Patent Publication Number: US-2023140212-A1

Title: Gas turbine rotor component and method of manufacture

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a division of U.S. Application No. 16/793,950 filed Feb. 18, 2020, the disclosure of which is incorporated by reference herein in its entirety. 
    
    
     BACKGROUND 
     Exemplary embodiments pertain to the art of ceramic matrix composites. 
     In gas turbine engines, disks which support turbine blades rotate at high speeds in a high temperature environment. In modern engines, operating temperatures can exceed 1500° F. (816° C.) in the exterior or rim portion of disks, and about 1000° F. (538° C.) at the inner or hub portions. In addition to this radial temperature gradient, there is also a stress gradient, with higher stresses occurring in the lower temperature hub region, while lower stresses occur in the higher temperature rim region in a typical disk. These differences in operating conditions radially across a disk result in different mechanical property requirements in the different disk regions, with the rim portion subjected to severe creep and hold time fatigue crack growth conditions, and the hub portion subjected to severe fatigue and high stress conditions. In order to achieve the maximum operating conditions in terms of efficiency and performance in an advanced turbine engine, it is desirable to utilize disk alloys having excellent hold time fatigue crack growth resistance and high temperature creep resistance in the rim portion while having high tensile strength and fatigue crack resistance at moderate temperatures in the hub portion. 
     BRIEF DESCRIPTION 
     A turbine rotor disk is disclosed. The turbine rotor disk includes a radially inner portion comprising a wrought nickel alloy having a yield strength of at least 126 ksi at 1,000° F. The turbine rotor disk also includes a radially outer portion bonded to the radially inner portion, said radially outer portion comprising a cast nickel alloy configured as a single crystal or with a grain size of ASTM 2 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can include the cast nickel alloy configured as a single crystal. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can include the cast nickel alloy configured with a grain size of ASTM 2 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can have a grain size of ASTM 1 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can have a grain size of ASTM 0 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first portion can include a bore of the turbine rotor disk, and the second portion can include a rim of the turbine rotor disk. 
     Also disclosed is a gas turbine engine comprising a compressor, a combustor, and a turbine disposed along an air flow path including the turbine rotor disk of one or more of the features described above. 
     A method of making a gas turbine rotor component is also disclosed. According to the method, a first preform is provided corresponding to a radially inner portion of the turbine rotor disk. The first preform comprises a wrought nickel alloy having a yield strength of at least 126 ksi at 1,000° F. A second preform is provided corresponding to a radially outer portion of the turbine rotor disk. The second preform comprising a cast nickel alloy configured as a single crystal or with a grain size of ASTM 2 or larger. The first and second preforms are solid-state bonded together under heat and pressure to form a turbine rotor disk including a radially inner portion comprising the first preform and a radially outer portion comprising the second preform. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, bonding the first and second preforms together can include inertia bonding the first and second preforms together. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, bonding the first and second preforms together can include diffusion bonding the first and second preforms together. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can include the cast nickel alloy configured as a single crystal. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can include the cast nickel alloy configured with a grain size of ASTM 2 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, providing the first preform can include working a billet comprising the nickel alloy to form the first preform. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, providing the second preform can include casting the nickel alloy under conditions to form the nickel alloy of the second preform configured as a single crystal. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, providing the second preform can include casting the nickel alloy under conditions to form the nickel alloy of the second preform configured with a grain size of ASTM 2 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can have a grain size of ASTM 1 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second preform can have a grain size of ASTM 0 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, providing the first preform can include working a billet comprising the nickel alloy to form the first preform, and providing the second preform can include casting the nickel alloy under conditions to form the nickel alloy of the second preform as a single crystal or with a grain size of ASTM 2 or larger. 
     In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first preform can include a portion corresponding to a bore of the turbine rotor disk, and the second preform can include a portion corresponding to a rim of the turbine rotor disk. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: 
         FIG.  1    is a schematic illustration of an aircraft that can incorporate various embodiments of the present disclosure; 
         FIG.  2    is a partial cross-sectional view of a gas turbine engine; 
         FIG.  3    schematically shows a turbine disk; and 
         FIGS.  4 A and  4 B  schematically show a cross-sectional view components of a dual alloy disk, and of an assembled dual alloy disk. 
     
    
    
     DETAILED DESCRIPTION 
     A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures. 
     Although shown and described above and below with respect to an aircraft, embodiments of the present disclosure are applicable to turbines used for any type of vehicle or for on-site installation in fixed systems. For example, military vehicles, heavy machinery vehicles, sea craft, ships, submarines, etc., as well as numerous stationary power systems such as electricity generation or other applications where power is generated or used. As such, the present disclosure is not limited to application to aircraft, but rather aircraft are illustrated and described as example and explanatory embodiments for implementation of embodiments of the present disclosure. 
     With respect now to  FIG.  1   , an aircraft includes an aircraft body  101 , which can include one or more bays  103  beneath a center wing box. The bay  103  can contain and/or support one or more components of the aircraft  101 . Also shown in  FIG.  1   , the aircraft includes one or more engines  111 . The engines  111  are typically mounted on the wings  112  of the aircraft and are connected to fuel tanks (not shown) in the wings, but may be located at other locations depending on the specific aircraft configuration. 
       FIG.  2    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 other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct, 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. 
     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 bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a 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 high pressure compressor  52  and 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 . An engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The engine static structure  36  further supports 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. 
     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 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 . 
     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 disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     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 (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), 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 1150 ft/second (350.5 m/sec). 
     Turbines in the turbine section  28  such as the low pressure turbine  46  or the high pressure turbine  54  typically include radially-extending turbine blades attached to a radially central disk. An example embodiment of a turbine rotor  200  is schematically shown in  FIG.  3   . As shown in  FIG.  3   , the turbine rotor  200  includes a disk  210  made up of a radially inner portion  212  extending radially outward from an inner circumference that defines a bore  230 . The radially inner portion  212  is also commonly referred to as a bore or bore portion. The disk  210  also includes a radially outer portion  214  extending inwardly from a rim at the radially outer periphery of the disk  210 . The radially outer portion  214  is also commonly referred to as a rim or rim portion. The portion of the disk  210  between the bore portion and the rim portion is commonly referred to as a web  215 . The turbine rotor  200  also includes a plurality of attachments  300  for blades attached to the radially outer portion  214  at the rim of the disk  210 . The blades can be attached to the disk rim with various type of attachments, including but not limited to mechanical attachment or welded attachment. 
     As mentioned above, a turbine rotor disk is formed by joining preforms of different metals. An example embodiment of a disk  210  is schematically shown in  FIGS.  4  and  4 B  with a cross-section cut-away to illustrate the joining of a radially inner preform  312  and a radially outer preform  314  along a joint  316 . As shown in  FIG.  4 A  (which carries forward numbering from  FIG.  3   ), the inner and outer preforms  312 / 314  are arranged to be brought together along as shown in  FIG.  4 B . The joint  316  is disposed at an angle to (i.e., not parallel with) the rotational axis of the disk  210  (and of course is also at an angle to the radius of the disk  210 ). An angled joint can facilitate contact between the radially inner and outer preforms  312 / 314  during bonding, although it is not required and in some aspects the joint  316  can be parallel with the rotational axis of the disk  210 . 
     As further mentioned above, the radially inner preform  312  can be formed from a wrought nickel alloy having a yield strength (defined in ASTM E8 - 13) of at least 125 ksi at 1,000° F. In some aspects, the radially outer preform  314  can have ayield strength of 135 ksi at 1,000° F. In some aspects, the radially outer preform  314  can have a yield strength of 150 ksi at 1,000° F. Examples of alloys for the radially inner preform  312  include but are not limited to Inconel  718 , Waspaloy, or powder based alloys such as IN-100. The radially inner preform  312  can be formed by any metal forming process, including wrought processing of billets (which in turn can be formed by known techniques such as casting, extrusion, or hot rolling) or nickel starting preforms formed by powder metallurgy. The alloy out of which the radially inner preform  312  is formed can be configured to have properties such as high strength, with fatigue resistance and high fracture toughness. 
     As further mentioned above, the radially outer preform  314  can be formed from a cast nickel alloy configured as a single crystal or an equiaxed alloy with a grain size of ASTM 2 (defined in ASTM E112 - 12) or larger. Examples of alloys for the radially outer preform  314  include but are not limited to Mar-M-200, Mar-M-247, Rene  80 , Rene  125 , or CMSX-4. In some aspects, the alloy out of which the radially outer preform  314  is formed can have a grain size of ASTM 1 or larger, or of ASTM 0 or larger. The radially outer preform  314  can be formed by any casting under conditions to produce the target grain size or single-crystal structure. A single-crystal grain structure can be provided by gradual directional solidification in a ceramic mold in which a helical channel with smooth continuous turning a short distance above a knurled chill plate surface (i.e., “starter chamber”) provides a filtering effect to reduce the number of crystals exiting the channel. A seed crystal can be used to further promote formation of a single crystal grain structure. Coarse grain sizes of ASTM 2 or greater in cast metals can be promoted by higher mold temperatures, greater melt temperature, and slower cooling rates. The alloy out of which the radially outer preform  314  is formed can be configured to have properties such as creep resistance, thermo-mechanical fatigue resistance. 
     The first and second preforms  312 / 314  can be fused together by solid state bonding, also known as thermocompression bonding. Examples of solid state bonding techniques include inertia bonding and diffusion bonding. In some aspects, forge bonding can be uses; however, in some other aspects, forge bonding is avoided in order to avoid trapping of flash inside the forge, and to avoid potential reduction of grain size in the second preform  314 . 
     Inertia bonding is a solid-state bonding technique performed by rotating one or both of the preforms  312  and  314  with respect to each other about the disk axis. In some aspects, the outer preform  314  can be held stationary while the inner preform  312  is rotated. This approach can facilitate engagement of the stationary outer preform  314  with a press or other source of compressive force for application of pressure between the preforms  312  and  314  while allowing for relative rotation of the preform provided by the rotating inner preform  312 . Relative rotation of the preforms  312 / 314  generates heat from friction between the contacted surfaces of the preforms  312  and  314 , and the combination of heat and pressure creates conditions for thermocompression bonding. In inertia bonding-one part is stationary and the other is moving, for linear inertia bonding (linear friction welding) external motion is applied between the two parts while pressure is also applied to cause heating, flow, and joining very rapidly. For rotational friction bonding one part is typically brought up to rotating speed, the external for of rotation removed, and the parts brought into contact under pressure to promote heating, flow, and joining. 
     Diffusion bonding is a solid-state bonding technique performed by contacting the preforms  312  and  314  and applying heat and pressure. Compressive force can be applied with a press or die other source of compressive force to the outer rim surface of the preform  314  and/or to the inner circumference of the preform  312 . Heat can be applied externally, such as by placing the preforms  312 / 314  in a furnace or oven during bonding or internally such as by induction. The preforms  312 / 314  can be maintained at these conditions for a period of time sufficient to produce a bond (e.g., 1-12 hours). 
     Once bonded, the preforms  312  and  314  are joined together along the joint  316 . The joint  316  can be a solid state weld joint that contains elements from the metal alloys of each of the preforms, and can exhibit hybrid or blended properties of the blended alloys. The combination of different alloys provides a technical benefit of a robust rotor disk structure having customized properties for the hot conditions encountered by the outer portion  214  of the disk  210  and the high-stress conditions encountered by the inner portion  212  of the disk  210 . For example, traditional rotor alloys are limited to temperatures far below the gas path temperatures of turbine engines. Substantial cooling is required to keep the rotors at an acceptable temperature, this cooling air contributes to loss of efficiency (thrust-specific fuel consumption, also known as TSFC) in the engine. Engine efficient could be improved by increasing allowable rotor rim temperatures. 
     The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof. 
     While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.