Abstract:
A method of assembling a combustor assembly is provided, wherein the method includes providing a combustor liner having a centerline axis and defining a combustion chamber therein, and coupling an annular flowsleeve radially outward from the combustor liner such that an annular flow path is defined substantially circumferentially between the flowsleeve and the combustor liner. The method also includes orienting the flowsleeve such that a plurality of inlets formed within the flowsleeve are positioned to inject cooling air in a substantially axial direction into the annular flow path to facilitate cooling the combustor liner.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines and more particularly, to combustor assemblies for use with gas turbine engines.  
         [0002]     At least some known gas turbine engines use cooling air to cool a combustion assembly within the engine. Moreover, often the cooling air is supplied from a compressor coupled in flow communication with the combustion assembly. More specifically, in at least some known gas turbine engines, the cooling air is discharged from the compressor into a plenum extending at least partially around a transition piece of the combustor assembly. A first portion of the cooling air entering the plenum is supplied to an impingement sleeve surrounding the transition piece prior to entering a cooling channel defined between the impingement sleeve and the transition piece. Cooling air entering the cooling channel is discharged into a second cooling channel defined between a combustor liner and a flowsleeve. The remaining cooling air entering the plenum is channeled through inlets defined within the flowsleeve prior to also being discharged into the second cooling channel.  
         [0003]     Within the second cooling channel, the cooling air facilitates cooling the combustor liner. At least some known flowsleeves include inlets and thimbles that are configured to discharge the cooling air into the second cooling channel at an angle that is substantially perpendicular to the flow of the first portion of cooling air entering the second cooling chamber. More specifically, because of the different flow orientations, the second portion of cooling air loses axial momentum and may create a barrier to the momentum of the first portion of cooling air. The barrier may cause substantial dynamic pressure losses in the air flow through the second cooling channel.  
         [0004]     At least one known approach to decreasing the amount of pressure losses requires resizing the inlets in the existing system. However, this approach may require multiple inlets to be resized at multiple sections of the engine. As such, the economics of this approach may outweigh any potential benefits.  
       BRIEF DESCRIPTION OF THE INVENTION  
       [0005]     In one aspect, a method of assembling a combustor assembly is provided, wherein the method includes providing a combustor liner having a centerline axis and defining a combustion chamber therein, and coupling an annular flowsleeve radially outward from the combustor liner such that an annular flow path is defined substantially circumferentially between the flowsleeve and the combustor liner. The method also includes orienting the flowsleeve such that a plurality of inlets formed within the flowsleeve are positioned to inject cooling air in a substantially axial direction into the annular flow path to facilitate increasing dynamic pressure recovery.  
         [0006]     In another aspect, a combustor assembly is provided, wherein the combustor assembly includes a combustor liner having a centerline axis and defining a combustion chamber therein. The combustor liner also includes an annular flowsleeve coupled radially outward from the combustor liner such that an annular flow path is defined substantially circumferentially between the flowsleeve and the combustor liner. The flowsleeve includes a plurality of inlets configured to inject cooling air therefrom in a substantially axial direction into the annular flow path to facilitate increasing dynamic pressure recovery.  
         [0007]     In a further aspect, a gas turbine engine is provided, wherein the gas turbine engine includes a combustor assembly including a combustor liner having a centerline axis and defining a combustion chamber therein. The combustor assembly also includes an annular flowsleeve coupled radially outward from the combustor liner such that an annular flow path is defined substantially circumferentially between the flowsleeve and the combustor liner. The flowsleeve includes a plurality of inlets configured to inject cooling air therefrom in a substantially axial direction into the annular flow path to facilitate increasing dynamic pressure recovery. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0008]      FIG. 1  is a schematic cross-sectional illustration of an exemplary gas turbine engine;  
         [0009]      FIG. 2  is an enlarged cross-sectional illustration of a portion of an exemplary combustor assembly that may be used with the gas turbine engine shown in  FIG. 1 ;  
         [0010]      FIG. 3  is a perspective view of a known flowsleeve that may be used with the combustor assembly shown in  FIG. 2 ;  
         [0011]      FIG. 4  is a perspective view of an exemplary flowsleeve that may be used with the combustor assembly shown in  FIG. 2 ;  
         [0012]      FIG. 5  is a cross-sectional view of an exemplary flowsleeve and an impingement sleeve/flowsleeve interface that may be used with the combustor assembly shown in  FIG. 2 ; and  
         [0013]      FIG. 6  is a perspective view of an exemplary combustor liner that may be used with the combustor assembly shown in  FIG. 2 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0014]     As used herein, “upstream” refers to a forward end of a gas turbine engine, and “downstream” refers to an aft end of a gas turbine engine.  
         [0015]      FIG. 1  is a schematic cross-sectional illustration of an exemplary gas turbine engine  100 . Engine  100  includes a compressor assembly  102 , a combustor assembly  104 , a turbine assembly  106  and a common compressor/turbine rotor shaft  108 . It should be noted that engine  100  is exemplary only, and that the present invention is not limited to engine  100  and may instead be implemented within any gas turbine engine that functions as described herein.  
         [0016]     In operation, air flows through compressor assembly  102  and compressed air is discharged to combustor assembly  104 . Combustor assembly  104  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 a high temperature combustion gas stream. Combustor assembly  104  is in flow communication with turbine assembly  106 , and discharges the high temperature expanded gas stream into turbine assembly  106 . The high temperature expanded gas stream imparts rotational energy to turbine assembly  106  and because turbine assembly  106  is rotatably coupled to rotor  108 , rotor  108  subsequently provides rotational power to compressor assembly  102 .  
         [0017]      FIG. 2  is an enlarged cross-sectional illustration of a portion of combustor assembly  104 . Combustor assembly  104  is coupled in flow communication with turbine assembly  106  and with compressor assembly  102 . Compressor assembly  102  includes a diffuser  140  and a discharge plenum  142 , that are coupled to each other in flow communication to facilitate channeling air downstream to combustor assembly  104  as discussed further below.  
         [0018]     In the exemplary embodiment, combustor assembly  104  includes a substantially circular dome plate  144  that at least partially supports a plurality of fuel nozzles  146 . Dome plate  144  is coupled to a substantially cylindrical combustor flowsleeve  148  with retention hardware (not shown in  FIG. 2 ). A substantially cylindrical combustor liner  150  is positioned within flowsleeve  148  and is supported via flowsleeve  148 . A substantially cylindrical combustor chamber  152  is defined by liner  150 . More specifically, liner  150  is spaced radially inward from flowsleeve  148  such that an annular combustion liner cooling passage  154  is defined between combustor flowsleeve  148  and combustor liner  150 . Flowsleeve  148  includes a plurality of inlets  156  which provide a flow path into cooling passage  154 .  
         [0019]     An impingement sleeve  158  is coupled substantially concentrically to combustor flowsleeve  148  at an upstream end  159  of impingement sleeve  158 , and a transition piece  160  is coupled to a downstream end  161  of impingement sleeve  158 . Transition piece  160  facilitates channeling combustion gases generated in chamber  152  downstream to a turbine nozzle  174 . A transition piece cooling passage  164  is defined between impingement sleeve  158  and transition piece  160 . A plurality of openings  166  defined within impingement sleeve  158  enable a portion of air flow from compressor discharge plenum  142  to be channeled into transition piece cooling passage  164 .  
         [0020]     In operation, compressor assembly  102  is driven by turbine assembly  106  via shaft  108  (shown in  FIG. 1 ). As compressor assembly  102  rotates, it compresses air and discharges compressed air into diffuser  140  as indicated in  FIG. 2  with a plurality of arrows. In the exemplary embodiment, the majority of air discharged from compressor assembly  102  is channeled through compressor discharge plenum  142  towards combustor assembly  104 , and a smaller portion of air discharged from compressor assembly  102  is channeled downstream for use in cooling engine  100  components. More specifically, a first flow leg  168  of the pressurized compressed air within plenum  142  is channeled into transition piece cooling passage  164  via impingement sleeve openings  166 . The air is then channeled upstream within transition piece cooling passage  164  and discharged into combustion liner cooling passage  154 . In addition, a second flow leg  170  of the pressurized compressed air within plenum  142  is channeled around impingement sleeve  158  and injected into combustion liner cooling passage  154  via inlets  156 . Air entering inlets  156  and air from transition piece cooling passage  164  is then mixed within passage  154  and is then discharged from passage  154  into fuel nozzles  146  wherein it is mixed with fuel and ignited within combustion chamber  152 .  
         [0021]     Flowsleeve  148  substantially isolates combustion chamber  152  and its associated combustion processes from the outside environment, for example, surrounding turbine components. The resultant combustion gases are channeled from chamber  152  towards and through a transition piece combustion gas stream guide cavity  160  that channels the combustion gas stream towards turbine nozzle  174 .  
         [0022]      FIG. 3  is a perspective view of a known flowsleeve  200  that may be used with combustor assembly  104 . Flowsleeve  200  is substantially cylindrical and includes an upstream end  202  and a downstream end  204 . Upstream end  202  is coupled to dome plate  144  (shown in  FIG. 2 ) and downstream end  204  is coupled to impingement sleeve  158  (shown in  FIG. 2 ). Combustor liner  150  (shown in  FIG. 2 ) is coupled radially inward from flowsleeve  200  such that cooling passage  154  (shown in  FIG. 2 ) is defined between flowsleeve  200  and combustor liner  150 .  
         [0023]     Flowsleeve  200  also includes a plurality of inlets  206  and thimbles  208  defined adjacent downstream end  204 . Inlets  206  and thimbles  208  are substantially circular and are oriented substantially perpendicular to a flowsleeve center axis  210 . Furthermore, thimbles  208  extend substantially radially inward from flowsleeve  200  such that airflow is discharged from thimbles  208  and inlets  206  from around impingement sleeve  158 , radially inward through flowsleeve  200 , and into combustion liner cooling passage  154 . The radial flow direction of airflow entering passage  154  through inlets  206  and thimbles  208  substantially reduces the axial momentum of airflow and creates a barrier to air flowing within passage  154  from transition piece cooling passage  164 . Furthermore, the radial length of thimbles  208  creates an obstruction to airflow channeled from transition piece cooling passage  164 . As such, a pressure drop of the airflow results within combustion cooling passage  154 . The resulting pressure drop may cause disproportional cooling around combustor liner  150 .  
         [0024]      FIG. 4  is a perspective view of an exemplary embodiment of a flowsleeve  250  that may be used with combustor assembly  104 . Flowsleeve  250  is substantially cylindrical and includes an upstream end  252  and a downstream end  254 . Upstream end  252  is coupled to dome plate  144  (shown in  FIG. 2 ) and downstream end  254  is coupled to impingement sleeve  158  (shown in  FIG. 2 ). Combustor liner  150  (shown in  FIG. 2 ) is coupled radially inward from flowsleeve  250  such that combustion liner cooling passage  154  (shown in  FIG. 2 ) is defined between flowsleeve  250  and combustor liner  150 .  
         [0025]     Flowsleeve  250  also includes a plurality of injectors  256  spaced circumferentially about flowsleeve  250  at a distance  258  upstream from downstream end  254 . In the exemplary embodiment, injectors  256  are substantially circular and each has a large length/diameter ratio. In an alternative embodiment, injectors  256  are substantially rectangular slots having a width that is larger than a slot height. Moreover, injectors  256  are configured to substantially axially eject airflow from around impingement sleeve  158  through flowsleeve  250  and into combustion liner cooling passage  154 . More specifically, airflow ejected from injectors  256  enters passage  154  in a generally axial direction that is substantially tangential to a direction of flow discharged into passage  154  from airflow channeled into passage  154  from passage  164 , and in substantially the same direction as airflow channeled into passage  154  from passage  164 . Furthermore, injectors  256  are configured to accelerate airflow ejected therefrom. An annular gap (not shown) is defined between flowsleeve  250  and combustor liner  150  within distance  258 . Injectors  256  and the annular gap facilitate regulating pressure in airflow entering combustion liner cooling passage  154 .  
         [0026]      FIG. 5  is a cross-sectional view of flowsleeve  250  and an impingement sleeve/flowsleeve interface  300 . Specifically,  FIG. 5  illustrates the interface  300  defined between the coupling of flowsleeve  250  and impingement sleeve  158 . Furthermore  FIG. 5  illustrates a cross-sectional view of the axial injection geometry of injectors  256 . Specifically, flowsleeve  250  is oriented such that injectors  256  are positioned an axial distance  302  upstream from interface  300 . As such, an annular gap  304  defined at the intersection region of flowsleeve  250  and impingement sleeve  158  has an axial length  302 . Annular gap  304  facilitates regulating air flow from transition piece cooling passage  164 .  
         [0027]      FIG. 6  is a perspective view of an exemplary combustor liner  350  that may be used with combustor assembly  104 . Combustor liner  350  is substantially cylindrical and includes an upstream end  352  and a downstream end  354 . In the exemplary embodiment, upstream end  352  has a radius R 1  that is substantially larger than a radius R 2  of downstream end  354 . Upstream end  352  receives a fuel/air mixture from fuel nozzles  146  and discharges the fuel/air mixture into transition piece  160 . Combustor liner  350  is oriented within flowsleeve  250  such that flowsleeve  250  and combustor liner  350  define combustion liner cooling passage  154 . Cooling air received in combustion liner cooling passage  154  is channeled upstream and across a surface  356  of combustor liner  350  to facilitate cooling combustor liner  350 .  
         [0028]     Combustor liner surface  356  is configured with a plurality of grooves  358  defined thereon that facilitate circumferentially distributing the airflow from injectors  256  across liner surface  356 . In the exemplary embodiment, grooves  358  are configured in a criss-crossed pattern across a length L 1  of combustor liner surface  356  such that diamond shaped raised portions  359  are defined between grooves  358 . In alternative embodiments, grooves  358  may be configured in other geometrical patterns.  
         [0029]     During operation of engine  100  cooling air is discharged from plenum  142  such that it substantially surrounds impingement sleeve  158 . First flow leg  168  enters transition piece cooling passage  164  through openings  166 . First flow leg  168  cools transition piece  160  by traveling upstream through transition piece cooling passage  164 . First flow leg  168  continues through annular gap  304  and discharges into combustion liner cooling passage  154 . Second flow leg  170  flows around impingement sleeve  158  and enters combustion liner cooling passage  154  through injectors  256 . Within combustion liner cooling passage  154 , the first and second flow legs  168  and  170  mix and continue upstream to facilitate cooling combustor liner  350 .  
         [0030]     The configuration of injectors  256  increases the velocity of cooling air within second flow leg  170 . The increased velocity facilitates enhanced heat transfer between the cooling air and combustor liner  350 . Annular gap  304  facilitates regulating flow of first flow leg  168  into combustion cooling passage  154 . As such, injectors  256  and annular gap  304  facilitate balancing the pressure and velocity of the two flow legs  168  and  170  such that a balanced flow path results from the mixing of the two flow paths.  
         [0031]     Furthermore, due to the axial configuration of injectors  256 , the second flow leg  170  does not create an air darn which restricts the flow of first flow leg  168 . As a result, the axial configuration of injectors  256  facilitates increasing dynamic pressure recovery within the resultant flow path. By balancing pressure loss and velocity within combustion liner cooling passage  154 , injectors  256  and annular gap  304  facilitate substantially uniform heat transfer between combustor liner  350  and the cooling air.  
         [0032]     Moreover, grooves  358  of combustor liner surface  356  facilitate enhancing the heat transfer between cooling air and combustor liner  350 . Specifically, grooves  358  facilitate circumferentially distributing cooling air from injectors  256  and facilitate creating a uniform heat transfer coefficient distribution across the length and circumference of combustor liner  350 . In addition, grooves  358  facilitate allowing high velocity cooling air to facilitate improving heat transfer.  
         [0033]     The above-described apparatus and methods facilitate providing constant heat transfer between cooling air and a combustor liner, while maintaining an overall pressure of the gas turbine engine. Specifically, the injectors facilitate reducing pressure losses by injecting the cooling air of the second flow leg axially such that dynamic pressure recovery is increased between the first and second flow leg. Furthermore, the enhancements to the combustor liner facilitate greater heat exchange between the combustor liner and the cooling air.  
         [0034]     As used herein, an element or step recited in the singular and proceeded with the word “a” or “an” should be understood as not excluding plural said elements or steps, unless such exclusion is explicitly recited. Furthermore, references to “one embodiment” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features.  
         [0035]     Although the apparatus and methods described herein are described in the context of a combustor assembly for a gas turbine engine, it is understood that the apparatus and methods are not limited to combustor assemblies or gas turbine engines. Likewise, the combustor assembly components illustrated are not limited to the specific embodiments described herein, but rather, components of the combustor assembly can be utilized independently and separately from other components described herein.  
         [0036]     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.