Patent Publication Number: US-11661855-B2

Title: Low density hybrid knife seal

Description:
BACKGROUND 
     The present disclosure is directed to the improved low density hybrid knife seal. 
     Knife edge seals present issues during rub events as heat generated between the rotor tip and the seal material dramatically increases the temperature of the seal and associated rotor tip. Attempts to reduce the heat generated during rub events have produced abradable seals with lower density materials. Also, the abradable seal includes the use of precut trenches to allow for better heat transfer. The precut trenches along with the lower density materials have created a low elastic moduli in the cross section of the seal material. During operation, engine backpressure can deflect the abraded seal material causing the material to make further contact with the rotor. 
     What is needed is a structure that prevents the low density abradable seal materials from deflecting during engine operation. 
     SUMMARY 
     In accordance with the present disclosure, there is provided a hybrid abradable seal comprising a stator substrate having an external surface; a casing coupled to the external surface, the casing including radial walls extending radially from the external surface; an abradable material disposed within the casing; the abradable material and the casing being coupled together and configured to resist a deflection responsive to engine gas loads. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing includes a floor directly coupled to the exterior surface and coupled to the abradable material. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the abradable material comprises a silicone material with imbedded hollow microspheres. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the abradable material comprises a density of from about 0.5 to about 0.65 grams/cubic centimeter. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the hybrid abradable seal further comprising a mechanical fastener configured to attach at least one of the casing and the abradable material to the stator substrate. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing comprises radial walls extending radially from the exterior surface and configured to contain the abradable material within the radial walls, relative to an axis A. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing and the abradable material comprise a ratio of elastic modulus of casing to abradable of about 50-5000 X. 
     In accordance with the present disclosure, there is provided a hybrid abradable seal for a gas turbine engine rotor and stator comprising a stator substrate having an external surface; a casing coupled to the external surface, the casing including radial walls extending radially from the external surface; an abradable material disposed within the casing radial walls relative to an axis A; the abradable material and the casing being coupled together and configured to resist a deflection responsive to gas turbine engine gas loads. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing includes a floor directly coupled to the exterior surface and coupled to the abradable material. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing comprises a material selected from the group consisting of polyether ketone, polyether ether ketone, polyetherimide, polyamide imide, polyphenylene sulfide or polyphenylsulfone and a reinforced thermoset organic matrix composite such as an epoxy or imide-based resin reinforced with carbon or glass fibers or fabric. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the reinforced thermoset organic matrix composite is selected from the group consisting of an epoxy or imide-based resin reinforced with at least one of a carbon fiber, a glass fiber and a fabric. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing comprises neat or reinforced thermoplastic. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the abradable material comprises a silicone material with imbedded hollow microspheres to form an abradable material density of from about 0.5 to about 0.65 grams/cubic centimeter. 
     In accordance with the present disclosure, there is provided a gas turbine engine abradable seal deflection reduction process comprising providing a stator substrate having an external surface; coupling a casing to the external surface, the casing including radial walls extending radially from the external surface; disposing an abradable material within the casing radial walls relative to an axis A; coupling the abradable material and the casing together being configured to resist a deflection responsive to gas turbine engine gas loads. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing includes a floor directly coupled to the exterior surface; and coupling the floor to the abradable material. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the abradable material comprises a silicone material with imbedded hollow microspheres to form an abradable material density of from about 0.5 to about 0.65 grams/cubic centimeter. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing and the abradable material comprise a ratio of elastic modulus of casing to abradable of about 50-5000 X. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the process further comprising fabricating the abradable seal in-situ with low modulus abradable material loaded and cured into the casing following installation of the abradable seal. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include the casing comprises neat or reinforced thermoplastic. 
     A further embodiment of any of the foregoing embodiments may additionally and/or alternatively include process further comprising forming channels into the abradable material. 
     Other details of the low density hybrid knife seal are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a schematic longitudinal sectional view of a turbofan engine. 
         FIG.  2    is a cross sectional schematic of an exemplary abradable seal. 
         FIG.  3    is a cross sectional schematic of an exemplary abradable seal. 
     
    
    
     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  may include a single-stage fan  42  having a plurality of fan blades  43 . The fan blades  43  may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan  42  drives air along a bypass flow path B in a bypass duct  13  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 . A splitter  29  aft of the fan  42  divides the air between the bypass flow path B and the core flow path C. The housing  15  may surround the fan  42  to establish an outer diameter of the bypass duct  13 . The splitter  29  may establish an inner diameter of the bypass duct  13 . 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 the 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 inner shaft  40  may interconnect the low pressure compressor  44  and low pressure turbine  46  such that the low pressure compressor  44  and low pressure turbine  46  are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine  46  drives both the fan  42  and low pressure compressor  44  through the geared architecture  48  such that the fan  42  and low pressure compressor  44  are rotatable at a common speed. Although this application discloses geared architecture  48 , its teaching may benefit direct drive engines having no geared architecture. 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 the 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. 
     Airflow in the core flow 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 through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core flow 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 low pressure compressor  44 , high pressure compressor  52 , high pressure turbine  54  and low pressure turbine  46  each include one or more stages having a row of rotatable airfoils. Each stage may include a row of static vanes adjacent the rotatable airfoils. The rotatable airfoils and vanes are schematically indicated at  47  and  49 . 
     The engine  20  may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to 10.0 and less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture  48  may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan  42 . A gear reduction ratio may be greater than or equal to 2.3, or more narrowly greater than or equal to 3.0, and in some embodiments the gear reduction ratio is greater than or equal to 3.4. The gear reduction ratio may be less than or equal to 4.0. The fan diameter is significantly larger than that of the low pressure compressor  44 . The low pressure turbine  46  can have a pressure ratio that is greater than or equal to 8.0 and in some embodiments is greater than or equal to 10.0. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. Low pressure turbine  46  pressure ratio is pressure measured prior to an 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. 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. All of these parameters are measured at the cruise condition described below. 
     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. The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade  43  alone, without a Fan Exit Guide Vane (“FEGV”) system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct  13  at an axial position corresponding to a leading edge of the splitter  29  relative to the engine central longitudinal axis A. The low fan pressure ratio is a span-wise average of the pressure ratios measured across the fan blade  43  alone over radial positions corresponding to the distance. The low fan pressure ratio can be less than or equal to 1.45, or more narrowly greater than or equal to 1.25, such as between 1.30 and 1.40. “Low corrected fan tip speed” is the actual fan tip speed in feet/second divided by an industry standard temperature correction of [(Tram ° R) / (518.7° R)] 0.5 . The “low corrected fan tip speed” can be less than or equal to 1150.0 feet/second (350.5 meters/second), and greater than or equal to 1000.0 feet/second (304.8 meters/second). 
     Referring also to  FIG.  2    and  FIG.  3   , a seal  60  is shown. The seal  60  can be an abradable seal  60  proximate a knife edge rotor  62 . The rotor  62  rotates around axis A. The abradable seal  60  is configured to interact with the rotor  62  and provide a sealing function. The knife edge rotor  62  can interact with the abradable seal  60  and can cause wear to the abradable seal  60 . 
     The abradable seal  60  includes a stator substrate  64 . The substrate  64  has an exterior surface  66 . A casing  68  can be disposed on the exterior surface  66 . The casing  68  can be a material that includes a stiffness that resists deflection. In an exemplary embodiment, the casing  68  material can include a thermoplastic. In another exemplary embodiment, the casing  68  can include neat or reinforced thermoplastic such as but not limited to polyether ketone, polyether ether ketone, polyetherimide, polyamide imide, polyphenylene sulfide or polyphenylsulfone or a reinforced thermoset organic matrix composite such as an epoxy or imide-based resin reinforced with carbon or glass fibers or fabric. The casing  68  can include an optional floor  69 , as shown in  FIG.  2   . The casing  68  can include radial walls  70 . The radial walls  70  can be range from about  0 . 050  to about  0 . 25  inches thick. This thickness provides the necessary support to withstand the deflection induced by the internal air flow forces  72 . 
     A low density abradable material  74  can be disposed within the casing  68 . The abradable material  74  can be partially disposed within the casing  68  and axially contained within the radial walls  70 , relative to axis A. The abradable material  74  can be exposed to the rotor  62  with no casing  68  between the rotor  62  and abradable material  74 . The abradable material  74  can be disposed directly onto the exterior surface  66 , in an exemplary embodiment, when there is no casing floor  69 , as shown in  FIG.  3   . The abradable material  74  can include low elastic moduli. The abradable material  74  can comprise a lightweight abradable with imbedded hollow carbon or glass microspheres creating a material with a density of from about 0.5 to about 0.65 grams/cubic centimeter. In an exemplary embodiment, the casing  68  can relate to the abradable material  74  with a ratio of elastic modulus of casing to abradable of about 50-5000 X. 
     In an exemplary embodiment, the abradable material  74  consisting of silicone  76  filled to a high volume percentage of hollow microspheres  78  creates an overall low density material which has shown to significantly reduce the evolved temperature during rub events. Higher concentrations of microspheres  78  perform the best as elimination of compressible elastomer reduces heat generation. The high volume or high concentration of microspheres can be the range of 50-70% microspheres by volume. 
     In an exemplary embodiment, the abradable seal  60  can include channels  80  cut into the abradable material  74 . The depth D and width W of the channels  80  can be predetermined dimensions. In an exemplary embodiment, the width D can range from 0.15-0.30 inches; the depth D can range from 0.125-0.50 inches. The channels  80  can be cut to prevent any rub of the seal  60  which prevents heat build-up. 
     The use of the low density abradable material  74  can avoid the use of pre-cut channels  80 . As the seal  60  is abraded, trenches will be formed by the rotor  62  however these trenches will be much smaller than pre-cut channels  80 , as the trenches are a perfect fit for the rotor  62 . Pre-cut channels  80  have to take up all the build tolerances and so they are quite a bit wider. This causes inefficiencies in the engine. The low density material  74  could also have pre-cut channels but overall may not be necessary, as the abradable material  74  can accommodate in-situ rotor  62  trenching. 
     With the extreme low density of the abradable material  68 , channeled seals may not withstand engine gas loads  72  causing the abradable seal  60  to deflect axially into a rotor tip (not shown) causing additional rub. The combination of the radial walls  70  and the abradable material  74  coupled together provide a load  72  resistant structure. In an exemplary embodiment the casing walls  70  need not be as high as the contained abradable material  74 , as shown in  FIG.  3   , and thus radially extend along a portion of the abradable material  74 . 
     In an exemplary embodiment, at least one of the abradable material  74  and the casing  68  can be attached to the stator substrate  64  via an imbedded mechanical fastener  82 . The abradable seal  60  can be fabricated in-situ with low modulus material  74  loaded and cured into the casing  68  following installation of the seal  60 . 
     A technical advantage of the disclosed abradable seal includes the hybrid seal uses a thermoplastic casing which provides necessary stiffness to resist deflection and prevent any additional rub. 
     A technical advantage of the disclosed abradable seal includes the elastomeric nature of the material which provides toughness to the bonded element preventing brittle crack growth due to thermal expansion mismatch. 
     A technical advantage of the disclosed abradable seal includes a hybrid design which provides a way to reduce seal weight and to permit rub in seals which previously were pre-trenched to eliminate the issue of overheating. 
     A technical advantage of the disclosed abradable seal includes a rub system that gains efficiency over current pre-trenched designs. 
     A technical advantage of the disclosed abradable seal includes the walled casing being configured to bond the low density filler to the casing. The adhesive bond creates additional benefit from deflection. 
     There has been provided a low density hybrid knife seal. While the low density hybrid knife seal has been described in the context of specific embodiments thereof, other unforeseen alternatives, modifications, and variations may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations which fall within the broad scope of the appended claims.