Abstract:
A gas turbine combustor includes a combustor liner enclosing a combustion chamber; at least one fuel nozzle arranged to provide fuel to the combustion chamber; a flow sleeve surrounding the combustor liner forming a passage radially between the combustor liner and the flow sleeve for supplying air to the combustion chamber, the flow sleeve configured to permit air to flow substantially axially into the passage via a substantially annular flow sleeve inlet. A downstream end of the flow sleeve is formed to include an annular manifold provided with plural outlets about a circumference of the downstream end of the flow sleeve and adapted to supply supplemental air from an external variable air source substantially radially into the passage to thereby maintain axial air flow boundary layer attachment at the flow sleeve inlet.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to gas turbine engines and more particularly, to gas turbine combustor assemblies. 
         [0002]    At least some known gas turbine engines use cooling air to cool a combustion assembly within the engine. Moreover, oftentimes the cooling air is supplied from a compressor coupled in flow communication with the combustion assembly. For example, 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 radially between the impingement sleeve and the transition piece. Cooling air entering the cooling channel is discharged into a second, aligned cooling channel radially between a combustor liner and a flow sleeve. The remaining cooling air entering the plenum is channeled through inlets formed within the flow sleeve and also discharged into the second cooling channel. 
         [0003]    In combustion systems for industrial-turbines generally as described above, multiple independent combustor cans are equally spaced around the centerline of the machine. These are referred to “can-annular” systems. There are many benefits to these types of systems, but one of the drawbacks is that each combustion can will experience a different amount of inlet airflow due to factors such as part tolerances, obstructions in the inlet, non-integral stage-one nozzle counts downstream of the combustors, etc. Also, the amount of fuel entering each combustor can will be slightly different due to variation in fuel nozzle effective areas, obstacles in piping and manifolds, etc. As a result, each can will have a different fuel/air ratio, which can drive variations in performance of each can. In most cases, it is desirable to have all of the cans perform as close to the design point as possible. 
         [0004]    As mentioned above, most combustion systems use a reverse-flow cooling arrangement meaning that the air that will eventually be used for combustion is first used for cooling. Impingement jets, turbulators and other types of structures are used to augment heat transfer at the expense of pressure drop. In a good combustion system design, most of the pressure drop is directly used for cooling, and very little is due to flow separations and other sources of pressure loss that do not influence cooling. 
         [0005]    In state-of-the-art air cooled combustion systems, an axial-injection-type flow sleeve surrounds each combustor liner and air that is not used to cool the transition piece (for example, by impingement cooling) is injected along the centerline of the combustor so as not to waste any of the potential energy used to force the air inside the liner/flow sleeve annulus as occurs with radial-injection flow sleeves. 
         [0006]    The air is intended to enter the annulus between the flow sleeve and the combustor liner cleanly; however, it has been found that flow separation occurs at the inlet, along the inside surface of the flow sleeve, causing increased pressure loss. One solution to this problem utilizes surface contouring on the inside surface of the flow sleeve which allows the flow to stay attached and then diffuse over some length within the annulus. 
         [0007]    While this solution is satisfactory in some respects, it has been determined that it would be desirable to control the flow of air at the entry to the annulus between the combustor liner and flow sleeve so as to modify flow behavior at the inlet anywhere between fully-separated and fully-attached conditions. 
       BRIEF SUMMARY OF THE INVENTION 
       [0008]    In accordance with an exemplary but nonlimiting embodiment, the invention provides a gas turbine combustor comprising a combustor liner enclosing a combustion chamber; at least one fuel nozzle arranged to provide fuel to the combustion chamber; a flow sleeve surrounding the combustor liner forming a passage radially between the combustor liner and the flow sleeve for supplying air to the combustion chamber, the flow sleeve configured to permit air to flow substantially axially into the passage via a substantially annular flow sleeve inlet; a downstream end of the flow sleeve formed to include an annular manifold provided with plural outlets about a circumference of the downstream end of the flow sleeve and adapted to supply supplemental air from an external variable air source substantially radially into the passage to thereby maintain axial air flow boundary layer attachment at the flow sleeve inlet. 
         [0009]    In another aspect, the invention provides a can-annular combustor arrangement for a gas turbine comprising plural combustors arranged in an annular array about a turbine rotor, the plural combustors adapted to supply combustion gases to a first stage of the gas turbine; each combustor comprising a combustor liner enclosing a combustion chamber, at least one fuel nozzle arranged to provide fuel to the combustion chamber, a flow sleeve surrounding the combustor liner forming a passage radially between the combustor liner and the flow sleeve for supplying air to the combustion chamber; wherein a downstream end of the flow sleeve is formed to include an annular manifold provided with plural outlets about a circumference of the downstream end of the flow sleeve and adapted to supply supplemental air from an external variable air source to the passage. 
         [0010]    In still another aspect, the invention provides a method of controlling flow of air to any one or all of a plurality of combustors in a can-annular array of combustors about a gas turbine rotor, where each combustor includes a liner enclosing a combustion chamber and supports at least one nozzle for supplying fuel to the combustion chamber; and a flow sleeve surrounding the combustor liner, with an annular passage extending between the liner and the flow sleeve for supplying compressor discharge air to the combustion chamber via an axially-oriented inlet at the downstream end of the flow sleeve, the method comprising supplying supplemental air under pressure selectively to the annular passage of each of the plurality of combustors; and modulating flow of the supplemental air to control a fuel/air ratio for any one or all of the plurality of combustors. 
         [0011]    The invention will now be described in detail in connection with the drawings identified below. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0012]      FIG. 1  is a schematic cross-sectional illustration of an exemplary gas turbine engine. 
           [0013]      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 ; 
           [0014]      FIG. 3  is an enlarged partial cross section illustrating a known combustor flow sleeve inlet configuration showing an idealized flow pattern for air flowing axially into an annulus between the flow sleeve and combustor liner; 
           [0015]      FIG. 4  is a view similar to  FIG. 3  but showing an actual flow pattern for the air entering the annulus; 
           [0016]      FIG. 5  is a partial cross section of a combustor flow sleeve in accordance with an exemplary but nonlimiting embodiment of the invention supplying supplemental air to the annulus and its effect on the air flowing axially into the annulus; 
           [0017]      FIG. 6  is a view similar to  FIG. 5  but showing the flow pattern when the supplemental air supply system is in the “off” position; and 
           [0018]      FIG. 7  is a schematic diagram illustrating a supplemental air supply system for a can-annular array of combustors in accordance with an exemplary but nonlimiting embodiment of the invention. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0019]    At the outset, it is noted that, as used herein, “upstream” refers to a forward end of a gas turbine engine or other component in the combination gas flow path, and “downstream” refers to an aft end of a gas turbine engine or other component in the combustion gas flow path. 
         [0020]      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 may instead be implemented within any gas turbine engine that functions generally as described herein. 
         [0021]    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 the compressor assembly  102  and the 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 . 
         [0022]      FIG. 2  is an enlarged cross-sectional illustration of a portion of the compressor assembly  102  and the combustor assembly  104 . 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 the combustor assembly  104 . 
         [0023]    In the exemplary embodiment, combustor assembly  104  includes a substantially circular endcover or cover plate  144  that at least partially supports a plurality of fuel nozzles  146 . The cover plate  144  is coupled to a substantially cylindrical combustor flow sleeve  148  with retention hardware (not shown). A substantially cylindrical combustor liner  150  is positioned within the flow sleeve  148  and is supported via the flow sleeve. A substantially cylindrical combustor chamber  152  is defined by liner  150 . More specifically, liner  150  is spaced radially inward from flow sleeve  148  such that an annular combustion liner cooling passage or annulus  154  is defined between combustor flow sleeve  148  and combustor liner  150 . Flow sleeve  148  includes a plurality of inlets  156  which provide an axially-oriented flow path into the cooling passage or annulus  154 . 
         [0024]    In some turbine configurations, an impingement sleeve  158  is coupled to the combustor flow sleeve  148  at an upstream end  159  of the impingement sleeve  158 , and substantially surrounds a transition piece or duct  160  that channels the combustion gases generated in chamber  152  to the turbine, represented by the turbine nozzle  174 . A transition piece cooling passage or annulus  164  is thus defined between the impingement sleeve  158  and the 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 directed radially into transition piece cooling passage or annulus  164  where it flows along and through the annulus  164  and continues into the passage or annulus  154 . 
         [0025]    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 the diffuser  140  as indicated in  FIG. 2  by a plurality of flow 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 components of the engine  100 . More specifically, a first flow leg  168  of the pressurized compressed air within plenum  142  is channeled into transition piece cooling passage or annulus  164  via radially-oriented impingement sleeve openings  166 . The air is then channeled upstream within transition piece cooling passage or annulus  164  and discharged into combustion liner cooling passage or annulus  154 . In addition, a second flow leg  170  of the pressurized compressed air within plenum  142  is channeled around impingement sleeve  158  and injected substantially axially 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 . For turbine models or configurations that do not employ an impingement sleeve about the transition piece, the compressor discharge air flows axially into the annulus  154  between the flow sleeve  148  and the combustor liner  150  via inlets  156 . It will be appreciated that the invention described herein is applicable to both arrangements. 
         [0026]      FIG. 3  is an enlarged detail showing the introduction of compressor discharge cooling air into the liner cooling passage or annulus  154  via the circumferential opening or inlet  156  (opening  156  may be sectioned by struts (not shown) that support the aft end of the flow sleeve  148 ). The air flow streams are shown in an “idealized” pattern, where the boundary flow remains attached to the inside surface of the flow sleeve  148 . 
         [0027]      FIG. 4  illustrates actual air flow introduced through the inlets  156  where separation and recirculation of the air (i.e., turbulence) along the inside surface of the flow sleeve  148  causes an undesirable increase in pressure drop. 
         [0028]      FIG. 5  illustrates an exemplary but nonlimiting embodiment of this invention. The enlarged aft end  176  of the flow sleeve  178  has been reconfigured to incorporate an annular manifold  180  that distributes additional or supplemental air under pressure into the annular passage  154  at the inlet  156 . Air distribution holes  182  are located about the periphery of the flow sleeve aft end  176 , in communication with the manifold  180 , and preferably angled variably to introduce the air generally radially but with a flow component in the direction of flow of the cooling air in passage  154 . By supplying additional air under higher pressure than the cooling air flowing axially through the passage  154  in this manner, the separation and recirculation shown in  FIG. 4  is substantially eliminated. 
         [0029]    The supplemental air flow fed through the manifold  180  and blown into the annulus  154  can be adjusted from “full-on” to “full-off” positions and anywhere in between, as will be described in greater detail below.  FIG. 6  shows the flow in the full-off condition, such that the flow separation and recirculation reoccurs, similar to the flow pattern illustrated in  FIG. 4 . 
         [0030]    Referring now to  FIG. 7 , an exemplary air distribution system is illustrated for supplying supplemental air to the liner aft manifolds  180  of each of eight (8) combustors (numbered 1-8) in a can-annular array of combustors about a turbine  184 . Thus, in an exemplary eight-combustor arrangement, an air distribution box  186  can selectively supply supplemental air via individual conduits  188  to any or all of the manifolds  180 . It will be appreciated that the invention also contemplates employing multiple conduits to supply air to a single manifold  180 , thus permitting air to be selectively routed to specific locations within a single combustor. The air distribution box, in turn, is controlled by a control system  190  which also receives input from certain subsystems of the gas turbine  184 , for example, a dynamics monitoring system, emissions monitoring system, turbine exhaust temperature monitoring system and/or the turbine main controller. 
         [0031]    Air may be supplied to the air distribution box  186  from a high pressure, external source  192 . 
         [0032]    This arrangement permits independent and selective control of the additional or supplemental air supplied to the manifolds  164  of the respective combustors 1-8 in order to optimize the performance of each. 
         [0033]    In other words, this so-called “boundary layer blowing” is used to control the aerodynamic performance of the inlet to the combustor liner/flow sleeve passage  154  of each combustor by modulating the amount of supplemental air introduced via the manifolds  180  and air distribution holes  182 , and thus not only modulating the pressure loss at each combustor, but also permitting adjustments to the fuel/air ratio in each combustor. 
         [0034]    With respect to the manifolds  180 , they may be cast in place integrally with the flow sleeve  178 , or formed as split flow sleeve ends welded in place (as indicated by the weld line  194  shown in phantom in  FIG. 5 ). The distribution holes  182  may be uniformly distributed about the flow sleeve periphery, or they may be arranged in discrete groups located symmetrically about the periphery, or at locations dictated by the arrangement of fuel nozzles within the combustor. The air supplied to the discrete groups of outlet holes  182  could also be individually controlled within a single combustor by subdividing the manifold into discrete segments. 
         [0035]    The supplemental air introduced via the outlet holes  182  could also be taken from a substantially constant volume of air within the manifold  180  by means of piezoelectric devices that can be switched to “pump” air into the stream or passage  154  and then switched to “pull” air back into the manifold. 
         [0036]    It can thus be appreciated that the invention as described may be altered in any of several ways to control air flow at the inlets  156  to the passage  154  and thus selectively control combustor performance and emissions for each of the several combustors in a can-annular array. 
         [0037]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.