Patent Publication Number: US-9423132-B2

Title: Ultra low emissions gas turbine combustor

Description:
FIELD OF THE INVENTION 
     The present invention relates to can combustors. In particular, the present invention relates to gaseous fuel-fired, impingement cooled, dry low emission can combustors for gas turbine engines. 
     BACKGROUND OF THE INVENTION 
     Gas turbine combustion systems utilizing can type combustors are often prone to air flow mal-distribution. The problems caused by such anomalies are of particular concern in the development of low NOx systems. The achievement of low levels of oxides of nitrogen in combustors is closely related to flame temperature and its variation through the early parts of the reaction zone. Flame temperature is a function of the effective fuel-air ratio in the reaction zone which depends on the applied fuel-air ratio and the degree of mixing achieved before the flame front. These factors are obviously influenced by the local application of fuel and associated air and the effectiveness of mixing. Uniform application of fuel typically is under control in well designed injection systems but the local variation of air flow is often not, unless special consideration is given to correct mal-distribution. 
     The achievement of current levels of oxides of nitrogen set by regulations in some areas of the world calls for effective fuel-air ratio to be controlled to low standard deviations on the order of 10%. The cost of development of such combustion systems is high but can be significantly influenced by the right choice of configuration. However, the use of film cooling in these low flame temperature combustors generates high levels of carbon monoxide emissions. External impingement cooling of the flame tube (liner) can curtail such high levels. Moreover, in systems where high exit temperature is a performance requirement in addition to low NOx, the air flow to swirler/reaction zone is a large proportion of total air flow and therefore cooling and dilution air flows are limited. Hence there is considerable advantage in controlling these flows to optimize the overall flow conditions. 
     One such recent combustor design is that shown in U.S. Pat. No. 7,617,684 to Norster, assigned to the assignee of the present invention, the disclosure of which is hereby incorporated by reference. In the subject Norster combustor, essentially all the air flow for combustion is first separated from the dilution air stream and used for impingement cooling the portion of a combustor liner defining the combustion zone, and then channeled to swirl vanes for mixing with fuel. While the features of the Norster combustor may provide better control of the amount of air delivered to the swirl vanes, and thus the bulk fuel/air ratio, compared to previous impingement cooled combustors, further improvements in the aerodynamics of the combustion air flow to the swirl vanes may minimize local deviations in the fuel/air ratio. Improvements are also possible in the control of other cooling air flows in the combustor, which affect the level of emissions and the thermal efficiency of the combustor. Such improvements are set forth hereinafter. 
     SUMMARY OF THE INVENTION 
     In one aspect of the present invention, a gaseous fuel-fired can combustor for use with a gas turbine, for example in a gas turbine engine, includes a generally cylindrical housing having an interior, an axis, and a closed axial end. A generally cylindrical combustor liner is disposed coaxially within the housing interior and is configured to define with the housing a radial outer flow passage for combustion air. The liner also defines respective radially inner volumes for a combustion zone and a dilution zone, the dilution zone being axially distant the closed housing end relative to the combustion zone, and the combustion zone being axially adjacent the closed housing end. Mixing apparatus is disposed at the closed housing end and in flow communication with the combustion air passage. The mixing apparatus includes a plurality of vanes for mixing the gaseous fuel to be combusted with at least a part of the combustion air, and a mixing apparatus outlet for admitting the resulting fuel/air mixture to the combustion zone. An impingement cooling sleeve is coaxially disposed in the combustion air passage between the housing and the liner, the sleeve having a plurality of apertures sized and distributed to direct the combustion air against a radially outer surface of a portion of the liner defining the combustion zone, for impingement cooling the liner portion. Channeling apparatus is disposed in the combustion air passage for channeling the combustion air from an impingement cooling sleeve exit region to the inlet of the mixing apparatus. The channeling apparatus is configured to prevent flow separation and includes a diffuser section with an inlet flow area and an outlet flow area, wherein a ratio of the outlet flow area to the inlet flow area is in the range 1.3-1.5. 
     In another aspect of the present invention, the gaseous fuel can combustor for a gas turbine includes a generally cylindrical outer housing having an interior, an axis, and a closed end. A generally cylindrical combustor liner is disposed coaxially within the housing interior and is configured to define with the housing a radially outer flow passage for combustion air, with the liner having an interior defining a radially inner volume for a combustion zone proximate the housing closed end. Mixing apparatus including a plurality of swirl vanes is disposed at the housing closed end. The mixing apparatus has an inlet in flow communication with the combustion air flow passage and an axially directed outlet in flow communication with the combustion zone. The swirl vanes are arranged circumferentially spaced apart about the housing axis in a plane generally perpendicular to the axis. A gaseous fuel supply system is operatively connected to deliver gaseous fuel to the mixing apparatus in the vicinity of the swirl vanes for mixing with combustion air received from the combustion air flow passage. Adjacent ones of the circumferentially spaced apart vanes partly define generally radially inwardly directed mixing flow passages, wherein each the mixing flow passages has a substantially constant cross-sectional flow area and an increasing aspect ratio along a flow direction between the swirl vanes. 
     The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate several embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic cross-sectional view of a gas turbine can combustor in accordance with the present invention; 
         FIG. 2  is a detail of the mixing apparatus of the  FIG. 1  combustor, including swirl vanes; 
         FIGS. 3 and 4  are, respectively, axial and side schematic views showing the design characteristics of the swirl vanes of the  FIG. 1  combustor; and 
         FIG. 5  is a detail of the combustor in  FIG. 1  showing holes for admitting air to minimize flow separation in the diffuser section. 
     
    
    
     DESCRIPTION OF THE EMBODIMENTS 
     The can combustor of the present invention, generally designated by the numeral  10  in the figures, is intended for use in combusting gaseous fuel with compressed air from compressor  6 , and delivering combustion gases to gas turbine  8 , e.g., for work-producing expansion such as in a gas turbine engine. See  FIG. 1 . Compressor  6  may be a centrifugal compressor and gas turbine  8  may be a radial inflow turbine, but these are merely preferred and are not intended to limit the scope of the present invention, which is defined by the appended claims and their equivalents. 
     In accordance with the present invention, as embodied and broadly described herein, the can combustor may include a generally cylindrical housing having an interior, an axis, and a closed axial end. As embodied herein, and with reference to  FIG. 1 , can combustor  10  includes outer housing  12  having interior  14 , longitudinal axis  16 , and closed axial end  18 . Housing  12  is generally cylindrical in shape about axis  16 , but can include tapered and/or step sections of a different diameter in accordance with the needs of the particular application and to accommodate certain features of the present invention to be discussed hereinafter. 
     In accordance with the present invention, the combustor also includes a generally cylindrical combustor liner disposed coaxially within the housing and configured to define with the housing respective radial outer passage for combustion air. The liner also defines respective radially inner volumes for a combustion zone and a dilution zone. The dilution zone is axially distant the closed housing end relative to the combustion zone, and the combustion zone is axially adjacent the closed housing end. 
     As embodied herein, and with continued reference to  FIG. 1 , combustor  10  includes combustor liner  20  disposed within housing  12  generally concentrically with respect to axis  16 . Liner  20  may be sized and configured to define with housing  12  outer passage  26  for compressed air supplied from engine compressor  6  to be used for impingement cooling and combustion air. Liner  20  also partially defines dilution air path  28 . In the  FIG. 1  embodiment, path  28  for the dilution air includes a plurality of dilution ports  30  distributed about the circumference of liner  20 . 
     The interior of liner  20  also defines combustion zone  32  axially adjacent closed end  18 , where the swirling combustion air and fuel mixture is combusted to produce hot combustion gases. In conjunction with mixing apparatus  40  at closed end  18  (to be discussed hereinafter) liner portion  20   a  is configured to provide stable recirculation in region  34  of combustion zone  32 , in a manner known to those skilled in the art. The interior of liner  20  further defines dilution zone  36  where combustion gases are mixed with dilution air from dilution ports  30  to lower the temperature of the combustion gases, before work-producing expansion in turbine  8 . 
     Also, in accordance with the present invention, the combustor includes apparatus having a plurality of vanes for mixing at least a part of the combustion air with gaseous fuel, the mixing apparatus having an outlet for admitting the resulting fuel/air mixture to the combustion zone. As embodied herein, and with continued attention to  FIG. 1 , mixing apparatus  40  includes swirl plate  42  with a plurality of swirl vanes  44  disposed about the circumference of swirl plate  42 , and mixing apparatus inlet  46  and outlet  48 . Each vane  44  has a leading edge  68 , trailing edge  70 , top  72 , and bottom  74 . See  FIG. 4 . Mixing apparatus  40  further includes a plurality of nozzles  50 , each preferably having multiple orifices  52  for injecting the gaseous fuel. Nozzles  50  are controllably fed from fuel supply  54  via appropriate valved connections and channels, as one skilled in the art would understand. 
     With reference now to  FIGS. 2-4 , swirl vanes  44  preferably are aerodynamically shaped with a taper angle of α 2  and are spaced apart circumferentially to provide combustion air passages  60  with good fuel/air mixing without separation. Specifically, the passages  60  are configured to have a constant cross section flow area  62  between adjacent vanes but with a varying aspect ratio of passage height H to passage width W along the vane length from passage inlet  64  to passage outlet  66 , respectively proximate vane leading edge  68  and vane trailing edge  70  (see  FIG. 3 ). Preferably, the aspect ratio ranges from about 1.5 at passage inlet  64  to about 4.5 at passage outlet  66 . 
     Further, and as best seen in  FIG. 2 , each vane  44  has a pair of nozzles  50  recessed into opposing sides  44   a ,  44   b  of the vane, each nozzle being proximate vane leading edge  68  and having a plurality of orifices  52  directed into a respective passage  60 . Nozzles  50  can be configured to be replaceable e.g., with nozzles having different orifice sizes to accommodate different gaseous fuels, or for repair. Also, and as best seen in  FIG. 4 , leading vane edge  68  is preferably set at an angle β relative to the axial direction  16   a , to better receive the incoming combustion air. The angle β may be set to be at right angles to the direction of the incoming air as depicted in  FIG. 4 . 
     Table 1 presents a particularly preferred set of design parameter ranges for the profile and orientation of vanes  44 , in relation to the depiction in  FIGS. 3 and 4 . 
     
       
         
           
               
               
               
               
             
               
                   
                 TABLE 1 
               
               
                   
                   
               
               
                   
                 Parameter 
                 Min. value 
                 Max. value 
               
               
                   
                   
               
             
            
               
                   
                 L 1 /L 2   
                 1.2 
                 1.4 
               
               
                   
                 R 1 /L 2   
                 2.5 
                 2.6 
               
               
                   
                 H 2 /L 2   
                  0.35 
                  0.45 
               
               
                   
                 H 1 /L 1   
                  0.65 
                  0.75 
               
               
                   
                 α 2   
                 20°      
                 25°      
               
               
                   
                 H 2 /W 2   
                 1.4 
                 1.6 
               
               
                   
                 H 1 /W 1   
                 4.4 
                 4.6 
               
               
                   
                   
               
            
           
         
       
     
     Still further in accordance with the present invention, as embodied and broadly described herein, the can combustor may further include an impingement cooling sleeve coaxially disposed between the housing and the combustion liner and extending axially from the closed housing end for a substantial length of the combustion zone. The impingement cooling sleeve may have a plurality of apertures sized and distributed to direct combustion air against the radially outer surface of the portion of the combustor liner defining the combustion zone, for impingement cooling. 
     As embodied herein, and with reference to  FIG. 1 , impingement cooling sleeve  80  is depicted coaxially disposed between housing  12  and liner  20 . Impingement cooling sleeve  80  extends axially along a portion of liner  20  defining combustion zone  32  from a location adjacent closed end  18  to a location proximate but upstream of dilution ports  30  relative to the axial flow of the combustion gases. Sleeve  80  includes a plurality of impingement cooling orifices  82  distributed circumferentially around sleeve  80  and configured and oriented to direct combustion air in passage  26  against the outer surface of liner  20  in the vicinity of combustion zone  32 . It is preferred that the shape of the impingement cooling sleeve  80  be axially tapered, to achieve a frusto-conical shape with an increasing diameter from sleeve end  84  to sleeve end  86  which comprises the exit region for the combustion air flow after it has traversed sleeve  80  and has impingement cooled liner surface  88 . The sleeve end  84  preferably is configured to seal the combustion/impingement cooling air in passage  26  from dilution air path  28  after the combustion air his traversed impingement cooling orifices  82 . 
     Significantly, in the embodiment depicted in  FIG. 1 , essentially all of the combustion air eventually admitted to combustion zone  32  first passes through orifices  82  of impingement sleeve  80  to provide cooling, that is, all except possibly unavoidable leakage. Combustion air may comprise between about 45-55% of the total air supplied to the can combustor (combustion air plus dilution air) for low NOx configurations. 
     Still further in accordance with the invention, as embodied and broadly described herein, the can combustor includes apparatus for channeling the combustion air from an exit region downstream of the impingement cooling sleeve to an inlet of the mixing apparatus. The channeling apparatus is configured to prevent flow separation and includes a diffuser section with an inlet flow area and an outlet flow area, with the ratio of the outlet flow area to the inlet flow area being in the range 1.3-1.5 or greater. 
     As embodied herein, and with reference to  FIG. 1 , channeling apparatus  90  includes diffuser section  92  and a guide section  94 , both comprising sequential parts of the combustion air flow passage  26 . Diffuser section  92  extends between a location “A” downstream of sleeve exit region  86  to a location “B” which is the beginning of inwardly curved guide section  94 . Guide section  94 , in turn, extends from location “B” to inlet  46  of mixing apparatus  40  proximate leading edges  68  of swirl vanes  44 . Guide section  94  serves to turn the combustion air inwardly toward axis  16  and mixing apparatus inlet  46  with a minimum of flow separation using smoothly curved inner surface  96  of housing  1  and surface  42   a  of swirl plate  42 , with a large radius of curvature. As depicted in  FIG. 1 , guide section surface  96  should preferably be configured to have the same O.D. and curvature at the location of leading edge  68  as swirl plate surface  42   a , to avoid an abrupt step and possible flow separation. 
     It may specifically be preferred to use a radius of curvature r that satisfies the following relations: 
               1.15   ≤     r     H   1       ≤   1.35     ,       and   ⁢           ⁢   0.35     ≤     r     R   1       ≤   0.45     ,         
where H 1  is the height of vane  44  at trailing edge  70 , and R 1 , is the radial distance from axis  16  to inner surface  96  of housing  18  at the beginning of guide section  94  (location B). See  FIGS. 1 and 4 . Also, it may specifically be preferred that vanes  44 , as well as swirl plate  42 , be configured such that the air and fuel mixture leaves the swirl vanes  44  in the tangential direction relative to axis  16  (within ±3°). This provides the longest flow path for the air and fuel mixture, which gives a more homogenous mixture. This feature has been made possible due to the varying aspect ratio in the swirl vane passages.
 
     Returning to diffuser section  92 , diffuser flow area  98  in the depicted embodiment is the space between the conical inside surface  100  of housing  14  between locations “A” and “B”, and the conical outside surface  104  of wall  114  of toroidal spacer member  102 . These two conical surfaces are sized and configured to provide a continuously increasing annular diffuser flow area from the diffuser section inlet (location “A”) to diffuser section outlet (location “B”) to provide an expansion ratio of the outlet flow area to the inlet flow area in the range of 1.3-1.5, via a smooth, continuous expansion. The consequent lowering of the average velocity may provide a more optimum velocity ratio between the combustion air entering mixing apparatus  40  and the fuel injected from nozzles  50 , thus providing more uniform mixing. 
     One skilled in the art would understand from the above that the configuration of the surfaces defining diffuser section  92  need not both be conical to provide the desired expansion ratio. That is, wall  114  with outer surface  104  of toroidal spacer member  102  could be cylindrical while inner surface  100  of diffuser section  42  of housing  14  could be conical, or vice versa. While each of these alternatives may result in a more radially compact combustor, each would increase the severity of hydraulic losses in guide section  94  due to the sharper turn (smaller radius of curvature) proximate mixing apparatus inlet  46 , and hence may not be preferred. In the  FIG. 1  embodiment, the bulk combustion air flow through diffuser section  92  is slightly away from axis  16 , while the flow through guide section  94  is toward axis  16 , allowing most of the turning to be accomplished smoothly over an extended guide section length and not abruptly at the mixing apparatus inlet. Dish-shaped curved mixing plate surface  42   a , which provides the upper boundary of swirl vane passages  60 , also helps turning the combustion air. 
     It may also be preferred that a small fraction (˜14%) of the combustion air from the diffuser section  92  be used to cool the “head” end of liner  20 , namely, liner part  20   a  surrounding portion  34  of the combustion zone, where the recirculated combustion gases can create high heat loading. In the  FIG. 1  embodiment, toroidal member  102  can be configured with inner wall  106  spaced from liner portion  20   a  and provided with directed impingement cooling apertures  108 . In the  FIG. 1  embodiment, the combustion air for impingement cooling liner portion  20   a  enters toroidal member  102  through apertures  112  in outer wall  114 . 
     Still further and as best seen in  FIG. 1 , top wall  116  of toroidal member  102  abuts swirl vanes  44  and defines the bottom portions of swirl vane passages  60 . 
     It may be further preferred to use another small fraction (˜1%) of the combustion air to prevent flow separation at the diffuser inlet A. As best seen in  FIG. 5 , impingement sleeve  80  is captured to housing  14  via a flanged connection that causes step  120 . To prevent flow separation due to the sudden expansion in the flow area at step  120 , bleed holes  122  are provided in step  120  and are supplied with combustion air from passage  26  upstream of impingement sleeve  80 . 
     As a consequence of the features of the can combustor described above, and in addition to the advantage of the more uniform air flow to the swirl vanes discussed previously, the can combustor may provide more uniform pre-mixing in the swirl vanes and, consequently, a higher effective fuel-air ratio for a given NOx and CO requirement. Also, the above-described can combustor may provide a higher margin of stable burning, in terms of providing a more stable recirculation pattern and may also minimize temperature deviations (“spread”) in the combustion products delivered to the turbine. Finally, the can combustor disclosed above may also maximize the effectiveness of the cooling air and provide optimum liner wall metal temperatures. 
     It will be apparent to those skilled in the art that various modifications and variations can be made in the disclosed impingement cooled can combustor, without departing from the teachings contained herein. Although embodiments will be apparent to those skilled in the art from consideration of this specification and practice of the disclosed apparatus, it is intended that the specification and examples be considered as exemplary only, with the true scope being indicated by the following claims and their equivalents.