Abstract:
A seal for use with a static guide vane of a gas turbine engine has a strap member extending to a generally cylindrical bulb. The bulb forms a diameter, and is spaced a distance defined between an end of the strap remote from the bulb, to a point on the bulb furthest from the end. A ratio of the diameter to the distance is between 0.2 and 0.5. A static guide vane and a gas turbine engine are also disclosed.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/708,306, which was filed Oct. 1, 2012. 
     
    
     BACKGROUND 
       [0002]    This application relates to a seal to seal a gap between adjacent platform edges of guide vanes for use in a gas turbine engine. 
         [0003]    Gas turbine engines are known, and typically include a fan delivering air into a compressor section. The air 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. The turbine rotors drive compressor and fan rotors. Historically, a common turbine rotor has driven a compressor rotor and the fan. 
         [0004]    More recently, a gear reduction has been placed to drive the fan, such that the fan and compressor rotors can rotate at different speeds. This has allowed the size of the fan to increase dramatically. The fan typically delivers air into the compressor, but also delivers a portion of air as bypass flow into a bypass duct. With the increase in fan size, a bypass ratio, or the ratio of the air delivered into the bypass duct compared to the air delivered into the compressor has increased. 
         [0005]    With the increase in bypass flow, the air delivered into the compressor becomes more valuable, and all air delivered into the compressor is desirably utilized efficiently. Thus, it becomes more important to limit leakage. 
         [0006]    Within the compressor and fan sections, there are guide vanes positioned intermediate blade rows on the fan or compressor rotors. These guide vanes have platform edges which extend circumferentially towards an adjacent guide vane. There is often a gap between these edges. 
         [0007]    Seals have been proposed to cover this gap, and prevent air from leaking inwardly or outwardly of the guide vanes. The seals to date have required relatively complex structure formed into the platforms, and also require complex assembly techniques. 
       SUMMARY 
       [0008]    In a featured embodiment, a seal for use with a static guide vane of a gas turbine engine has a strap member extending to a generally cylindrical bulb. The generally cylindrical bulb forms to a diameter, and is spaced a distance defined between an end of the strap remote from the bulb, to a point on the bulb furthest from the end. A ratio of the diameter to the distance is between 0.2 and 0.5. 
         [0009]    In another embodiment according to the previous embodiment, the bulb includes a hollow opening. 
         [0010]    In another embodiment according to any of the previous embodiments, the seal is formed of a silicone rubber. 
         [0011]    In another featured embodiment, a static guide vane for use in a gas turbine engine has an airfoil extending between two radial platforms, with a suction side and a pressure side to both the airfoil and platforms. A seal is secured to a side of each of the platforms remote from the airfoil. The seal has a strap secured to the side of the platforms. An enlarged bulb sits outwardly of an edge of the platform. 
         [0012]    In another embodiment according to the previous embodiment, a vane is designed for use in a compressor section of a gas turbine engine. 
         [0013]    In another embodiment according to any of the previous embodiments, there is a seal on only one of the suction and pressure sides at each of the two platforms. 
         [0014]    In another embodiment according to any of the previous embodiments, the bulb has a diameter, and a distance defined between an end of the strap remote from the bulb to a point on the bulb furthest spaced from the distance. A ratio of the diameter to the distance is between 0.2 and 0.5. 
         [0015]    In another embodiment according to any of the previous embodiments, the seal is formed of a silicone rubber. 
         [0016]    In another featured embodiment, a gas turbine engine has a fan, a compressor, a combustor, and a turbine. One of the fan, compressor and turbine is provided with a row of static guide vanes, which have an airfoil extending between two radial platforms, with a suction side and a pressure side to both the airfoil and platforms. There are circumferentially adjacent static guide vanes, with a seal secured to a side of each of the platforms remote from the airfoil on one of the circumferentially adjacent vanes. No seal is secured to the other. The seal has a strap secured to the side of the platforms. An enlarged bulb sits outwardly of an edge of the platform and in engagement with the other of the static guide vanes. 
         [0017]    In another embodiment according to the previous embodiment, the bulb has a diameter, and a distance defined between an end of the strap remote from the bulb to a point on the bulb furthest spaced from the end. A ratio of the diameter to the distance is between 0.2 and 0.5. 
         [0018]    In another embodiment according to any of the previous embodiments, the is in the compressor. 
         [0019]    In another embodiment according to any of the previous embodiments, one of the circumferentially adjacent guide vanes is the same one at both of the platforms. 
         [0020]    In another embodiment according to any of the previous embodiments, the seal is formed of a silicone rubber. 
         [0021]    These and other features may be best understood from the following specification and drawings, the following which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]      FIG. 1  schematically shows a gas turbine engine. 
           [0023]      FIG. 2  shows a detail of a guide vane. 
           [0024]      FIG. 3A  shows one side of a guide vane. 
           [0025]      FIG. 3B  shows an opposed of the guide vane. 
           [0026]      FIG. 4  shows adjacent guide vanes and seals. 
           [0027]      FIG. 5  shows a detail of a seal. 
       
    
    
     DETAILED DESCRIPTION 
       [0028]      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 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. 
         [0029]    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. 
         [0030]    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  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. 
         [0031]    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. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
         [0032]    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. 
         [0033]    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 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. 
         [0034]      FIG. 2  shows a guide vane  80  which is illustrated as a static compressor guide vane. It should be understood that guide vanes utilized in a fan section or even the turbine section may also benefit from teachings of this application. In guide vane  80 , an airfoil  81  extends away from an inner platform  78 . As shown at  88 , an edge of the platform  78  curves relative to the center line A of the engine such as shown in  FIG. 1 . 
         [0035]    A seal  82  has a strap portion  86  secured to an underside  79  of the platform  78 . 
         [0036]    A bulb  84  is positioned outward of the edge  88 . The bulb  84  is shown to have a hole  110 . The seal  82  may be an extrusion, and may be formed of any elastomer that may be appropriate for the environmental conditions that the seal  82  will see during use. Silicone rubber may be used. 
         [0037]    As shown in  FIG. 3A , the vane  80  has a radially inner platform  102  and a radially outer platform  101  on one side. As shown in  FIG. 3B , there is also the platform  78  at an inner end, and the platform  100  at a radially outer end of an opposed side. A seal  82  is placed on the  FIG. 3B  side of the platform  78 , and a seal  188  at the radially outer platform  100 . It should be understood that the platforms  100  and  101  are actually a single platform, as are the platforms  78  and  102  and each wrap around leading and trailing edges of the airfoils  81 . 
         [0038]    The combination of  FIGS. 3A and 3B  simply show that the seals  82  and  188  are placed on only one side of a vane  80 . It should be understood that the seals at the radially inner and outer ends could be placed on opposed sides of the airfoil  81 . That is, a seal could be placed on the pressure side of the airfoil  81  at, say, the radially outer location, and on the suction side of the airfoil  81  at a radially inner location. However, as shown in combination, there is typically a seal at only one side of the vane  80  at each of the radially inner and outer locations. 
         [0039]      FIG. 4  shows the seal  82  having the bulb  84  secured to platform  78 , but also abutting an edge of the platform side  102 . Thus, the seal  82  provides an effective seal at the radially inner edge. Similarly, the seal  188  is shown bonded to the platform  112 , and abutting the platform  101 . Again, the seal  188  will provide an effective seal at the radially outer location. It would also be possible to place a seal on all four platforms of a guide vane  80 , and simply not have any on adjacent vanes  80 . 
         [0040]      FIG. 5  shows a detail of the seal  82 , but would this would also be true of the seal  188 . A diameter D of the bulb  84  is defined, and in one embodiment was 0.125 inch (0.3175 cm). A length d from an outermost point  200  of the bulb to an innermost location  201  of the strap  86  is defined. In one embodiment, d was 0.4375 inch (1.111 cm). In embodiments, a ratio of D to d was between 0.2 and 0.5. 
         [0041]    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.