Abstract:
A gas turbine engine comprises a fan, a compressor section, a turbine section, and a gear reduction for driving the fan through the turbine section. A rotating element and at least one bearing compartment includes a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment, and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. The seal has a plurality of sealing members extending radially toward a sealing surface.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Patent Application No. 62/010,486, filed Jun. 11, 2014. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    This application relates to a geared turbofan having unique seals at at least one bearing compartment. 
         [0003]    Gas turbine engines are known, and typically include a fan rotor delivering air into a bypass duct as propulsion air. Air is also delivered into a compressor as core airflow. The air in the compressor is compressed and delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
         [0004]    Historically, the fan rotor rotated at a single speed with a fan drive turbine. This limited the speed of the fan drive turbine, as the fan rotor speed was constrained by a number of factors. More recently, it has been proposed to include a gear reduction between the fan rotor and the fan drive turbine. 
         [0005]    There are a number of bearing compartments in a geared gas turbine engine. It is important to seal each bearing compartment by having a seal at each axial end. The seals ensure that oil does not leak outwardly of the bearing compartment and may receive a supply of pressurized air at an opposed side to resist the flow of oil across the seal. 
         [0006]    While brush seals and labyrinth seals have been proposed at a number of locations in direct drive gas turbine engines, they have not been proposed in a geared gas turbine engine. 
       SUMMARY OF THE INVENTION 
       [0007]    In a featured embodiment, a gas turbine engine comprises a fan, a compressor section, a turbine section, and a gear reduction for driving the fan through the turbine section. A rotating element and at least one bearing compartment includes a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment, and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. The seal has a plurality of sealing members extending radially toward a sealing surface. 
         [0008]    In another embodiment according to the previous embodiment, the seal is a labyrinth seal having a plurality of knife edges. 
         [0009]    In another embodiment according to any of the previous embodiments, a first radius is defined to a radial extent of the knife edges and a second radius may be defined on a drive shaft associated with the fan drive turbine at a location in a plane defined by a fuel nozzle in a combustor in the gas turbine engine. A diameter ratio of the first radius to the second radius is less than or equal to about 2.0. 
         [0010]    In another embodiment according to any of the previous embodiments, the diameter radius is less than or equal to about 1.75. 
         [0011]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with the gear reduction. 
         [0012]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with the fan. 
         [0013]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with a compressor rotor. 
         [0014]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with a turbine rotor in the turbine section. 
         [0015]    In another embodiment according to any of the previous embodiments, the seal is a brush seal. 
         [0016]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with the gear reduction. 
         [0017]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with the fan. 
         [0018]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with a compressor rotor. 
         [0019]    In another embodiment according to any of the previous embodiments, the bearing compartment is associated with a turbine rotor in the turbine section. 
         [0020]    In another embodiment according to any of the previous embodiments, the gear reduction has a gear ratio greater than or equal to about 2.6. 
         [0021]    In another embodiment according to any of the previous embodiments, the fan delivers air into a bypass duct as propulsion air and into the compressor section as core air. A bypass ratio of the bypass air to the core air is greater than or equal to about 6.0. 
         [0022]    In another embodiment according to any of the previous embodiments, the bypass ratio is greater than or equal to about 10.0. 
         [0023]    In another embodiment according to any of the previous embodiments, the bypass air is greater than or equal to about 12.0. 
         [0024]    In another embodiment according to any of the previous embodiments, the fan delivers air into a bypass duct as propulsion air and into the compressor section as core air. A bypass ratio of the bypass air to the core air is greater than or equal to about 6.0. 
         [0025]    In another embodiment according to any of the previous embodiments, the bypass ratio is greater than or equal to about 10.0. 
         [0026]    In another embodiment according to any of the previous embodiments, the bypass air is greater than or equal to about 12.0. 
         [0027]    These and other features may be best understood from the following drawings and specification. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]      FIG. 1  schematically shows a gas turbine engine. 
           [0029]      FIG. 2  shows bearing compartment locations as may be found on the geared gas turbine engine of  FIG. 1 . 
           [0030]      FIG. 3  shows a first type of seal. 
           [0031]      FIG. 4  shows a location of a seal. 
           [0032]      FIG. 5  shows a second type of seal. 
           [0033]      FIG. 6  shows details of an embodiment. 
       
    
    
     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 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 should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
         [0035]    The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
         [0036]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  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]    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 fan drive gear system  48  may 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 . 
         [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 about 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 five. 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 five 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 2.3:1. 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 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 (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf 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 (350.5 meters/second). 
         [0040]    In such a geared gas turbine engine, there are more bearing compartments than there were found in the direct drive gas turbine engine. In addition, the bearing compartments, particularly as associated with a gear reduction, become critical. It is important to ensure that oil is maintained in the bearing compartments 
         [0041]    As shown in  FIG. 2 , several bearing compartments  100  associated with a gas turbine engine, such as the gas turbine engine  20  illustrated in  FIG. 1 , include seals. A bearing compartment  102  is associated with a low speed shaft  92  at a location associated with the low pressure turbine. Bearings  106  are shown schematically as is a seal  104 . 
         [0042]    A bearing compartment  108  is associated with the high speed rotor  90  and the high pressure turbine of  FIG. 1 . Bearing compartment  108  includes seals  110  at each axial end and a central bearing  112 . 
         [0043]    A second bearing compartment  114  is also associated with the high speed rotor  90  and the high pressure compressor and includes a bearing  118  and seals  116 . 
         [0044]    Finally, a third bearing compartment  120 / 123  is associated with a fan drive gear system  122 , or the gear reduction of  FIG. 1 . The third bearing  120 / 123  compartment is also associated with a fan bearing  130 , forward of the fan drive gear system  122 . Seals  126  and  128  mechanically seal axial ends of the bearing compartment  120 / 123  and are associated with a fan rotor  127  and the low speed rotor  92 . Seals  126  and  128  are also respectively associated with bearings  124  and  130  that are positioned within the bearing compartment  120 / 123 . 
         [0045]    The locations of the seals and the bearing compartments, as mentioned above, are exemplary and this disclosure extends to any number of other bearing component locations. 
         [0046]    In the past, particular types of seals have been provided in a geared gas turbine engine. Contact seals have been utilized and complex non-contact seals have been proposed. While these seals may operate efficiently, they are prone to wear and must be repaired or replaced periodically. Replacing these seals may require shut down of the engine, which is undesirable. 
         [0047]    Thus, a labyrinth seal  80 , such as shown in  FIG. 3 , may be utilized. In a labyrinth seal, a base  82  has knife edges  84 . The  FIG. 3  embodiment has the knife edges  84  associated with a static component. That is, base  82  may be fixed to housing structure 
         [0048]      FIG. 4  shows an embodiment  90  where the knife edges  96  are associated with a shaft  94 , which is positioned inwardly and facing a static structure  92 . It should be understood that this disclosure extends to labyrinth seals  90  which rotate ( FIG. 4 ) or are associated with the static structure ( FIG. 3 ). 
         [0049]    A wear surface  99  is positioned to face the knife edges  96  as shown in  FIG. 4 . In some applications, it may be ensured that there is a gap between the radial extent of the knife edges and wear surfaces  99 , such that there is no wear. However, it is also known to include an abradable material at surface  99 . As shown schematically, lubricant L from a portion  101  of the bearing chamber may tend to flow outwardly of the bearing chamber portion  101 . The knife edges  96  resist this flow. A supply of pressurized air P is supplied to a chamber  98  to further assist in resisting this lubricant flow, as would be understood by one of ordinary skill. 
         [0050]    Labyrinth seals provide benefits, particularly, when utilized in a geared gas turbine engine. 
         [0051]    In embodiments, there are at least two knife edges associated with the seal. The knife edges may have different diameters. 
         [0052]      FIG. 5  shows an alternative seal  140  which may be a brush seal. In a brush seal  140 , a ring  142  secures a plurality of brush bristles  144 . These brush bristles provide a seal much like the knife edges  96 , as would be appreciated by one of ordinary skill. 
         [0053]    Speaking generically, the illustrated seal  80  is a seal member having a plurality of distinct sealing members  84  extending towards a facing surface. 
         [0054]      FIG. 6  shows an engine  200  having a rotating shaft  202 . A labyrinth seal  210  may be associated with a bearing compartment #?. A location  208  of the shaft  202  may be defined as being in a plane of a fuel nozzle  206  of a combustor  204 . A radius R 1  may be defined to the outer tip of the knife edges at labyrinth seal  210 . A second radius R 2  is defined at portion  208 . 
         [0055]    In embodiments, R 1  may be less than or equal to about twice R 2 . Further, R 1  may be less than or equal to about one and three quarters (1.75) R 2 . In the prior art, labyrinth seals have typically been much larger. 
         [0056]    A gas turbine engine incorporating seals, such as disclosed in this application, may be provided in an engine with a bypass ratio greater than or equal to about 12. A gear ratio for gear reduction  122  may be greater than or equal to about 2.6. 
         [0057]    Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.