Patent Publication Number: US-10309251-B2

Title: Interlocking rotor assembly with thermal shield

Description:
BACKGROUND 
     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 combustion 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. 
     The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
     A speed reduction device, such as an epicyclical gear assembly, may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed. 
     SUMMARY 
     A rotor assembly according to an example of the present disclosure includes a first rotor, a second rotor mounted on the first rotor and co-rotatable there with, and a thermal shield interlocked with the second rotor for co-rotation there with. 
     In a further embodiment of any of the foregoing embodiments, the first rotor has a first outer diameter and the second rotor has a second outer diameter that is smaller than the first outer diameter. 
     In a further embodiment of any of the foregoing embodiments, the second rotor includes at least one radially-extending tab and the thermal shield includes at least one radially-extending tab circumferentially interlocked with the at least one radially-extending tab of the second rotor. 
     In a further embodiment of any of the foregoing embodiments, the one radially-extending tab of the second rotor includes a step. 
     In a further embodiment of any of the foregoing embodiments, the second rotor includes a plurality of radially-extending circumferentially-spaced tabs and the thermal shield includes a plurality of radially-extending circumferentially-spaced tabs circumferentially interlocked with the plurality of radially-extending circumferentially-spaced tabs of the second rotor. 
     In a further embodiment of any of the foregoing embodiments, the first rotor includes a plurality of radially-extending circumferentially-spaced tabs and the plurality of radially-extending circumferentially-spaced tabs of the second rotor are circumferentially interlocked with the plurality of radially-extending circumferentially-spaced tabs of the first rotor. 
     In a further embodiment of any of the foregoing embodiments, the plurality of radially-extending circumferentially-spaced tabs of the thermal shield are circumferentially aligned with the plurality of radially-extending circumferentially-spaced tabs of the first rotor. 
     In a further embodiment of any of the foregoing embodiments, the plurality of radially-extending circumferentially-spaced tabs of the thermal shield are axially trapped between the second rotor and the plurality of radially-extending circumferentially-spaced tabs of the first rotor. 
     In a further embodiment of any of the foregoing embodiments, the second rotor includes a plurality of radially outwardly-extending circumferentially-spaced tabs and the thermal shield includes a plurality of radially inwardly-extending circumferentially-spaced tabs circumferentially interlocked with the plurality of radially outwardly-extending circumferentially-spaced tabs of the second rotor. 
     In a further embodiment of any of the foregoing embodiments, the second rotor includes an axially-facing pocket, and a portion of the thermal shield is seated in the axially-facing pocket. 
     In a further embodiment of any of the foregoing embodiments, the thermal shield is a continuous ring. 
     A gas turbine engine according to an example of the present disclosure includes a first rotor, a second rotor mounted on the first rotor and co-rotatable there with, and a thermal shield interlocked with the second rotor for co-rotation there with. 
     A method of assembling a rotor assembly according to an example of the present disclosure includes interlocking a thermal shield with a second rotor for co-rotation there with. The second rotor is mounted on a first rotor and co-rotatable there with. 
     In a further embodiment of any of the foregoing embodiments, the interlocking includes mounting the thermal shield on the second rotor and rotating the thermal shield to circumferentially misalign at least one tab on the thermal shield with at least one tab on the second rotor. 
     In a further embodiment of any of the foregoing embodiments, the tab on the thermal shield is axially offset with respect to the tab on the second rotor. 
     In a further embodiment of any of the foregoing embodiments, the interlocking moves the second rotor relative to the thermal shield to axially align, and circumferentially interlock, the tab on the thermal shield with the tab on the second rotor. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The various features and advantages of the present 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. 
         FIG. 1  illustrates an example gas turbine engine. 
         FIG. 2  illustrates an axial cross-section of a rotor assembly of the gas turbine engine of  FIG. 1 . 
         FIG. 3  illustrates a perspective view of a portion of the rotor assembly of  FIG. 2 . 
         FIG. 4  illustrates an axial view of a portion of the rotor assembly of  FIG. 2 . 
         FIG. 5  illustrates a magnified view of a portion of the rotor assembly of  FIG. 2 . 
         FIG. 6  illustrates an axial cross-sectional view of a tab of a second rotor of the rotor assembly of  FIG. 2 . 
         FIG. 7  illustrates a thermal shield of a rotor assembly as a continuous ring. 
         FIG. 8  illustrates mounting of a second rotor onto a first rotor of a rotor assembly. 
         FIG. 9A  illustrates an axial cross-sectional view of mounting a thermal shield onto a second rotor of a rotor assembly. 
         FIG. 9B  illustrates an axial view according to the section line shown in  FIG. 9A . 
         FIG. 10  illustrates an axial view of rotating a thermal shield during assembly of a rotor assembly. 
         FIG. 11  illustrates an axial cross-sectional view of moving a second rotor axially forward during assembly of a rotor assembly. 
     
    
    
     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 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 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 is to be understood that the concepts described herein are not limited to use with two-spool turbofans and the teachings can be applied to other types of turbine engines, including three-spool architectures and ground-based engines. 
     The engine  20  includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central axis A relative to an engine static structure  36  via several bearing systems, shown at  38 . It is to be understood that various bearing systems at various locations may alternatively or additionally be provided, and the location of bearing systems may be varied as appropriate to the application. 
     The low speed spool  30  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 this example is a gear system  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 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 between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing system  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via, for example, bearing systems  38  about the engine central 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 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 gear system  48  can 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 engine. In a further example, the engine  20  has a bypass ratio that is greater than about six (6), with an example embodiment being greater than about ten (10), the gear system  48  is an epicyclic gear train, such as a planet or star 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 (5). In one disclosed embodiment, the 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). 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  can be an epicycle gear train, such as a planet or star gear system, with a gear reduction ratio of greater than about 2.3:1. It is to be understood, however, that the above parameters are only exemplary and that the present disclosure is applicable to other gas turbine engines. 
     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 (&#39;TSFC&#39;)”—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. 
     The engine  20  includes a rotor assembly  60  (shown schematically) that is rotatable about the engine central axis A. In this example, the rotor assembly  60  is in the turbine section  28  and is a first stage rotor of the high pressure turbine  54 . Turbine blades  62  are mounted on the rotor assembly  60 . It is to be understood that although the examples herein are described with reference to the rotor assembly  60  being in the turbine section  28 , the examples are not limited to the turbine section  28 , high pressure turbine  54  or first stage rotor. 
       FIG. 2  illustrates an isolated axial cross-sectional view of the rotor assembly  60 . The rotor assembly  60  includes a first rotor  64 , a second rotor  66  that is mounted on the first rotor  64  and co-rotatable there with, and a thermal shield  68  interlocked with the second rotor  66  for co-rotation there with. That is, rotation of the first rotor  64  causes the second rotor  66  and thermal shield  68  to co-rotate with each other and with the first rotor  64 . The first rotor  64 , the second rotor  66  and the thermal shield  68  can be formed of superalloy materials, such as but not limited to nickel- or cobalt-based alloys, ceramics, composites and the like. 
     In this example, the second rotor  66  (which can alternatively be termed a “mini-rotor” or “mini-disk”) is generally smaller in mass than the first rotor  64 , which carries the turbine blades  62 . The first rotor  64  has an outer diameter D 1  (with respect to engine central axis A) and the second rotor  66  has a second diameter D 2  (with respect to engine central axis A) that is smaller than the first diameter D 1 . The first rotor  64  serves to carry the turbine blades  62 , while the second rotor  66  serves to provide secondary functions, such as but not limited to sealing. In this regard, the second rotor  66  can include one or more sealing features (not shown), such as knife seals. 
     Referring also to  FIG. 3  showing a perspective view of a portion of the rotor assembly  60 ,  FIG. 4  showing an axial view of a portion of the rotor assembly  60  and  FIG. 5  showing a magnified perspective view of a portion of the rotor assembly  60 , the second rotor  66  includes at least one tab  66   a  and the thermal shield  68  includes at least one tab  68   a.  As described in further detail below, the at least one tab  66   a  of the second rotor  66  circumferentially interlocks with the at least one tab  68   a  of the thermal shield  68  to secure the thermal shield  68  in place. It is to be understood that relative positional terms, such as “forward,” “aft,” “circumferential,” “radial” and the like are with reference to the normal operational attitude of the engine  20  and engine central axis A, unless otherwise indicated. 
     In the illustrated example, the second rotor  66  includes a plurality of the tabs  66   a  and the thermal shield  68  includes a plurality of the tabs  68   a,  although only one of each of the tabs  66   a / 68   a  is needed for circumferential interlocking. The tabs  66   a  extend radially outwards and are circumferentially-spaced. The tabs  68   a  extend radially inwards and are also circumferentially-spaced. The tabs  66   a  are circumferentially interlocked with the tabs  68   a  such that the second rotor  66  and thermal shield  68  are locked for co-rotation. 
       FIG. 6  shows an axial cross-section of a portion of the second rotor  66  and a representative one of the tabs  66   a.  As shown, the tab  66   a  has a step  70 . The circumferential sides of the step  70  are proximate the respective circumferential sides of the tabs  68   a  on the thermal shield  68 . Upon rotation of the rotor assembly  60 , the circumferential sides of the step  70  can abut the respective circumferential sides of the tabs  68   a  on the thermal shield  68 . 
     The second rotor  66  includes an axially-facing pocket  72  ( FIG. 2 ) formed in a radially outer portion thereof relative to the tabs  66   a.  The thermal shield  68  includes a corresponding curved portion  74  that nests or seats in the pocket  72  such that the curved portion  74  contacts the wall of the pocket  72 . In this example, in a fully nested or seated position, there is a gap between the portion and the walls of the pocket  72 , which can facilitate thermal shielding. 
     The first rotor  64  also includes at least one tab  64   a  ( FIG. 3 ). In this example, the first rotor  64  includes a plurality of the tabs  64   a.  The tabs  64   a  extend radially outwards and are circumferentially-spaced. In a fully assembled state of the rotor assembly  60 , the tabs  64   a  are circumferentially aligned with, and axially offset from, the tabs  68   a  of the thermal shield  68 . The tabs  68   a  are circumferentially offset from, and axially aligned with, the steps  70  of the tabs  66   a  of the second rotor  66 . The lower forward portions of the tabs  66   a  circumferentially interlock with the tabs  64   a  such that the second rotor  66  co-rotates with the first rotor  64 . Further, the tabs  68   a  of the thermal shield  68  are axially trapped between the second rotor  66  and the tabs  64   a  of the first rotor  64 . The thermal shield  68  is thus axially and circumferentially locked in position by the first rotor  64  and the second rotor  66 , without the use of an additional fastener or fastening device. In this regard, the thermal shield  68  need not be split (i.e., a split ring) and can instead be a continuous ring, as schematically shown in  FIG. 7 . 
       FIG. 8 ,  FIGS. 9A / 9 B,  FIG. 10  and  FIG. 11  illustrate various views of the rotor assembly  60  through a method of assembling the rotor assembly  60 . In a broad example, the method includes interlocking the thermal shield  68  with the second rotor  66  for co-rotation there with. 
     The following further examples describe assembly of the second rotor  66  onto the first rotor  64 , and assembly of the thermal shield  68  onto the second rotor  66 . Referring to  FIG. 8 , the second rotor  66  is mounted onto the first rotor  64 . For example, the second rotor  66  is mounted such that the tabs  66   a  are circumferentially misaligned with, and axially offset from, the tabs  64   a  of the first rotor  64 . That is, the second rotor  66  is axially “over-seated” from its fully assembled position in which the tabs  66   a  are axially aligned with the tabs  64   a  of the first rotor  64 . 
     Referring to  FIGS. 9A / 9 B, the thermal shield  68  is then mounted onto the second rotor  68 . To mount the thermal shield  68 , the tabs  68   a  are circumferentially aligned with the tabs  66   a  of the second rotor  66  and circumferentially misaligned with the tabs  64   a  of the first rotor  64  such that the tabs  68   a  are received between and through the spaces between the tabs  64   a.  The thermal shield  68  is moved such that the tabs  68   a  axially clear the tabs  64   a . As shown in  FIG. 10 , the thermal shield  68  is then rotated (either clockwise or counterclockwise) such that the tabs  68   a  move out of circumferential alignment with the tabs  66   a  of the second rotor  66  and into circumferential alignment with the tabs  64   a  of the first rotor  64 . 
     Referring to  FIG. 11 , the second rotor  66  is then moved axially forward (to the left in  FIG. 11 ) into a fully seated position relative to the thermal shield  68  and first rotor  64 . The axial forward movement moves the lower forward portions of the tabs  66   a  of the second rotor  66  into axial alignment with the tabs  64   a  of the first rotor  64 , thus circumferentially interlocking the second rotor  66  and the first rotor  64 . The axial forward movement also moves the steps  70  of the tabs  66   a  of the second rotor  66  into axial alignment with the tabs  68   a  of the thermal shield  68 , thus circumferentially interlocking the thermal shield  68  and the second rotor  68 . The axial forward movement additionally moves the curved portion  74  of the thermal shield  68  into a nested or seated position in the pocket  72  of the second rotor  66 . 
     Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure.