Abstract:
A method for combustion in a combustor in a gas turbine including: fueling the center fuel nozzle with a fuel-rich mixture of gaseous fuel and air and fueling the outer fuel nozzles with a fuel-lean mixture of fuel and air; igniting the fuel-rich mixture injected by the center fuel nozzle while the fuel-lean mixture injected by the outer combustors is insufficient to sustain ignition; stabilizing a flame on the center fuel nozzle using the bluff body and while the outer fuel nozzles inject the fuel-lean mixture; staging fuel to the outer nozzles by increasing a fuel ratio of the fuel-lean mixture, and after the outer nozzles sustain ignition, reducing fuel applied to the center nozzle.

Description:
RELATED APPLICATION 
   This application is a divisional of U.S. patent application Ser. No. 10/821,975, filed Apr. 12, 2004, and is incorporated by reference herein in its entirety. 

   BACKGROUND OF THE INVENTION 
   The invention relates to a multi-nozzle combustor for a gas turbine and to limiting dynamic flame oscillations in such a combustor. 
   Industrial gas turbines have a combustion section typically formed by an annular array of combustors. Each combustor is a cylindrical chamber which receives gas and/or liquid fuel and combustion air which are combined into a combustible mixture. The air-fuel mixture burns in the combustor to generate hot, pressurized combustion gases that are applied to drive a turbine. 
   The combustors are generally dual mode, single stage multi-burner units. Dual mode refers to the ability of the combustor to burn gas or liquid fuels. Single stage refers to a single combustion zone defined by the cylindrical lining of each combustor. Conventional combustors are shown in U.S. Pat. Nos. 5,722,230 and 5,729,968. 
   Stabilizing a flame in a combustor assists in providing continuous combustion, efficient generation of hot combustion gases and reduced emissions from combustion. The flames of combustion tend to oscillate due to dynamic pressure fluctuations in the combustors especially during combustion transition operations to lean fuel-air mixtures. These oscillations can extinguish the flame in a combustor and fatigue the combustor. There is a long felt need for combustors to have good flame stabilization, combustor performance, and reduced emissions. 
   BRIEF DESCRIPTION OF THE INVENTION 
   The invention may be embodied as a single stage combustor for a gas turbine comprising: an annular array of outer fuel nozzles arranged about a center axis of the combustor; a center fuel nozzle aligned with the center axis, wherein the center fuel nozzle is substantially smaller than each of the outer fuel nozzles, wherein said combustor further comprises a pre-mix combustor operating mode in which the center nozzle receives a fuel rich air-fuel mixture as compared to a fuel mixture applied to the outer fuel nozzles, and a diffusion mode wherein the outer fuel nozzles receive a lean fuel-air mixture and the center fuel nozzle receives no more than the lean fuel-air mixture. 
   The invention may be further embodiment as a single stage combustor for a gas turbine comprising: an annular array of outer fuel nozzles arranged about a center axis, wherein each of said outer fuel nozzle comprises a gaseous fuel passage and a liquid fuel passage, a center fuel nozzle aligned with the center axis and having a fuel passage consisting of at least one gaseous fuel passage, wherein the center fuel nozzle is substantially smaller than each of the outer fuel nozzles, and a pre-mix combustor operating mode in which the center nozzle receives a fuel rich air-fuel mixture as compared to a fuel mixture applied to the outer fuel nozzles. 
   The invention may be further embodied as a method for combustion in a dual mode, single stage combustor in a gas turbine, wherein said combustor comprises an annular array of outer fuel nozzles arranged about a center axis and a small center fuel nozzle aligned with the center axis, said method comprising: fueling the center fuel nozzle with a fuel-rich mixture of gaseous fuel and air and fueling the outer fuel nozzles with a fuel-lean mixture of fuel and air, wherein a fuel rate applied to the center fuel nozzle and a fuel rate applied to each of the outer fuel nozzles are substantially similar; igniting the fuel-rich mixture injected by the center fuel nozzle while the fuel-lean mixture injected by the outer combustors is insufficient to sustain ignition; stabilizing a flame on the center fuel nozzle while the outer fuel nozzles inject the fuel-lean mixture; staging fuel to the outer nozzles by increasing a fuel ratio of the fuel-lean mixture, and after the outer nozzles sustain ignition, reducing fuel applied to the center nozzle. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a partial section of a dual mode, single stage, multi-burner combustor showing an outer fuel injection nozzle and a smaller center fuel nozzle. 
       FIG. 2  is a sectional view of a fuel injection nozzle, such as either or both the outer and center fuel nozzle shown in  FIG. 1 . 
       FIG. 3  is an enlarged end partial cross-sectional view of the nozzle shown in  FIG. 2 . 
       FIG. 4  is a front end view of the nozzle shown in  FIG. 2 . 
       FIG. 5  is a simplified end view of a multi-burner combustor of the type shown in  FIG. 2 . 
       FIG. 6  is an enlarged partial cross-sectional view of a second embodiment of a center fuel nozzle. 
       FIG. 7  is a simplified end view of a multi-burner combustor of the type shown in  FIG. 6 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   In a gas turbine having a plurality of combustors, each having an annular array of outer fuel nozzles arranged about a center axis, an improved range of operability (such as flame stability, emissions, peak, fire, turndown, and fuel composition) may be achieved by locating a small-sized center fuel nozzle on the center axis of each combustor and within an annular array of larger outer fuel nozzles. The center fuel injection nozzle is substantially smaller than outer nozzles in a surrounding nozzle array of the combustor. The rate of fuel flow (which may be solely a pre-mixture of gaseous fuel and air) through the center nozzle is substantially less than the fuel flow rate through the larger outer nozzles so as to reduce emissions from the combustor when the center fuel nozzle operates on a fuel-rich mixture. 
   The small center nozzle provides a flame stabilization structure that creates a flame front anchored by a re-circulation zone in the front face of the center nozzle. The reduced size center fuel injection nozzle provides a pilot flame and recessed bluff body (which forms the recirculation zone) to initiate and stabilize a flame in a combustor, and to reduce the dynamic pressure oscillations that occur in lean dry NOx premix combustion. By reducing oscillations, the center nozzle enhances flame stability. Moreover, the small sized center nozzle has a reduced rate of fuel flow (as compared to the fuel flow through the outer nozzles) to avoid excessive emissions, especially when the center nozzle operates on a fuel rich air-fuel mixture as compared to the fuel mixture flowing to the outer nozzles. The center fuel injection nozzle may also provide improved peak fire turndown and allow for varying fuel composition. Furthermore, the reduced size center nozzle allows the combustor to operate in overall combustor fuel lean conditions that were otherwise not accessible due to dynamic pressure oscillators. 
     FIG. 1  shows a combustor  14 , in partial cross-section, for a gas turbine  10  having a compressor  12  (partially shown), a plurality of combustors  14  (one shown), and a turbine represented here by a single turbine blade  16 . The turbine is drivingly connected to the compressor blading along a common axis. Compressor air is reverse flowed to the combustor  14  where it is used to cool the combustor and to provide air to the combustion process. 
   The gas turbine includes a plurality of combustors  14  arranged in an annular array about the periphery of the gas turbine casing. High pressure air from the compressor  10  flows (see flow arrows in  FIG. 1 ) into the combustor through a compressed air inlet  19  near the outlet end  21  of the combustor. The compressed air flows through an annular passage defined by the combustor flow sleeve  34  and the combustor liner  38  to a combustor inlet end  23  where there is arranged a plurality of air-fuel injectors  32 ,  132 . At the inlet end of each combustor, compressed air and fuel mix and flow into a combustion burning zone  70 . Ignition is achieved in the combustors  14  by spark plug(s)  20  in conjunction with cross fire tubes  22  (one shown). At the opposite end of the burning zone, hot combustion gases flow into a double-walled transition duct  18  that connects the outlet end  21  of each combustor with the inlet end of the turbine (see blade  16 ) to deliver the hot combustion gases to the turbine. 
   Each combustor  14  includes a substantially cylindrical combustion casing  24  which is secured at an open forward end  19  to the gas turbine casing  26  by bolts  28 . The inlet end  21  of the combustion casing is closed by an end cover assembly  30  which may include conventional fuel and air supply tubes, manifolds and associated valves for feeding gas, liquid fuel and air (and water if desired) to the combustor as described in greater detail below. The end cover assembly  30  receives a plurality (for example, five) outer fuel nozzle assemblies  32  arranged in an annular array about a longitudinal axis of the combustor (see  FIG. 5 ). The array of outer fuel nozzle assemblies  32  is arranged around a center fuel nozzle assembly  132  that is small (in terms of size and fuel flow) relative to the outer nozzle assemblies  32 . 
   A substantially cylindrical flow sleeve  34  is concentrically mounted in the casing  24 . The flow sleeve connects at its forward end to the outer wall  36  of a double walled transition duct  18 . Compressor air flows through the duct  18 , over and through the flow sleeve, and to the inlet end  21  of the combustor. 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  where fore and aft sections of the combustor casing  24  are joined. 
   The flow sleeve  34  is concentrically arranged with a combustion liner  38  which is connected at one end with the inner wall  40  of the transition duct  18 . The opposite 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 associated mounting flange assembly  41  ( FIG. 5 ). The outer wall  36  of the transition duct  18 , as well as that portion of flow sleeve  34  extending forward of the location where the combustion casing  24  is bolted to the turbine casing 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 pre-mix tubes  46 , one for each fuel nozzle assembly  32 ,  132 . Each pre-mix tube  46  is supported within the combustion liner cap assembly  42  at its forward and rearward ends by front and rear plates  47 ,  49 , respectively, each provided with openings aligned with the open-ended pre-mix tubes  46 . This arrangement is best seen in  FIG. 5 , with openings  43  (for nozzle  32  and their premix tube),  143  (for nozzle  132  and its premix tube) shown in the front plate  47 . The front plate  47  (such as an impingement plate provided with an array of cooling apertures) may be shielded from the thermal radiation of the combustor flame by shield plates  45 . 
   The rear plate  49  mounts to a plurality of rearwardly extending floating collars  48  (one for each pre-mix 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 corresponding fuel nozzle assembly  32 ,  132 . The arrangement is such that air flowing in the annular space between the liner  38  and flow sleeve  34  is forced to reverse direction at the combustor inlet end  23  of the combustor (between the end cover assembly  30  and liner cap assembly  42 ) and to flow through the swirlers  50  and pre-mix tubes  46  before entering the burning zone  70  within the liner  38 , downstream of the pre-mix tubes  46 . 
     FIGS. 2 ,  3 , and  4  show an exemplary conventional fuel nozzle assembly  32  that includes a rearward supply section  52  with inlets for receiving liquid fuel, atomizing air, diffusion gas fuel and pre-mix gas fuel. The fuel nozzle assembly shown in  FIGS. 2 ,  3  and  4  is most suitable for the outer nozzles  32 . It may also be used as a center nozzle  32 , in a scaled down form. The fuel nozzle assembly includes suitable connecting passages for supplying these fluids to a respective passage in a forward delivery section  54  of the fuel nozzle assembly. The forward delivery section  54  of the fuel nozzle assembly is comprised of a series of concentric tubes. 
   In the fuel nozzle  32 , the two radially outermost concentric tubes  56 ,  58  provide a pre-mix gas passage  60  which receives pre-mix gas fuel from an inlet  62  connected to fuel passage  60  by means of conduit  64 . The pre-mix gas passage  60  also communicates with a plurality (for example, eleven) radial fuel injectors spokes  66 , each of which is provided with a plurality of fuel injection ports or holes  68  for discharging gas fuel into a pre-mix zone  69  ( FIG. 1 ) located within the pre-mix tube  46 . The injected fuel mixes with air reverse flowed from the compressor  12 , and swirled by means of the annular swirler  50  surrounding the fuel nozzle assembly upstream of the radial injectors  66 . The pre-mix passage  60  is sealed by an O-ring  72  ( FIG. 3 ) at the forward or discharge end of the fuel nozzle assembly, so that pre-mix fuel may exit only via the radial fuel injectors  66 . 
   The next adjacent passage  74  is formed between concentric tubes  58  and  76 , and supplies diffusion gas to the burning zone  70  ( FIG. 1 ) of the combustor via orifice  78  ( FIG. 3 ) at the forwardmost end of the fuel nozzle assembly  32 ,  132 . The forwardmost (or discharge) end of the nozzle is located within the pre-mix tube  46 , but relatively close to the forward end thereof. The diffusion gas passage  74  receives diffusion gas from an inlet  80  via conduit  82 . 
   A third passage  84  is defined between concentric tubes  76  and  86 , and supplies atomizing air to the burning zone  70  via orifice  88  where it then mixes with diffusion fuel exiting the orifice  78 . The atomizing air is supplied to passage  84  from an inlet  90  via conduit  92 . 
   The fuel nozzle assembly  32  is also provided with an optional water passage  94  for supplying water to the burning zone  70  to effect NOx reductions in a manner understood by those skilled in the art. The water passage may be included in the outer fuel nozzles and not included in the center fuel nozzle. The water passage  94  is defined between tube  86  and adjacent concentric tube  96 . Water exits the nozzle via an orifice  98 , radially inward of the atomizing air orifice  88 . 
   Tube  96 , the innermost of the series of concentric tubes forming the fuel injector nozzle, forms a central passage  100  for liquid fuel which enters the passage by means of inlet  102 . The liquid fuel exits the nozzle by means of a discharge orifice  104  in the center of the nozzle. The liquid fuel capability may be provided as a back-up system. Passage  100  may be normally shut off while the gas turbine is in its normal gas fuel mode. 
     FIGS. 6 and 7  show an enlarged cross-sectional view and end view of a center fuel nozzle  132  that is smaller and has fewer passages than the outer fuel nozzle  32  (shown in  FIGS. 2 ,  3  and  4 ). The center nozzle may be structurally similar to the outer fuel nozzle assemblies  32 , except that the center nozzle (may not have passages (or their associated tubes) for water (see passage  94  in  FIG. 3 ), liquid fuel (see central passage  100 ), or air (see passage  84 ). 
   The center nozzle  132  may include a bluff body  160  formed at the nozzle discharge face and recessed, e.g., by one and one half inches, from the outer lip  162  of the outer nozzle tube  56 . The bluff body may be formed from the end of cylinder  164  that is coaxial to the other tubes  56  and  58 . Immediately downstream of the bluff body is formed an aerodynamic stagnant zone which stabilizes combustion of the fuel mixture from the center nozzle. The center nozzle  132  includes a gas fuel passage  60  similar to he passage  60  shown in  FIGS. 2 ,  3  and  4  and having an associated radial fuel injector. The center nozzle also includes a diffusion gas passage  74  similar to the passage  74  shown in  FIGS. 2 ,  3  and  4 . The recessed bluff body  160  and tube lip  162  define a flame recirculation zone  166  to stabilize a flame in front of the center nozzle. 
   The combustor functions in a dual mode, single stage manner. However, center nozzle  132  as shown in  FIG. 6  does not operate on liquid fuel. At low turbine loads, diffusion gas fuel is fed to each nozzle  32 ,  132  and dedicated pre-mix tube assembly  46  through inlet  80 , conduit  82  and passage  74  for discharge via orifice  78  into the combustion zone  70  of the combustor. In burning zone  70 , the diffusion gas mixes with atomizing air discharged from passage  84  via orifice  88 . The mixture of atomizing air and diffusion gas is ignited by the spark plug  20  and burned in the combustion zone  70  within the liner  38 . At higher turbine loads, pre-mix gas fuel is supplied to passage  60  via inlet  62  and conduit  64  for discharge through orifices  68  in radial injectors  66 . The pre-mix fuel mixes with air entering the pre-mix tube  46  by means of swirlers  50 , the mixture igniting by the pre-existing flame in burning zone  70  in liner  38 . During pre-mix operation, fuel to the diffusion passage  74  is shut down. 
   Referring again to  FIGS. 1 and 2 , there are provided a plurality of circumferentially spaced, radially extending fuel injector pegs  105  upstream of the swirlers  50  for each nozzle. As illustrated in  FIG. 2 , the pegs  105  lie in communication with a manifold  137  about the outer tube  56  of each nozzle. Pre-mix fuel is supplied to manifold  137  for injection into the reverse flow of air from the compressor for flow with the air through the swirler and past the downstream injectors  66 . The two axial locations for injection of the pre-mix fuel are upstream and downstream of the swirlers  50 . 
   The center pre-mix tube  150  (see  FIG. 5 ), and center fuel nozzle  132  (see  FIGS. 1 and 5 ) are located on the center axis of the combustor in opening  107  (see  FIG. 5 ), and surrounded concentrically by an annular array of nozzle assemblies  32 . The center pre-mix tube  150  is supported within the combustion liner cap assembly  42  at its forward and rearward ends by front and rear plates  47  and  49  respectively, with opening  107  aligned with center pre-mix tube  150 . 
   The outer diameter (OD) of the center nozzle assembly  132  may be 85% or less than the OD of an outer fuel injection nozzle  32 . The center fuel nozzle  132  may operate as a premix pilot burner that receives a rich fuel-air mixture as compared to the mixture received by the outer nozzles  32 , at least during a combustor transition to a lean low NOx combustion condition. The center nozzle  132  stabilizes the flame in the combustion zone  70 , especially during certain transition operations when flame pressure oscillations tend to be severe in the burning zone. For example, a rich air-fuel mixture may be supplied to the center nozzle  132  to anchor a flame just downstream of the center nozzle in the burning zone  70 . At the same time, the fuel-air mixture applied to the outer nozzles  32  may be excessively lean and inadequate to independently support combustion. As the fuel-air mixture to the outer fuel nozzles transitions to a combustible mixture the flame anchored by the center nozzle propagates axisymetrically to the outer nozzles. In this way the fuel is staged from the center nozzle to the outer nozzles to the combustor to a lean fuel-air mixture. The fuel staging from the center nozzle to the outer nozzle may be performed axisymetrically to control the propagation of the flame from the center nozzle  132  to the outer nozzle  32 . When the outer nozzles are operating stably with a lean-mixture, the center nozzle  132  may also be operated at the same lean mixture. 
   The reduced size of the center fuel nozzle assembly  132  results in relatively smaller fuel flow through the center nozzle assembly as compared to the fuel flowing through one of the larger outer fuel nozzle assemblies  32  though the fuel-air ratio may be higher. The fuel flow through a center fuel nozzle assembly may be 80% to 95% of the fuel flow through one of the outer fuel nozzle assemblies at the low emissions baseload design point. The center nozzle  132  thus has limited adverse impact on emission levels even when the center nozzle operates at a rich air-fuel mixture. 
   In full pre-mixed mode (all six nozzles fueled), the presence of the center nozzle  132  allows turndown of the combustor to a lower temperature, and hence lower load, than is possible without the center nozzle, while still maintaining low NOx, CO and UHC emissions. The extended turndown in pre-mixed operation is achieved as described below. 
   During high temperature operation, low NOx emissions are achieved by maintaining the same fuel/air ratio in all nozzles, both center nozzle  132  and outer nozzles  32 . The proportion of fuel to the center nozzle will be the same as its proportion of total burner air flow. At constant air flow, as total fuel is reduced while still maintaining equal fuel/air ratio in all nozzles, the nozzles begin to approach a lean blow-out limit. 
   The use of the center nozzle  132  also improves flame stability for low load operation. Typically, operation between full speed, no load and 40% of load occurs with some or all of the outer nozzles running in diffusion mode. At very low loads in diffusion mode, the low fuel/air ratio in the outer nozzles  32  makes the flame unstable and prone to blow-out. By running rich nozzle  132  at full speed, no-load (FSNL) the center conditions, in either diffusion or pre-mixed mode, the flame stability can be greatly enhanced. As the load is increased, fuel can be gradually added to the outer nozzles  32 , in stages, or to all nozzles at once. Even when the fuel/air ratio in the outer nozzles is very lean, a high fuel/air ratio in the center nozzle  132  can be used to maintain flame stability and provide an ignition source for the fuel in the outer nozzles  32 . 
   The reduced size center nozzle  132  has good flame stabilization that limits dynamic oscillations and enhances the combustor capability to run with reduced emissions. The small center nozzle  132  also allows for increased firing temperatures with low combustion pressure dynamics; reduces the need for additional fuel circuits to manage high fuel temperatures; increases the range of acceptable fuel compositions/temperatures (as expressed by the Modified Wobbe Index); expands the range of load turndown with acceptable emissions; increases system reliability by providing a common method of flame stabilization from no-load to full-load (mode transfer robustness). The above-described procedure that has been developed allows the center nozzle in a multi-nozzle combustor to be transitioned to and operated with a single (center-body) or dual-concentric (center-body and burner tube) flame front, in a manner that creates a central axis-symmetric flame front stabilization point. 
   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.