Patent Publication Number: US-11391172-B2

Title: Piloted retaining plate for a face seal arrangement

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application is a continuation of U.S. patent application Ser. No. 15/030,768 filed on Apr. 20, 2016. U.S. patent application Ser. No. 15/030,768 is a National Phase application of International Application No. PCT/US2014/060927 filed on Oct. 16, 2014. International Application No. PCT/US2014/060927 claims priority to U.S. Provisional Application No. 61/894,015 filed on Oct. 22, 2013. 
    
    
     BACKGROUND OF THE INVENTION 
     The present disclosure relates generally to a face seal for a gas turbine engine and more particularly to a retaining plate for the same. 
     High performance gas turbine engines include main shaft bearing compartments and employ seals to prevent cooling/lubricating oil from escaping from the bearing compartments into the rest of the engine. One type of seal commonly used for this function is a face seal. Face seals enable the engine, and bearing compartments within the engine, to function properly with minimal impact on a Thrust Specific Fuel Consumption (TSFC), the Thermal Management System (TMS), and the lubrication system of the gas turbine engine. 
     In some face seals, a component, such as a retaining plate, is utilized to maintain the sealing element in position within the face seal. Existing retaining plates include a large overlap between the retaining plate and a seal element in order to ensure that a full 360 degree arc of the seal element is covered by the retaining plate. The overlap between the retaining plate and the seal element is referred to as a “seal overlap”. When the gas turbine engine includes particularly small gapping and/or physical space and geometry requirements, a large seal overlap can disqualify the existing retaining ring based carbon seal designs. 
     SUMMARY OF THE INVENTION 
     A face seal assembly according to an exemplary embodiment of this disclosure, among other possible things includes a seal carrier defining an axis and having a pilot flange disposed circumferentially about the axis, a seal element disposed at least partially in the seal carrier, a retaining plate connected to one axial end of the seal carrier, the retaining plate having an axial step disposed circumferentially about the retaining plate, and the axial step is interfaced with the pilot flange. 
     In a further embodiment of the foregoing face seal, the seal element is a graphitic carbon face seal. 
     A further embodiment of the foregoing face seal includes a gap between the axial step and the pilot flange, such that the seal element is loose fit to the pilot flange. 
     In a further embodiment of the foregoing face seal, the retaining plate further includes a plurality of radially outward retainer flanges and the seal carrier comprises a plurality of radially outward carrier flanges, and each of the retainer flanges corresponds to one of the carrier flanges. 
     A further embodiment of the foregoing face seal includes a coil spring guide connecting each carrier flange to the corresponding retainer flange. 
     In a further embodiment of the foregoing face seal, the axial step includes a pilot surface facing radially inward and facing the pilot flange. 
     In a further embodiment of the foregoing face seal, the retaining plate further comprises a seal overlap, and the seal overlap radially covers the seal element on a full 360 degree arc. 
     In a further embodiment of the foregoing face seal, a radial length of the seal overlap is greater than a maximum possible radial misalignment of retaining plate relative to an axis defined by the seal element. 
     In a further embodiment of the foregoing face seal, the radial length of the seal overlap is approximately equal to the maximum possible radial misalignment of the retaining plate relative to the axis defined by the seal element. 
     A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible thing includes a compressor section connected to a first turbine spool, a combustor fluidly connected to the compressor section, a turbine section fluidly connected to the combustor and connected to a second turbine spool, an engine static structure connected to the first spool and the second spool via a plurality of bearing systems, at least one face seal assembly sealing at least one of the bearing systems, the at least one face seal assembly includes a seal carrier defining an axis and having a pilot flange disposed circumferentially about the axis, a seal element disposed at least partially in the seal carrier, a retaining plate connected to one axial end of the seal carrier, the retaining plate having an axial step disposed circumferentially about the retaining plate, and the axial step is interfaced with the pilot flange. 
     In a further embodiment of the foregoing gas turbine engine, the axial step includes a pilot surface facing radially inward and facing the pilot flange. 
     In a further embodiment of the foregoing gas turbine engine, the retaining plate includes a seal overlap, and the seal overlap radially covers the seal element on a full 360 degree arc. 
     In a further embodiment of the foregoing gas turbine engine, a radial length of the seal overlap is greater than a maximum possible radial misalignment of retaining plate relative to the axis defined by the seal element. 
     A retaining plate for a face seal according to an exemplary embodiment of this disclosure, among other possible things includes a main retaining plate body including a ring defining an axis, an axial step in the retaining plate body, a seal overlap extending radially inward from the axial step, and a plurality of retainer flanges extending radially outward from the retaining plate. 
     In a further embodiment of the foregoing retaining plate, the axial step is an axial shift in the main retaining plate body. 
     In a further embodiment of the foregoing retaining plate, the axial step comprises a radially inward facing pilot surface. 
     In a further embodiment of the foregoing retaining plate, a radial length of the seal overlap is defined by a tolerance of the axial step. 
     In a further embodiment of the foregoing retaining plate, the plurality of flanges is disposed circumferentially evenly about the retaining plate. 
     In a further embodiment of the foregoing retaining plate, the plurality of flanges is disposed circumferentially unevenly, and circumferentially balanced about the retaining plate. 
     In a further embodiment of the foregoing retaining plate, each flange in the plurality of flanges includes a through hole operable to receive a coil spring guide. 
     The foregoing features and elements may be combined in any combination without exclusivity, unless expressly indicated otherwise. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically illustrates an example gas turbine engine. 
         FIG. 2  schematically illustrates a partial isometric view of a face seal for use in the gas turbine engine of  FIG. 1 . 
         FIG. 3A  schematically illustrates a partial cross sectional view of a face seal in an installed position. 
         FIG. 3B  Schematically illustrates a full cross sectional view of the face seal of  FIG. 3A . 
         FIG. 4  schematically illustrates a fore view of a retaining plate. 
         FIG. 4A  illustrates an alternate retaining plate. 
         FIG. 5  schematically illustrates a cross sectional view of the retaining plate of  FIG. 4 . 
     
    
    
     DETAILED DESCRIPTION OF AN EMBODIMENT 
       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. 
     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 . A face seal  39  is positioned adjacent to one or more of the bearing systems  38  and seals the bearing systems  38 . Although only illustrated in a single position, it is understood that face seals  39  can be utilized throughout the gas turbine engine  20  in alternate positions. 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 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 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 high pressure compressor  52  and 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. 
     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 . 
     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. 
     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. 
       FIG. 2  schematically illustrates an isometric view of an assembled face seal  100  for use in a gas turbine engine such as the gas turbine engine  20  of  FIG. 1 . The face seal  100  includes a seal carrier  110 , alternately referred to as a “carrier” or “housing”. The seal carrier  110  defines an axis which is collinear with the engine centerline axis A. Disposed partially within the seal carrier  110  is a sealing element  130 . In one example, the sealing element  130  is a graphitic carbon face seal. The sealing element  130  is retained in position axially by a retaining plate  120  and radially by an interference fit between an outer diameter of the sealing element  130  and the seal carrier  110 . The face seal  100  is axially aligned with an engine shaft, that protrudes through a central opening  150  defined by the face seal  100 . The sealing element  130  interfaces with an axially adjacent seal seat  250  (illustrated in  FIG. 3B ) to seal a bearing compartment of the gas turbine engine  20 . In alternate examples, the face seal  100  can be utilized to seal any similar configuration and is not limited to a bearing compartment. 
     Each of the seal carrier  110  and the retaining plate  120  includes multiple radially outward flanges  112 ,  122 . When the face seal  100  is assembled, each of the seal carrier flanges  112  corresponds directly to one of the retaining plate flanges  122 . The flanges  112 ,  122  are maintained in position relative to each other via a coil spring guide  140  that protrudes through an opening  124  in the retaining plate flange  122  and a hidden opening in the seal carrier flange  112 . 
     In prior retaining plate designs, the retaining plate is approximately straight along a radial line drawn from an axis defined by the central opening, and includes a substantial seal overlap that overlaps the sealing element. The seal overlap maintains the sealing element in position axially by ensuring that a full 360 degree arc of the sealing element is overlapped by the retaining plate. The flanges and coil spring guides approximately radially center the retaining plate relative to the sealing element. The amount of seal overlap required in any given face seal is determined by the tolerances of the centering mechanisms. Because the previous designs use the flanges and the coil spring guides as the centering mechanism for the retaining plate, there are multiple tolerances (at each of the flanges and the coil spring guide holes) that compound to require a substantial seal overlap. 
     The retaining plate  120  of the instant disclosure includes a radially shortened seal overlap  126  relative to prior retaining plates. The seal overlap  126  extends radially inward from an axial step region  128  and a pilot portion hidden. The pilot portion is a radially inward surface of the axial step  128  and operates in conjunction with a pilot flange  218  (illustrated in  FIGS. 3A and 3B ) of the seal carrier  110  to radially center the retaining plate  120  and provide a full circumferential radial constraint on the sealing element  130  via a single centering feature. 
       FIG. 3A  provides a cross sectional view of a face seal arrangement with a face seal  200 , such as the face seal  100  illustrated in  FIG. 2 , positioned in a gas turbine engine  20  (illustrated in  FIG. 1 ).  FIG. 3B  illustrates a full cross sectional view of the face seal arrangement of  FIG. 3A . As with the face seal  100  of  FIG. 2 , the face seal  200  includes a seal carrier  210  with multiple circumferentially positioned carrier flanges  214  extending radially outwards from the seal carrier  210 . A sealing element  230 , such as a graphitic carbon ring, is retained in the seal carrier  210  via a retaining plate  220 . The retaining plate  220  has multiple circumferentially positioned retaining flanges  222  extending radially outward, with each retaining flange  222  corresponding to a carrier flange  214 . Each pair of corresponding flanges  214 ,  222  includes a coil spring guide  240  (hidden in  FIG. 3B ) passing through both flanges  214 ,  222  and maintaining a connection of the flanges  214 ,  222 . The sealing element  230  interfaces with a seal seat  250  (removed from  FIG. 3A  for illustrative purposes) to seal a bearing compartment  280 . 
     In contrast to prior face seals, the seal carrier  210  and the retaining plate  220  each include an axial step region  260 . The axial step regions  260  interact with each other, thereby centering the retaining plate  220  relative to the sealing element  230 . The axial step regions  260  include two tolerances (a pilot surface  262  of the retaining plate  220  and a pilot flange  218  of the seal carrier  210 ). As such, a seal overlap  270  need only account for the variation of the two tolerances. Furthermore, there is no interference between the retaining plate  220  and the pilot flange  218  as a result of a gap between the pilot surface  262  and the pilot flange  218 . The lack of interference between the pilot portion  262  and the pilot flange  218  is alternately referred to as a loose fit. The tighter radial alignment in turn reduces the radial length of the seal overlap  270  that is required to be present between the retaining plate  220  and the seal element  230 . 
       FIG. 4  schematically illustrates a fore view of a simplified retaining plate  300  for use in a seal, such as the face seals  100 ,  200  illustrated in  FIGS. 2 and 3 . The retaining plate  300  includes a main retaining plate body  310 , from which radially protruding flanges  320  protrude outward. Each of the retaining flanges  320  includes an opening  322  for receiving a coil spring guide.  FIG. 4  illustrates four flanges  320  disposed circumferentially evenly about the retaining plate  300 . In alternate examples different numbers of flanges  320  are utilized. In yet further alternate examples, as shown in  FIG. 4A , the flanges  320  are disposed circumferentially unevenly about the retaining plate  300  in a circumferentially balanced manner. In a practical embodiment, the number and position of the flanges  320  correspond to flanges of the seal carrier  210  (illustrated in  FIG. 3 ). 
     Radially inward of the flanges  320 , on the main retaining plate body  310  is an axial step  330 . As described above, at the axial step  330  the retaining plate body  310  shifts axially via an axial aligned portion  334  (illustrated in  FIG. 5 ). In the illustrated example, the axial step  330  is a full ring of the retaining plate main body  310 . A seal overlap portion  340  extends radially inward from the axial step  330 . 
     With continued reference to  FIG. 4 , and with like numerals indicating like elements,  FIG. 5  schematically illustrates a cross section of the retaining plate  300  along view line x. The axial step  330  includes a radially inward pilot surface  332 . As described above with regard to  FIG. 3 , the pilot surface  332  is disposed about a pilot flange  218  of a seal carrier (such as seal carrier  110  illustrated in  FIG. 2 ) in a loose fit arrangement. The pilot surface  332  and the pilot flange  218  of the seal carrier form the axial step region described above. The interfacing between the pilot flange  218  and the pilot surface  332  in the axial step region radially centers the retaining plate  300  relative to the sealing element  230 , such that the retaining plate  300  to sealing element  230  relative radial position is primarily controlled by the differences in tolerances between the pilot flange  218  of the seal carrier and the pilot surface  332 . The tolerances define a gap between the pilot flange  218  of the seal carrier and the pilot surface  332 . In an ideal face seal assembly, this difference is designed to have a minimal length, thereby minimizing the potential radial misalignment that the retaining plate  300  can have. As described above, the minimized potential radial misalignment corresponds directly to a minimized seal overlap  340  required to ensure that the retaining plate  300  provides a full 360 degree arc of radial coverage of the seal element. 
     It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. 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.