Patent Abstract:
A unitary one-piece hub has first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members. The windows include a lip configured to surround the cross members. A gas turbine engine and a method of providing a hub for a gas turbine engine are also disclosed.

Full Description:
CROSS REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/835,871 filed Jun. 17, 2013. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0002]    This invention was made with government support under Contract No. N00019-02-C-3003, awarded by the United States Navy. The Government has certain rights in this invention. 
     
    
     BACKGROUND 
       [0003]    This disclosure is related to a hub for a gas turbine engine, particularly a one-piece cast or forged hub. 
         [0004]    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. 
         [0005]    Gas turbine engines may include various hubs such as turbine exhaust cases, mid-turbine frames, transition or intermediate ducts, stator sections, or engine mounts. Hub assemblies may include inner and outer portions with airfoils or struts arranged in between the two portions. Current hubs are typically cast or forged in multiple pieces which then must be assembled, increasing cost and processing time. 
         [0006]    Engine manufacturers continue to develop methods to ease engine manufacture and assembly, and improve engine efficiency. 
       SUMMARY 
       [0007]    A unitary one-piece hub according to an exemplary embodiment of this disclosure, among other possible things includes first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members, and the windows each include a lip configured to surround the cross members. 
         [0008]    In a further embodiment of the foregoing hub, at least one the first and second rings include a stiffening element. 
         [0009]    In a further embodiment of any of the foregoing hubs, the stiffening element is located on one of a radially inner surface and a radially outward surface of the first ring. 
         [0010]    In a further embodiment of any of the foregoing hubs, the stiffening element is a third ring. 
         [0011]    In a further embodiment of any of the foregoing hubs, the plurality of cross members are airfoils. 
         [0012]    In a further embodiment of any of the foregoing hubs, the plurality of cross members are struts. 
         [0013]    In a further embodiment of any of the foregoing hubs, the plurality of cross members are welded to the lips. 
         [0014]    In a further embodiment of any of the foregoing hubs, the lips are disposed on one of a radially inward side and a radially outward side of the midsection. 
         [0015]    In a further embodiment of any of the foregoing hubs, the first and second rings and the midsection are cylindrical or conical in shape. 
         [0016]    In a further embodiment of any of the foregoing hubs, the hub is formed by a casting process. 
         [0017]    In a further embodiment of any of the foregoing hubs, the hub is formed by a forging process. 
         [0018]    A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a turbine, an exhaust arranged downstream from the turbine, and a case surrounding the turbine and exhaust. The case includes an inner case and an outer case wherein at least one of the inner and outer cases includes first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members, and the windows each include a lip configured to surround the cross members. 
         [0019]    In a further embodiment of the foregoing gas turbine engine, the inner case includes the includes first and second rings and the midsection arranged between the first and second rings, the midsection including the plurality of windows configured to receive the plurality of cross members, and the windows each include a lip configured to surround the cross members. 
         [0020]    In a further embodiment of any of the foregoing gas turbine engines, at least one of the first and second rings include a stiffening element. 
         [0021]    In a further embodiment of any of the foregoing gas turbine engines, the plurality of cross members are airfoils. 
         [0022]    In a further embodiment of any of the foregoing gas turbine engines, the plurality of cross members are welded to the lips. 
         [0023]    A method of providing a hub for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes the step of casting a first hub as one piece, the hub including first and second rings and a midsection arranged between the first and second rings. The midsection includes a plurality of windows configured to receive a plurality of cross members, and the windows each include a plurality of lips, respectively, configured to surround the cross members. 
         [0024]    In a further embodiment of the foregoing method of providing a hub for a gas turbine engine, the method further includes the step of attaching the plurality of cross members to the plurality of lips. 
         [0025]    In a further embodiment of any of the foregoing methods of providing a hub for a gas turbine engine, the attaching step comprises welding the plurality of cross members to the plurality of lips. 
         [0026]    In a further embodiment of any of the foregoing methods of providing a hub for a gas turbine engine, the method further includes the steps of providing a second hub and installing the first hub into the second hub. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0028]      FIG. 1  schematically illustrates a gas turbine engine. 
           [0029]      FIG. 2  schematically illustrates one-piece cast hub. 
           [0030]      FIG. 3  schematically illustrates a cutaway view of the hub of  FIG. 2 . 
           [0031]      FIG. 4  schematically illustrates an alternate cutaway view of the hub of  FIG. 2 . 
           [0032]      FIG. 5  schematically illustrates a detail cutaway view of the hub of  FIGS. 2-4 . 
           [0033]      FIG. 6  schematically illustrates a hub assembly. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      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 augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath B while the compressor section  24  drives air along a core flowpath C 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 including three-spool architectures. 
         [0035]    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. 
         [0036]    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  is connected to the fan  42  through a geared architecture  48  (shown schematically) 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 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 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0037]    Airflow through the core airflow path C 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. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the previously mentioned expansion. 
         [0038]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other 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 5. In one disclosed embodiment, the engine  20  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 5:1. 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 geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0039]    A significant amount of thrust is provided by the 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”). TSFC 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 [(Tambient deg 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. 
         [0040]    Referring to  FIG. 2 , an example hub  70  is schematically shown. The example hub  70  may be axially aligned on an engine axis A. The hub  70  in this example is an inner hub for a turbine  46 ,  54 , however, in another example the hub  70  may be an outer hub for a turbine  46 ,  54 , a mid-turbine frame  57 , a transition or intermediate duct  71  within the engine  20 , a stator section for the compressor  44 ,  52  or turbine  46 ,  54 , or an engine  20  mount. In another example, the hub  70  may be incorporated into any static component in the engine  20 . 
         [0041]    The hub  70  is a one-piece cast or forged structure. The hub  70  includes a forward ring  72  (a first ring), a midsection  74 , and an aft ring  76  (a second ring). The forward ring  72  may be cylindrical or conical in shape. One of the radially inner and the radially outer surfaces  78 ,  80  of the forward ring  72  provides a flowpath for upstream air entering the hub  70 . The other of the radially inner and outer surfaces 78 ,  80  of the forward ring  72  may include structural supports or stiffening elements. In another example, the forward ring  72  may be cantilevered off of the hub  70 . 
         [0042]    The aft ring  76  is similar to the forward ring  72 . One of the radially inner and outer surfaces  84 ,  86  of the aft ring  76  may include stiffening elements. For example, the stiffening element may be a cast or forged ring  87  on the radially inner side  86  of the aft ring  76 . The aft ring  76  may also be cantilevered off of the hub  70 . The other of the radially inner and outer surfaces  84 ,  86  may provide a flowpath for downstream air exiting the hub  70 . The aft ring  76  may include one or more flanges  82  for connecting to other parts of the engine  20 . The aft ring  76  may also include flanges (not shown) and openings  89 . The midsection  74  includes windows  88  to accommodate cross members  106 . 
         [0043]    Referring to  FIGS. 3-4 , a schematic cutaway view along the line  3 - 3  ( FIG. 2 ) of the hub  70  is shown. In the example shown in  FIG. 3 , the windows  88  are configured to receive cross members  106  such as airfoils (not shown). In the example shown in  FIG. 4 , the windows are configured to receive cross members  106  such as struts (not shown). 
         [0044]    As is shown in  FIG. 5 , the windows  88  may include a lip  90  extending radially inward from the window  88 . The lip  90  allows for attachment of the cross members  106  to the hub  70  by conventional fastening or bonding means. In one example, the cross members  106  may be welded to the lip  90 . In another example, the lip  90  may extend radially outward from the hub  70 . 
         [0045]      FIG. 6  shows a hub assembly  100 . The hub assembly  100  may include an outer member  102  and an inner member  104  with the cross members  106  arranged therebetween. One of the inner and outer members  102 ,  104  may be a hub  70  as described above. That is, if the hub  70  is the inner member  102 , the cross members  106  extend radially outward from the hub  70 . If the hub  70  is the outer member  104 , the cross member  106  extend radially inward from the hub  70 . 
         [0046]    Accordingly, casting or forging the hub  70  as one piece may provide a hub  70  with enhanced properties, such as improved directional uniformity. Additionally, the potential to employ a near-net casting process allows for limited machining after casting as one piece. A one-piece casting process may provide significant cost saving by eliminating the need for many complex fabrications and assemblies. 
         [0047]    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.

Technology Classification (CPC): 5