second mold.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a gas turbine engine airfoil, for example. More particularly, the disclosure relates to a method of manufacturing a component with a thin ceramic core feature. 
         [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]    Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. 
         [0004]    Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages. Typically the internal cooling passages are formed using ceramic cores and/or refractory metal cores. Ceramic cores become increasingly fragile as the thickness and width decrease. As a result, thin cooling passage features cannot be formed using ceramic cores. Instead, a refractory metal core, which includes molybdenum for example, is glued into a slot in a thicker ceramic core to provide, for example, an airfoil trailing edge cooling passage. Using multiple core materials can be relatively expensive. 
       SUMMARY 
       [0005]    In one exemplary embodiment, a method of manufacturing a gas turbine engine component includes providing a core having a brittle feature, supporting the feature with a first meltable material, arranging the core with the first meltable material in a first mold, and surrounding the core and the first meltable material with a second meltable material to provide a component shape. The method also includes coating the second meltable material with a refractory material to produce a second mold, removing the first and second meltable material, and casting a component in the second mold. 
         [0006]    In a further embodiment of the above, the core and feature are constructed from ceramic. 
         [0007]    In a further embodiment of any of the above, the feature has a thickness of less than 0.013 inch and a width of greater than 0.100 inch. 
         [0008]    In a further embodiment of any of the above, the core is an airfoil trailing edge core. 
         [0009]    In a further embodiment of any of the above, the trailing edge core has a thickness of less than 0.013 inch and a width of greater than 0.100 inch. The core includes an integral adjacent core structure that has a thickness of greater than 0.013 inch. 
         [0010]    In a further embodiment of any of the above, the airfoil trailing edge core has multiple holes. The supporting step includes having the first meltable material extend through the holes. 
         [0011]    In a further embodiment of any of the above, the supporting step includes having the first meltable material adjoin the adjacent core structure. 
         [0012]    In a further embodiment of any of the above, the arranging step includes assembling multiple core structures relative to one another within the mold. The core structures are configured to provide cooling passages in an airfoil. 
         [0013]    In a further embodiment of any of the above, the first and second meltable materials are wax. 
         [0014]    In a further embodiment of any of the above, the supporting step includes dipping the feature in molten wax. 
         [0015]    In a further embodiment of any of the above, the surrounding step includes injecting molten wax into the first mold. 
         [0016]    In a further embodiment of any of the above, the coating step includes dipping the second meltable material in ceramic slurry, and providing the second mold with a hardened ceramic exterior. 
         [0017]    In a further embodiment of any of the above, the casting step includes pouring molten metal into the second mold. 
         [0018]    In a further embodiment of any of the above, the molten metal is a nickel alloy. 
         [0019]    In a further embodiment of any of the above, the component is one of a blade and a vane. 
         [0020]    In a further embodiment of any of the above, the method of manufacturing a gas turbine engine component includes the step of removing the core and the second mold from the component. 
         [0021]    In another exemplary embodiment, a core for a gas turbine engine component includes a ceramic core structure with a feature extending from the core structure. The feature has a thickness of less than 0.013 inch and a width of greater than 0.100 inch. A first meltable material coats the feature and adjoins the core structure. 
         [0022]    In a further embodiment of any of the above, a second meltable material surrounds the core structure, the feature and the first meltable material. 
         [0023]    In a further embodiment of any of the above, the first and second meltable materials are wax. 
         [0024]    In a further embodiment of any of the above, the feature is integral with the core structure. The feature is an airfoil trailing edge core. 
         [0025]    In another exemplary embodiment, a method of manufacturing a gas turbine engine component core includes providing a core having a brittle feature and supporting the feature with a first meltable material. 
         [0026]    In a further embodiment of any of the above, the core is an airfoil trailing edge core. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0028]      FIG. 1  schematically illustrates a gas turbine engine embodiment. 
           [0029]      FIG. 2A  is a perspective view of the airfoil having the disclosed cooling passage. 
           [0030]      FIG. 2B  is a plan view of the airfoil illustrating directional references. 
           [0031]      FIG. 3  is a cross-sectional view of the airfoil taken along line  3 - 3  in  FIG. 2A . 
           [0032]      FIG. 4  is an enlarged view of a ceramic core structure including a support for a brittle ceramic feature. 
           [0033]      FIG. 5  is a flow chart depicting an example method of manufacturing a gas turbine engine component, such as an airfoil. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      FIG. 1  schematically illustrates an example gas turbine engine  20  that includes 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 or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
         [0035]    Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
         [0036]    The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis X 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. 
         [0037]    The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis X. 
         [0038]    A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
         [0039]    The example low pressure turbine  46  has a pressure ratio that is greater than about five ( 5 ). The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
         [0040]    A mid-turbine frame  57  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  57  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0041]    The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  59  of the mid-turbine frame  57  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  57 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
         [0042]    The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
         [0043]    In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
         [0044]    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 pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
         [0045]    “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 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
         [0046]    “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)] 0.5 . The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
         [0047]    The disclosed cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade  64  is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example. 
         [0048]    Referring to  FIGS. 2A and 2B , a root  74  of each turbine blade  64  is mounted to the rotor disk. The turbine blade  64  includes a platform  76 , which provides the inner flow path, supported by the root  74 . An airfoil  78  extends in a radial direction R from the platform  76  to a tip  80 . It should be understood that the turbine blades may be integrally formed with the rotor such that the roots are eliminated. In such a configuration, the platform is provided by the outer diameter of the rotor. The airfoil  78  provides leading and trailing edges  82 ,  84 . The tip  80  is arranged adjacent to a blade outer air seal (not shown). 
         [0049]    The airfoil  78  of  FIG. 2B  somewhat schematically illustrates exterior airfoil surface extending in a chord-wise direction C from a leading edge  82  to a trailing edge  84 . The airfoil  78  is provided between pressure (substantially concave) and suction (substantially convex) wall  86 ,  88  in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple turbine blades  64  are arranged circumferentially in a circumferential direction A. The airfoil  78  extends from the platform  76  in the radial direction R, or spanwise, to the tip  80 . 
         [0050]    The airfoil  78  includes a cooling passage  90  provided between the pressure and suction walls  86 ,  88 . The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage  90 . The example cooling passages  90  illustrated in  FIG. 2A  is shown in more detail in  FIG. 3 . 
         [0051]      FIG. 3  illustrates one example arrangement of cooling passages  90  in the turbine blade  64 . One of the cooling passages  90  includes a trailing edge main passage  92  that communicates with a relatively thin trailing edge cooling passage  94  that extends to the trailing edge  84 . The trailing edge cooling passage  94  may include pins  96  that extend laterally between the pressure and suction side walls and promote turbulence. 
         [0052]    The trailing edge cooling passages  92 ,  94  are provided by a ceramic core structure  98 , shown in  FIG. 4 . The ceramic core structure  98  includes a main core structure  100  from which a trailing edge core  102  is integral with and extends from. The trailing edge core  102  includes holes  104 , which form the pins  96  during the casting process. Since the ceramic core structure  98  is constructed from a brittle material, the trailing edge core  102  is susceptible to breaking away from the core  100  during handling and casting. In the example, the trailing edge core  102  has a width  110  greater than 0.100 in (2.54 mm) and a thickness  112  of less than 0.013 in (0.33 mm) The main core structure  100  includes a thickness  108  of greater than 0.013 in (0.33 mm). 
         [0053]    A support  106  is arranged on either side of the trailing edge core  102 . The support  106  adjoins the core  100  to support the trailing edge core  102  relative to the main core structure  100  to resist breakage. In one example, the support  106  is provided by a meltable material such as wax or a water-soluble material. 
         [0054]    Referring to  FIG. 5 , a method  114  of manufacturing a gas turbine engine component, such as an airfoil, is depicted in the flow chart. A ceramic core is manufactured having a brittle feature, for example a thickness of less than 0.013 in (0.33 mm) and a width of greater than 0.10 in (2.54 mm), as schematically indicated at block  116 . The feature, which may be a trailing edge core  102 , is supported with a first meltable material, such as wax or a water-soluble material, as indicated at block  118 . The feature may be dipped into a molten wax to provide the support  106 , which extends through the holes  104  of the trailing edge core  102 . The core structure  98  along with any other cores may be assembled into a mold, as indicated at block  120 . The mold provides an exterior shape of a component, such as a turbine blade, and the cores provide the shape of the internal cooling passages  90 . 
         [0055]    The cores and support  106  are surrounded by a second meltable material, as indicated at block  122 . In one example, the first mold is injected with molten wax, which is solidified to provide a component shape. The solidified wax is removed from the first mold and coated in a refractory material, such as ceramic slurry, as indicated at block  124 . Once the ceramic slurry has solidified, the first and second meltable materials are removed. Molten metal may be poured into the second mold provided by the hardened ceramic, to provide the cast component, as indicated at block  126 . In the example of a turbine blade, the blade is cast from a nickel alloy. The hardened ceramic is broken away from the cast component, and the cores are removed by a chemical leaching process, for example. 
         [0056]    Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that and other reasons, the following claims should be studied to determine their true scope and content.