Abstract:
A method of operating a gas turbine engine includes providing an inlet flow conditioner (IFC). The IFC has an annular chamber defined therein by at least one wall wherein the wall includes a plurality of perforations extending therethrough. The perforations are spaced in at least two axially-spaced rows that extend circumferentially about the wall. The method also includes channeling a fluid into the IFC and discharging the fluid from the IFC with a substantially uniform flow profile.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to rotary machines and more particularly, to gas turbine engines and methods of operation. 
         [0002]    At least some gas turbine engines ignite a fuel-air mixture in a combustor and generate a combustion gas stream that is channeled to a turbine via a hot gas path. Compressed air is channeled to the combustor by a compressor. Combustor assemblies typically have fuel nozzles that facilitate fuel and air delivery to a combustion region of the combustor. The turbine converts the thermal energy of the combustion gas stream to mechanical energy that rotates a turbine shaft. The output of the turbine may be used to power a machine, for example, an electric generator or a pump. 
         [0003]    Some known fuel nozzles include at least one inlet flow conditioner (IFC). Typically, an IFC includes a plurality of perforations and is configured to channel air from the compressor into a portion of the fuel nozzle to facilitate mixing of fuel and air. One known engine channels air into the fuel nozzle to facilitate mitigating air turbulence and to produce a radial and circumferential air flow velocity profile that is substantially uniform within the IFC. Some known IFCs include at least one flow vane that facilitates the generation of a non-uniform radial air flow velocity profile within some portions of the IFC. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0004]    In one aspect, a method of operating a gas turbine engine is provided. The method includes providing an inlet flow conditioner (IFC) having an annular chamber defined therein by at least one wall that is formed with a plurality of perforations extending therethrough. The plurality of perforations are spaced in at least two axially-spaced rows that extend substantially circumferentially about the wall. The method also includes channeling a fluid into the IFC and discharging the fluid from the IFC with a substantially uniform flow profile 
         [0005]    In another aspect, an inlet flow conditioner (IFC) is provided. The IFC includes an annular chamber at least partially defined therein by a first wall that includes a plurality of perforations extending therethrough. The plurality of perforations are spaced equidistantly circumferentially from each other and are configured to channel a fluid such that a substantially uniform flow profile of the fluid is discharged from the at least one chamber. 
         [0006]    In a further aspect, a gas turbine engine is provided. The engine includes a compressor and a combustor in flow communication with the compressor. The combustor includes a fuel nozzle assembly that includes an inlet flow conditioner (IFC). The IFC includes an annular IFC chamber at least partially defined therein by a first wall that includes a plurality of perforations extending therethrough. The plurality of perforations are spaced equidistantly circumferentially from each other and are configured to channel a fluid such that a substantially uniform flow profile discharges from the annular IFC chamber. 
     
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  is a schematic view of an exemplary gas turbine engine; 
           [0008]      FIG. 2  is a cross-sectional schematic view of an exemplary combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0009]      FIG. 3  is a cross-sectional schematic view of an exemplary fuel nozzle assembly that may be used with the combustor shown in  FIG. 2 ; 
           [0010]      FIG. 4  is a fragmentary view of an exemplary inlet flow conditioner (IFC) that may be used with the fuel nozzle assembly shown in  FIG. 3 ; and 
           [0011]      FIG. 5  is an axial cross-sectional view of the IFC shown in  FIG. 4  facing downstream and illustrating a first axial flow stream; 
           [0012]      FIG. 6  is an axial cross-sectional view of the IFC shown in  FIG. 4  facing downstream and illustrating a second axial flow stream; and 
           [0013]      FIG. 7  is an axial cross-sectional view of the IFC shown in  FIG. 4  facing downstream and illustrating a third axial flow stream. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0014]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine  100 . Engine  100  includes a compressor  102  and a plurality of combustors  104 . Combustor  104  includes a fuel nozzle assembly  106 . Engine  100  also includes a turbine  108  and a common compressor/turbine shaft  110  (sometimes referred to as rotor  110 ). In one embodiment, engine  100  is a MS9001H engine, sometimes referred to as a 9H engine, commercially available from General Electric Company, Greenville, S.C. 
         [0015]    In operation, air flows through compressor  102  and compressed air is supplied to combustors  104 . Specifically, the compressed air is supplied to fuel nozzle assembly  106 . Fuel is channeled to a combustion region wherein the fuel is mixed with the air and ignited. Combustion gases are generated and channeled to turbine  108  wherein gas stream thermal energy is converted to mechanical rotational energy. Turbine  108  is rotatably coupled to, and drives, shaft  110 . 
         [0016]      FIG. 2  is a cross-sectional schematic view of combustor  104 . Combustor assembly  104  is coupled in flow communication with turbine assembly  108  and with compressor assembly  102 . Compressor assembly  102  includes a diffuser  112  and a compressor discharge plenum  114  that are coupled in flow communication to each other. 
         [0017]    In the exemplary embodiment, combustor assembly  104  includes a endcover  120  that provides structural support to a plurality of fuel nozzles  122 . Endcover  120  is coupled to combustor casing  124  with retention hardware (not shown in  FIG. 2 ). A combustor liner  126  is positioned within and is coupled to casing  124  such that a combustion chamber  128  is defined by liner  126 . An annular combustion chamber cooling passage  129  extends between combustor casing  124  and combustor liner  126 . 
         [0018]    A transition portion or piece  130  is coupled to combustor casing  124  to facilitate channeling combustion gases generated in chamber  128  towards turbine nozzle  132 . In the exemplary embodiment, transition piece  130  includes a plurality of openings  134  formed in an outer wall  136 . Piece  130  also includes an annular passage  138  defined between an inner wall  140  and outer wall  136 . Inner wall  140  defines a guide cavity  142 . 
         [0019]    In operation, compressor assembly  102  is driven by turbine assembly  108  via shaft  110  (shown in  FIG. 1 ). As compressor assembly  102  rotates, compressed air is discharged into diffuser  112  as the associated arrows illustrate. In the exemplary embodiment, the majority of air discharged from compressor assembly  102  is channeled through compressor discharge plenum  114  towards combustor assembly  104 , and a smaller portion of compressed air may be channeled for use in cooling engine  100  components. More specifically, the pressurized compressed air within plenum  114  is channeled into transition piece  130  via outer wall openings  134  and into passage  138 . Air is then channeled from transition piece annular passage  138  into combustion chamber cooling passage  129 . Air is discharged from passage  129  and is channeled into fuel nozzles  122 . 
         [0020]    Fuel and air are mixed and ignited within combustion chamber  128 . Casing  124  facilitates isolating combustion chamber  128  and its associated combustion processes from the outside environment, for example, surrounding turbine components. Combustion gases generated are channeled from chamber  128  through transition piece guide cavity  142  towards turbine nozzle  132 . 
         [0021]      FIG. 3  is a cross-sectional schematic view of fuel nozzle assembly  122 . In the exemplary embodiment, an air atomized liquid fuel nozzle (not shown) coupled to assembly  122  to provide dual fuel capability has been omitted for clarity. Assembly  122  has a centerline axis  143  and is coupled to endcover  120  (shown in  FIG. 2 ) via fuel nozzle flange  144 . 
         [0022]    Fuel nozzle assembly  122  includes a convergent tube  146  that is coupled to flange  144 . Tube  146  includes a radially outer surface  148 . Assembly  122  also includes a radially inner tube  150  that is coupled to flange  144  via a tube-to-flange bellows  152 . Bellows  152  facilitates compensating for varying rates of thermal expansion between tube  150  and flange  144 . Tubes  146  and  150  define a substantially annular first premixed fuel supply passage  154 . Assembly  122  also includes a substantially annular inner tube  156  that defines a second premixed fuel supply passage  158  in cooperation with radially inner tube  150 . Inner tube  156  partially defines a diffusion fuel passage  160  and is coupled to flange  144  via an air tube-to-flange bellows  162  that facilitates compensating for varying rates of thermal expansion between tube  156  and flange  144 . Passages  154 ,  158 , and  160  are coupled in flow communication to fuel sources (not shown in  FIG. 3 ). In one embodiment, passage  160  receives the air atomized liquid fuel nozzle therein. 
         [0023]    Assembly  122  includes a substantially annular inlet flow conditioner (IFC)  164 . IFC  164  includes a radially outer wall  166  that includes a plurality of perforations  168 , and an end wall  170  that is positioned on an aft end of IFC  164  and extends between wall  166  and surface  148 . Walls  166  and  170  and surface  148  define a substantially annular IFC chamber  172  therein. Chamber  172  is in flow communication with cooling passage  129  (shown in  FIG. 2 ) via perforations  168 . Assembly  122  also includes a tubular transition piece  174  that is coupled to wall  166 . Transition piece  174  defines a substantially annular transition chamber  176  that is substantially concentrically aligned with respect to chamber  172  and is positioned such that an IFC outlet passage  178  extends between chambers  172  and  176 . 
         [0024]    Assembly  122  also includes an air swirler assembly or swozzle assembly  180  for use with gaseous fuel injection. Swozzle  180  includes a substantially tubular shroud  182  that is coupled to transition piece  174 , and a substantially tubular hub  184  that is coupled to tubes  146 ,  150 , and  156 . Shroud  182  and hub  184  define an annular chamber  186  therein wherein a plurality of hollow turning vanes  188  extend between shroud  182  and hub  184 . Chamber  186  is coupled in flow communication with chamber  176 . Hub  184  defines a plurality of primary turning vane passages (not shown in  FIG. 3 ) that are coupled in flow communication with premixed fuel supply passage  154 . A plurality of premixed gas injection ports (not shown in  FIG. 3 ) are defined within hollow turning vanes  188 . Similarly, hub  184  defines a plurality of secondary turning vane passages (not shown in  FIG. 3 ) that are coupled in flow communication with premixed fuel supply passage  158  and a plurality of secondary gas injection ports (not shown in  FIG. 3 ) that are defined within turning vanes  188 . Inlet chamber  186 , and the primary and secondary gas injection ports, are coupled in flow communication with an outlet chamber  190 . 
         [0025]    Assembly  122  further includes a substantially annular fuel-air mixing passage  192  that is defined by a tubular shroud extension  194  and a tubular hub extension  196 . Passage  192  is coupled in flow communication with chamber  190  and extensions  194  and  196  are each coupled to shroud  182  and hub  184 , respectively. 
         [0026]    A tubular diffusion flame nozzle assembly  198  is coupled to hub  184  and partially defines annular diffusion fuel passage  160 . Assembly  198  also defines an annular air passage  200  in cooperation with hub extension  196 . Assembly  122  also includes a slotted gas tip  202  that is coupled to hub extension  196  and assembly  198 , and that includes a plurality of gas injectors  204  and air injectors  206 . Tip  202  is coupled in flow communication with, and facilitates fuel and air mixing in, combustion chamber  128 . 
         [0027]    In operation, fuel nozzle assembly  122  receives compressed air from cooling passage  129  (shown in  FIG. 2 ) via a plenum (not shown in FIG.  3 ) surrounding assembly  122 . Most of the air used for combustion enters assembly  122  via IFC  164  and is channeled to premixing components. Specifically, air enters IFC  164  via perforations  168  and mixes within chamber  172  and air exits IFC  164  via passage  178  and enters swozzle inlet chamber  186  via transition piece chamber  176 . A portion of high pressure air entering passage  129  is also channeled into an air-atomized liquid fuel cartridge (not shown in  FIG. 3 ) inserted within diffusion fuel passage  160 . 
         [0028]    Fuel nozzle assembly  122  receives fuel from a fuel source (not shown in  FIG. 3 ) via premixed fuel supply passage  154  and  158 . Fuel is channeled from premixed fuel supply passage  154  to the plurality of primary gas injection ports defined within turning vanes  188 . Similarly, fuel is channeled from premixed fuel supply passage  158  to the plurality of secondary gas injection ports defined within turning vanes  188 . 
         [0029]    Air channeled into swozzle inlet chamber  186  from transition piece chamber  176  is swirled via turning vanes  188  and is mixed with fuel, and the fuel/air mixture is channeled to swozzle outlet chamber  190  for further mixing. The fuel and air mixture is then channeled to mixing passage  192  and discharged from assembly  122  into combustion chamber  128 . In addition, diffusion fuel channeled through diffusion fuel passage  160  is discharged through gas injectors  204  into combustion chamber  128  wherein it mixes and combusts with air discharged from air injectors  206 . 
         [0030]      FIG. 4  is a fragmentary view of IFC  164 . Centerline axis  143 , transition piece  174  and swozzle shroud  182  are illustrated for perspective.  FIG. 5  is an axial cross-sectional view of exemplary IFC  164  facing downstream and illustrating a first axial flow stream  212 . Centerline axis  143 , diffusion fuel passage  160 , tube  156 , premixed fuel supply passage  158 , radially inner tube  150 , premixed fuel supply passage  154 , convergent tube  146 , and convergent tube radially outer surface  148  are illustrated for perspective. Only six circumferentially spaced perforations  168  are illustrated in  FIG. 5 . Alternatively, IFC  164  may include any number of perforations  168 . IFC  164  includes radially outer wall  166  that defines plurality of substantially circular perforations  168 . In the exemplary embodiment, IFC  164  includes six axially spaced rows  207  of perforations  168 . For example, in  FIG. 4 , first, second and third circumferential perforation rows  208 ,  214  and  220 , respectively, are identified. Alternatively, IFC  164  may include any number of axially-spaced rows  207  of perforations  168 . 
         [0031]    In the exemplary embodiment, perforations  168  are each formed substantially identical in diameter D 1  and the axially-spaced rows  207  are oriented such that six perforations are substantially axially aligned. Moreover, in the exemplary embodiment, perforations  168  are spaced substantially equally circumferentially and axially. The exemplary orientation of perforations  168  facilitates mitigating a pressure drop across IFC  164  that subsequently facilitates improving engine efficiency. Alternatively, IFC  164  may include any number of perforations  168  arranged in any orientation that enables IFC  164  to function as described herein. 
         [0032]    IFC  164  may also include an end wall  170  that is positioned on an aft end of IFC  164  extending between wall  166  and surface  148 . IFC  164  may be coupled to tube  146  such that walls  166  and  170 , and surface  148  define an annular IFC chamber  172  therein. Chamber  172  is coupled in flow communication with combustion chamber cooling passage  129  (shown in  FIG. 2 ) via perforations  168 . 
         [0033]    In operation, compressed air from passage  129  flows around IFC  164 . Perforations  168  facilitate increasing the backpressure around an outer periphery of IFC  164  by restricting air flow into IFC  164 . The increased backpressure facilitates substantially equalizing air flow through perforations  168 . For example, air flows through perforations  208  and enters chamber  172  in a plurality of radial air streams  210  (only three illustrated in  FIG. 4  and only six illustrated in  FIG. 5 ). A substantial portion of each air stream  210  impinges against surface  148  and change direction to substantially fill that portion of chamber  172  defined between row  208  and end cap  170 . As such, static pressure is generated within that portion of chamber  172 . Another portion of radial air streams  210  that impinge surface  148  change direction and are channeled towards transition piece  174 . Radial air streams  210  form a boundary layer of air over a portion of surface  148  such that a plurality of axial air streams  212  (only six illustrated in  FIG. 5 ) are formed and are defined with a first radial and circumferential velocity profile within chamber  172 . Axial air streams  212  that are formed tend to flow substantially parallel to the row of perforations  208  that admitted the first radial air streams  210 . A lesser portion of air streams  212  flow into that portion of chamber  172  defined between perforations  208 . Air streams  212  tend to expand in the radial and circumferential directions as they travel towards transition piece  174 . As such, the radial and circumferential velocity profile of air streams  212  is substantially non-uniform. 
         [0034]      FIG. 6  is an axial cross-sectional view of IFC  164  facing downstream, and illustrating a second axial flow stream  218 . Centerline axis  143 , diffusion fuel passage  160 , inner tube  156 , premixed fuel supply passage  158 , radially inner tube  150 , premixed fuel supply passage  154 , convergent tube  146 , and convergent tube radially outer surface  148  are illustrated for perspective. For clarity, only six perforations  168  are illustrated in  FIG. 6 . Air flows through second row  214  and enters chamber  172  in a plurality of radial air streams  216  (only three are illustrated in  FIG. 4  and only six are illustrated in  FIG. 6 ). A substantial portion of air streams  216  impinges against surface  148  and air streams  212  such that a plurality of second axial air streams  218  are formed that have a second radial and circumferential velocity profile within chamber  172 . Axial air streams  218  tend to form such that circumferential regions of chamber  172  defined between axial perforations  208  and  214  fill in with flowing air. This action thereby decreases the difference in mass flow between the portion of air streams  218  directly under perforations  168  and the portion of air streams  218  between circumferentially adjacent perforations  168 . Air streams  218  flowing towards transition piece  174  tend to expand in the radial and circumferential directions. Therefore, in general, the radial and circumferential velocity profile of air streams  218  is more uniform than the velocity profile of air streams  212 . 
         [0035]      FIG. 7  is an axial cross-sectional view of IFC  164  facing downstream and illustrating a third axial flow stream  224 . Centerline axis  143 , diffusion fuel passage  160 , inner tube  156 , premixed fuel supply passage  158 , radially inner tube  150 , premixed fuel supply passage  154 , convergent tube  146 , and convergent tube radially outer surface  148  are illustrated for perspective. For clarity, only six perforations  168  are illustrated in  FIG. 7 . Air flows through third row  220  and enters chamber  172  in a plurality of radial air streams  222  (only three are illustrated in  FIG. 4  and only six are illustrated in  FIG. 7 ). A first portion of each air stream  222  impinges against surface  148  and a second portion of each air stream  222  impinges air streams  218  such that a plurality of third axial air streams  224  are formed that have a third radial and circumferential velocity profile within chamber  172 . Axial air streams  224  tend to form such that circumferential regions of chamber  172  defined between perforations  208 ,  214  and  220  fill in with flowing air. This action thereby further decreases the difference in mass flow between the portion of air streams  224  directly under perforations  168  and the portion of air streams  224  between circumferentially adjacent perforations  168 . Air streams  224  flowing towards transition piece  174  tend to expand in the radial and circumferential directions. In general, the radial and circumferential velocity profile of air streams  224  is more uniform than the velocity profile of air streams  218 . 
         [0036]    The iterative process of subsequent radial streams impinging on the composite axial streams induces a flow velocity profile into the air flowing within chamber  172  across IFC outlet passage  178  (shown in  FIG. 3 ) into transition piece  174  that is substantially constant in the radial direction across passage  178 . The substantially uniform velocity profile of air facilitates reducing pockets of rich, or excess, air within fuel nozzle  122  and combustion chamber  142  that subsequently facilitates a reduction in formation of undesirable combustion byproducts, such as NO x . Similarly, the substantially uniform velocity profile of air facilitates reducing pockets of lean air within fuel nozzle  122  and combustion chamber  142  thereby facilitating increased flame stability. 
         [0037]    The methods and apparatus for assembling and operating a combustor described herein facilitates operation of a gas turbine engine. More specifically, the inlet flow conditioner facilitates a more uniform air flow velocity profile being induced within the fuel nozzle assembly. Such air flow profile facilitates efficiency of combustion and a reduction in undesirable combustion by-products. Moreover, the inlet flow conditioner facilitates reducing capital and maintenance costs, as well as increasing operational reliability. 
         [0038]    Exemplary embodiments of inlet flow conditioners as associated with gas turbine engines are described above in detail. The methods, apparatus and systems are not limited to the specific embodiments described herein nor to the specific illustrated inlet flow conditioner. 
         [0039]    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.