Abstract:
A nozzle for assemblies and gas turbines is provided. The nozzle exhibits destabilized flame holding characteristics, i.e., the nozzle is unable to stabilize flame up to an equivalence ratio of about 0.65. As a result, flame heat release is delayed resulting in lower peak flame temperatures and correspondingly lower NOx levels. Flame stabilization capability is retained for higher equivalence ratios to support operation of the combustor in other regions of the load range.

Description:
BACKGROUND 
     The embodiments disclosed relate generally to gas and liquid fuel turbines, including both can-annular or annular combustion systems, and methods of operating such combustion systems. 
     Dry Low NOx technology is routinely applied for emissions control with gaseous fuel combustion in industrial gas turbines with can-annular combustion systems through utilization of premixing of fuel and air. The primary benefit of premixing is to provide a uniform rate of combustion resulting in relatively constant reaction zone temperatures. Through careful air management, these temperatures can be optimized to produce very low emissions of oxides of nitrogen (NOx), carbon monoxide (CO) and unburned hydrocarbons (UHC). Modulation of a center premix fuel nozzle can expand the range of operation by allowing the fuel-air ratio and corresponding reaction rates of the outer nozzles to remain relatively constant while varying the fuel input into the turbine. 
     Fuel staging is well-understood by those experienced in the art as a means of achieving higher turbine inlet temperatures with uniform heat release. Axially staged systems employ multiple planes of fuel injection along the combustor flow path. Utilization of axial fuel staging requires special design considerations to inject fuel into the high temperature products of combustion. The high temperature and pressure environment of the latter stages of an axially staged combustor have prevented development of robust designs suitable for commercial applications. 
     It would therefore be desirable to develop new gas turbines having a fuel system configuration and/or utilizes a method of staging fuel so that lower peak fuel temperatures are achieved. Such gas turbines would be expected to have correspondingly low NOx and CO emissions. The ability of a new gas turbine to exhibit of an increased range of operability within such “Emissions Compliant” regimes would provide further advantages. 
     BRIEF DESCRIPTION 
     A gas turbine fuel nozzle is provided. The fuel nozzle has a physical configuration so that the nozzle is unable to stabilize flame up to an equivalence ratio of about 0.65. 
     In another aspect, an assembly for a single stage gas turbine combustor is provided. The assembly comprises an array of outer nozzles arranged about a center axis, and a center nozzle located on said center axis, wherein said center nozzle has a physical configuration such that the center nozzle is unable to stabilize flame up to an equivalence ratio of about 0.65. 
     In another aspect, there is provided a gas turbine comprising a plurality of combustors. Each combustor has a plurality of outer fuel nozzles arranged about a longitudinal axis of the combustor, a center nozzle disposed substantially along said longitudinal axis, and a single combustion zone. The center nozzle is unable to stabilize flame up to an equivalence ratio of about 0.65. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       These and other features, aspects, and advantages of the present invention will become even better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
         FIG. 1  is a representation of combustor operability or flame stability for a gas turbine combustion system; 
         FIG. 2  is a graphical depiction of the fuel air stoichiometric ratio (x-axis) versus the NOx levels at 15% O 2  (y-axis) showing the benefit of late lean combustion; 
         FIG. 3  shows the regions of flame stability for a premixed combustion system, region “ 1 ” is the range where conventional fuel nozzles are unable to stabilize a flame (conventional lean blow out), region “ 2 ” is the range where this improved fuel nozzle is unable to stabilize a flame (extended lean blow out), and region “ 3 ” is a region where all fuel nozzles can stabilize flame; 
         FIG. 4  is a schematic cross-sectional view of a can-annular combustor of a turbine in accordance with one embodiment; 
         FIG. 5  is a schematic front end view of an end cover and fuel nozzle assembly in accordance with one embodiment; 
         FIG. 6  is a schematic cross-sectional view of an outer fuel nozzle in accordance with some embodiments; 
         FIG. 7  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 8  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 9  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 10  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 11  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 12  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 13  is a schematic cross-sectional view of a center fuel nozzle in accordance with one embodiment; 
         FIG. 14A  is a representation of the flame shapes for a conventional can-annular combustor; 
         FIG. 14B  is a representation of the flame shapes for a can-annular combustor according to one embodiment; and 
         FIG. 14C  is a representation of the flame shapes for a can-annular combustor according to one embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     The above brief description sets forth features of the various embodiments of the present invention in order that the detailed description that follows may be better understood, and in order that the present contributions to the art may be better appreciated. There are, of course, other features of the invention that will be described hereinafter and which will be for the subject matter of the appended claims. 
     In this respect, before explaining several embodiments of the invention in detail, it is understood that the various embodiments of the invention are not limited in their application to the details of the construction and to the arrangements of the components set forth in the following description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced and carried out in various ways. Also, it is to be understood that the phraseology and terminology employed herein are for the purpose of description and should not be regarded as limiting. 
     The terms “first,” “second,” and the like, as used herein do not denote any order, quantity, or importance, but rather are used to distinguish one element from another. The terms “a” and “an” herein do not denote a limitation of quantity, but rather denote the presence of at least one of the referenced items. The modifier “about” used in connection with a quantity is inclusive of the stated value, and has the meaning dictated by context, (e.g., includes the degree of error associated with measurement of the particular quantity). The suffix “(s)” as used herein is intended to include both the singular and the plural of the term that it modifies, thereby including one or more of that term. 
     Reference throughout the specification to “one embodiment” or “an embodiment” means that a particular feature, structure, or characteristic described in connection with an embodiment is included in at least one embodiment. Thus, the appearance of the phrases “in one embodiment” or “in an embodiment” in various places throughout the specification is not necessarily referring to the same embodiment. Further, the particular features, structures or characteristics may be combined in any suitable manner in one or more embodiments. 
     Provided herein is a fuel nozzle, and assembly and gas turbine comprising the nozzle, that utilizes fuel staging to achieve very low emissions on gaseous fuel. Nozzles, assemblies and combustors incorporate physical configurations so that flame stabilization is avoided without utilizing down-stream fuel injection. The desired low emissions are thus provided. 
       FIG. 1  is a graphical depiction of flame stability for a conventional gas turbine combustion system. As shown, flame stability is a function of fuel/air ratio and air flow. There is a region of stable burning, the size of which potentially being impacted by several variables including fuel type. The nozzles, assemblies and combustors provided herein are physically configured so that the region of stable burning is decreased, and the region of flame stability is increased. 
     Avoidance of flame stabilization, in turn, allows unburned fuel to propagate downstream, beyond the primary reaction zones ( FIG. 4, 43 ) of adjacent fuel nozzles. That is, a flame supported by the present nozzle will not burn right away, but will burn within the combustor zone of the assembly and/or combustor. The result is similar to that provided by axial fuel staging, without the conventional requirement for downstream fuel injection. 
     The benefits of axial fuel staging, or late lean injection, on NOx emissions from a premixed flame are graphically depicted in  FIG. 2 . The conventional NOx versus fuel/air relationship is shown with the solid line, while the NOx versus fuel/air relationship that occurs in nozzles, assemblies and combustors employing axial fuel staging is indicated by the dashed line (also referred to from time to time by those of ordinary skill in the art as late lean fuel injection). As shown, an enhanced area of operating fuel/air ratios is provided, that is yet capable of operating within the desired NOx emissions. Late introduction of a portion of the fuel enables extension of the overall flame zone, which in turn results in a lowering of peak temperatures and a reduction in NOx emissions. However, it has not yet been possible to demonstrate a practical means of placing the late or down-stream fuel nozzles in the path of the high temperature combustion gas. In the present nozzles, assemblies and combustors, this enhanced operating area is provided by the physical configuration of the nozzle so that late lean injection can be avoided, while yet the same effect can be seen. Even though the present nozzle is capable of providing the benefits of late lean injection, without requiring such a configuration, the nozzle is yet also capable of use at high fuel/air ratios, i.e., of greater than 0.65 for operation in low-power modes. 
       FIG. 3  is a graphical depiction of NOx emissions versus fuel air ratio. The right-hand region of the graph shows the normal range of lean blow out for a premixed fuel nozzle. The center region shows a range of extended lean blow out that can be achieved using embodiments of the present center fuel nozzle. The left region shows the area where flame stabilization could not occur for the center fuel nozzle due to insufficient fuel flow or excessively low fuel-air ratio. 
     And so, provided herein is a nozzle that may desirably be part of a combustor assembly, arranged in annular or can-annular configuration on an industrial gas turbine. The present nozzles, assemblies and combustors are advantageously employed at low to moderate fuel/air ratios, e.g., at fuel/air ratios of less than 0.65, as may typically be utilized in high-power modes. 
       FIG. 4  is a schematic cross-sectional view through one of the combustors of a turbine comprising a can-annular combustor configuration. Gas turbine  10  includes a compressor  12  (partially shown), a plurality of combustors  14  (one shown), and a turbine represented here by a single blade  16 . Although not specifically shown, the turbine is drivingly connected to the compressor  12  along a common axis. The compressor  12  pressurizes inlet air which is then flows in reverse to the combustor  14  where it is used to cool combustor  14  and to provide air to the combustion process. 
     As noted above, the gas turbine includes a plurality of combustors  14  located about the periphery of the gas turbine. A double-walled transition duct  18  connects the outlet end of each combustor with the inlet end of the turbine to deliver the hot products of combustion to the turbine. Ignition is achieved in the various combustors  14  by means of spark plug  20  in conjunction with cross fire tubes  22  in the usual manner. 
     Each combustor  14  includes a substantially cylindrical combustor casing  24  which is secured at an open forward end to the turbine casing  26  by means of bolts  28 . The rearward or proximal end of the combustion casing is closed by an end cover assembly  30  which includes supply tubes, manifolds and associated valves for feeding gaseous fuel, liquid fuel, air and water to the combustor  14  as described in greater detail below. The end cover assembly  30  receives a plurality (for example, three to six) “outer” fuel nozzle assemblies  32  (one shown) arranged in a circular array about a longitudinal axis of combustor  14 , and one center nozzle  33 . 
     Within the combustor casing  24 , there is mounted, in substantially concentric relation thereto, a substantially cylindrical flow sleeve  34  which connects at its forward end to the outer wall  36  of the double walled transition duct  18 . The flow sleeve  34  is connected at its rearward end by means of a radial flange  35  to the combustor casing  24  at a butt joint  37  wherein fore and aft sections of the combustor casing  24  are joined. 
     Within the flow sleeve  34 , there is a concentrically arranged combustion liner  38  which is connected at its forward end with the inner wall  40  of the transition duct  18 . The rearward end of the combustion liner  38  is supported by a combustion liner cap assembly  42 , which is, in turn, supported within the combustor casing by a plurality of struts  39  and an associated mounting assembly. Outer wall  36  of the transition duct  18  and that portion of flow sleeve  34  extending forward of the location where the combustor casing  24  is bolted to the turbine case are formed with an array of apertures  44  over their respective peripheral surfaces to permit air to reverse flow from the compressor  12  through the apertures  44  into the annular space between the flow sleeve  34  and the liner  38  toward the upstream or rearward end of the combustor (as indicated by the flow arrows shown in  FIG. 1 ). 
     The combustion liner cap assembly  42  supports a plurality of premix tubes  46 , one for each of “outer” fuel nozzle assemblies  32  and center nozzle  33 . More specifically, each premix tube  46  is supported within the combustion liner cap assembly  42  at its forward and rearward ends by front and rear plates  47  and  49  respectively, each provided with openings aligned with the open-ended premix tubes  46 . The front plate  47  (an impingement plate provided with an array of cooling apertures) may be shielded from the thermal radiation of the combustor flame by shield plates (not shown). 
     The rear plate  49  mounts on a plurality of rearwardly extending floating collars  48  (one for each premix tube  46 , arranged in substantial alignment with the openings in the rear plate), each of which supports an air swirler  50  in surrounding relation to a radially outermost wall of the respective nozzle assembly. The arrangement is such that air flowing in the annular space between the liner  38  and flow sleeve  34  is forced to again reverse direction in the rearward end of the combustor (between the end cover assembly  30  and sleeve aperture  44 ) and to flow through the swirlers  50  and premix tubes  46 . The construction details of the combustion liner cap assembly  42 , the manner in which the liner cap assembly is supported within the combustion casing, and the manner in which the premix tubes  46  are supported in the liner cap assembly is the subject of U.S. Pat. No. 5,259,184, hereby incorporated herein by reference in its entirety. 
       FIG. 5  schematically shows a front end view and fuel nozzle assembly of one embodiment of an endcover arrangement of the can-annular combustor shown in  FIG. 4 . As noted above, outer fuel nozzle assemblies  32  and one center nozzle  33  are attached to endcover  30 . The endcover  30  comprises internal passages which supply the gaseous and liquid fuel, water, and atomizing air to the nozzles as detailed below. Piping and tubing for supply of the various fluids are, in turn, connected to the outer surface of the endcover assembly. 
     Outer fuel nozzle assemblies  32  and center nozzle  33  may conventionally be configured to supply premix gaseous fuel, liquid fuel, water injection, atomizing air and/or diffusion fuel. In some embodiments, outer fuel nozzle assemblies  32  and center nozzle  33  are configured to provide premixed gaseous fuel. 
     Referring to  FIG. 6 , each outer fuel nozzle assembly  32  includes a proximal end or rearward supply section  72 , with inlets for receiving liquid fuel, water injection, atomizing air, and premixed gas fuel, and with suitable connecting passages for supplying each of the above-mentioned fluids. As mentioned above, outer fuel nozzle assemblies  32  are each configured to receive premixed gaseous fuel, and to supply it to a respective passage in a forward or distal delivery section  74  of the fuel nozzle assembly. Outer fuel nozzle assemblies may be configured so as to be substantially parallel to the longitudinal axis (axis of symmetry) of center fuel nozzle assembly  33 , or may be tilted outward relative to this axis so that their flames are angled toward the wall of the liner. Such a configuration enables the center nozzle fuel to progress further downstream before igniting. Although the particular angle is not critical so long as the foregoing objective is achieved, the tilt angle may be limited by the wall of the liner. Useful tilt angles, relative to the longitudinal axis of center fuel nozzle  33  are expected to range from about 1° to about 7 degrees. 
     In the embodiment shown, the forward delivery section of the outer fuel nozzle assembly  32  is comprised of a series of concentric tubes. Tubes  76  and  78  define premix gas passage(s)  80  which receive(s) premix gas fuel from premix gas fuel inlet(s)  82  in rearward supply section  72  via conduit  84 . The premix gas passages  80  communicate with a plurality of radial fuel injectors  86 , each of which is provided with a plurality of fuel injection ports or holes  88  for discharging gas fuel into the premix zone located within the premix tube  46 . The injected premix fuel mixes with air reverse flowed from compressor  12 . 
     A second passage  90  is defined between concentric tubes  78  and  92  and is used to supply atomizing air from atomizing air inlet  94  to the burning zone  70  of the combustor  14  via orifice  96 . A third passage  98  is defined between concentric tubes  92  and  100  and is used to supply water from water inlet  102  to the burning zone  70  to effect NOx reductions in the manner understood by those skilled in the art. 
     Tube  100 , the innermost of the series of concentric tubes forming the outer nozzles  32 , itself forms a central passage via liquid fuel inlet  106 . The liquid fuel exits the nozzle by means of a discharge orifice  108  in the center of outer nozzle assembly  32 . Thus, all outer nozzles  32  and center gas nozzle  33  provide premix gaseous fuel. The center nozzle  33 , but not the outer nozzles  32  provides a passive air purge, and each of the outer nozzles  32 , but not the center nozzle  33 , is configured for delivering liquid fuel, water for emissions abatement, and atomizing air. A number of quaternary pegs (not shown) are located circumferentially around the forward combustion casing distributing fuel through 8 holes per peg. 
     Center fuel nozzle  33  is provided with a physical configuration that minimizes turbulence and flow recirculation such that flame stability is poor. Center nozzle  33  is thus capable of providing such flame destabilization at equivalence ratios lower than about 0.65. Non-limiting examples of physical configurations that provide such ability to center nozzle  33  include one or more aerodynamic features, such as, e.g., a streamlined nozzle tip, nozzle tip air purge, streamlined swirler, dual swirler, dual counter-rotating swirler, combined swirler and nozzle, inlet flow conditioner, burner tube exit bell-mouth and/or diverging burner tube wall. 
     For example, in some embodiments, center fuel nozzle  33  may be provided with streamlined tip, alone or in combination with a nozzle tip air purge that both cools the aft region of the nozzle and quenches the remaining recirculation zones. As a result, the flame has difficulty attaching in this region, i.e., center fuel nozzle  33  exhibits reduced flame stability as compared to a conventional center fuel nozzle. And so, premixed fuel dispensed from center fuel nozzle  33  will travel, or convect, downstream, prior to igniting. The result is similar to the affect of axial fuel staging but does advantageously not require downstream fuel injection. 
     In some embodiments, center fuel nozzle  33  may comprise any number of swirlers, in any configuration. For example, center nozzle  33  may be provided with a streamlined swirler, dual swirler, dual counter-rotating swirler, a swirler combined with a nozzle or fuel peg, etc. Any such swirlers may provide for rotating or counter-rotating flow of fluids dispensed there, and may act to destabilize the flame provided at the tip of center fuel nozzle  33 . Or, center fuel nozzle  33  may be disposed within a burner tube having a “bell shaped” exit. Inlet flow conditioners may also be utilized to achieve the desired flame destabilization, or, the same may be provided by a different outer nozzle configuration. Any of these may be used alone or in any combination. Several embodiments of such configurations of center fuel nozzle  33  are shown in  FIGS. 7-13 . 
     One embodiment of center fuel nozzle  33  is shown in  FIG. 7 . As shown, center fuel nozzle assembly  33  includes a proximal end or rearward supply section  52  with passage  56  that extends through center nozzle assembly  33  and for receiving a passive air purge. Inlet  54  is operatively disposed to receive air via extraction port  112  from compressor discharge region  114 , both of which are shown in  FIG. 4 . Central passage  56  passively supplies air to burning zone  70  of combustor  14  ( FIG. 4 ) via nozzle tip air purge orifices  58  defined at the forwardmost end  60  of the center fuel nozzle assembly  33 . In the context of turbine  10 , the distal or forward discharge end  60  of center fuel nozzle assembly  33  is located within premix tube  46 , and close to the distal or forward end thereof. 
     Inlets  62  are also defined in the rearward supply section  52  of the nozzle for premix gas fuel. The premix gas passage(s)  64  communicate with a plurality of radial fuel injectors  66 , each of which is provided with a plurality of fuel injection ports or holes  68  for discharging premix gas fuel into a premix zone located within premix tube  46 . 
       FIGS. 8 and 9  show two additional embodiments of center fuel nozzle  33 . More particularly, in the embodiments shown in  FIGS. 8 and 9 , center fuel nozzle  33  is provided with a streamlined nozzle tip  116 , as well as nozzle tip air purge ports  114  to cool the tip of center fuel nozzle  33 , and to prevent attachment of a flame thereto. The embodiments shown in  FIGS. 8 and 9  also employ swirlers in order to destabilize the flame, the embodiment of  FIG. 8  showing single streamlined annular swirlers  118 , and the embodiment of  FIG. 9  utilizing dual annular swirlers  118 . 
       FIG. 10  shows an additional embodiment of center fuel nozzle  33 , wherein the burner tube  120  is provided with bell-mouth exit  122 . While not wishing to be bound by any theory, it is believed that providing burner tube  120  with such an exit can reduce turbulence and flow recirculation that, in turn, can enhance flame stability.  FIG. 11  shows an embodiment of center fuel nozzle  33  wherein swirler  118  and fuel injection pegs  124  are combined to form “swozzle”  126 . This embodiment thus advantageously provides a more aerodynamic configuration with less opportunity for development of turbulence, vortex generation, or recirculation.  FIG. 11  also shows bell mouth exit  122  on burner tube  120 , although as mentioned above, this is not necessarily the case, and any single configuration that allows center fuel nozzle  33  to provide a destabilized flame at fuel/air ratios of lower than 0.65 may be utilized alone, or in combination with one or more of any other such configuration. 
       FIG. 12  shows a further embodiment of center fuel nozzle  33 , wherein inlet flow conditioner  128  is provided proximal to combined swirler  118  and fuel pegs  124 , or “swozzle”  126 . Inlet flow conditioners  128  can be considered analogous to flow straighteners and serve to provide a uniform and one-dimensional inlet flow to the swirler or swozzle. The benefit is that less turbulence, vortex generation, or recirculation occurs.  FIG. 13  shows an embodiment of center fuel nozzle  33  wherein burner tube  120  is provided with bellmouth  122 , wherein bellmouth  122  is divergent from a plane parallel to the longitudinal axis of center fuel nozzle  33 . 
     Flame shapes for conventional nozzle combustion systems as compared to the inventive combustion systems are shown in  FIGS. 14A-14C . More particularly, a conventional passive late lean combustion system is shown in  FIG. 14A , and shows the flame stabilized on all fuel injectors. In contrast, an inventive combustion system comprising embodiments of the present center nozzle is shown in  FIG. 14B , and shows a destabilized flame on the center nozzle, that only ignites farther downstream from the center nozzle.  FIG. 14C  shows another embodiment, wherein the outer fuel nozzles are tilted outward, resulting in the convection of unburned fuel even farther downstream prior to ignition. 
     The turbine operates on gaseous fuel in a number of modes. The first mode supplies premix gas fuel to two of outer nozzles  32  and to center nozzle  33 , for acceleration of the turbine. From ignition and completion of cross-firing thereof and until approximately 95% speed, the flow of premix fuel to center nozzle  33  is turned off, and that percentage of fuel is redirected to two of outer fuel nozzles  32 . From approximately 95% speed and very low load operation, the flow of premix fuel to outer fuel nozzles  32  is turned off, and that percentage of fuel is to premix gaseous fuel is supplied to center nozzle  33 . As the unit load is further raised, premix gaseous fuel is supplied to two of outer fuel nozzles  32  and center fuel nozzle  33 . At approximately 20% load, flow is diverted from two of outer fuel nozzles  32  to three of outer fuel nozzles  32 , while flow is maintained through center fuel nozzle  33 . At approximately 30% load, the flow of premix gaseous fuel through center fuel nozzle  33  is turned off and that percentage of the premix gas fuel is delivered through the two of outer fuel nozzles  32  so that all outer fuel nozzles  32  are delivering premix gaseous fuel. For a brief period of time, fuel is supplied exclusively to the outer premixed and quaternary nozzles. As approximately 30% load, the center nozzle  33  is turned on again to deliver premix gaseous fuel through the premix gas fuel passages ( 64 ). This mode is applied with controlled fuel percentages to the premix gas nozzles up to 100% of the rated load. In order to operate in modes  1 ,  3  and  4 , the center fuel nozzle generally must be capable of stabilizing flame at equivalence ratios of greater than 0.65. 
     Those skilled in the art will appreciate that the conception, upon which the disclosure is based, may readily be utilized as a basis for designing other structures, methods, and/or systems for carrying out the several purposes of the present invention. It is important, therefore, that the claims be regarded as including such equivalent constructions insofar as they do not depart from the spirit and scope of the present invention. 
     While the disclosed embodiments of the subject matter described herein have been shown in the drawings and fully described above with particularity and detail in connection with several exemplary embodiments, it will be apparent to those of ordinary skill in the art that many modifications, changes, and omissions are possible without materially departing from the novel teachings, the principles and concepts set forth herein, and advantages of the subject matter recited in the appended claims. Hence, the proper scope of the disclosed innovations should be determined only by the broadest interpretation of the appended claims so as to encompass all such modifications, changes, and omissions. In addition, the order or sequence of any process or method steps may be varied or re-sequenced according to alternative embodiments.