Abstract:
A gear assembly support for a gas turbine engine includes a first portion engageable to a case of the gas turbine engine and a second portion configured for supporting a gear assembly. The support includes a torque reacting portion for transferring torque from the second portion to the first portion, a forward flange disposed forward of the torque reacting portion, the forward flange defining a first interface to the case and an aft flange disposed aft of the torque reacting portion, the aft flange defining a second interface to the case.

Description:
REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/843,418 filed on Jul. 7, 2013. 
     
    
     BACKGROUND 
       [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 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. 
         [0003]    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 so as to increase the overall propulsive efficiency of the engine. 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 such that both the turbine section and the fan section can rotate at closer to optimal speeds. 
         [0004]    The gear assembly is attached to a static structure through a flexible support. The flexible support orientates the gear assembly within the engine and also accommodates generated torque during operation. The support function includes a desired fit with static structure and is balanced against the torque transfer function. The configuration of the flexible support is therefore balanced against the desire to reduce cost and weight along with the separate functions. Accordingly, engine manufacturers continue to seek improvements in the support structure that balance the functional requirements against cost and weight. 
       SUMMARY 
       [0005]    A gear assembly support for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a first portion engageable to a case of the gas turbine engine. A second portion is configured for supporting a gear assembly. A torque reacting portion transfers torque from the second portion to the first portion. A forward flange is disposed forward of the torque reacting portion. The forward flange defines a first interface to the case. An aft flange is disposed aft of the torque reacting portion. The aft flange defines a second interface to the case. The forward flange and the aft flange include separately adjustable features for modifying the first interface and the second interface. 
         [0006]    In a further embodiment of the foregoing gear assembly support, the forward flange includes a first diameter disposed forward of first portion the aft flange includes a second diameter aft of the first portion. The separately adjustable features include a first thickness between a forward undercut and the first diameter and a second thickness between an aft undercut and the second diameter. 
         [0007]    In a further embodiment of any of the foregoing gear assembly supports, the first portion includes a spline engageable with features defined within the case of the engine. 
         [0008]    In a further embodiment of any of the foregoing gear assembly supports, the forward undercut and the aft undercut are continuous about a circumference of the support. 
         [0009]    In a further embodiment of any of the foregoing gear assembly supports, the torque reacting portion includes an axial width between the first portion and the second portion for defining torque transmitted to the case of the engine. 
         [0010]    In a further embodiment of any of the foregoing gear assembly supports, includes a flex portion disposed between the second portion and the torque reacting portion. 
         [0011]    In a further embodiment of any of the foregoing gear assembly supports, the forward flange includes an annular lip extending radially outward that fits within the case. 
         [0012]    In a further embodiment of any of the foregoing gear assembly supports, includes a plurality of flanges extending axially forward of the first portion. 
         [0013]    A turbofan engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan including a plurality of fan blades rotatable about an engine axis, a turbine section;, a geared architecture driven by the turbine section for rotating the fan about the engine axis, and a support member that supports the geared architecture. The support member includes a first portion engageable to a case of the turbofan engine, a second portion configured for supporting a gear assembly, a torque reacting portion for transferring torque from the second portion to the first portion, a forward flange disposed forward of the torque reacting portion, the forward flange defining a first interface to the case, and an aft flange disposed aft of the torque reacting portion. The aft flange defines a second interface to the case. The forward flange and the aft flange include separately adjustable features for adjusting the first interface and the second interface. 
         [0014]    In a further embodiment of the foregoing turbofan engine, the forward flange includes a first diameter disposed forward of first portion the aft flange includes a second diameter aft of the first portion. The separately adjustable features include a first thickness between a forward undercut and the first diameter and a second thickness between an aft undercut and the second diameter. 
         [0015]    In a further embodiment of any of the foregoing turbofan engines, the first portion includes a spline engageable with features defined within the case of the engine. 
         [0016]    In a further embodiment of any of the foregoing turbofan engines, the forward undercut and the aft undercut are continuous about a circumference of the support. 
         [0017]    In a further embodiment of any of the foregoing turbofan engines, the torque reacting portion includes an axial width between the first portion and the second portion for defining torque transmitted to the case of the engine. 
         [0018]    In a further embodiment of any of the foregoing turbofan engines, includes a flex portion disposed between the second portion and the torque reacting portion. 
         [0019]    In a further embodiment of any of the foregoing turbofan engines, the forward flange includes an annular lip extending radially outward that fits within the case. 
         [0020]    A method of supporting a gear assembly within a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes defining a forward interface between a support member and a static structure within the gas turbine engine with a flange portion, defining an aft interface between the support member and the static structure, defining a torque transfer path between a first portion between the forward interface and the aft interface and a second portion configured for attachment to the gear assembly independent of the flange portion, attaching the gear assembly to the second portion, and attaching the support member to the static structure such that the flange portion positions the support member relative to the static structure. 
         [0021]    In a further embodiment of the foregoing method, defining the fit between the support member and the static structure includes defining a thickness between at least one of a first diameter in the forward interface and a first undercut and a second diameter and a second undercut. 
         [0022]    In a further embodiment of any of the foregoing methods, includes defining a plurality of undercuts, spacing the plurality of undercuts circumferentially apart and defining the torque transfer path at least partially between the plurality of undercuts. 
         [0023]    In a further embodiment of any of the foregoing methods, defining the torque transfer path includes defining an axial thickness of the torque transfer path. 
         [0024]    Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0025]    These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]      FIG. 1  is a schematic view of an example gas turbine engine. 
           [0027]      FIG. 2  is a sectional view of a portion of an example flex support for a geared architecture. 
           [0028]      FIG. 3  is a cross-sectional view through a flange portion of the example flex support. 
           [0029]      FIG. 4  is a front view of the example flex support. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      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 . 
         [0031]    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. 
         [0032]    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. 
         [0033]    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. 
         [0034]    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. 
         [0035]    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. 
         [0036]    A mid-turbine frame  58  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  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
         [0037]    Airflow through the core airflow path 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  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  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  58 . 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. 
         [0038]    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. 
         [0039]    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. 
         [0040]    A significant amount of thrust is provided by airflow through the bypass flow path 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. 
         [0041]    “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. 
         [0042]    “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) 0.5 ]. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
         [0043]    The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about  26  fan blades. In another non-limiting embodiment, the fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about three (3) turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
         [0044]    A support member referred to in this disclosure as a flex support  62  is provided to mount the geared architecture  48  to a static structure or case  64  of the gas turbine engine  20 . The flex support  62  supports the geared architecture  48  in a manner that provides flexibility to compensate for operational torque. 
         [0045]    Referring to  FIG. 2  with continued reference to  FIG. 1 , the example flex support  62  is fit within the case  64 . The case  64  includes an inner surface  66  that includes a spline  68 . The flex support  62  includes a first portion or mating spline  70  that engages the spline  68  of the case  64 . 
         [0046]    The flex support  62  includes a forward flange  72  disposed forward of the spline  70  and an aft flange  74  aft of the spline  70 . The forward and aft flanges  72  and  74  define an interference fit sometimes referred to as a snap or snap fit. The forward flange  72  defines a first interface with the case  64  and the aft flange  74  defines a second interface with the case  64 . The forward and aft flanges  72 ,  74  define interference fits between the flex support  62  and the case  64 . The interference fit of the forward and aft flanges  72 ,  74  along with the spline  70  orientate the flex support  62  relative to the case  64 . Flanges  104  extend forward and are attached to a portion of the case  64  to further secure the flex support  62  within the case  64 . Orientation of the flex support  62  relative to the case  64  provides an orientation of the geared architecture  48  relative to the fixed case structure  64 . 
         [0047]    The flex support  62  grounds torque generated by the geared architecture  48  to the fixed case structure  64 . The geared architecture  48  is attached to a thin-walled, contoured flexible portion  78 . A thicker contoured portion is a torque reacting portion also referred to as a torque portion  76  transfers torque from the flex portion  78  radially outward through the spline  70  to the case  64 . 
         [0048]    Referring to  FIGS. 3 and 4  with continued reference to  FIGS. 2 , the flex support  62  grounds torque in a direction indicated by arrow  102  between the geared architecture  48  and the case  64 . The torque encountered by the flex support  62  is communicated through the interface between the spline  70  of the flex support  62  and the spline  68  of the case  64 . 
         [0049]    The flex support  62  provides the desired interference fit with the case  64  by providing a desired interference fit at the forward and aft flanges  72 ,  74 . The forward flange  72  is partially defined by a first diameter  80  and the aft flange  74  is partially defined by a second diameter  82 . The forward flange  72  is further defined by a forward thickness  87  of a forward tab  96 . The aft flange  74  is further defined by an aft thickness  85  of the aft tab  96 . 
         [0050]    The forward flange  72  includes a first thickness  88  having a forward undercut  84 . The aft flange  74  similarly includes an aft undercut  86  having a second-aft thickness  90 . Each of the forward and aft undercuts  84 ,  86  are continuous annular channels or grooves that extend along the inner diameter of the outer rim  106  of the flex support  62 . 
         [0051]    The first interference fit defined by the forward flange  72  is tailored by setting the first diameter  80  of the tab  96  and the forward thickness  87  of the forward tab  96  to provide the desired fit. The second interference fit is defined by the aft flange  74  by sizing the diameter  82  and the aft thickness  85  of the aft tab  95 . The first thickness  88  of the forward undercut  84  along with an axial width  92  is adjusted to tailor the moment loading aspect of the flex support  62 . 
         [0052]    The second interference fit defined by the aft flange  74  is disposed at the second diameter  82  and defined by the aft thickness  85  of the aft tab  95 . The size of the aft undercut  86  is formed by providing a desired thickness  90  and axial width  94 . The size of the aft undercut  86  is utilized to tailor moment loading of the flex support  62 . 
         [0053]    The fit of the flex support  62  to the case  64  is tailored by defining the forward and aft thicknesses  87  and  85  of the forward and aft tabs  96 ,  95  along with the first and second outer diameters  80  and  82 . Further adjustments to the flex support relating to moment loading are made by adjusting the size of the forward and aft undercuts  84  and  86 . 
         [0054]    Combinations of the aft thickness  85  of the aft tab  95  and the second diameter  82  provide for tailoring of the aft flange  74  independent of the forward flange  72  and of the torque portion  76 . 
         [0055]    In the illustrated example, the forward undercut  84  and the aft undercut  86  are similar in size and thickness, however, it is within the contemplation of this disclosure, that each of the forward and aft undercuts  84 ,  86  may be much different to separately and independently tailor the forward and aft flanges  72 ,  74  to meet application specific requirements. The first flange  72  is also adjustable independent and separate from aft flange  74  and torque portion  76 . 
         [0056]    Torque  102  is transferred between an inner flange  108  of the flex portion  78  to the outer rim  106  and the spline  70  through the torque portion  76 . It should be understood, that the torque indicated by arrow  100  could be in an opposite direction and remain within the contemplation of this disclosure. The transmission of the torque  102  proceeds through the flex portion  78  and the torque portion  76  through the spline  70  to the spline  68  to be grounded to the case  64 . 
         [0057]    The example flex support  62  separates torque transmission from the forward and aft flange  72 , 74  such that each of the forward and aft flange  72 ,  74  and the torque portion  76  can be individually sized to accommodate application specific fitting requirements. Adjustment or tuning of the torque portion  76 , to modify its vibration and flexure responsiveness, is provided by varying an axial thickness of material within the torque portion between the flex portion  78  and the outer rim  106 . The torque portion  76  is a full annulus of material and transfers torque between flex portion  78  to the outer rim  106 . 
         [0058]    The example torque portion  76  includes a radially outer width  98  and a radially inner width  100 . Both of the radially outer width  98  and the radially inner width  100  are in the axial direction with relative to the engine axis A. In this example the overall axial width decreases in a direction radially outward from the inner width  100  toward the outer width  98 . This provides for the inner width  100  to be greater than the outer width  98 . The decrease in width from the inner width  100  to the outer width  98  is a taper that provides a smooth transition between the inner and outer widths  100 ,  98 . 
         [0059]    Accordingly the example flex support  62  provides for the support of the geared architecture  48  and also for the separate adjustments of forward and aft flange fits  72 ,  74  between the case  64  that is independent of the torque path provided by the torque portion  76 . 
         [0060]    The separation of the flange fits from the torque portion  76  provides for the independent tailoring of material thicknesses through the flex support  62  to adjust the interference fit with the case  64  separately and independent of the material required to define a desired torque transmission path through the torque portion  76 . The independent adjustment provides for a lighter and more economically robust flex support  62 . 
         [0061]    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 this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.