Patent Publication Number: US-11391163-B1

Title: Vane arc segment with seal

Description:
BACKGROUND 
     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 may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines. 
     Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils. 
     SUMMARY 
     A vane arc segment according to an example of the present disclosure includes an airfoil fairing that has an airfoil wall defining first and second fairing platforms and a hollow airfoil section extending there between. A spar has a spar leg that extends through the hollow airfoil section. The spar leg has an end portion that protrudes from the hollow airfoil section. The spar leg is spaced from the airfoil wall in the hollow airfoil section such that there is a first gap there between. A support platform is adjacent the second fairing platform such that there is a second gap there between. The support platform is secured with the end portion of the spar leg. A baffle is disposed in the first gap. The baffle is spaced apart from the airfoil wall and the spar leg so as to divide the first gap into a plenum space between the spar leg and the baffle and an impingement space between the baffle and the airfoil wall. The baffle has impingement holes directed toward the airfoil wall and connects the plenum space with the impingement space. A seal is disposed between the airfoil wall and the spar leg. The seal seals the impingement space from the second gap. 
     In a further embodiment of any of the foregoing embodiments, the seal is a rope seal. 
     In a further embodiment of any of the foregoing embodiments, the seal is radially offset from the baffle. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes a scallop and the seal is seated against the scallop. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes a protrusion that has a weld land at which the baffle is welded thereto, and the seal is radially offset from the protrusion. 
     In a further embodiment of any of the foregoing embodiments, the seal seats against the baffle. 
     In a further embodiment of any of the foregoing embodiments, the support platform includes a radially upstanding lip against which the seal seats. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip is adjacent the baffle and includes at least one through-hole connecting the plenum space and the second gap. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip includes a shank portion and a cup portion, the seal seating in the cup portion. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip includes a shank portion and a band portion, and the band portion wraps around the seal. 
     In a further embodiment of any of the foregoing embodiments, the hollow airfoil section includes first and second cavities, the spar leg extends through the first cavity, and there is an additional baffle disposed in the second cavity, with an additional seal disposed between the airfoil wall and the additional baffle. The additional seal seals the second cavity from the second gap. 
     A gas turbine engine according to an example of the present disclosure includes compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has vane arc segments disposed about a central axis of the gas turbine engine. Each of the vane arc segments includes an airfoil fairing has an airfoil wall defining first and second fairing platforms and a hollow airfoil section that extends there between. A spar has a spar leg that extends through the hollow airfoil section. The spar leg has an end portion that protrudes from the hollow airfoil section. The spar leg is spaced from the airfoil wall in the hollow airfoil section such that there is a first gap there between. A support platform is adjacent the second fairing platform such that there is a second gap there between. The support platform is secured with the end portion of the spar leg. A baffle is disposed in the first gap. The baffle is spaced apart from the airfoil wall and the spar leg so as to divide the first gap into a plenum space between the spar leg and the baffle and an impingement space between the baffle and the airfoil wall. The baffle has impingement holes directed toward the airfoil wall and connect the plenum space with the impingement space. A seal is disposed between the airfoil wall and the spar leg. The seal seals the impingement space from the second gap. 
     In a further embodiment of any of the foregoing embodiments, the seal is a rope seal and is radially offset from the baffle. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes a scallop and the seal is seated against the scallop. 
     In a further embodiment of any of the foregoing embodiments, the spar leg includes a protrusion that has a weld land at which the baffle is welded thereto, and the seal is radially offset from the protrusion. 
     In a further embodiment of any of the foregoing embodiments, the seal seats against the baffle. 
     In a further embodiment of any of the foregoing embodiments, the support platform includes a radially upstanding lip against which the seal seats. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip is adjacent the baffle and includes at least one through-hole connecting the plenum space and the second gap. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip includes a shank portion and a cup portion, the seal seating in the cup portion. 
     In a further embodiment of any of the foregoing embodiments, the radially upstanding lip includes a shank portion and a band portion. The band portion wraps around the seal. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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 description can be briefly described as follows. 
         FIG. 1  illustrates a gas turbine engine. 
         FIG. 2  illustrates a vane arc segment from the engine. 
         FIG. 3  illustrates a seal between a spar leg and an airfoil fairing of the vane arc segment. 
         FIG. 4  illustrates another example in which the seal seats against a baffle adjacent the spar leg and is supported by a radially upstanding lip. 
         FIG. 5  illustrates another example in which a radially upstanding lip has one or more through-holes for cooling. 
         FIG. 6  illustrates another example in which there is a separable radially upstanding lip. 
         FIG. 7  illustrates another example in which the separable radially upstanding lip has a cup portion. 
         FIG. 8A  illustrates another example in which the separable radially upstanding lip has a band portion. 
         FIG. 8B  illustrates another example in which the band portions are at spaced intervals along the seal. 
         FIG. 9  illustrates a support platform that includes a radially upstanding lip. 
     
    
    
     DETAILED DESCRIPTION 
       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 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a housing  15  such as a fan case or nacelle, and also 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. 
     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. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects, 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 a 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  may be 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. 
     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 through 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 the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  may be positioned forward or aft of the location of gear system  48 . 
     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 and less than about 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. 
     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 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 (350.5 meters/second). 
       FIG. 2  illustrates a line representation of an example of a vane arc segment  60  from the turbine section  28  of the engine  20  (see also  FIG. 1 ). It is to be understood that although the examples herein are discussed in context of a vane from the turbine section, the examples can be applied to vanes in other portions of the engine  20 . 
     The vane arc segment  60  includes an airfoil fairing  62  that is formed by an airfoil wall  63 . The airfoil fairing  62  is comprised of a hollow airfoil section  64  and first and second platforms  66 / 68  between which the airfoil section  64  extends. The airfoil section  64  generally extends in a radial direction relative to the central engine axis A. Terms such as “inner” and “outer” used herein refer to location with respect to the central engine axis A, i.e., radially inner or radially outer. Moreover, the terminology “first” and “second” used herein is to differentiate that there are two architecturally distinct components or features. It is to be further understood that the terms “first” and “second” are interchangeable in that a first component or feature could alternatively be termed as the second component or feature, and vice versa. 
     The airfoil wall  63  is continuous in that the platforms  66 / 68  and airfoil section  64  constitute a unitary body. As an example, the airfoil wall  63  is formed of a ceramic matrix composite, an organic matrix composite (OMC), or a metal matrix composite (MMC) or homogeneous polymer, metallic or ceramic material. For instance, the ceramic matrix composite (CMC) is formed of ceramic fiber tows that are disposed in a ceramic matrix. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fiber tows are disposed within a SiC matrix. Example organic matrix composites include, but are not limited to, glass fiber tows, carbon fiber tows, and/or aramid fiber tows disposed in a polymer matrix, such as epoxy. Example metal matrix composites include, but are not limited to, boron carbide fiber tows and/or alumina fiber tows disposed in a metal matrix, such as aluminum. A fiber tow is a bundle of filaments. As an example, a single tow may have several thousand filaments. The tows may be arranged in a fiber architecture, which refers to an ordered arrangement of the tows relative to one another, such as, but not limited to, a 2D woven ply or a 3D structure. 
     The airfoil section  64  circumscribes an interior cavity  70 , which in this example is subdivided by a rib  70   a  into first and second sub-cavities  71   a / 71   b . Alternatively, the airfoil section  64  may have a single cavity  70 , or the cavity  70  may be divided by additional ribs. The vane arc segment  60  further includes a spar  72  that mechanically supports the airfoil fairing  62 . The spar  72  includes a spar platform  72   a  and a spar leg  72   b  that extends from the spar platform  72   a  into the cavity  70  (through the first sub-cavity  71   a ). Although not shown, the radially outer side of the spar platform  72   a  may include attachment features that secure it to a fixed support structure, such as an engine case. The spar leg  72   b  defines an interior through-passage  72   c.    
     The end of the spar leg  72   b  extends past the platform  68  so as to protrude from the fairing  62 . There is a support platform  74  adjacent the platform  68  of the airfoil fairing  62 . The support platform  74  includes a first through-hole  74   a  through which the end of the spar leg  72   b  extends. In this example, the end of the spar leg  72   b  includes a clevis mount  76 , although other mounting schemes could alternatively be used. The clevis mount  76  may include one or more prongs that protrude from the support platform  74 . The prong or prongs include a pinhole through which a pin  76   a  extends. The pin  76   a  is wider than the through-hole  74   a  of the support platform  74 . The ends of the pin  76   a  thus abut the face of the support platform  74  and thereby prevent the spar leg  72   b  from being retracted from the through-hole  74   a . The pin  76   a  thus locks the support platform  74  to the spar leg  72   b  such that the airfoil fairing  62  is mechanically trapped between the spar platform  72   a  and the support platform  74 . It is to be appreciated that the example configuration could be used at the outer end of the airfoil fairing  62 , with the spar  72  being inverted such that the spar platform  72   a  is adjacent the (inner) platform  68  and the support platform  74  is adjacent the (outer) platform  66 . The spar  72  may be formed of a relatively high temperature resistance, high strength material, such as a single crystal metal alloy (e.g., a single crystal nickel- or cobalt-alloy). 
     The spar leg  72   b  is spaced from the airfoil wall  63  such that there is a first gap  78  there between. The walls of the spar leg  72   b  are solid and continuous. There is a baffle  80  disposed in the gap  78 . The baffle  80  generally circumscribes the spar leg  72   b . The baffle  80  is spaced apart from the airfoil wall  63  and the spar leg  72   b  so as to divide the gap  78  into a plenum space  78   a  between the spar leg  72   b  and the baffle  80  and an impingement space  78   b  between the baffle  80  and the airfoil wall  63 . The baffle  80  has impingement holes (represented at unnumbered flow arrows) that are directed toward the airfoil wall  63  and connect the plenum space  78   a  and the impingement space  78   b . The baffle  80  is formed of sheet metal but may alternatively be formed from an alloy using additive manufacturing. 
     The baffle  80  may not extend entirely through the airfoil section  62  to the support platform  74 . Rather, the end of the baffle  80  is joined to the spar leg  72   b  prior to the clevis mount  76 . In this regard, the impingement holes in the baffle  80  may be the exclusive exit from the plenum space  78   a  into the impingement space  78   b.    
     Cooling air, such as bleed air from the compressor section  24 , is conveyed into and through the through-passage  72   c . This cooling air is destined for a downstream cooling location, such as a tangential onboard injector (TOBI). Cooling air is also conveyed into the plenum space  78   a . The cooling air in the plenum space  78   a  is emitted through the impingement holes in the baffle  80  onto the airfoil wall  63  for cooling thereof. 
     In the illustrated example, there is an additional, second baffle  82  that extends through the second sub-cavity  71   b . The baffle  82  is also provided with cooling air and may have cooling holes therein for directing the cooling air at portions of the airfoil wall  63 . In this example, like the spar leg  72   b , the baffle  82  protrudes from the airfoil fairing  62  and through a second through-hole  74   b  in the support platform  74 . 
     The support platform  74  is radially spaced from the platform  68  of the airfoil fairing  62  such that there is a second gap  84  there between. There is a first seal  86  disposed between the airfoil wall  63  and the spar leg  72   b . The seal  86  seals the impingement space  78   b  from the second gap  84  such that cooling air in the impingement space  78   b  cannot escape into the second gap  84 . There is a second seal  88  disposed between the airfoil wall  63  and the baffle  82 . The seal  88  seals the space in the second sub-cavity  71   b  between the baffle  82  and the airfoil wall  63  from the second gap  84  such that cooling air in the second sub-cavity  71   b  cannot escape into the second gap  84 . Alternatively, if the second gap  84  is pressurized, the seals  86 / 88  prevent flow from the second gap  84  into the sub-cavities  71   a / 71   b.    
     In a further example, the seals  86 / 88  are rope seals. For example, the rope seals are formed of fibers, such as ceramic fibers, metallic fibers, graphite fibers, or polymer fibers. The fibers may be braided, knitted, or woven. Example ceramic fibers include, but are not limited to, oxide fibers. For instance, the ceramic fibers are NEXTEL fibers, which are composed of Al 2 O 3 , SiO 2 , and B 2 O 3 . Example metallic fibers include, but are not limited to, nickel alloy or a cobalt alloy fibers. Example polymer fibers include, but are not limited to, meta-aramid or para-aramid fibers. For instance, the polymer fibers are NOMEX fibers, which are composed of m-phenylenediamine isophthalamide. Optionally, the rope seals may include a sheath surrounding a fiber core. The sheath can be an overbraid or foil that surrounds the core. In one example, the sheath comprises a high-temperature metallic material, such as a single crystal nickel alloy or a cobalt alloy. For instance, in the overbraid example, the sheath comprises an overbraid of metallic wire. In other examples, the sheath comprises a ceramic-based material. 
       FIG. 3  illustrates a further example in which the support platform  174  and spar leg  172   b  are configured to facilitate positioning of the seal  86 . In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements. Here, the spar leg  172   b  includes a protrusion  90  that extends around the periphery of the spar leg  172   b . The protrusion has a weld land  90   a  at which the end of the baffle  80  is welded to. The spar leg  172   b  further includes a scallop  91  that is radially offset from the protrusion  90 . The scallop  91  is a sloped portion of the spar leg  172   b  that curves outwards. The curvature of the scallop  91  provides a seat against which the seal  86  is positioned (the seal  86  is thus radially offset from the protrusion  90 ). The curvature cradles the seal  86  and thereby serves to keep the seal  86  from shifting radially inwards (toward the support platform  174 ). The protrusion  90  serves to keep the seal from shifting radially outwards. 
     The support platform  174  in the illustrated example also includes a radially upstanding lip  92 . The lip  92  extends around the periphery of the first through-hole  74   a  of the support platform  174 . The lip  92  includes a radially-facing surface  92   a  against which the seal  86  seats. The lip  92  further prevents the seal  86  from shifting radially (inwards in this example). 
     In the example of  FIG. 3 , the seal  86  seals against the airfoil wall  63  and the spar leg  172   b .  FIG. 4  illustrates a modified configuration in which the baffle  80  extends radially farther (inwards) such that the seal  86  seals against the airfoil wall  63  and the surface of the baffle  80 . 
       FIG. 5  illustrates a further example that is the same as the example of  FIG. 4  except that the lip  92  includes one or more through-holes  92   b  that connect the plenum space  78   a  and the second gap  84 . Cooling air from the plenum space  78   a  is emitted through the cooling hole(s)  92   b  toward the fillet, platform  68 , and/or airfoil section  64  of the fairing  62 . As shown, the cooling hole(s)  92   b  may be sloped (relative to the engine axis A) such that the cooling air impinges on the fillet region or other regions which are susceptible to extreme temperatures. 
     In the example of  FIG. 4 , the radially upstanding lip  92  is integral with the support platform  174 .  FIG. 6  illustrates an additional example in which the support platform  74  has a separable radially upstanding lip  192 . For instance, the lip  192  is a separate piece from the body of the support platform  74 . In this regard, the support platform  74  includes a slot  93  into which a portion of the lip  192  is received. the slot  93  facilitates retaining and positioning the lip  192 . Like the lip  92 , the lip  192  includes a radially-facing surface  192   a  against which the seal  86  seats to prevent the seal  86  from shifting radially (inwards in this example). 
     The example in  FIG. 7  is the same as in  FIG. 6  except that here the separable radially upstanding lip  292  has a shank portion  292   b  and a cup portion  292   c . The shank portion  292   b  is received into the slot  93 . The cup portion  292   c  is located at the radial end of the shank portion  292   b . The cup portion  292   c  is curved and the seal  86  is seated against the curved surface of the cup portion  292   c  to retain and position the seal  86  in place. 
     The example in  FIG. 8A  is the similar to the example of  FIG. 7  except that here instead of a cup portion the separable radially upstanding lip  392  has a band portion  392   c . The shank portion  292   b  is received into the slot  93 . The band portion  392   c  is located at the radial end of the shank portion  292   b . The band portion  392   c  wraps around the seal  86  to retain and position the seal  86  in place. The band portion  392   c  may initially be “open” so as to enable the seal  86  to be received therein, and then subsequently bent to wrap around the seal  86 . 
     The shank portion  292   b  and band portion  392   c  may be coextensive with the seal  86  or provided at intervals along the seal  86 . For example, as shown in  FIG. 8B , the shank portion  292   b  and band portion  392   c  are provided at spaced intervals along the seal  86 . The portion of the seal  86  between the band portions  392   c  is not directly supported by the shank portion  292   b  and the band portion  392   c , however, the intervals may be relatively close together to facilitate the elimination of sagging of the seal  86  there between. 
     As mentioned above, the seal  88  seals the space in the second sub-cavity  71   b  between the baffle  82  and the airfoil wall  63  from the second gap  84  such that cooling air in the second sub-cavity  71   b  cannot escape into the second gap  84 . As shown in  FIG. 9 , the support platform  174  may include an additional radially upstanding lip  92  around the periphery of the second through-hole  74   b  of the support platform  174  to support the seal  88 . Additionally, similar to the spar leg  72   b , the baffle  82  may have a scallop to further facilitate positioning. 
     As indicated, the seals  86 / 88  facilitate sealing from the second gap  84 . In particular, sealing against composite materials that form the airfoil fairing  62  is challenging because such composites may have higher surface roughness in comparison to traditional metallic alloy surfaces that are machined. Rope seals, which are flexible and conform to surface contours, facilitate sealing against such surfaces but must be maintained in proper position. In this regard, the seals  86 / 88  are trapped between the airfoil fairing  62 , the support platform  74 / 174  and the respective spar leg  72   b  or baffle  82 . One or more spring members may be provided between the platform  66  of the airfoil fairing  62  and the spar platform  72   a  to bias the airfoil fairing towards the support platform  74 / 174 . Such a biasing facilitates providing a constant clamping force of the airfoil fairing  62  against the seals  86 / 88  to thereby further maintain position of the seals  86 / 88 . 
     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. 
     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 this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.