Patent Publication Number: US-11041391-B2

Title: Conformal seal and vane bow wave cooling

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application is a continuation in-part of U.S. application Ser. No. 16/751,298 filed on Jan. 24, 2020, that is a continuation of U.S. application Ser. No. 15/690,615 filed on Aug. 30, 2017, now U.S. Pat. No. 10,584,601 granted Mar. 10, 2020. 
    
    
     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-energy exhaust gas flow. The high-energy 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. 
     An interface between the combustor exit and the first vane stage can experience elevated temperatures at localized areas near a leading edge of each vane. The interface between the combustor exit and the first vane stage includes a gap. Bow wave phenomena at the leading edge of each vane in combination with the gap can result in elevated temperatures within and near the gap at this location. 
     Turbine engine manufacturers continue to seek improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. 
     SUMMARY 
     In a featured embodiment, a gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. 
     In another embodiment according to the previous embodiment, the first vane stage includes a plurality of vanes with each of the plurality of vanes including a leading edge and the seal assembly includes a plurality of slots disposed at circumferential positions corresponding with the leading edge of each of the plurality of vanes. 
     In another embodiment according to any of the previous embodiments, the first plurality of openings extend through the seal assembly to communicate cooling airflow into a first set of the plurality of slots and the second plurality of openings extend through the first vane stage to communicate cooling airflow into a second set of the plurality of slots. 
     In another embodiment according to any of the previous embodiments, the first plurality of slots and the second plurality of slots alternate circumferentially about the circumference of the first vane stage. 
     In another embodiment according to any of the previous embodiments, the first plurality of slots and the second plurality of slots are the same such that cooling air is communicated to each of the first and second plurality of slots from cooling holes in both the first vane stage and the seal assembly. 
     In another embodiment according to any of the previous embodiments, the seal assembly includes a radially outer surface and the first plurality of openings are angled relative to the radially outer surface. 
     In another embodiment according to any of the previous embodiments, the seal assembly includes aft face that seals against a forward rib of the first vane stage. 
     In another embodiment according to any of the previous embodiments, the seal assembly includes an alignment slot that aligns the seal assembly circumferentially with the first vane stage. 
     In another featured embodiment, a first vane stage assembly for a gas turbine engine includes a first vane stage including an axial face. A seal assembly abuts the axial face and extends axially across a gap between a combustor and the first turbine vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into the gap. 
     In another embodiment according to any of the previous embodiments, the seal assembly includes a plurality of slots disposed at circumferential positions corresponding with the leading edge of vanes of the first turbine stage. 
     In another embodiment according to any of the previous embodiments, the first plurality of openings and the second plurality of openings open into a corresponding one of the plurality of slots. 
     In another embodiment according to any of the previous embodiments, the first plurality of openings and the second plurality of openings are disposed in groups spaced apart circumferentially to correspond with the circumferential positions of the plurality of slots. 
     In another embodiment according to any of the previous embodiments, the first plurality of openings are in communication with a first set of the plurality of slots and the second plurality of openings are in communication with a second set of the plurality of slots that is different than the first set of the plurality of slots. 
     In another embodiment according to any of the previous embodiments, the second plurality of openings extend at an angle through the axial face of the first vane stage. 
     In another embodiment according to any of the previous embodiments, the seal assembly includes an alignment slot that aligns the seal assembly circumferentially with the first vane stage. 
     In another featured embodiment, a method of cooling an interface between a combustor and a turbine vane stage includes assembling a seal across a gap between a combustor and a turbine vane stage aft of the combustor. Cooling air flow is communicated into the gap through a first plurality of openings in the seal and a second plurality of openings in the turbine vane stage. 
     In another embodiment according to any of the previous embodiments, includes forming the seal to include a plurality of circumferential slots and aligning the plurality of circumferential slots with a leading edge of turbine vanes within the turbine vane stage. 
     In another embodiment according to any of the previous embodiments, includes grouping the first plurality of openings and the second plurality of openings circumferentially to correspond with the location of the plurality of circumferential slots and the leading edge of the turbine vanes. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     These and other features disclosed herein 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  is a schematic view of an example gas turbine engine. 
         FIG. 2  is a cross-sectional view of a portion of the gas turbine engine. 
         FIG. 3  is a front view of a first turbine vane stage. 
         FIG. 4  is a perspective view of an interface between a combustor and an example first turbine stage. 
         FIG. 5  is another perspective view of the interface between the combustor and the example first turbine stage. 
         FIG. 6  is an axial front view of a conformal seal. 
         FIG. 7  is a front view of a portion of the example conformal seal. 
         FIG. 8  is an enlarged view of a slot of the example conformal seal. 
         FIG. 9A  is a schematic illustration of an example cooling air hole grouping. 
         FIG. 9B  is another schematic illustration of another example cooling air hole grouping. 
         FIG. 10  is a perspective view of an interface between a combustor another example turbine vane stage. 
         FIG. 11  is an enlarged perspective view of the interface between the combustor and the example turbine vane stage of  FIG. 10 . 
         FIG. 12  is a perspective view radially outward of the example turbine vane stage and seal. 
         FIG. 13  is front view of the example 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  drives air along a bypass flow path B in a bypass duct defined within a nacelle  18 , 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 . 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 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 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 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  58  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  58  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  58  includes airfoils  60  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 disclosure is applicable to other gas turbine engines including direct drive turbofans, land based turbine engines utilized for power generation as well as turbine engines for use in land based vehicles and sea going vessels. 
     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). 
     The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, the fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment, the low pressure turbine  46  includes about three (3) turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Referring to  FIG. 2  with continued reference to  FIG. 1 , the example combustor  56  includes an axially aft end  62  that is adjacent to an axially forward face  72  of a first turbine vane stage  64 . The first turbine vane stage  64  includes an upper platform  70  that defines the forward face  72  and has a radially outward extending rib  92 . The combustor  56  includes a rib  84  extending radially outward and spaced apart from the end of the combustor  56 . 
     A conformal seal  76  is disposed between the rib  84  and the forward face  72  on the radially outer surface  74  of the combustor  56 . The conformal seal  76  extends axially aft from the rib  84  over a radially extending gap  78  between the combustor  56  and the first turbine vane stage  64 . 
     Referring to  FIG. 3  with continued reference to  FIG. 2 , the first turbine vane stage  64  includes a plurality of turbine vanes  65  that extend between the upper platform  70  and a lower platform  75 . Each vane  65  includes a leading edge  68  facing toward the combustor  56 . The leading edge  68  encounters the high-energy gas flow generated in the combustor  56  and directs that gas flow into the turbine section  28 . The leading edge  68  of each vane  65  can cause undesired distortions in gas flow that generate non-uniform temperature variations within the gap  78 . Bow wave flow phenomena is one such flow distortion that may cause undesired discreet temperature increases. Other flow disruptions that result in gas flow entering the gap  78  may also result in undesired localized temperature variations and also will benefit from this disclosure. 
     The example turbine stage  64  includes a plurality of doublets  66  that are arranged circumferentially about the engine axis A. Each of the doublets  66  includes two vanes  65  with common upper and lower platforms  70 ,  75 . It is within the contemplation of this disclosure to utilize other turbine vane stage configurations with the disclosed seal  76 . 
     Referring to  FIG. 4  with continued reference to  FIGS. 2 and 3 , the disclosed example conformal seal  76  includes a first plurality of cooling holes  80  that extend from a radially outer surface  96  to a radially inner surface  98  that is in communication with the gap  78 . Each of the cooling holes  80  are disposed at an angle  100  relative to the radially outer surface  96  such that cooling air schematically shown at  102  exits into the gap  78 . The conformal seals  76  includes a wearing end portion  82  that wears down during initial operation to provide a desired seal against the axial face  72 . 
     Referring to  FIG. 5  with continued reference to  FIGS. 2, 3 and 4 , the axial face  72  includes a second plurality of cooling air holes  88  that communicate cooling air into the gap  78 . The cooling air holes  88  are disposed at an angle  105  relative to the axial face  72  to direct airflow radially inward into the gap  78 . 
     The first and second plurality of cooling air holes  80 ,  88  are sized to provide a desired pressure and cooling airflow into the gap  78 . In one disclosed embodiment, the first plurality of cooling air holes  80  are 0.025 inch (0.635 mm) in diameter and the second plurality of cooling air holes  88  are 0.023 inch (0.5842 mm) in diameter. According to another embodiment, the cooling air holes may vary from between 0.015 inch (0.381 mm) and 0.080 inch (2.032 mm) in diameter. It should be understood that although an example size of hole is disclosed by way of example, other sizes and combinations of cooling hole structures are within the contemplation of this disclosure. 
     Referring to  FIGS. 6 and 7  with continued reference to  FIGS. 2-5 , the example conformal seal  76  includes a plurality of slots  86  arranged circumferentially about the engine axis A. Each of the slots  86  is aligned with a corresponding leading edge  68  of the vanes  65  within the first turbine vane section  64 . The cooling holes  80  open on the radial inner surface  98  of the seal  76  within each of the plurality of slots  86 . The cooling holes  80  communicate cooling airflow to the gap  78  at a circumferential location that corresponds with the leading edge  68  of each of the vanes  65 . 
     The first and second plurality of cooling air holes  80  and  88  are grouped at the circumferential location that corresponds with the leading edge  68 . In one disclosed embodiment, each grouping includes between  1  and  10  holes. In other disclosed embodiment, each grouping of cooling air holes includes  8  holes. While specific grouping counts are disclosed, other grouping counts are within the contemplation of this disclosure. 
     The conformal seal  76  includes a tab  104  with a slot  106 . The slot  106  corresponds with slots  95  defined in rib  92  of the vane stage  64  and slot  85  defined as part of the combustor rib  84 . An alignment member  94  extends through the slots  85 ,  106  and  95  to align the slots  86  and the cooling holes  80  with the leading edge  68  of each vane  65 . 
     Referring to  FIG. 8  with continued reference to  FIGS. 4, 5, 6 and 7 , each of the slots  86  provides for communication of cooling airflow  102  into the gap  78  in a location corresponding with the leading edge  68  of each of the vanes  65 . The cooling holes  80  are grouped circumferentially about the circumference of the conformal seal  76  to correspond with each of the slots  86 . 
     The first and second plurality of cooling air holes  80  and  88  may both communicate air into each of the slots  86  defined within the seal  76 . The first and second plurality of cooling air holes  80 ,  88  may also be incremented such that a first circumferential position includes cooling air holes  80  from the seal  76  and a second circumferential position includes cooling air provide by the second plurality of cooling air holes  88 . The orientation of cooling air holes  80 ,  88  can be incremented, and combined to provide cooling air flow through one or both the first and second plurality of cooling air holes  80 ,  88  to provide desired airflow required to maintain the gap  78  within desired temperature ranges. 
     Referring to  FIG. 9A , a set of the plurality of slots  86  indicated at  110  receives cooling air from the first plurality of cooling air holes  80  and a second set of the plurality of slots  86  indicated at  112  receives cooling air from the second plurality of cooling air holes  88 . Referring to  FIG. 9B , another example disclosed embodiment is schematically illustrated to show another circumferential orientation of the first set of slots  110  and the second set of slots  112 . It should be appreciated that the sets of slots need not be symmetric about the engine axis and can include different combination of cooling air holes communicating cooling air into the plurality of slots  86 . Cooling airflow  102  is communicated through the conformal seal  76  and vane stage  64  into the gap  78  at the specific circumferential location that corresponds with the leading edge  68  of each of the vanes  65 . 
     The example conformal seal  76  provides a seal between the end of the combustor and the first turbine stage while also providing directed cooling airflow to prevent or substantially limit hot gas flow into the gap  78 . 
     Referring to  FIG. 10 , a turbine vane stage  114  includes forward flange  122  with a forward face  124 . The seal  76  includes a seal face  140  that abuts directly against the forward face  124  to maintain core gas flow with the core flow path C. The seal  76  is also abutted directly against an aft wall  116  of the combustor  56 . The aft wall  116  includes an aft facing side  118  that is abutted directly against the seal  76 . The seal  76  covers a gap  78  between the combustor  56  and the turbine vane stage  114 . 
     The turbine vane stage  114  includes a plurality of vane cooling passages  130 . Each of the vane cooling passages  130  includes an inner opening  132  and an outer opening  144 . Cooling air is communicated from the outer opening  144  through the cooling passages and out the inner opening  132  into the core flow path C forward of a leading edge  138 . The outer opening  144  is disposed within the forward flange  122  and the inner opening  132  breaks through on a radial inner surface  128  of the turbine vane stage  114 . The cooling passage  130  is angled radially inward in a direction forward from the outer opening  144  to the inner opening  132 . 
     Referring to  FIG. 11 , with continued reference to  FIG. 10 , the inner opening  132  includes a diameter  146  that corresponds with the diameter of the passage  130 . The inner opening  132  includes a centerline  134  that is spaced a distance  136  from the forward face  124 . In one disclosed example, the distance  136  is greater than one diameter  136  from the forward face  125 . As appreciated, spacing the centerline  134  one diameter  136  from the front face  124  provides for an edge of the inner opening  132  to be spaced at least ½ the diameter from the forward face  124 . In another disclosed embodiment, the centerline  134  of the inner opening  132  is spaced no more than three times the diameter  146  from the forward face  124 . 
     Referring to  FIGS. 12 and 13 , the vane cooling passages  130 , and thereby the inner openings  132  are aligned with the leading edge  138  of each vane in the turbine vane stage  114 . The slots  86  of the seal  76  are also circumferentially aligned with each leading edge  138  of the turbine vane stage  114 . The slots  86  are disposed within circumferentially spaced apart annular sectors  148  that are a portion of the complete circumference of the seal  76 . Only two annular sectors  148  are shown in  FIG. 13  for clarity purposes, however an annular sector is present that corresponds with each of the slots  86 . Accordingly, each slot  86  is disposed within a corresponding one of the annular sectors  148 . 
     The seal passages  80  and the inner openings  132  for the vane cooling passages  130  are grouped together within the annular sectors  145 . The annular sectors  145  are disposed at an angle  150 . In one disclosed embodiment, the angle  150  is  5 . 9  degrees. An arc length  152  of the annular sector is between 1% and 2% of the complete circumference of the seal  76  and vane stage. It should be appreciated, that the angle  150  and arc length  152  may vary depending on the number and orientation of vanes in the turbine vane stage  114  and such variations are within the contemplation and scope of this disclosure. 
     The inner openings  132  are spaced a distance  148  apart from each other within each annular sector  145 . In one example embodiment, the distance  148  is no more than six times one hole diameter  136  and no less than four hole diameters  136 . In another disclosed embodiment, the distance  148  is no more than about five times one hole diameter  136 . In each of the disclosed embodiments, the distance between openings  132  within each of the annular sectors  145  is much less than a distance between cooling openings that may be outside of the annular sectors  145 . 
     In one example embodiment, an annular area of the annular sector  145  is defined between radially outer edges of the openings  132  and by the outer most edge of openings  132  through the turbine vane stage  114 . An open area within the annular sector  145  is defined by the number and size of the openings  132 . In one disclosed embodiment, the open area is between 18 percent and 22 percent of the annular area. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.