Patent Publication Number: US-2012031099-A1

Title: Combustor assembly for use in a turbine engine and methods of assembling same

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to turbine engines and more particularly, to combustor assemblies for use with turbine engines. 
     At least some known gas turbine engines use cooling air to cool a combustion assembly included within the engine. Often the cooling air is supplied from a compressor coupled in flow communication upstream from the combustion assembly. More specifically, in at least some known turbine engines, cooling air is discharged from the compressor into a plenum that extends at least partially around a transition piece of the combustor assembly. A portion of the cooling air entering the plenum is supplied to an impingement sleeve circumscribing the transition piece prior to being channeled into a cooling channel defined between the impingement sleeve and the transition piece. Cooling air entering the cooling channel is discharged downstream into a second channel defined between a combustor liner and a flowsleeve. Any remaining cooling air entering the plenum is channeled through inlets defined within the flowsleeve prior to being discharged downstream into the second channel. 
     Cooling air flowing through the second channel cools an exterior of the combustor liner. At least some known flowsleeves include inlets and thimbles that discharge the cooling air into the second channel. The inlets channel the cooling air in a non-uniform air flow pattern circumferentially about an outer surface of the combustor liner. The non-uniform distribution may cause temperature variations across the combustor liner outer surface and may cause an uneven heat transfer between the combustor liner and the cooling air. Overtime, the uneven heat transfer may result in thermal cracking and/or damage to the combustor liner, both of which may reduce the overall useful life of the combustor liner and/or increase the cost of maintaining and operating the turbine engine. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a combustor assembly is provided. The combustor assembly includes a combustor liner having a centerline axis and defining a combustion chamber there within. A plurality of fuel nozzles extends through the combustion liner. An annular flowsleeve is coupled radially outward from the combustor liner such that an annular flow path is defined between the flowsleeve and the combustor liner. The flowsleeve includes a forward surface that extends between an upper endwall and a lower endwall. The upper endwall is positioned a first distance from the plurality of fuel nozzles. The lower endwall is positioned a second distance from the plurality of fuel nozzles that is different than the first distance. 
     In another aspect, a turbine engine is provided. The turbine engine includes a compressor and a combustor in flow communication with the compressor to receive at least some of the air discharged by the compressor. The combustor includes a plurality of combustor assemblies. At least one combustor assembly of the plurality of combustor assemblies includes a combustor liner having a centerline axis and defining a combustion chamber there within. A plurality of fuel nozzles extends through the combustion liner. An annular flowsleeve is coupled radially outward from the combustor liner such that an annular flow path is defined between the flowsleeve and the combustor liner. The flowsleeve includes a forward surface that extends between an upper endwall and a lower endwall. The upper endwall is positioned a first distance from the plurality of fuel nozzles. The lower endwall is positioned a second distance from the plurality of fuel nozzles that is different than the first distance. 
     In a further aspect, a method of assembling a combustor assembly is provided. The method includes coupling a combustor liner to a plurality of fuel nozzles, wherein the combustor liner includes a combustion chamber defined therein, the combustion liner extending along a centerline axis. An annular flowsleeve is coupled radially outwardly from the combustor liner such that an annular flow path is defined between the flowsleeve and the combustor liner. The annular flowsleeve includes a forward surface that extends between an upper endwall and a lower endwall. The upper endwall is positioned a first distance from the plurality of fuel nozzles. The lower endwall is positioned a second distance from the plurality of fuel nozzles that is different than the first distance. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic cross-sectional illustration of an exemplary turbine engine. 
         FIG. 2  is an enlarged cross-sectional illustration of a portion of an exemplary combustor assembly that may be used with the turbine engine shown in  FIG. 1 . 
         FIG. 3  is a partial cross-sectional view of an exemplary flowsleeve that may be used with the combustor assembly shown in  FIG. 2 . 
         FIGS. 4-9  are cross-sectional views of alternative flowsleeves that may be used with the combustor assembly shown in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The exemplary methods and systems described herein overcome disadvantages of known combustor assemblies by providing a flowsleeve that discharges a substantially uniform flow distribution of cooling fluid about a combustor liner to facilitate enhanced heat transfer between the cooling fluid and the combustor liner outer surface. More specifically, the embodiments described herein provide a flowsleeve that includes an inlet opening that is oriented obliquely to a centerline axis of the combustor liner to enable a flow of cooling fluid having a uniform circumferential pressure distribution to be defined about the combustor liner outer surface. The uniform distribution of cooling fluid facilitates substantially evenly reducing a temperature of the combustor liner outer surface, which facilitates increasing the operating life of the combustor liner. 
     As used herein, the term “upstream” refers to a forward end of a turbine engine, and the term “downstream” refers to an aft end of a turbine engine. 
       FIG. 1  is a schematic view of an exemplary turbine engine  10 . Turbine engine  10  includes an intake section  12 , a compressor section  14  that is downstream from intake section  12 , a combustor section  16  downstream from compressor section  14 , a turbine section  18  downstream from combustor section  16 , and an exhaust section  20  downstream from turbine section  18 . Turbine section  18  is coupled to compressor section  14  via a rotor assembly  22  that includes a shaft  28 . Combustor section  16  includes a plurality of combustor assemblies  30  that are each coupled in flow communication with the compressor section  14 . A fuel nozzle assembly  26  is coupled to each combustor assembly  30 . Turbine section  18  is rotatably coupled to compressor section  14  and to a load (not shown) such as, but not limited to, an electrical generator and/or a mechanical drive application. In one embodiment, turbine engine  10  is a MS9001E engine, commercially available from General Electric Company, Schenectady, N.Y. It should be noted that turbine engine  10  is exemplary only, and that the present invention is not limited to being used only with turbine engine  10 , but rather may instead be implemented within any turbine engine that functions as described herein. 
     In operation, air flows through compressor section  14  and compressed air is discharged into combustor section  16 . Combustor assembly  30  injects fuel, for example, natural gas and/or fuel oil, into the air flow, ignites the fuel-air mixture to expand the fuel-air mixture through combustion, and generates high temperature combustion gases. Combustion gases are discharged from combustor assembly  30  towards turbine section  18  wherein thermal energy in the gases is converted to mechanical rotational energy. Combustion gases impart rotational energy to turbine section  18  and to rotor assembly  22 , which subsequently provides rotational power to compressor section  14 . 
       FIG. 2  is an enlarged cross-sectional illustration of a portion of combustor assembly  30 . In the exemplary embodiment, combustor assembly  30  is coupled in flow communication with turbine section  18  and with compressor section  14 . Moreover, in the exemplary embodiment, compressor section  14  includes a diffuser  32  coupled in flow communication with a discharge plenum  34  that enables air to be channeled downstream from compressor section  14  towards combustor assembly  30 . 
     In the exemplary embodiment, combustor assembly  30  includes a substantially circular dome plate  36  that at least partially supports a plurality of fuel nozzles  38 . Dome plate  36  is coupled to a substantially cylindrical combustor flowsleeve  40  that includes an outer surface  42  that extends between a forward section  44  and an aft section  46 . A combustor casing  48  is coupled to outer surface  42 , and flowsleeve  40  is at least partially positioned within a chamber  50  defined by an inner surface  52  of combustor casing  48 . More specifically, combustor casing  48  is coupled to flowsleeve  40  between forward section  44  and aft section  46 . Forward section  44  is coupled to dome plate  36 , such that chamber  50  is in flow communication with plenum  34  to enable a flow of air from compressor section  14  to be channeled to flowsleeve  40 . A substantially cylindrical combustor liner  54  positioned within flowsleeve  40  is coupled to, and is supported by, flowsleeve  40 . More specifically, in the exemplary embodiment, flowsleeve  40  is coupled radially outwardly from combustor liner  54  such that an annular cooling passage  56  is defined between flowsleeve  40  and combustor liner  54 . Flowsleeve  40  and combustor casing  48  substantially isolate combustor liner  54  and its associated combustion processes from surrounding turbine components. 
     In the exemplary embodiment, combustor liner  54  includes a substantially cylindrically-shaped inner surface  58  that defines an annular combustion chamber  60  that has a centerline axis  62  extending through combustor chamber  60 . Combustor liner  54  is also coupled to fuel nozzles  38  that channels fuel into combustion chamber  60 . Annular cooling passage  56  channels cooling fluid across an outer surface  64  of combustor liner  54  towards fuel nozzles  38 . In the exemplary embodiment, flowsleeve  40  includes an inlet opening  66  that defines a flow path into cooling passage  56 . 
     A transition piece  68  is coupled to combustor liner  54  for use in channeling combustion gases from combustor liner  54  towards turbine section  18 . In the exemplary embodiment, transition piece  68  includes an inner surface  70  that defines a guide cavity  72  that channels combustion gases from combustion chamber  60  downstream to a turbine nozzle  74 . Combustor liner inner surface  58  defines a combustion gas flow path  76  that is substantially parallel to centerline axis  62 . Combustion gases generated within combustion chamber  60  are channeled along path  76  towards transition piece  68 . An upstream end  78  of transition piece  68  is coupled to a downstream end  80  of combustor liner  54 . In one embodiment, combustor liner  54  is at least partially inserted into upstream end  78  such that combustion chamber  60  is positioned in flow communication with guide cavity  72 , and such that combustion chamber  60  and guide cavity  72  are substantially isolated from plenum  34 . 
     An impingement sleeve  82  is spaced radially outwardly from transition piece  68 . More specifically, a downstream end  84  of impingement sleeve  82  is coupled to transition piece  68  such that impingement sleeve  82  is positioned radially outwardly from transition piece  68 , and such that a transition piece cooling passage  86  is defined between impingement sleeve  82  and transition piece  68 . A plurality of openings  88  extending through impingement sleeve  82  enable a portion of air flow from compressor discharge plenum  34  to be channeled into cooling passage  86 . In the exemplary embodiment, an upstream end  90  of impingement sleeve  82  is aligned substantially concentrically with respect to flowsleeve  40  to enable cooling fluid to be channeled from cooling passage  86  into cooling passage  56 . 
     During operation, compressor section  14  is driven by turbine section  18  via shaft  28  (shown in  FIG. 1 ). As compressor section  14  rotates, compressed air  92  is discharged into diffuser  32 . In the exemplary embodiment, the majority of compressed air  92  discharged from compressor section  14  into diffuser  32  is channeled through compressor discharge plenum  34  towards combustor assembly  30 . A smaller portion of compressed air  92  discharged from compressor section  14  is channeled downstream for use in cooling turbine engine  10  components. More specifically, a first flow  94  of pressurized compressed air  92  within plenum  34  is channeled into cooling passage  86  through impingement sleeve openings  88 . The air  94  is then channeled through cooling passage  86  prior to being discharged into cooling passage  56 . In addition, a second flow  96  of pressurized compressed air  92  within plenum  34  is channeled around impingement sleeve  82  and is discharged into cooling passage  56  through inlet opening  66 . Air  96  entering inlet opening  66  and air  94  from transition piece cooling passage  86  is then mixed within cooling passage  56  prior to being discharged from cooling passage  56  towards fuel nozzles  38 . The air  92  is mixed with fuel discharged from fuel nozzles  38  and is ignited within combustion chamber  60  to form a combustion gas stream  98 . Combustion gases  98  are channeled from chamber  60  through transition piece guide cavity  72  towards turbine nozzle  74 . 
       FIG. 3  is a cross-sectional view of an exemplary flowsleeve  100  that may be used with combustor assembly  30 . Identical components shown in  FIG. 3  are labeled with the same reference numbers used in  FIG. 2 . Flowsleeve  100  is substantially cylindrical and includes an inner surface  102  that extends between an upstream end  104  and a downstream end  106 . Upstream end  104  is coupled to dome plate  36  (shown in  FIG. 2 ), and downstream end  106  extends from upstream end  104  towards impingement sleeve  82 . Combustor liner  54  is coupled radially inward from flowsleeve  100  such that cooling passage  56  is defined between flowsleeve inner surface  102  and combustion liner outer surface  64 . Downstream end  106  includes a forward surface  110  that defines an inlet opening  112  that is in flow communication with cooling passage  56  to enable air  96  from combustor plenum  34  (shown in  FIG. 2 ) to cooling passage  56 . 
     In the exemplary embodiment, forward surface  110  includes an upper endwall  114 , a lower endwall  116 , and an inlet plane  119  that extends between upper and lower endwalls  114  and  116 , respectively. Upper endwall  114  is positioned a first distance  117  from fuel nozzles  38 . Lower endwall  116  is positioned a second distance  118  from fuel nozzles  38  that is different than first distance  117  such that inlet plane  119  is oriented obliquely with respect to centerline axis  62 . More specifically, an angle α 1  is defined between an intersection of centerline axis  62  and inlet plane  119 . In the exemplary embodiment, lower endwall  116  is positioned closer to fuel nozzles  38  than upper endwall  114  is, such that angle α 1  is defined between about 90° and about 155° as measured clockwise from centerline axis  62 . In one embodiment, angle α 1  is approximately equal to 135°. Impingement sleeve upstream end  90  includes an upstream edge  120  that defines an upstream opening  122 . Upstream opening  122  enables cooling fluid to be channeled from transition piece cooling passage  86  into cooling passage  56 . In the exemplary embodiment, upstream edge  120  defines an impingement plane  124  that is oriented substantially perpendicularly to centerline axis  62 . Flowsleeve forward surface  110  is positioned with respect to upstream edge  120  such that an annular gap  126  is defined between forward surface  110  and upstream edge  120 . Gap  126  enables air flow from transition piece cooling passage  86  and plenum  34  to cooling passage  56  to be regulated. In the exemplary embodiment, flowsleeve upper endwall  114  is positioned a first distance  130  from upstream edge  120 . Flowsleeve lower endwall  116  is positioned a second distance  132  from upstream edge  120  that is greater than first distance  130 . 
     During operation of turbine engine  10 , cooling air is discharged from plenum  34  such that it substantially circumscribes impingement sleeve  82  and flowsleeve  100 . More specifically, cooling air is channeled from plenum  34  into combustor casing chamber  50  with a non-uniform pressure distribution about flowsleeve  100  and impingement sleeve  82 . Moreover, first flow  94  enters transition piece cooling passage  86  through openings  88  and facilitates cooling transition piece  68  by traveling through transition piece cooling passage  86 . As such, first flow  94  facilitates reducing a temperature of transition piece  68 . First flow  94  flows through annular gap  126  into combustor liner cooling passage  56  to facilitate reducing a temperature of combustor liner  54 . A first portion  134  of second flow  96  flows around impingement sleeve  82  and enters combustor liner cooling passage  56  near lower endwall  116  of inlet opening  112 . A second portion  136  of second flow  96  enters cooling passage  56  near upper endwall  114  of inlet opening  112 . The orientation of inlet opening  112  ensures that first portion  134  and second portion  136  are channeled through cooling passage  56  such that second flow  96  has a substantially uniform flow distribution about combustor liner  54 . Within liner cooling passage  56 , first and second flows  94  and  96  mix and facilitate reducing a temperature of combustor liner  54 . 
     The orientation of flowsleeve inlet opening  112  ensures a substantially uniform flow distribution of second flow  96  is channeled through cooling passage  56 . The uniform flow distribution facilitates enhancing heat transfer between first and second flows  94  and  96  channeled through cooling passage  56  and combustor liner  54 . Annular gap  126  enables first flow  94  to enter combustor cooling passage  56  in a regulated flow. As such, inlet opening  112  and annular gap  126  facilitate a uniform pressure distribution being developed circumferentially about combustor liner outer surface  64 . 
       FIGS. 4-9  are cross-sectional views of various alternative embodiments of flowsleeve  100 . Identical components shown in  FIGS. 4-9  are identified with the same reference numbers used in  FIG. 3 . Referring to  FIG. 4 , in one embodiment, upper endwall  114  is positioned closer to fuel nozzles  38  than lower endwall  116  is such that angle α 1  is defined to be between about 25° and about 90°. In one embodiment, angle α 1  is approximately equal to about 45°. In such an embodiment, impingement sleeve upstream edge  120  is oriented such that impingement plane  124  is oriented obliquely with respect to centerline axis  62  such that first distance  130  is approximately equal to second distance  132 . Moreover, in one embodiment, impingement plane  124  forms an angle α 2  between centerline axis  62  and impingement plane  124  that is approximately equal to inlet plane angle α 1 . Alternatively, angle α 2  may be greater than, or less than, inlet plane angle α 1 . In the exemplary embodiment, a plurality of openings  138  defined in flowsleeve  100  are positioned adjacent to flowsleeve downstream end  106 . Openings  138  are substantially circular and are oriented to facilitate reducing the pressure of air entering cooling passage  56  through openings  138 . 
     Referring to  FIG. 5 , in one embodiment, combustor assembly  30  does not includes impingement sleeve  82 , but rather, combustor liner  54  is coupled to transition piece  68  at a transition section  140 . Flowsleeve  100  extends from dome plate  36  towards transition piece  68  such that flowsleeve inner surface  102  overlaps a portion of an outer surface  142  of transition piece  68 . More specifically, forward surface  110  extends over transition piece upstream end  78  such that cooling passage  56  is at least partially defined between flowsleeve inner surface  102  and transition piece outer surface  142 . In one embodiment, forward surface  110  includes an arcuate surface  144  that extends between upper endwall  114  and lower endwall  116 , such that forward surface  110  forms a substantially concave surface  144  that extends between upper endwall  114  and lower endwall  116 . Alternatively, forward surface  110  may include a substantially convex surface  144  (shown in phantom lines). In one embodiment, flowsleeve  100  extends over an entire length of transition piece  68 , such that flowsleeve  100  extends from dome plate  36  to turbine nozzle  74 . 
     Referring to  FIG. 6 , in one embodiment, flowsleeve forward surface  110  includes an upper portion  146  and a lower portion  148 . In one embodiment, upper portion  146  is coupled to lower portion  148  along centerline axis  62 . In such an embodiment, upper portion  146  extends a distance  150  downstream from lower portion  148 , such that lower portion  148  is positioned closer to fuel nozzles  38  than upper portion  146  is positioned. Moreover, in such an embodiment, upper portion  146  includes an outer edge  152  that is oriented substantially perpendicular to centerline axis  62 . In one embodiment, outer edge  152  is oriented obliquely (shown in phantom lines) with respect to centerline axis  62 . 
     Referring to  FIG. 7 , in one embodiment, upper portion  146  includes an arcuate surface  154 , that extends between upper endwall  114  and lower portion  148 , such that upper portion  146  forms a substantially concave surface  154  that extends between upper endwall  114  and lower portion  148 . In this embodiment, lower portion  148  includes an arcuate surface  156 , that extends between upper portion  146  and lower endwall  166 , such that lower portion  148  forms a substantially convex surface  156  that extends between upper portion  146  and lower endwall  116 . Alternatively, upper portion  146  may include a substantially convex surface  154  (shown in phantom lines), and lower portion  148  may include a substantially concave surface  156  (shown in phantom lines). 
     Referring to  FIG. 8 , in one embodiment, flowsleeve  100  is spaced radially outward from combustor liner  54 , such that upper endwall  114  is spaced a first distance  158  from liner outer surface  64  and lower endwall  116  is spaced a second distance  160  from outer surface  64 . In such an embodiment, second distance  160  is longer than first distance  158 . Moreover, in one embodiment, flowsleeve  100  is positioned such that first distance  158  is longer than second distance  156 . 
     Referring to  FIG. 9 , in one embodiment, flowsleeve  100  includes an outer surface  162  that has an arcuate shape that extends radially outwardly from combustor liner  54  at, or near, forward surface  110 . In such an embodiment, flowsleeve  100  includes a diverging inner surface  102  that defines inlet opening  112  with a bell-shape. A plurality of openings  164  extend through flowsleeve outer surface  162  at, or near, inlet opening  112 . 
     The above-described apparatus and methods overcome disadvantages of known combustor assemblies by providing a flowsleeve that discharges a substantially uniform flow distribution of cooling fluid about a combustor liner to facilitate enhanced heat transfer between the cooling fluid and the combustor liner outer surface. More specifically, by providing a flowsleeve that includes an inlet opening oriented obliquely with respect to a combustor liner centerline axis, a uniform pressure distribution about the combustor liner is facilitated to be increased. In addition, the embodiments described herein facilitate uniformly reducing a temperature across an outer surface of the combustor liner outer surface, which facilitates increasing the operating life of the combustor liner. As such, the cost of maintaining the gas turbine engine system is facilitated to be reduced. 
     Exemplary embodiments of a combustor assembly for use in a turbine engine and methods for assembling the same are described above in detail. The methods and apparatus are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the method may be utilized independently and separately from other components and/or steps described herein. For example, the methods and apparatus may also be used in combination with other combustion systems and methods, and are not limited to practice with only the turbine engine assembly as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other combustion system applications. 
     Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. Moreover, references to “one embodiment” in the above description are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.