Abstract:
A blade assembly for a gas turbine engine includes a rim seal including leading and trailing edge seal portions joined to one another by an axial portion. The leading and trailing edge seal portions and the axial portion together provide a notch. A blade has a root received in the notch. A rotating stage of a gas turbine engine includes a rotor including a slot. A rim seal includes leading and trailing edge seal portions adjoined to one another by an axial portion and providing a notch. A blade has a root received in the notch. A method of assembling a rotor stage includes inserting a root of a blade into a notch of a rim seal, and sliding the rim seal and blade into a rotor slot.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a multi-piece blade for a gas turbine engine. In one example, a two-piece turbine blade is provided. 
         [0002]    A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
         [0003]    A typical gas engine includes turbine blades that are a single piece. The turbine blade includes a root, which may be a fir-tree shape, received in a correspondingly shaped rotor slot. A platform is supported on the root and provides an aerodynamic inner flow path through the stage. An airfoil extends outward radially from the platform. The platform may provide complex geometries and includes the seal geometry that seals with adjacent structure to the rotor. 
         [0004]    It may be desirable to provide at least a portion of the platform that is separate from the airfoil. In one example, a composite platform has been provided, which includes forward and aft portions of the root. The platform includes an aperture through which the airfoil extends. The composite platform entirely surrounds the airfoil. 
       SUMMARY 
       [0005]    In one exemplary embodiment, a blade assembly for a gas turbine engine includes a rim seal including leading and trailing edge seal portions joined to one another by an axial portion. The leading and trailing edge seal portions and the axial portion together provide a notch. A blade has a root received in the notch. 
         [0006]    In a further embodiment of any of the above, the blade has a first circumferential edge and the rim seal has a second circumferential edge. The first and second circumferential edges are aligned with one another in generally the same plane. 
         [0007]    In a further embodiment of any of the above, the rim seal and the blade root together provide a root contour. 
         [0008]    In a further embodiment of any of the above, the rim seal and the blade together provide a platform defining an inner flow path. 
         [0009]    In a further embodiment of any of the above, the blade includes an airfoil extending from the platform. The rim seal does not circumscribe the airfoil. 
         [0010]    In a further embodiment of any of the above, the axial portion is arranged beneath the root opposite the airfoil. 
         [0011]    In a further embodiment of any of the above, the rim seal and blade are constructed from metallic alloys. 
         [0012]    In a further embodiment of any of the above, the axial portion and the leading and trailing edge seal portions are integral with one another. 
         [0013]    In a further embodiment of any of the above, the axial portion and the leading and trailing edge seal portions are discrete from and secured to one another. 
         [0014]    In another exemplary embodiment, a rotating stage of a gas turbine engine includes a rotor including a slot. A rim seal includes leading and trailing edge seal portions adjoined to one another by an axial portion and providing a notch. A blade has a root received in the notch. 
         [0015]    In a further embodiment of any of the above, the rotating stage includes a retainer secured to the rotor configured to maintain the rim seal and blade within the slot. 
         [0016]    In a further embodiment of any of the above, the rotating stage includes a seal structure adjacent to the rim seal. The rim seal includes seal geometry interleaved with the seal structure. 
         [0017]    In a further embodiment of any of the above, the blade has a first circumferential edge and the rim seal has a second circumferential edge. The first and second circumferential edges are aligned with one another in generally the same plane. 
         [0018]    In a further embodiment of any of the above, the rim seal and the blade root together provide a root contour. The rim seal and the blade together provide a platform defining an inner flow path. The blade includes an airfoil extending from the platform. The rim seal does not circumscribe the airfoil and the axial portion is arranged beneath the root opposite the airfoil. 
         [0019]    In a further embodiment of any of the above, the axial portion and the leading and trailing edge seal portions are integral with one another. 
         [0020]    In a further embodiment of any of the above, the axial portion and the leading and trailing edge seal portions are discrete from and secured to one another. 
         [0021]    In another exemplary embodiment, a method of assembling a rotor stage includes inserting a root of a blade into a notch of a rim seal, and sliding the rim seal and blade into a rotor slot. 
         [0022]    In a further embodiment of any of the above, the method includes securing a retainer to the rotor to abut the rim seal. 
         [0023]    In a further embodiment of any of the above, the method includes arranging a seal geometry of the rim seal in an interleaved relationship with a seal structure. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0025]      FIG. 1  schematically illustrates a gas turbine engine embodiment. 
           [0026]      FIG. 2  is a perspective view of a portion of a rotor supporting multiple two-pieced turbine blades. 
           [0027]      FIG. 3  is a perspective view of an example multi-piece turbine blade. 
           [0028]      FIG. 4  is an exploded view depicting a rim seal separate from the turbine blade. 
       
    
    
     DETAILED DESCRIPTION 
       [0029]      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 . 
         [0030]    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. 
         [0031]    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 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. 
         [0032]    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 A. 
         [0033]    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. 
         [0034]    The example low pressure turbine  46  has a pressure ratio that is greater than about 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. 
         [0035]    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 . 
         [0036]    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. 
         [0037]    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. 
         [0038]    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. 
         [0039]    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. 
         [0040]    “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. 
         [0041]    “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. 
         [0042]    An example turbine stage is illustrated in  FIG. 2 . The stage includes a rotor  60  having multiple slots  62  circumferentially arranged about an outer perimeter of the rotor  60 . A multi-piece blade assembly is provided, which is mounted within each slot  62 . In the example, one piece provides a rim seal  64  and another piece provides a blade  76 , which is a turbine blade in the example. The rim seal  64  and blade  76  are constructed from metallic alloys. The rim seal  64  and blade  76  slide together as an assembly into the slot  62  of the rotor  60  during installation. 
         [0043]    The rim seal  64  is provided by leading and trailing edge seal portions  68 ,  70  that are axially spaced apart from one another to provide a notch  66 , or gap, between the portions, best shown in  FIG. 4 . In the example, an axial portion  72  is integral with and interconnects the leading and trailing edge seal portions  68 ,  70  to provide a unitary, cast structure that forms a cradle that receives the blade  76 . The axial portion  72  and leading and trailing edge seal portions  68 ,  70  may be discrete components secured to one another, if desired. The rim seal  64  may be constructed from any suitable material for the given application. The leading and trailing edge seal portions  68 ,  70  provide inner axial surfaces  74  that are spaced apart from and face one another. The blade  76  includes spaced apart outer axial surfaces  84  that are adjacent to and engage the inner axial surfaces  74  with the blade  76  received in the notch  66 . The integral arrangement of the rim seal  64  maintains tight clearances between the inner and outer axial surfaces  74 ,  84 , which minimize leakage through the stage. 
         [0044]    The blade  76  is received in the notch  66  of the rim seal  64 , as shown in  FIG. 3 . The blade  76  includes a root  78  that together with the rim seal  64  provides a root contour  98  having a shape corresponding to the shape of the slot  62 . In the example, the root contour  98  corresponds to a firtree shape. 
         [0045]    The blade  76  includes a platform  80  that supports an airfoil  82 , which extends radially outward from the platform  80 . The platform  80  together with outer surfaces of the leading and trailing edge seal portions  68 ,  70  provide an inner flow path through the stage. 
         [0046]    The leading and trailing edge seal portions  68 ,  70  respectively provide forward and aft seal geometries  86 ,  88 . Referring to  FIG. 2 , the forward and aft seal geometries  86 ,  88  cooperate with forward and aft seal structures  90 ,  92  to provide an air seal along the inner flow path. Forward and aft retainers  94 ,  96  are secured to the rotor  60  to retain the rim seal  64  and blade  76  axially within the slots  62 . 
         [0047]    The blade  76  and rim seal  64  respectively include circumferential edges  100 ,  102  that adjoin and align with one another in the assembled position with the rim seals and blades  64 ,  76  installed in the rotor  60 . In the example, the circumferential edges  100 ,  102  are generally in the same plane as one another. The rim seal  64  does not circumscribe the airfoil  82 . 
         [0048]    Having a rim seal  64  that is separate from the blade  76  enables the seal geometry to be more easily changed without creating new blades  76 , which is a complex and expensive component to manufacture. However, the platform  80  and root  78  remains a unitary, cast structure with the airfoil  82  to provide a structurally robust design. 
         [0049]    Although an example embodiment has 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 reason, the following claims should be studied to determine their true scope and content.