Abstract:
An assembly for a gas turbine engine includes a minidisk that includes an axial extension extending from a disc. The axial extension includes an inner diameter surface and a recess arranged radially opposite the inner diameter surface. The recess provides a radially outwardly extending flange and a bumper extending radially inward from and proud of the inner diameter surface. A method of working on a gas turbine engine section includes inserting a tool into a cavity beneath a seal assembly, and engaging a flange of a minidisk with the tool to manipulate first and second rotors with respect to one another.

Description:
TECHNICAL FIELD 
       [0001]    This disclosure relates to a minidisk for a gas turbine engine turbine section, and more particularly, the disclosure relates to a feature on the minidisk for assembly and disassembly of the turbine section. 
       BACKGROUND OF THE INVENTION 
       [0002]    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. 
         [0003]    Early minidisks were used as windage covers disposed upon rotating gas turbine engine rotors. More modern minidisks are also used to cool the turbine rotor. An axial extension of the minidisk may extend into an area having a seal assembly and bearing. 
         [0004]    A typical turbine section includes multiple turbine rotors that are secured to one another using a very large press fit. The minidisk has used an annular recess on the axial extension. A tool cooperates with the annular recess to apply a press load to assemble or disassemble the turbine rotors for service. One prior art arrangement required the entire bearing and seal assembly to be removed to gain access to the annular recess, which made servicing the turbine section considerably more costly. 
       SUMMARY 
       [0005]    In one exemplary embodiment, an assembly for a gas turbine engine includes a minidisk that includes an axial extension extending from a disc. The axial extension includes an inner diameter surface and a recess arranged radially opposite the inner diameter surface. The recess provides a radially outwardly extending flange and a bumper extending radially inward from and proud of the inner diameter surface. 
         [0006]    In a further embodiment of any of the above, the recess is axially elongated compared to a radial depth of the recess. 
         [0007]    In a further embodiment of any of the above, the recess axially overlaps the bumper. 
         [0008]    In a further embodiment of any of the above, the assembly includes a turbine rotor having a hub. The minidisk is mounted on the hub and the hub is operatively supported relative to an engine static structure by a bearing. 
         [0009]    In a further embodiment of any of the above, the assembly includes a seal assembly arranged between the minidisk and the bearing to create a bearing compartment. 
         [0010]    In a further embodiment of any of the above, the assembly includes a sleeve supported by the hub. The bearing is mounted to the sleeve. 
         [0011]    In a further embodiment of any of the above, the hub includes circumferentially spaced radially outwardly extending first tabs. The axial extension includes circumferentially spaced radially inwardly extending second tabs. The first and second tabs are axially aligned with one another and fingers of the sleeve are received in circumferential gaps provided between the first and second tabs to prevent relative circumferential movement between the first and second tabs. 
         [0012]    In a further embodiment of any of the above, the sleeve includes an outer diameter surface. A clearance is provided between the bumper and the outer diameter surface of 0.000-0.005 inch (0.000-0.127 mm). 
         [0013]    In a further embodiment of any of the above, the seal assembly is secured to the engine static structure. The seal assembly includes a seal support and a carbon seal axially slidable relative to the seal support. The flange extends axially beyond the seal support. 
         [0014]    In a further embodiment of any of the above, the flange provides forward and aft tool engagement faces configured to be accessible by a tool with at least a portion of the seal assembly mounted the engine static structure. 
         [0015]    In a further embodiment of any of the above, the turbine rotor provides a second turbine rotor. A first turbine rotor is secured to the second turbine rotor at a joint by an interference fit. The flange is configured to be manipulated by a tool to alter the interference fit at the joint. 
         [0016]    In another exemplary embodiment, a method of working on a gas turbine engine section includes inserting a tool into a cavity beneath a seal assembly, and engaging a flange of a minidisk with the tool to manipulate first and second rotors with respect to one another. 
         [0017]    In a further embodiment of any of the above, the method includes removing a portion of a bearing prior to performing the inserting step. 
         [0018]    In a further embodiment of any of the above, the method includes removing a seal land prior to the performing the inserting step, with portions of the seal assembly remaining mounted to an engine static structure during the engaging step. 
         [0019]    In a further embodiment of any of the above, the engaging step includes closing the tool radially inward to engage the flange. 
         [0020]    In a further embodiment of any of the above, the method includes the step of separating the first and second rotors at a joint having an interference fit. 
         [0021]    In a further embodiment of any of the above, the method includes the step of joining the first and second rotors at a joint in an interference fit. 
         [0022]    In a further embodiment of any of the above, the engaging step includes deflecting a bumper of the minidisk into engagement with a surface of one of the first and second rotors. 
         [0023]    In a further embodiment of any of the above, the first and second rotors are first and second turbine rotors. 
         [0024]    In a further embodiment of any of the above, the method includes a sleeve locking the minidisk to one of the first and second rotors. 
         [0025]    In a further embodiment of any of the above, the engaging step includes deflecting a bumper of the minidisk into engagement with a surface of an assembly tool. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]    The various features and advantages of the disclosed examples 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. 
           [0027]      FIG. 1  schematically illustrates a gas turbine engine embodiment. 
           [0028]      FIG. 2  is a schematic cross-sectional view through a turbine section. 
           [0029]      FIG. 3  is a partial cross-sectional view of the turbine section shown in  FIG. 2  in more detail. 
           [0030]      FIG. 4  is an enlarged view of a portion of the turbine section during a disassembly procedure. 
           [0031]      FIG. 5  is a cross-sectional perspective view of a minidisk, turbine hub and sleeve shown in  FIGS. 3 and 4 . 
           [0032]      FIG. 6  is a cross-sectional view taken along line  6 - 6  in of  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION 
       [0033]      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 . 
         [0034]    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 and other turbo machinery; 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. 
         [0035]    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 X 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. 
         [0036]    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 X. 
         [0037]    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. 
         [0038]    The example low pressure turbine  46  has a pressure ratio that is greater than about five (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. 
         [0039]    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 . 
         [0040]    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. 
         [0041]    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. 
         [0042]    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 and other turbo machinery. 
         [0043]    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. 
         [0044]    “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. 
         [0045]    “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. 
         [0046]    Referring to  FIG. 2 , a cross-sectional view through the turbine section  28  is illustrated. In the example turbine section  28 , first and second arrays  66 ,  68  of circumferentially spaced fixed vanes  70 ,  72  are axially spaced apart from one another. A first stage array  74  of circumferentially spaced turbine blades  76 , mounted to a first rotor disk  78 , is arranged axially between the first and second fixed vane arrays  70 ,  72 . A second stage array  80  of circumferentially spaced turbine blades  82 , mounted to a second rotor disk  94 , is arranged aft of the second array  68  of fixed vanes  72 . 
         [0047]    The turbine blades  76 ,  82  each include a tip  84 ,  86  adjacent to a blade outer air seals  88 ,  90  of a case structure  92 . The first and second stage arrays  66 ,  68  of turbine vanes and first and second stage arrays  74 ,  80  of turbine blades are arranged within a flow path F and are operatively connected to the shaft  32 , which is rotatable about an axis X. 
         [0048]    One of ordinary skill in the art will recognize that the teachings of disclosed arrangement may be used for either the high pressure turbine section  54  or the low pressure turbine section  46 . Moreover, one of ordinary skill will recognize that the teachings herein can be used wherever high press fits are used and may include other parts of the engine like the high pressure compressor section  52 , more turbine stages or other types of engines besides the gas turbine engine  20  shown herein. 
         [0049]    Referring to  FIG. 3 , the first and second rotor discs  78 ,  94  are secure to one another at a joint  96  using a very high press fit load. This press fit load must be overcome during assembly and disassembly procedures of the turbine section  28 , for example, during service. 
         [0050]    A minidisk  100  is mounted to the aft side of the second turbine rotor  94 , which is the last stage in the example turbine section  28 . The minidisk  100  provides a plate  101  that creates a cavity in the turbine section  28  that is cooled to lower turbine rotor temperatures. An axial extension  99  is provided by the minidisk  100  and is supported by hub  98  of the second turbine rotor  94 . The hub  98  operatively supports the bearing  38  to support the turbine section  28  for rotation relative to the engine static structure  36 . 
         [0051]    In the example, a sleeve  102  is provided radially between the hub  98  and an inner race  120  of the bearing  38 . An air seal assembly  106  is provided between the turbine section  28  and the bearing  38 , which is arranged within a bearing compartment that is separated from hot gases by the seal assembly  106 . Heat shields  108  are used to further insulate the bearing compartment and seal assembly  106  from hot gases. 
         [0052]    The example seal assembly  106  includes a seal support  110  mounted to the engine static structure  36 . Multiple circumferentially spaced pins  112  are secured to the seal support  110 . A carrier  114  having a carbon seal  116  is slidably supported by the pins  112  for axial movement. A seal land  118  is mounted to the sleeve  102  and arranged adjacent to the inner race  120 . The seal land  118  engages the carbon seal  116  during rotation of the seal land  118  with the second turbine rotor  94  and hub  98  to seal the bearing compartment from hot gases. A retainer  122  secures the hub  98 , minidisk  100  and sleeve  102  to one another in a stack to maintain assembly loads. 
         [0053]    Referring to  FIGS. 4 and 6 , the hub  98  includes circumferentially spaced first tabs  124  that extend radially outward. The axial extension  99  includes circumferentially spaced second tabs  126  that extend radially inward. During assembly, the second tabs  126  are slid through circumferential gaps  128  between the first tabs  124  as the minidisk  100  is mounted onto the hub  98 . The minidisk  100  and second turbine rotor  94  are then rotated relative to one another to position the second tabs  126  in alignment with and behind the first tabs  124 , as shown in  FIGS. 4 and 6 . The minidisk  100  is axially retained to the hub  98  in this assembled position. Fingers  130  of the sleeve  102  are inserted into the circumferential gaps  128  to maintain the circumferential position of the first and second tabs  124 ,  126  and lock the minidisk  100  to the second turbine rotor  94 . 
         [0054]    An outer diameter surface  132  of the sleeve  102  supports the axial extension  99  during disassembly of the turbine section  28 . A radius  134  adjoins the second tabs  126  and an inner diameter surface  136  of the axial extension  99 . A bumper  142  extends radially inward from and proud of the inner diameter surface  136 , which is cylindrical in shape in the example. In one example, a clearance between the outer diameter surface  132  and the bumper  142  is about 0.005 inch (0.127 mm). 
         [0055]    Referring to  FIGS. 4 and 5 , recess  138  is provided in the axial extension  99  on a surface radially opposite the inner diameter surface  136  to provide a flange  140 . The recess is elongated to reduce the size and weight of the axial extension  99 . However, this increases the flexibility of the axial extension during assembly and disassembly procedures. In the example, the recess  138  begins in an area that axially overlaps the second tabs  126 . Portions of the flange  140  and bumper  142  axially overlap one another. In the example, the axial length of the recess  138  is at least four times that of the radial depth of the recess  138 . The flange  140  is axially outboard of portions of the seal assembly  106 , such as the heat shield  108  and seal support  110 , such that the flange  140  is readily accessible. 
         [0056]    During assembly or disassembly, such as during a service procedure, the inner bearing race  120  and seal land  118  are removed exposing a cavity  146 , as shown in  FIG. 4 . The remaining portions of the bearing  38  and seal assembly  106 , including the carrier  114  and carbon seal  116 , can remain mounted to the engine static structure  36 . A tool  148 , similar to a collet arrangement, is inserted into the cavity  146  at position P 1 . Lips  150  of the tool  148  are deflected radially inward to position P 2  to engage the flange  140 . During assembly, the lips  150  engage an aft face  154  to apply a pushing load on the minidisk  100  through the flange  140 . During disassembly, the lips engage a forward face  152  to apply a pulling load on the flange  140 , which can be used to separate the first and second turbine rotors  78 ,  94  at the joint  96 . 
         [0057]    During assembly and disassembly procedures, in particular during disassembly, the axial extension  99  is deflected radially inward such that the bumper  142  contacts the outer diameter surface  132 . The bumper  142  and its tight clearance with respect to the outer diameter surface  132  prevents the axial extension from plastically deforming or breaking, which enables a smaller, lighter axial extension to be used. In this example, surface  132  is part of sleeve  102  but in another embodiment can be part of tool  148  (shown in the same position as the sleeve in  FIG. 4 ). 
         [0058]    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. The scope of legal protection given to this disclosure can only be determined by studying the following claims.