Patent Publication Number: US-10760423-B2

Title: Spoked rotor for a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     The present disclosure is a divisional of U.S. patent application Ser. No. 13/283,689, filed Oct. 28, 2011. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine, and more particularly to a rotor system therefor. 
     Gas turbine rotor systems include successive rows of blades, which extend from respective rotor disks that are arranged in an axially stacked configuration. The rotor stack may be assembled through a multitude of systems such as fasteners, fusion, tie-shafts and combinations thereof. 
     Gas turbine rotor systems operate in an environment in which significant pressure and temperature differentials exist across component boundaries which primarily separate a core gas flow path and a secondary cooling flow path. For high-pressure, high-temperature applications, the components experience thermo-mechanical fatigue (TMF) across these boundaries. Although resistant to the effects of TMF, the components may be of a heavier-than-optimal weight for desired performance requirements. 
     SUMMARY 
     A rotor for a gas turbine engine according to an exemplary aspect of the present disclosure includes a plurality of blades which extend from a rotor disk, each of the plurality of blades extend from the rotor disk at an interface, the interface defined along a spoke. 
     A spool for a gas turbine engine according to an exemplary aspect of the present disclosure includes a compressor rotor disk defined along an axis of rotation. A plurality of compressor blades which extend from the compressor rotor disk, each of the plurality of compressor blades extend from compressor rotor disk at an interface, said interface defined along a spoke. 
     A spool for a gas turbine engine according to an exemplary aspect of the present disclosure includes a rotor disk defined along an axis of rotation. A plurality of blades which extend from the rotor disk, each of the plurality of blades extend from the rotor disk at a blade interface, the blade interface defined along a spoke radially inboard of a blade platform. A rotor ring defined about the axis of rotation, the rotor ring axially adjacent to the rotor disk. A plurality of core gas path seals which extend from the rotor ring, each of the plurality of core gas path seals extend from the rotor ring at a seal interface, the seal interface defined along a spoke, the plurality of core gas path seals axially adjacent to the blade platform. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-sectional view of a gas turbine engine; 
         FIG. 2  is an exploded view of the gas turbine engine separated into primary build modules; 
         FIG. 3  is an enlarged schematic cross-sectional view of a high pressure compressor section of the gas turbine engine; 
         FIG. 4  is a perspective view of a rotor of the high pressure compressor section; 
         FIG. 5  is an expanded partial sectional perspective view of the rotor of  FIG. 4 ; 
         FIG. 6  is an expanded partial sectional perspective view of a portion of the high pressure compressor section; 
         FIG. 7  is a top partial sectional perspective view of a portion of the high pressure compressor section with an outer directed inlet; 
         FIG. 8  is a top partial sectional perspective view of a portion of the high pressure compressor section with an inner directed inlet; 
         FIG. 9  is an expanded partial sectional view of a portion of the high pressure compressor section; 
         FIG. 10  is an expanded partial sectional perspective view of a portion of the high pressure compressor section illustrating a rotor stack load path; 
         FIG. 11  is a RELATED ART expanded partial sectional perspective view of a portion of the high pressure compressor section illustrating a more tortuous rotor stack load path; 
         FIG. 12  is an expanded partial sectional perspective view of a portion of the high pressure compressor section illustrating a wire seal structure; 
         FIG. 13  is an expanded schematic view of the wire seal structure; 
         FIG. 14  is an expanded partial sectional perspective view of a high pressure turbine section; 
         FIG. 15  is an expanded exploded view of the high pressure turbine section; and 
         FIG. 16  is an expanded partial sectional perspective view of the rotor of  FIG. 15 . 
     
    
    
     DETAILED DESCRIPTION 
       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 augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a 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 turbofans as the teachings may be applied to other types of turbine engines, such as three-spool architectures. 
     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. 
     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  may be connected to the fan  42  directly or through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30  which in one disclosed non-limiting embodiment includes a gear reduction ratio of, for example, at least 2.4:1. The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor (HPC)  52  and high pressure turbine (HPT)  54 . A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate 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  54 ,  46  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     The gas turbine engine  20  is typically assembled in build groups or modules ( FIG. 2 ). In the illustrated embodiment, the high pressure compressor  52  includes eight stages and the high pressure turbine  54  includes two stages in a stacked arrangement. It should be appreciated, however, that any number of stages will benefit hereform as well as other engine sections such as the low pressure compressor  44  and the low pressure turbine  46 . Further, other gas turbine architectures such as a three-spool architecture with an intermediate spool will also benefit herefrom as well. 
     With reference to  FIG. 3 , the high pressure compressor (HPC)  52  is assembled from a plurality of successive HPC rotors  60 C which alternate with HPC spacers  62 C arranged in a stacked configuration. The rotor stack may be assembled in a compressed tie-shaft configuration, in which a central shaft (not shown) is assembled concentrically within the rotor stack and secured with a nut (not shown), to generate a preload that compresses and retains the HPC rotors  60 C with the HPC spacers  62 C together as a spool. Friction at the interfaces between the HPC rotor  60 C and the HPC spacers  62 C is solely responsible to prevent rotation between adjacent rotor hardware. 
     With reference to  FIG. 4 , each HPC rotor  60 C generally includes a plurality of blades  64  circumferentially disposed around a rotor disk  66 . The rotor disk  66  generally includes a hub  68 , a rim  70 , and a web  72  which extends therebetween. Each blade  64  generally includes an attachment section  74 , a platform section  76  and an airfoil section  78  ( FIG. 5 ). 
     The HPC rotor  60 C may be a hybrid dual alloy integrally bladed rotor (IBR) in which the blades  64  are manufactured of one type of material and the rotor disk  66  is manufactured of different material. Bi-metal construction provides material capability to separately address different temperature requirements. For example, the blades  64  are manufactured of a single crystal nickel alloy that are transient liquid phase bonded with the rotor disk  66  which is manufactured of a different material such as an extruded billet nickel alloy. Alternatively, or in addition to the different materials, the blades  64  may be subject to a first type of heat treat and the rotor disk  66  to a different heat treat. That is, the Bi-metal construction as defined herein includes different chemical compositions as well as different treatments of the same chemical compositions such as that provided by differential heat treatment. 
     With reference to  FIG. 5 , a spoke  80  is defined between the rim  70  and the attachment section  74 . The spoke  80  is a circumferentially reduced section defined by interruptions which produce axial or semi-axial slots which flank each spoke  80 . The spokes  80  may be machined, cut with a wire EDM or other processes to provide the desired shape. An interface  801  that defines the transient liquid phase bond and or heat treat transition between the blades  64  and the rotor disk  66  are defined within the spoke  80 . That is, the spoke  80  contains the interface  801 . Heat treat transition as defined herein is the transition between differential heat treatments. 
     The spoke  80  provides a reduced area subject to the thermo-mechanical fatigue (TMF) across the relatively high temperature gradient between the blades  64  which are within the relatively hot core gas path and the rotor disk  66  which is separated therefrom and is typically cooled with a secondary cooling airflow. 
     With reference to  FIG. 6 , the HPC spacers  62 C provide a similar architecture to the HPC rotor  60 C in which a plurality of core gas path seals  82  are bonded or otherwise separated from a rotor ring  84  at an interface  861  defined along a spoke  86 . In one example, the seals  82  may be manufactured of the same material as the blades  64  and the rotor ring  84  may be manufactured of the same material as the rotor disk  66 . That is, the HPC spacers  62 C may be manufactured of a hybrid dual alloy which are transient liquid phase bonded at the spoke  86 . Alternatively, the HPC spacers  62 C may be manufactured of a single material but subjected to the differential heat treat which transitions within the spoke  86 . In another disclosed non-limiting embodiment, a relatively low-temperature configuration will benefit from usage of a single material such that the spokes  86  facilitate a weight reduction. In another disclosed non-limiting embodiment, low-temperature bi-metal designs may further benefit from dissimilar materials for weight reduction where, for example, low density materials may be utilized where load carrying capability is less critical. 
     The rotor geometry provided by the spokes  80 ,  86  reduces the transmission of core gas path temperature via conduction to the rotor disk  66  and the seal ring  84 . The spokes  80 ,  86  enable an IBR rotor to withstand increased T3 levels with currently available materials. Rim cooling may also be reduced from conventional allocations. In addition, the overall configuration provides weight reduction at similar stress levels to current configurations. 
     The spokes  80 ,  86  in the disclosed non-limiting embodiment are oriented at a slash angle with respect to the engine axis A to minimize windage and the associated thermal effects. That is, the spokes are non-parallel to the engine axis A. 
     With reference to  FIG. 7 , the passages which flank the spokes  80 ,  86  may also be utilized to define airflow paths to receive an airflow from an inlet HPC spacer  62 CA. The inlet HPC spacer  62 CA includes a plurality of inlets  88  which may include a ramped flow duct  90  to communicate an airflow into the passages defined between the spokes  80 ,  86 . The airflow may be core gas path flow which is communicated from an upstream, higher pressure stage for use in a later section within the engine such as the turbine section  28 . 
     It should be appreciated that various flow paths may be defined through combinations of the inlet HPC spacers  62 CA to include but not limited to, core gas path flow communication, secondary cooling flow, or combinations thereof. The airflow may be communicated not only forward to aft toward the turbine section, but also aft to forward within the engine  20 . Further, the airflow may be drawn from adjacent static structure such as vanes to effect boundary flow turbulence as well as other flow conditions. That is, the HPC spacers  62 C and the inlet HPC spacer  62 CA facilitate through-flow for use in rim cooling, purge air for use downstream in the compressor, turbine, or bearing compartment operation. 
     In another disclosed non-limiting embodiment, the inlets  88 ′ may be located through the inner diameter of an inlet HPC spacer  62 CA′ ( FIG. 8 ). The inlet HPC spacer  62 CA′ may be utilized to, for example, communicate a secondary cooling flow along the spokes  80 ,  86  to cool the spokes  80 ,  86  as well as communicate secondary cooling flow to other sections of the engine  20 . 
     In another disclosed non-limiting embodiment, the inlets  88 ,  88 ′ may be arranged with respect to rotation to essentially “scoop” and further pressurize the flow. That is, the inlets  88 ,  88 ′ include a circumferential directional component. 
     With reference to  FIG. 9 , each rotor ring  84  defines a forward circumferential flange  92  and an aft circumferential flange  94  which is captured radially inboard of the associated adjacent rotor rim  70 . That is, each rotor ring  84  is captured therebetween in the stacked configuration. In the disclosed tie-shaft configuration with multi-metal rotors, the stacked configuration is arranged to accommodate the relatively lower-load capability alloys on the core gas path side of the rotor hardware, yet maintain the load-carrying capability between the seal rings  84  and the rims  70  to transmit rotor torque. 
     That is, the alternating rotor rim  70  to seal ring  84  configuration carries the rotor stack preload—which may be upward of 150,000 lbs—through the high load capability material of the rotor rim  70  to seal ring  84  interface, yet permits the usage of a high temperature resistant, yet lower load capability materials in the blades  64  and the seal surface  82  which are within the high temperature core gas path. Divorce of the sealing area from the axial rotor stack load path facilitates the use of a disk-specific alloy to carry the stack load and allows for the high-temp material to only seal the rotor from the flow path. That is, the inner diameter loading and outer diameter sealing permits a segmented airfoil and seal platform design which facilitates relatively inexpensive manufacture and highly contoured airfoils. The disclosed rotor arrangement facilitates a compressor inner diameter bore architectures in which the reduced blade/platform pull may be taken advantage of in ways that produce a larger bore inner diameter to thereby increase shaft clearance. 
     The HPC spacers  62 C and HPC rotors  60 C of the IBR may also be axially asymmetric to facilitate a relatively smooth axial rotor stack load path ( FIG. 10 ). The asymmetry may be located within particular rotor rims  70 A and/or seal rings  84 A. For example, the seal ring  84 A includes a thinner forward circumferential flange  92  compared to a thicker aft circumferential flange  94  with a ramped interface  84 Ai. The ramped interface  84 Ai provides a smooth rotor stack load path. Without tangentially slot assembled airfoils in an IBR, the load path along the spool may be designed in a more efficient manner as compared to the heretofore rather torturous conventional rotor stack load path ( FIG. 11 ; RELATED ART). 
     With reference to  FIG. 12 , the blades  64  and seal surface  82  may be formed as segments that include tangential wire seals  96  between each pair of the multiple of seal surfaces  82  and each pair of the multiple of blades  64  as well as axial wire seals  98  between the adjacent HPC spacers  62 C and HPC rotors  60 C. The tangential wire seals  96  and the axial wire seals  98  are located within teardrop shaped cavities  100  ( FIG. 13 ) such that centrifugal forces increase the seal interface forces. 
     Although the high pressure compressor (HPC)  52  is discussed in detail above, it should be appreciated that the high pressure turbine (HPT)  54  ( FIG. 14 ) is similarly assembled from a plurality of successive respective HPT rotor disks  60 T which alternate with HPT spacers  62 T ( FIG. 15 ) arranged in a stacked configuration and the disclosure with respect to the high pressure compressor (HPC)  52  is similarly applicable to the high pressure turbine (HPT)  54  as well as other spools of the gas turbine engine  20  such as a low spool and an intermediate spool of a three-spool engine architecture. That is, it should be appreciated that other sections of a gas turbine engine may alternatively or additionally benefit herefrom. 
     With reference to  FIG. 14 , each HPT rotor  60 T generally includes a plurality of blades  102  circumferentially disposed around a rotor disk  124 . The rotor disk  124  generally includes a hub  126 , a rim  128 , and a web  130  which extends therebetween. Each blade  102  generally includes an attachment section  132 , a platform section  134 , and an airfoil section  136  ( FIG. 16 ). 
     The blades  102  may be bonded to the rim  128  along a spoke  136  at an interface  1361  as with the high pressure compressor (HPC)  52 . Each spoke  136  also includes a cooling passage  138  generally aligned with each turbine blade  102 . The cooling passage  138  communicates a cooling airflow into internal passages (not shown) of each turbine blade  102 . 
     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 in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.