Abstract:
A gas turbine engine includes a core engine, a fan coupled to be driven by the core engine, a first case structure around the core engine and a second case structure around the fan. The first case structure and the second case structure define a bypass passage there between. A vane structure includes an airfoil extending radially between the first case structure and the second case structure. A seal is configured to control leakage between the bypass passage and a core engine.

Description:
CROSS REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/705,772, which was filed 26 Sep. 2012 and is incorporated herein by reference. 
     
    
     BACKGROUND OF THE INVENTION 
       [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]    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. 
         [0004]    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. 
         [0005]    Turbine engine manufacturers continue to seek further improvements to engine performance and assembly including improvements to thermal, transfer, assembly and propulsive efficiencies. 
       SUMMARY 
       [0006]    A gas turbine engine according to an exemplary aspect of the present disclosure includes a core engine, a fan coupled to be driven by the core engine, a first case structure around the core engine, and a second case structure around the fan. The first case structure and the second case structure define a bypass passage there between. A vane structure includes an airfoil which extends radially between the first case structure and the second case structure, and a seal configured to control fluid leakage between the bypass passage and the core engine. 
         [0007]    In a further non-limiting embodiment of any of the foregoing examples, the seal is located at a radially inner side of the vane structure. 
         [0008]    In a further non-limiting embodiment of any of the foregoing examples, the seal is hollow. 
         [0009]    In a further non-limiting embodiment of any of the foregoing examples, the seal is flexible. 
         [0010]    In a further non-limiting embodiment of any of the foregoing examples, the seal is elastomeric. 
         [0011]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes reinforced elastomer. 
         [0012]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes a rounded, hollow bulb. 
         [0013]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes an attachment section which secures the seal to a panel that is located inwards of the vane structure. 
         [0014]    In a further non-limiting embodiment of any of the foregoing examples, the attachment section includes a pair of spaced apart legs. 
         [0015]    In a further non-limiting embodiment of any of the foregoing examples, the seal is within a gas passage between the core engine and the bypass passage, and further comprising wear member within the gas passage inwards of the seal. 
         [0016]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes a seal flap bearing against the vane structure. 
         [0017]    A gas turbine engine according to an exemplary aspect of the present disclosure includes a vane structure which includes an airfoil extending outwardly therefrom, a panel which includes a section arranged inwards of the vane structure and a seal arranged inwards of the vane structure, between the vane structure and the panel. 
         [0018]    In a further non-limiting embodiment of any of the foregoing examples, the panel bounds a fan bypass passage. 
         [0019]    In a further non-limiting embodiment of any of the foregoing examples, the seal is hollow. 
         [0020]    In a further non-limiting embodiment of any of the foregoing examples, the seal is flexible. 
         [0021]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes a rounded, hollow bulb. 
         [0022]    In a further non-limiting embodiment of any of the foregoing examples, the seal is within a gas passage between the core engine and the bypass passage, and further comprising a wear member within the gas passage inwards of the seal. 
         [0023]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes an attachment section securing the seal to a panel that is located inwards of the vane structure, the attachment section including a pair of spaced apart legs. 
         [0024]    In a further non-limiting embodiment of any of the foregoing examples, the seal includes a seal flap bearing against the vane structure. 
         [0025]    A method for use with a gas turbine engine according to an exemplary aspect of the present disclosure includes controlling fluid leakage between the bypass passage and the core engine using a seal. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]    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 descriptions can be briefly described as follows. 
           [0027]      FIG. 1  illustrates an example gas turbine engine. 
           [0028]      FIG. 2  illustrates a vane structure and a seal configured to control fluid leakage between a bypass passage and a core engine. 
           [0029]      FIG. 3  illustrates another example of a seal configured to control leakage between a bypass passage and a core engine. 
       
    
    
     DESCRIPTION 
       [0030]      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 . In this example, the compressor section  24 , the combustor section  26  and the turbine section  28  are sections of a core engine C. Alternative engines might include an augmentor 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 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. 
         [0031]    The engine  20  generally includes a first spool  30  and a second spool  32  mounted for rotation about an engine central 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. 
         [0032]    The first spool  30  generally includes a first shaft  40  that interconnects a fan  42 , a first compressor  44  and a first turbine  46 . The first shaft  40  is connected to the fan  42  through a gear assembly of a fan drive gear system  48  to drive the fan  42  at a lower speed than the first spool  30 . The second spool  32  includes a second shaft  50  that interconnects a second compressor  52  and second turbine  54 . The first spool  30  runs at a relatively lower pressure than the second spool  32 . It is to be understood that “low pressure” and “high pressure” or variations thereof as used herein are relative terms indicating that the high pressure is greater than the low pressure. An annular combustor  56  is arranged between the second compressor  52  and the second turbine  54 . The first shaft  40  and the second shaft  50  are concentric and rotate via bearing systems  38  about the engine central axis A which is collinear with their longitudinal axes. 
         [0033]    The core airflow is compressed by the first compressor  44  then the second compressor  52 , mixed and burned with fuel in the annular combustor  56 , then expanded over the second turbine  54  and first turbine  46 . The first turbine  46  and the second turbine  54  rotationally drive, respectively, the first spool  30  and the second spool  32  in response to the expansion. 
         [0034]    The engine  20  is a high-bypass geared aircraft engine that has a bypass ratio that is greater than about six (6), with an example embodiment being greater than ten (10), the gear assembly of the fan drive gear system  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 2.3:1 and the first turbine  46  has a pressure ratio that is greater than about 5. The first turbine  46  pressure ratio is pressure measured prior to inlet of first turbine  46  as related to the pressure at the outlet of the first turbine  46  prior to an exhaust nozzle. The first turbine  46  has a maximum rotor diameter and the fan  42  has a fan diameter such that a ratio of the maximum rotor diameter divided by the fan diameter is less than 0.6. It should be understood, however, that the above parameters are only exemplary. 
         [0035]    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. 
         [0036]    Referring also to  FIG. 2 , the engine  20  can include structural guide vanes  60  that extend between a first, outer case structure or panel  62  and a second, inner case structure or panel  64 . Each structural guide vane  60  includes an airfoil AF and a base  66  secured with the airfoil AF. If there are gaps G between the end of the inner panel  64 , such as at a forward end, and the structural guide vanes  60  (at base  66 ), the gaps G can debit the aerodynamic performance of the engine  20 . In this example, the panel  64  is rigidly secured at an aft end (not shown) to a bracket  68 , such as by a bolt or other fastener. However, the forward ends E/E′ of the panel  64  and bracket  68 , respectively, are free. The forward end E′ of the bracket  68  includes a wear member  70  that is non-rigidly received into another bracket  72  such that loads exerted on the brackets  68 / 72  are decoupled from each other. The wear member  70  provides a low-friction interface with the bracket  72  and facilitates damping vibrations. The size of the gap G can vary due to thermal growth of the components, dimensional tolerances or both. Similar gaps may also exist elsewhere in the engine  20 . 
         [0037]    A seal  74  is provided to prevent or limit air from traveling through the gap G between the inner panel  64  and the structural guide vanes  60 . For example, the seal  74  is flexible and includes a rounded bulb shaped portion  76 , but could alternatively be a flap seal. In this example, the rounded bulb shaped portion  76  is hollow and can have open ends such that the rounded bulb shaped portion  76  compresses against the base  66 , as indicated in phantom at  76 ′. In a further example, the seal  74  can be fabricated of an elastomeric material or a reinforced elastomeric material. The reinforced elastomeric material can be reinforced with a fabric or fiber material that is impregnated with a rubber material, such as silicone. The reinforcement provides stiffening and enhanced durability. 
         [0038]    In this example, the seal  74  includes an attachment section  78  that is configured to secure the seal  74  to the panel  64 . The attachment section  78  includes a pair of spaced-apart legs  80 / 82 , between which a section  64   a  of the panel  64  is received. The seal  74  can be press-fit onto the section  64   a.  Optionally, an adhesive can additionally be used to secure the seal  74 . 
         [0039]    The section  64   a  of the panel  64  slopes inwards towards the central axis A (shown schematically) of the engine  20 . In this example, the sloping of the section  64   a  locates the bulb shaped portion  76  of the seal  74  inwards of the base  66  of the structural guide vane  60 , such that the bulb shaped portion  76  bears against, and thus seals, the underside of the base  66 . 
         [0040]    The seal  74 , along with the radially inwardly-located wear member  70  on the bracket  68 , are arranged in a gas passage P between the core engine C and the bypass flowpath B. The seal  74  thus controls gas flow through the gas flow passage P. It is to be understood that although the flow through the gas passage P is shown in one direction, the flow can be reversed depending upon the design and operating pressures. 
         [0041]      FIG. 3  shows another example seal  174  that can alternatively be used to prevent or limit air from traveling through the gap G between the panel  64  and the structural guide vanes  60 . In this example, the seal  174  includes a flap  176  rather than the rounded bulb shaped portion  76  as shown in  FIG. 2 . The flap  176  is flexible and can be fabricated of elastomeric or reinforced elastomer material, as described. 
         [0042]    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. 
         [0043]    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.