Patent Publication Number: US-2022228501-A1

Title: Seal assembly in a gas turbine engine

Description:
TECHNICAL FIELD OF THE INVENTION 
     This invention relates generally to a seal assembly in a gas turbine engine, in particular, a seal assembly for reducing a cooling air leakage in a gas turbine engine. 
     DESCRIPTION OF THE RELATED ART 
     An industrial gas turbine engine typically includes a compressor section, a turbine section, and a mid-frame section disposed therebetween. The compressor section includes multiple stages of compressor blades and vanes and an outlet guide vane assembly aft of the last stage blade and vane. The mid-frame section typically includes a compressor exit diffusor and a combustor assembly. The compressor exit diffusor diffuses the compressed air from the compressor section into a plenum through which the compressed air flows to a combustor assembly which mixes the compressed air with fuel and ignites the mixture and transits the ignited mixture to the turbine section for mechanical power. The turbine section includes multiple stages of turbine blades and vanes. Due to the high temperature of the ignited mixture in the turbine section, cooling air is used to cool the turbine blades and vanes to maintain an adequate component life. 
     Typically, cooling air may be extracted by bleeding compressor air. However, bleeding air from the compressor may reduce the performance and efficiency of the gas turbine engine. Seals are typically arranged at the mid-frame section to reduce cooling air leakage. A reliable sealing system is important to the performance and efficiency of the gas turbine engine. 
     SUMMARY OF THE INVENTION 
     Briefly described, aspects of the present invention relate to a gas turbine engine, a seal assembly in a gas turbine engine, and a method for reducing a leakage in a gas turbine engine. 
     According to an aspect, a gas turbine engine is presented. The gas turbine engine comprises a compressor section comprising an outlet guide vane assembly. The gas turbine engine comprises a mid-frame section arranged downstream of the compressor section. The mid-frame section comprises an inner compressor exit diffuser. The outlet guide vane assembly interfaces with a forward side of the inner compressor exit diffuser. The gas turbine engine comprises a turbine section arranged downstream of the mid-frame section. The gas turbine engine comprises a seal assembly arranged at the forward side of the inner compressor exit diffusor. 
     According to an aspect, a seal assembly in a gas turbine is presented. The gas turbine engine comprises an outlet guide vane assembly interfacing with a forward side of an inner compressor exit diffuser. The seal assembly comprises at least a seal arranged at the forward side of the inner compressor exit diffusor. 
     According to an aspect, a method for reducing a leakage in a gas turbine engine is presented. The gas turbine engine comprises an outlet guide vane assembly interfacing with a forward side of an inner compressor exit diffuser. The method comprises arranging a seal assembly at the forward side of the inner compressor exit diffuser. 
     Various aspects and embodiments of the application as described above and hereinafter may not only be used in the combinations explicitly described, but also in other combinations. Modifications will occur to the skilled person upon reading and understanding of the description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Exemplary embodiments of the application are explained in further detail with respect to the accompanying drawings. In the drawings: 
         FIG. 1  is a schematic longitudinal section view of a portion of a gas turbine engine according to an embodiment of the present invention; 
         FIGS. 2 to 8  are schematic section views of a seal assembly in a gas turbine engine according to various embodiments of the present invention; 
         FIG. 9  is a schematic perspective portion view of a seal assembly in a gas turbine engine according to an embodiment of the present invention; and 
         FIG. 10  is a schematic view of a seal segment of a seal assembly according to an embodiment of the present invention. 
     
    
    
     To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. 
     DETAILED DESCRIPTION OF THE INVENTION 
     A detailed description related to aspects of the present invention is described hereafter with respect to the accompanying figures. 
     For illustration purpose, term “axial” or “axially” refers to a direction along a longitudinal axis of a gas turbine engine, term “radial” or “radially” refers to a direction perpendicular to the longitudinal axis of the gas turbine engine, term “downstream” or “aft” refers to a direction along a flow direction, term “upstream” or “forward” refers to a direction against the flow direction. 
       FIG. 1  illustrates a schematic longitudinal section view of a portion of a gas turbine engine  10  according to an embodiment of the present invention. As illustrated in  FIG. 1 , the gas turbine engine  10  includes a plurality of components along a longitudinal axis  18 . The plurality of components may include a compressor section  100 , a turbine section  300  located downstream of the compressor section  100  with respect to a flow direction A, and a mid-frame section  200  that is located there between. The gas turbine engine  10  also includes an outer casing  12  that encloses the plurality of components. A rotor  14  longitudinally connects the compressor section  100 , the mid-frame section  200  and the turbine section  300  and is circumferentially enclosed thereby. The rotor  14  may be partially or fully enclosed by a shaft cover  16 . 
     The compressor section  100  includes multiple stages of compressor rotating blades  111  and compressor stationary vanes  112 .  FIG. 1  only shows the last stage of compressor rotating blade  111  and compressor stationary vane  112 . An outlet guide vane assembly  120  is arranged downstream of the last stage compressor vane  112 . The compressor blades  111  are installed into the rotor  14 . The compressor vanes  112  and the outlet guide vane assembly  120  are installed into a compressor vane carrier  113 . The compressor vane carrier  113  interfaces with the outer casing  12 . The turbine section  300  includes multiple stages of turbine stationary vanes  312  and turbine rotating blades  311 .  FIG. 1  only shows the first stage of turbine stationary vane  312  and turbine rotating blade  311 . The turbine vanes  312  are installed into a turbine vane carrier  313 . The turbine vane carrier  313  interfaces with the outer casing  12 . The turbine blades  311  are installed into the rotor  14 . The mid-frame section  200  typically includes a combustor assembly  210  and a compressor exit diffuser  220 . The compressor exit diffuser  220  is located downstream of the outlet guide vane assembly  120 . 
     The compressor exit diffusor  220  typically includes an outer compressor exit diffusor  221  and an inner compressor exit diffusor  222 . The outer compressor exit diffusor  221  is connected to the inner compressor exit diffusor  222  by bolting to a strut  223 . The inner compressor exit diffusor  222  may enclose the shaft cover  16 . Forward side of the outer compressor exit diffusor  221  interfaces with the outer casing  12 . Forward side of the inner compressor exit diffusor  222  interfaces with the last stage compressor vane  112  and the outlet guide vane assembly  120 . 
     In operation of the gas turbine engine  10 , the compressor section  100  inducts air via an inlet duct (not shown). The air is compressed and accelerated in the compressor section  100  while passing through the multiple stages of compressor rotating blades  111  and compressor stationary vanes  112 , as indicated by the flow direction A. The compressed air passes through the outlet guide vane assembly  120  and enters the compressor exit diffuser  220 . The compressor exit diffuser  200  diffuses the compressed air to the combustor assembly  210 . The compressed air is mixed with fuel in the combustor assembly  210 . The mixture is ignited and burned in the combustor assembly  210  to form a combustion gas. The combustion gas enters the turbine section  300 , as indicated by the flow direction A. The combustion gas is expanded in the turbine section  300  while passing through the multiple stages of turbine stationary vanes  312  and turbine rotating blades  311  to generate mechanical power which drives the rotor  14 . The rotor  14  may be linked to an electric generator (not shown) to convert the mechanical power to electrical power. The expanded gas constitutes exhaust gas and exits the gas turbine engine  10 . 
     In operation, due to the high temperature of the combustion gas, cooling air is used to cool the turbine blades  311  and vanes  312  to maintain an adequate component life. Cooling air may leak at the forward side of the inner compressor exit diffusor  222 . Cooling air leakage may negatively affect the performance and efficiency of the gas turbine engine  10 . According to embodiments of the present invention, a seal assembly  400  may be arranged at the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage. 
       FIG. 2  is a schematic longitudinal section view of a seal assembly  400  in a gas turbine engine  10  according to an embodiment of the present invention. The seal assembly  400  is arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage therebetween. The seal assembly  400  may circumferentially surround the outlet guide vane assembly  120  or the forward side of the inner compressor exit diffusor  222 . As shown in the exemplary embodiment of  FIG. 2 , the outlet guide vane assembly  120  includes an outlet guide vane airfoil  121  extending radially between an airfoil root  122  and an inner platform  123 . The outlet guide vane airfoil  121 , the airfoil root  122  and the inner platform  123  may be manufactured as an integral piece. The airfoil root  122  is installed into the compressor vane carrier  113 . The inner platform  123  may be axially extended toward upstream by connecting to an inner shroud  124  via flanges  125  forming an extended inner shroud  126  in the axially direction. It is understood that the inner platform  123  and the inner shroud  124  may be connected to each other by any suitable means or may be integrally manufactured as one piece. 
     The seal assembly  400  includes at least one seal  410 . According to an exemplary embodiment, the at least one seal  410  may be a brush seal  410 . The brush seal  410  is mounted in the inner compressor exit diffusor  222  at the forward side and radially engages the outlet guide vane assembly  120 . The seal assembly  400  may circumferentially surround the forward side of the inner compressor exit diffusor  222 . The brush seal  410  has a housing  411  and a plurality of bristles  412 . The housing  411  may have a U-shape. The plurality of bristles  412  is secured within the housing  411 . The inner compressor exit diffusor  222  includes a seal groove  224 . The seal groove  224  may have a U-shape. The housing  411  of the brush seal  410  is installed into the seal groove  224  on the inner compressor exit diffusor  222 . The housing  411  and the seal groove  224  are dimensioned to form a tight fit against each other. A radial length of the plurality of bristles  412  is dimensioned to provide sufficient air tight sealing contact against a sealing surface  420  to reduce cooling air leakage therebetween. The sealing surface  420  may be a bottom surface  127  of the extended inner shroud  126  of the outlet guide vane assembly  120 . 
     According to an aspect of the invention, the seal assembly  400  may include a plurality of brush seals  410 . The plurality of brush seals  410  may be arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  and tandem along the axial direction of the gas turbine engine  10  at the forward side of the inner compressor exit diffusor  222 . As shown in the exemplary embodiment of  FIG. 2 , the seal assembly  400  includes dual brush seals  410  located upstream and downstream respectively. The dual brush seals  410  are mounted in the inner compressor exit diffusor  222  and radially engage the outlet guide vane assembly  120 . The inner compressor exit diffusor  222  includes dual seal grooves  422  located upstream and downstream respectively at the forward side. The housing  411  of the upstream brush seal  410  is installed into the upstream seal groove  224 . The housing  411  of the downstream brush seal  410  is installed into the downstream seal groove  224 . A radial length of the bristles  412  is dimensioned to provide sufficient air tight sealing contact against a sealing surface  420  to reduce cooling air leakage therebetween. The sealing surface  420  may be a bottom surface  127  of the extended inner shroud  126  of the outlet guide vane assembly  120 . 
       FIG. 3  is a schematic longitudinal section view of a seal assembly  400  in a gas turbine engine  10  according to an embodiment of the present invention. The seal assembly  400  is arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage therebetween. The seal assembly  400  may circumferentially surround the outlet guide vane assembly  120 . The seal assembly  400  includes at least one seal  410 . In the exemplary embodiment shown in  FIG. 3 , the at least one seal  410  may be a brush seal  410 . The brush seal  410  is mounted in the outlet guide vane assembly  120  and radially engages the inner compressor exit diffusor  222 . The outlet guide vane assembly  120  includes a seal groove  128  at bottom side of the extended inner shroud  126 . The brush seal  410  is mounted in the outlet guide vane assembly  120  by installing the housing  411  into the seal groove  128  on the outlet guide vane assembly  120 . A radial length of the plurality of bristles  412  is dimensioned to provide sufficient air tight sealing contact against a sealing surface  420  to reduce cooling air leakage therebetween. The sealing surface  420  may be a top surface  225  of the inner compressor exit diffusor  222 . 
     According to an aspect of the invention, the seal assembly  400  may include a plurality of brush seals  410 . The plurality of brush seals  410  may be arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  and tandem along an axial direction of the gas turbine engine  10  at the forward side of the inner compressor exit diffusor  222 . As shown in the exemplary embodiment of  FIG. 3 , the seal assembly  400  includes dual brush seals  410  located upstream and downstream respectively. The dual brush seals  410  are mounted in the outlet guide vane assembly  120  and radially engage the inner compressor exit diffusor  222 . The outlet guide vane assembly  120  includes dual seal grooves  128  located upstream and downstream respectively at bottom side of the extended inner shroud  126 . The housing  411  of the upstream brush seal  410  is installed into the upstream seal groove  128 . The housing  411  of the downstream brush seal  410  is installed into the downstream seal groove  128 . A radial length of the plurality of bristles  412  of the dual brush seals  410  is dimensioned to provide sufficient air tight sealing contact against a sealing surface  420  to reduce cooling air leakage therebetween. The sealing surface  420  may be a top surface  225  of the inner compressor exit diffusor  222 . 
       FIG. 4  is a schematic longitudinal section view of a seal assembly  400  in a gas turbine engine  10  according to an embodiment of the present invention. The seal assembly  400  is arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  reduce cooling air leakage therebetween. As shown in the exemplary embodiment of  FIG. 4 , the seal assembly  400  includes dual brush seals  410  located upstream and downstream respectively. The upstream brush seal  410  is mounted in the inner compressor exit diffusor  222  and radially engages the outlet guide vane assembly  120 . The downstream brush seal  410  is mounted in the outlet guide vane assembly  120  and radially engages the inner compressor exit diffusor  222 . A radial length of the bristles  412  of the dual brush seals  410  is dimensioned to provide sufficient air tight sealing contact against a sealing surface  420  to reduce cooling air leakage therebetween. In the illustrated embodiment of  FIG. 3 , the sealing surface  420  for the upstream brush seal  410  is a bottom surface  127  of the extended inner shroud  126  of the outlet guide vane assembly  120 . The sealing surface  420  for the downstream brush seal  410  is a top surface  225  of the inner compressor exit diffusor  222 .  FIG. 4  is for illustration purpose only. It is understood that the plurality of brush seals  410  may be arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  and tandem along the axial direction in any configurations. For example, the upstream brush seal  410  may be mounted in the outlet guide vane assembly  120  and radially engages the inner compressor exit diffusor  222 , and the downstream brush seal  410  may be mounted in the inner compressor exit diffusor  222  and radially engages the outlet guide vane assembly  120 . 
       FIGS. 5 to 7  illustrate schematic longitudinal section views of a seal assembly  400  in a gas turbine engine  10  according to various embodiments of the present invention. The seal assembly  400  is arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage therebetween. As shown in the exemplary embodiment of  FIG. 5 , the seal assembly  400  includes three brush seals  410  arranged tandem along the axial direction at the forward side of the inner compressor exit diffusor  222 . The three brush seals  410  are mounted in the inner compressor exit diffusor  222  and radially engage the outlet guide vane assembly  120 . As shown in the exemplary embodiment of  FIG. 6 , the seal assembly  400  includes three brush seals  410  arranged tandem along the axial direction. The three brush seals  410  are mounted in the outlet guide vane assembly  120  and radially engage the inner compressor exit diffusor  222 . As shown in the exemplary embodiment of  FIG. 7 , the seal assembly  400  includes three brush seals  410  arranged tandem along the axial direction at the forward side of the inner compressor exit diffusor  222 . Two of the three brush seals  410  are mounted in the inner compressor exit diffusor  222  and radially engages the outlet guide vane assembly  120 . One of the three brush seals  410  is mounted in the outlet guide vane assembly  120  and radially engages the inner compressor exit diffusor  222 . The figures are for illustration purpose only. It is understood the seal assembly  400  may include any number of brush seals  410 . The brush seals  410  may be mounted in the inner compressor exit diffusor  222  and/or the outlet guide vane assembly  120  in any configurations. 
       FIG. 8  is a schematic partial circumferential section view of a seal assembly  400  in a gas turbine engine  10  according to an embodiment of the present invention. The gas turbine engine  10  includes a plurality of outlet guide vane assemblies  120 . The plurality of outlet guide vane assemblies  120  are circumferentially arranged. The plurality of outlet guide vane assemblies  120  forms an annular shape that interfaces with the forward side of the inner compressor exit diffusor  222 . A tangential gap may exist between adjacent outlet guide vane assemblies  120 . As shown in the exemplary embodiment of  FIG. 8 , the seal assembly  400  is arranged circumferentially between adjacent outlet guide vane assemblies  120  to reduce cooling air leakage therebetween. The seal assembly  400  may extend axially along the adjacent outlet guide vane assemblies  120 . The seal assembly  400  includes at least one seal  410 . The at least one seal  410  is mounted in an extended inner shroud  126  of one outlet guide vane assembly  120  and circumferentially engages an extended inner shroud  126  of an adjacent outlet guide vane assembly  120 . The at least one seal  410  may be a brush seal  410 .  FIG. 8  is for illustration purpose only. It is understood that the at least one seal  410  may be mounted in a flange  125  of one outlet guide vane assembly  120  and circumferentially engages a flange  125  of an adjacent outlet guide vane assembly  120 . It is also understood that a plurality of seals  410  may be arranged circumferentially between adjacent outlet guide vane assemblies  120  and tandem along the axial direction. 
     The seal assembly  400  may include any combinations of various embodiments. For example, the seal assembly  400  may include at least one seal  410  arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage therebetween and at least another seal  410  arranged circumferentially between adjacent outlet guide vane assemblies  120  to reduce cooling air leakage therebetween. 
       FIG. 9  is a schematic perspective partial view of a seal assembly  400  installed in a gas turbine engine  10  according to the embodiment of the present invention illustrated in  FIG. 2 . As shown in the exemplary embodiment of  FIG. 9 , the seal assembly  400  includes dual brush seals  410  arranged radially at the forward side of the inner compressor exit diffusor  222 . The dual brush seals  410  are located upstream and downstream respectively. The dual brush seals  410  are mounted in the inner compressor exit diffusor  222  and radially engage the outlet guide vane assembly  120 , as shown in  FIG. 2 . The inner compressor exit diffusor  222  includes dual seal grooves  422  located upstream and downstream respectively at the forward side of the inner compressor exit diffusor  222 . The housing  411  of the upstream brush seal  410  is installed into the upstream seal groove  224 . The housing  411  of the downstream brush seal  410  is installed into the downstream seal groove  224 . The seal assembly  400  may circumferentially surround the forward side of the inner compressor exit diffusor  222 . 
     The at least one seal  410  may include a plurality of seal segments.  FIG. 10  illustrates a schematic view of a seal segment  413  of a seal  410  in a seal assembly  400  according to an embodiment of the present invention. The seal segment  413  may have a circular shape. The seal segment  413  may have any circular degrees, for example, 45 degree, 60 degree, 90 degree, or 180 degree. The plurality of the seal segments  413  may form an annular seal  410  surrounding the inner compressor exit diffusor  222 . For example, the seal  410  may include six 60 degree seal segments  413 . Clearance may exist between adjacent seal segments  413  for thermal expansion. 
     According to an aspect, the proposed seal assembly  400  may reduce cooling air leakage in a gas turbine engine  10 . The seal assembly  400  may be arranged at a forward side of an inner compressor exit diffusor  222  of the gas turbine engine  10 . The seal assembly  400  may be arranged radially between the outlet guide vane assembly  120  and the forward side of the inner compressor exit diffusor  222  to reduce cooling air leakage therebetween. The seal assembly  400  may be arranged circumferentially between adjacent outlet guide vane assemblies  120  to reduce cooling air leakage therebetween. The proposed seal assembly  400  may reduce cooling air leakage in a gas turbine engine  10  and thus improve efficiency and performance of the gas turbine engine  10 . 
     According to an aspect, the proposed seal assembly  400  may include at least one seal  410 . The at least one seal  410  may be a brush seal  410 . 
     Although various embodiments that incorporate the teachings of the present invention have been shown and described in detail herein, those skilled in the art can readily devise many other varied embodiments that still incorporate these teachings. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted,” “connected,” “supported,” and “coupled” and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical or mechanical connections or couplings. 
     REFERENCE LIST 
     
         
           10 : Gas Turbine Engine 
           12 : Outer Casing 
           14 : Rotor 
           16 : Shaft Cover 
           18 : Longitudinal Axis 
           100 : Compressor Section 
           111 : Compressor Blade 
           112 : Compressor Vane 
           113 : Compressor Vane Carrier 
           120 : Outlet Guide Vane Assembly 
           121 : Outlet Guide Vane Airfoil 
           122 : Airfoil Root 
           123 : Inner Platform 
           124 : Inner Shroud 
           125 : Flange 
           126 : Extended Inner Shroud 
           127 : Bottom Surface of Extended Inner Shroud 
           128 : Seal Groove on Extended Inner Shroud 
           200 : Mid-Frame Section 
           210 : Combustor Assembly 
           220 : Compressor Exit Diffusor 
           221 : Outer Compressor Exit Diffusor 
           222 : Inner Compressor Exit Diffusor 
           223 : Strut 
           224 : Seal Groove on Inner Compressor Exit Diffusor 
           225 : Top Surface of Inner Compressor Exit Diffusor 
           300 : Turbine Section 
           311 : Turbine Blade 
           312 : Turbine Vane 
           313 : Turbine Vane Carrier 
           400 : Seal Assembly 
           410 : Brush Seal 
           411 : Housing of Brush Seal 
           412 : Bristle of Brush Seal 
           413 : Seal Segment 
           420 : Sealing Surface