Abstract:
A component according to an exemplary aspect of the present disclosure includes, among other things, a body comprised of a first material, a cover attached to the body and comprised of a second material, and a braze alloy employable to braze the cover to the body and comprised of a third material. The first material, the second material and the third material are different materials.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/950,869 which was filed on Mar. 11, 2014. 
     
    
     BACKGROUND 
       [0002]    This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component, such as a blade, that includes a brazed cover. 
         [0003]    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. The compressor section and turbine section typically employ alternating rows of rotating blades and stationary vanes that drive the hot combustion gases along a core flow path. Blades and vanes are typically cast structures and may include internal cooling passages depending on their location within the engine. 
         [0004]    A casting core may be used to form an internal cooling passage inside of the component during a casting operation. The casting core must be properly positioned inside the casting and may include surfaces that extend through the cast part, thereby creating openings or holes at undesirable locations once the core has been removed after casting. These openings must be sealed in order to close-off the internal cooling passage. One common technique for sealing the openings includes welding. However, welding operations generally create extreme local heat inputs that can lead to cracking in the part being welded. 
       SUMMARY 
       [0005]    A component according to an exemplary aspect of the present disclosure includes, among other things, a body comprised of a first material, a cover attached to the body and comprised of a second material, and a braze alloy employable to braze the cover to the body and comprised of a third material. The first material, the second material and the third material are different materials. 
         [0006]    In a further non-limiting embodiment of the foregoing component, an internal cooling passage extends inside the body, the internal cooling passage coated with an internal coating. 
         [0007]    In a further non-limiting embodiment of either of the foregoing components, the first material is a nickel-based superalloy. 
         [0008]    In a further non-limiting embodiment of any of the foregoing components, the second material is Hastelloy-X. 
         [0009]    In a further non-limiting embodiment of any of the foregoing components, the third material is AMS 4777. 
         [0010]    A blade for a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, an airfoil that extends to a tip shroud. The airfoil includes a cooling passage. A cover is attached to the tip shroud and covers the cooling passage, and a braze alloy is applied around the cover to braze the cover to the tip shroud. The airfoil, the cover and the braze alloy each include different material compositions. 
         [0011]    In a further non-limiting embodiment of the foregoing blade, the cover is positioned and configured to adapt relative to an uneven surface of the tip shroud. 
         [0012]    In a further non-limiting embodiment of either of the foregoing blades, the airfoil is made of a nickel-based superalloy. 
         [0013]    In a further non-limiting embodiment of any of the foregoing blades, the cover is made of a nickel-based alloy. 
         [0014]    In a further non-limiting embodiment of any of the foregoing blades, the braze alloy is made of a nickel-based compound. 
         [0015]    In a further non-limiting embodiment of any of the foregoing blades, the airfoil is made of a rhenium-free, nickel-based superalloy. 
         [0016]    In a further non-limiting embodiment of any of the foregoing blades, the cover is made of Hastelloy-X. 
         [0017]    In a further non-limiting embodiment of any of the foregoing blades, the braze alloy is made of AMS 4777. 
         [0018]    In a further non-limiting embodiment of any of the foregoing blades, the airfoil includes an internal aluminide coating. 
         [0019]    In a further non-limiting embodiment of any of the foregoing blades, the cover includes a thickness of about 0.010 inches (0.254 mm) 
         [0020]    A gas turbine engine method according to another exemplary aspect of the present disclosure includes, among other things, brazing a cover to a blade using a braze alloy. The cover, the blade and the braze alloy each comprise different material compositions. 
         [0021]    In a further non-limiting embodiment of the foregoing gas turbine engine method, prior to the brazing step the gas turbine engine method includes casting the blade, positioning the cover relative to blade and applying the braze alloy around the cover. 
         [0022]    In a further non-limiting embodiment of either of the foregoing gas turbine engine methods, prior to the brazing step, the gas turbine engine method includes applying an internal coating to an internal cooling passage of the blade, positioning the cover over at least one opening in a tip shroud of the blade and applying the braze alloy around the cover. 
         [0023]    In a further non-limiting embodiment of any of the foregoing gas turbine engine methods, prior to the brazing step, the method includes positioning the cover at an uneven surface of a tip shroud of the blade and adapting the cover to conform to the uneven surface. 
         [0024]    In a further non-limiting embodiment of any of the foregoing gas turbine engine methods, the gas turbine engine method is a repair method for repairing a part having a defect. 
         [0025]    The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible. 
         [0026]    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 
         [0027]      FIG. 1  illustrates a schematic, cross-sectional view of a gas turbine engine. 
           [0028]      FIG. 2  illustrates a gas turbine engine component. 
           [0029]      FIG. 3  illustrates the gas turbine engine component of  FIG. 2  prior to removal of a casting core. 
           [0030]      FIG. 4  illustrates a top view of the gas turbine engine component of  FIG. 2 . 
           [0031]      FIGS. 5A ,  5 B and  5 C illustrate a blade of a gas turbine engine. 
           [0032]      FIG. 5D  illustrates a cover that may be brazed to a gas turbine engine component. 
           [0033]      FIG. 6  schematically illustrates a gas turbine engine manufacturing method. 
           [0034]      FIG. 7  schematically illustrates a gas turbine engine repair method. 
       
    
    
     DETAILED DESCRIPTION 
       [0035]    This disclosure is directed to a gas turbine engine component, such as a turbine blade, that includes an airfoil, a cover attached to a tip shroud of the airfoil, and a braze alloy used to affix the cover to the tip shroud. The airfoil, the cover and the braze alloy may each include different material compositions. In one embodiment, the airfoil, the cover and the braze alloy are made from different nickel-based alloys. By brazing the cover, extreme local heat inputs and associated thermal stresses are substantially removed, thereby reducing part susceptibility to cracking. These and other features are discussed in greater detail herein. 
         [0036]      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 . 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 in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. For example, the teachings of this disclosure also extend to ground-based gas turbine engines. 
         [0037]    The exemplary 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 the 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 . 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 in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . 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 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0039]    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  57  includes airfoils  59  which are in the core airflow path C. 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 . 
         [0040]    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 gear system  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. 
         [0041]    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 (10,668 meters). 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)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1,150 ft/second (350.5 meters/second). 
         [0042]    Each of the compressor section  24  and the turbine section  28  may include alternating rows of rotor assemblies and vane assemblies (shown schematically). 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  may either create or extract energy in the form of pressure from the core airflow as it is communicated along the core flow path C. The vanes  27  direct the core airflow to the blades  25  to either add or extract energy. 
         [0043]      FIGS. 2 ,  3  and  4  illustrate a gas turbine engine component (hereinafter “component”)  60 . The non-limiting embodiment depicted by  FIGS. 2 ,  3  and  4  illustrate the component  60  as a blade, such as a turbine blade. It should be understood, however, that this disclosure is not limited to blades. 
         [0044]    The component  60  may include a body  62  that defines both an external shape and an internal shape of the component  60 . In one non-limiting embodiment, the body  62  includes an airfoil  64 , a platform  66  and a root  68 . The airfoil  64  extends outwardly in a first direct from the platform  66 , and the root  68  extends from the platform  66  in an opposed, second direction away from the airfoil  64 . The root  68  is adapted for connecting the component  60  to a rotating disk of a rotor assembly (not shown). 
         [0045]    The airfoil  64  may extend between the platform  66  and a tip shroud  70 . The tip shroud  70  is positioned at a tip  71  of the airfoil  64  and includes an outer diameter surface  74  that faces away from the platform  66 . The tip shroud  70  may include rails  72  that project radially outwardly from the outer diameter surface  74 . The rails  72  define knife seals that interface relative to a stationary engine structure (not shown) that may circumscribe the component  60 . 
         [0046]    In one non-limiting embodiment, the component  60  is a cast part and includes an internal cooling passage  76  (shown in phantom in  FIG. 2 ) that extends inside of the body  62 . For example, the internal cooling passage  76  may extend at least partially inside of the airfoil  64 . In one non-limiting embodiment, the internal cooling passage  76  is a serpentine cooling passage. The internal cooling passage  76  may be formed during a casting process, such as an investment casting process, using a casting core  78  (shown in phantom lines in  FIG. 3 ). The casting core  78  is removed from  FIG. 2  in order to better illustrate the configuration of the internal cooling passage  76 . The casting core  78  could include a ceramic core, a refractory metal core or a combined ceramic/refractory metal core. 
         [0047]    As best illustrated in  FIG. 3 , the casting core  78  may include one or more print-out posts  79  that protrude through the outer diameter surface  74  of the tip shroud  70 . The print-out posts  79  aid in positioning the casting core  78 , setting the wall thickness of the airfoil  64 , and preventing breakage of the cast part during the casting operation. 
         [0048]    Referring now to  FIGS. 3 and 4 , removal of the casting core  78 , including the print-out posts  79 , subsequent to a casting operation may form one or more openings  80  (e.g., holes) at the outer diameter surface  74  of the tip shroud  70 . The openings  80  must be sealed against the ingress or egress of airflow in order to close-off the internal cooling passage  76  so it can function to cool the component  60 . Exemplary configurations for achieving such sealing are discussed in additional detail below. 
         [0049]      FIGS. 5A and 5B  illustrate a turbine blade  160 . The turbine blade  160  may employed within a turbine section of a gas turbine engine, including but not limited to, within a low pressure turbine, a high pressure turbine or any intermediate turbine. 
         [0050]    The exemplary turbine blade  160  includes an airfoil  164  that extends between a platform  166  and a tip shroud  170 . The tip shroud  170  defines an outer diameter surface  174  that faces away from the platform  166 . Rails  172  may extend radially outwardly from the outer diameter surface  174 . One or more openings  180  may be formed in the outer diameter surface  174 . The openings  180  are formed in a finished casting after a casting core has been removed from the casting. 
         [0051]    Referring to  FIG. 5C , with continued reference to  FIGS. 5A and 5B , a cover  82  may be attached to the outer diameter surface  174  of the tip shroud  170  in order to seal the openings  180 . In this embodiment, the cover  82  is positioned to cover and seal two openings  180 . However, the cover  82  may seal one or more openings  180 . Although only a single cover  82  is illustrated in  FIG. 4 , multiple covers could be utilized to seal a component that includes a multitude of openings. 
         [0052]    The cover  82  may include a thickness T (see  FIG. 5D ) of approximately 0.010 inches (0.254 mm), with a tolerance of +/−0.002 inches (0.051 mm). The relatively thin thickness T enables the cover  82  to conform to irregular surfaces during assembly. In one embodiment, the cover  82  may be positioned over an uneven surface  86  of the outer diameter surface  174  of the tip shroud  170  during assembly. The exemplary cover  82  may also provide weight benefits and net “pull” (i.e., centrifugal load stress) reductions. 
         [0053]    The turbine blade  160  may additionally include a braze alloy  84 . In one embodiment, the cover  82  is brazed to the tip shroud  170  using the braze alloy  84 . 
         [0054]    Each of the airfoil  164 , the cover  82  and the braze alloy  84  may include different material compositions. For example, the airfoil  164  may be made of a nickel-based superalloy (i.e., a first material). One non-limiting embodiment of a suitable nickel-based superalloy includes a rhenium-free, investment cast, nickel-based superalloy. 
         [0055]    The cover  82  may be made of a sheet metal form of a nickel-based alloy (i.e., a second material). One non-limiting embodiment of a suitable nickel-based alloy is Hastelloy-x. 
         [0056]    The braze alloy  84  may be made of a nickel-based compound (i.e, a third material). One non-limiting embodiment of a suitable nickel-based compound includes AMS 4777. 
         [0057]    The turbine blade  160  may additionally include an internal cooling passage  176  for internally cooling the part (see  FIG. 5A ). In one non-limiting embodiment, the internal walls of the turbine blade  160  that circumscribe the internal cooling passage  176  are coated with an internal coating  90 . The internal coating  90  provides corrosion protection. One non-limiting embodiment of a suitable internal coating  90  is an aluminide coating. 
         [0058]    In another embodiment, the external walls of the turbine blade  160  are coated with an external coating  92 . The external coating  92  provides oxidation protection. Suitable external coatings include aluminide coatings or diffused overlay/sprayed coatings. 
         [0059]      FIG. 6 , with continued reference to  FIGS. 5A-5D , schematically illustrates a gas turbine engine manufacturing method  100 . The method may begin at block  102  by casting the turbine blade  160 . Of course, this disclosure is not limited to manufacturing a turbine blade. The turbine blade  160  may be investment cast using a casting core to form the internal cooling passage  176  inside the airfoil  164 . The turbine blade  160  may optionally undergo machining operations at block  104 . 
         [0060]    Next, at block  106 , the turbine blade  160  is cleaned. In one non-limiting cleaning procedure, the turbine blade  160  is furnace cleaned for thirty minutes at 1300° F. (704° C.). The turbine blade  160  may additionally be silicon carbide blasted and degreased. Other cleaning techniques are also contemplated. 
         [0061]    An internal coating, such as an aluminide coating, may be applied to the internal cooling passage  176  of the turbine blade  160  at block  108 . The internal coating may be applied using openings  180  formed at the outer diameter surface  174  of the tip shroud  170 . 
         [0062]    The cover  82  is positioned relative to the turbine blade  160  at block  110 . In one non-limiting embodiment, the cover  82  is tack-welded to the outer diameter surface  174  of the tip shroud  170  of the turbine blade  160  to attach the cover  82 . The cover  82  may conceal one or more openings  80  formed through the outer diameter surface  174  during the casting process of block  102 . 
         [0063]    Next, at block  112 , the braze alloy  84  may be applied around the edges of the cover  82 . The braze alloy  84  may be applied as a slurry or a paste, in one embodiment. Alternatively, the cover  82  could be pre-alloyed using a sintering process, thereby eliminating the need apply the braze alloy  84  around the cover  82 . 
         [0064]    Stop-off may be applied around the cover  82  and the braze alloy  84  at block  114 . The stop-off is applied to prevent undesired flow of the braze alloy  84  away from the cover  82 . 
         [0065]    At block  116 , the cover  82  is brazed to the outer diameter surface  174  of the tip shroud  170 . In one non-limiting embodiment, the turbine blade  160  is vacuum furnace brazed at approximately 1925° F. (1052° C.) for around fourteen minutes to braze the cover  82  to the turbine blade  160 . Finally, at block  118 , an external coating may be applied to the turbine blade  160 . The turbine blade  160  could then be subjected to an additional furnace operation. 
         [0066]      FIG. 7  illustrates a gas turbine engine repair method  200 . The method  200  may be employed to repair a turbine blade  160  that has been damaged or otherwise includes a defect. First, at block  202 , the cover  82  is removed from the blade  160 . Removal of the cover  82  may expose openings  80  in an outer diameter surface  174  of the tip shroud  170 . Exemplary removal operations include electrical discharge machining (EDM), hand-blending, milling, grinding or other operations. 
         [0067]    Next, at block  204 , the blade  160  is cleaned. An internal cooling passage  176  of the blade  160  may be recoated at block  206 . The internal coating may include an aluminide coating, in one non-limiting embodiment. 
         [0068]    A new cover  82  may next be positioned and/or attached to the tip shroud at block  208 . The braze alloy  84  may be applied around the cover  82  at block  210 , such as in the form of a slurry or paste. A pre-sintered cover  82  may alternatively be used. Finally, the cover  82  may be brazed to the tip shroud  170  of the blade  160  at block  212 . 
         [0069]    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. 
         [0070]    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. 
         [0071]    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.