Abstract:
A combustor for a gas turbine engine includes a plurality of primary nozzles configured to diffuse or premix fuel into an air flow through the combustor; and a secondary nozzle configured to premix fuel with the air flow. Each premixing nozzle includes a center body, at least one vane, a burner tube provided around the center body, at least two cooling passages, a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube. The cooling passages prevent the walls of the center body, the vane(s), and the burner tube from overheating during flame holding events.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support under Contract No. DE-FC26-05NT42643 awarded by the Department of Energy. The Government has certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to a flame tolerant secondary fuel nozzle in a premixer that includes cooling. 
     BACKGROUND OF THE INVENTION 
     Secondary nozzles in a combustor of a gas turbine may be permanently damaged when a flame is held in the premixing section of the nozzle. The use of high reactivity fuels makes this possibility more likely and confines operability of the gas combustor in a limited fuel space. 
     Use of high reactivity fuels increases flame holding risk that causes hardware damage and makes it more difficult to operate these fuels under premix operation. This has been previously addressed by so-called partially premixed design concepts that compromise mixing versus flame holding risk and increases NOx emissions. 
     Referring to  FIG. 1 , an exemplary gas turbine  12  includes a compressor  14 , a dual stage, dual mode combustor  16  and a turbine  18  represented by a single blade. Although not specifically shown, the turbine  18  is drivingly connected to the compressor  14  along a common axis. The compressor  14  pressurizes inlet air which is then turned in direction or reverse flowed to the combustor  16  where it is used to cool the combustor and also used to provide air to the combustion process. The gas turbine  12  includes a plurality of the combustors  16  (one shown) which are located about the periphery of the gas turbine  12 . A transition duct  20  connects the outlet end of its particular combustor  16  with the inlet end of the turbine  18  to deliver the hot products of the combustion process to the turbine  18 . 
     Referring to  FIGS. 1 and 2 , each combustor comprises a primary or upstream combustion chamber  24  and a second or downstream combustion chamber  26  separated by a venturi throat region  28 . The combustor is surrounded by a combustor flow sleeve  30  which channels compressor discharge air flow to the combustor. The combustor is further surrounded by an outer casing  31  which is bolted to the turbine casing  32 . 
     Primary nozzles  36  provide fuel delivery to the upstream combustion chamber  24  and are arranged in an annular array around a central secondary diffusion nozzle  38 . Each combustor may include six primary nozzles and one secondary nozzle, although it should be appreciated that other arrangements may be provided. Fuel is delivered to the nozzles through plumbing  42 . Ignition in the primary combustor is caused by spark plug  48  and in adjacent combustors by crossfire tubes  50 . 
     Referring to  FIG. 2 , a primary diffusion nozzle  36  includes a fuel delivery nozzle  54  and an annular swirler  56 . The nozzle  54  delivers only fuel which is then subsequently mixed with swirler air for combustion. The centrally located secondary nozzle  38  contains a major fuel/air premixing passage and a pilot diffusion nozzle. 
     During base-load operation, the dual stage, dual mode combustor is designed to operate in a premix mode such that all of the primary nozzles  36  are simply mixing fuel and air to be ignited by the secondary premixed flame supported by the secondary nozzle  38 . This premixing of the primary nozzle fuel and ignition by the secondary pilot diffusion nozzle leads to a lower NOx output in the combustor. 
     Referring still to  FIG. 2 , a diffusion piloted premix nozzle  100  includes a diffusion pilot having a fuel delivery pipe. The diffusion pilot further includes an air delivery pipe coaxial with and surrounding the fuel delivery axial pipe portion. The air input into the air delivery pipe is compressor discharge air which is reverse flowed around the combustor  16  into the volume  76  defined by the flow sleeve  30  and the combustion chamber liner  78 . The diffusion pilot includes at its discharge end a first or diffusion pilot swirler for the purpose of directing air delivery pipe discharge air to the diffusion pilot flame. 
     A premix chamber  84  is defined by a sleeve-like truncated cone which surrounds the diffusion pilot and includes a discharge end (as shown by the flow arrows) terminating adjacent the diffusion pilot discharge end. Compressor discharge air is flowed into the premix chamber  84  from volume  76  in a manner similar to the manner in which air is supplied to the air delivery pipe. The plurality of radial fuel distribution tubes extend through the air delivery pipe and into the premix chamber  84  such that the injected fuel and air are mixed and delivered to a second or premix chamber swirler annulus between the diffusion pilot and the premix chamber truncated cone. Further details of the combustor and gas turbine engine shown in  FIGS. 1 and 2  are disclosed in, for example, U.S. Pat. No. 5,193,346 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to one embodiment of the invention, a combustor for a gas turbine engine comprises a plurality of primary nozzles configured to diffuse fuel into an air flow through the combustor; and a secondary nozzle configured to premix fuel with the air flow, the secondary nozzle comprising a fuel passage, a center body provided around the fuel passage, a burner tube provided around the center body and defining an annular air-fuel mixing passage between the center body and the burner tube, at least one vane in the annular air-fuel mixing passage configured to swirl the air flow, and at least two cooling passages comprising a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube. 
     According to another embodiment of the invention, a method of operating a combustor of a gas turbine engine is provided. The combustor comprises a plurality of primary nozzles provided in a primary combustion chamber and configured to diffuse fuel of a fuel supply to the combustor into an air flow through the combustor; and a secondary nozzle provided in a secondary combustion chamber and configured to premix fuel of the fuel supply with the air flow, the secondary nozzle comprising a fuel passage, a center body provided around the fuel passage, a burner tube provided around the center body and defining an annular air-fuel mixing passage between the center body and the burner tube, at least one vane in the annular air-fuel mixing passage configured to swirl the air flow, and at least two cooling passages comprising a fuel cooling passage to cool surfaces of the center body and the at least one vane, and an air cooling passage to cool a wall of the burner tube. The method comprises providing an air flow to the combustor; and providing a fuel supply to at least one of the plurality of primary nozzles and the secondary nozzle; diffusing any fuel supplied to the primary nozzles into the air flow; premixing any fuel supplied to the secondary nozzle with the air flow; cooling the center body and the at least one vane with a portion of the fuel in the fuel cooling passage; and cooling the burner tube with a portion of the air flow between the burner tube and an outer peripheral wall. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is an elevation view of a gas turbine engine according to the prior art shown in partial cross section; 
         FIG. 2  is an enlarged detail elevation view of a combustor section of the gas turbine engine of  FIG. 1 ; 
         FIG. 3  schematically depicts a combustor according to an exemplary embodiment of the invention; 
         FIG. 4  schematically depicts a combustor head end according to an exemplary embodiment of the invention and a combustion liner taken from  FIG. 3 ; 
         FIG. 5  schematically depicts the combustor head end of  FIG. 4  including a flame tolerant secondary fuel nozzle according to an exemplary embodiment of the invention; 
         FIGS. 6-9  schematically depict operation of a combustor according to an exemplary embodiment of the invention; and 
         FIGS. 10 and 11  disclose a flame tolerant secondary fuel nozzle according to an exemplary embodiment of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIG. 3 , a combustor  2  according to an embodiment includes a combustor head end  4  having an array of primary nozzles  6  and a secondary nozzle  102 . A combustion chamber liner  10  comprises a venturi  46  provided between a primary combustion chamber  40  and a secondary combustion chamber  44 . The combustion chamber liner  10  is provided in a combustor flow sleeve  8 . A transition duct  22  is connected to the combustion chamber liner  10  to direct the combustion gases to the turbine. Dilution holes  34  may be provided in the transition duct  22  for late lean injection. 
     Referring to  FIG. 4 , the combustor head end  4  comprises the array of primary nozzles  6  and the secondary nozzle  102 . As shown in  FIG. 4 , the primary nozzles  6  are provided in a circular array around the secondary nozzle  102 . It should be appreciated, however, that other arrays of the primary nozzles  6  may be provided. 
     The combustion chamber liner  10  comprises a plurality of combustion chamber liner holes  52  through which compressed air flows to form an air flow  54  for the primary combustion chamber  40 . It should also be appreciated that compressed air flows on the outside of the combustion chamber liner  10  to provide a cooling effect to the primary combustion chamber  40 . 
     The secondary nozzle  102  comprises a plurality of swirl vanes  108  that are configured to pre-mix fuel and air as will be described in more detail below. The secondary nozzle  102  extends into the primary combustion chamber  40 , but not so far as the venturi  46 . 
     Referring to  FIG. 5 , the combustor head end  4  comprises an end cover  60  having an end cover surface  62  to which the primary nozzles  6  are connected by sealing joints  64 . The secondary nozzle  102  comprises a fuel passage  66  that is supported by the end cover  60 . The secondary nozzle  102  further comprises an air flow inlet  68  for the introduction of air into the secondary nozzle  102 . 
     A nozzle center body  106  surrounds the end portion of the fuel passage  66 . The nozzle center body  106  comprises an end wall  114 . In the fuel passage  66 , the fuel flows downstream until it contacts the end wall  114 . The fuel flow then enters a reverse flow passage  116  and flows upstream as explained further below. As used herein, the term downstream refers to a direction of flow of the combustion gases through the combustor toward the turbine and the term upstream may represent a direction away from or opposite to the direction of flow of the combustion gases through the combustor. 
     The nozzle center body  106  may comprise annular ribs  118  to enhance heat transfer and cool the outer surface of the center body  106 . It should also be appreciated that the fuel passage  66  may comprise ribs, for example on the outer circumferential surface. The fuel passage  66  may comprise a plurality of holes  110  that bypass fuel directly to the swirling vanes  108  to control cooling and the pressure drop in the secondary nozzle  102 . 
     The fuel flows upstream in the reverse flow passage  116  into a cooling chamber  70 . The fuel then flows around a divider  74  into an outlet chamber  72 . The divider  74  may, for example, be a piece of metal that restricts the direction of flow of the fuel into the outlet chamber  72 , thus causing the fuel to internally cool all surfaces of the vanes  108 . The cooling chamber  70  and the outlet chamber  72  may be described as a non-linear coolant flow passage, e.g., a zigzag coolant flow passage, a U-shaped coolant flow passage, a serpentine coolant flow passage, or a winding coolant flow passage. A portion of the fuel may also flow directly from the cooling chamber  70  to the outlet chamber  72  through a by-pass hole  88  formed in the divider  74 . 
     The by-pass hole  88  may allow, for example, approximately 1-50%, 5-40%, or 10-20%, of the total fuel flow flowing from the cooling chamber  70  into the outlet chamber  72  to flow directly between the chambers  70 ,  72 . Utilization of the by-pass hole  88  may allow for adjustments to any fuel system pressure drops that may occur, adjustments for conductive heat transfer coefficients, or adjustments to fuel distribution to fuel injection ports  86 . The by-pass hole  88  may improve the distribution of fuel into and through the fuel injection ports  86  to provide more uniform distribution. The by-pass hole  88  may also reduce the pressure drop from the cooling chamber  70  to the outlet chamber  72 , thereby helping to force the fuel through the fuel injection ports  86 . Additionally, the use of the by-pass hole  88  may allow for tailored flow through the fuel injection ports  86  to change the amount of swirl that the fuel flow contains prior to injection into a fuel-air mixing passage  112  via the injection ports  86 . 
     The fuel is ejected from the outlet chamber  72  through the fuel injection ports  86  formed in the swirl vanes  108 . The fuel is injected from the fuel injection ports  86  into the fuel-air mixing passage  112  for mixing with the air flow from the air flow inlet  68  of the secondary nozzle  102 . The swirl vanes  108  swirl the air flow from the air flow inlet  68  to improve the fuel-air mixing in the passage  112 . 
     Referring still to  FIG. 5 , the secondary nozzle  102  includes a burner tube  122  that surrounds the nozzle center body  106 . The fuel-air mixing passage  112  is provided between the nozzle center body  106  and the burner tube  122 . An outer peripheral wall  104  is provided around the burner tube  122  and defines a passage  96  for air flow. The burner tube  122  includes a plurality of rows of air cooling holes  120  to provide for cooling by allowing the coolant to form a film on the burner tube, protecting it from hot combustion gases. Coolant is also directed axially upstream within an annular cavity formed between the burner tube  122  and the outer peripheral wall  104 , in order that coolant may exit the cooling holes  120  upstream of the leading half of vanes  108 . The holes  120  may be angled in the range of 0° to 45° degree with reference to a downstream wall surface. The hole size, the number of holes in a circular row, and/or the distance between the hole rows may be arranged to achieve the desired wall temperature during flame holding events. 
     Operation of the combustor will now be described with reference to  FIGS. 6-9 . As shown in  FIG. 6 , during primary operation, which may be from ignition up to, for example, 20% of the load of the gas turbine engine, all of the fuel supplied to the combustor is primary fuel  80 , i.e. 100% of the fuel is supplied to the array of primary nozzles  6 . Combustion occurs in the primary combustion chamber  40  through diffusion of the primary fuel  80  from the primary fuel nozzles  6  into the air flow  54  through the combustor  4 . 
     As shown in  FIG. 7 , a lean-lean operation of the combustor occurs when the gas turbine engine is operated at, for example, 20-50% of the load of the gas turbine engine. Primary fuel  80  is provided to the array of primary nozzles  6  and secondary fuel  82  is provided to the secondary nozzle  102 . For example, about 70% of the fuel supplied to the combustor is primary fuel  80  and about 30% of the fuel is secondary fuel  82 . Combustion occurs in the primary combustion chamber  40  and the secondary combustion chamber  44 . 
     As used herein, the term primary fuel refers to fuel supplied to the primary nozzles  6  and the term secondary fuel refers to fuel supplied to the secondary nozzle  102 . 
     In a second-stage burning, shown in  FIG. 8 , which is a transition from the operation of  FIG. 7  to a pre-mixed operation described in more detail below with reference to  FIG. 9 , all of the fuel supplied to the combustor is secondary fuel  82 , i.e. 100% of the fuel is supplied to the secondary nozzle  102 . In the second-stage burning, combustion occurs through pre-mixing of the secondary fuel  82  and the air flow from the inlet  68  of the secondary nozzle  102 . The pre-mixing occurs in the pre-mixing passage  112  of the secondary nozzle  102 . 
     As shown in  FIG. 9 , the combustor may be operated in a pre-mixed operation at which the gas turbine engine is operated at, for example, 50-100% of the load of the gas turbine engine. In the pre-mixed operation of  FIG. 9 , the primary fuel  80  to the primary nozzles  6  is increased from the amount provided in the lean-lean operation of  FIG. 7  and the secondary fuel  82  to the secondary nozzle  102  is decreased from the amount from provided in the lean-lean operation shown in  FIG. 7 . For example, in the pre-mixed operation of  FIG. 9 , about 80-83% of the fuel supplied to the combustor may be primary fuel  80  and about 20-17% of the fuel supplied to the combustor may be secondary fuel  82 . 
     As shown in  FIG. 9 , during the pre-mixed operation, combustion occurs in the secondary combustion chamber  44  and damage to the secondary nozzle  102  is prevented due to the cooling measures. Referring to  FIG. 4 , flashback may occur in the event that the flame speed  58  is greater than the velocity of the air flow  54  in the primary combustion chambers  40 . Control of the air-fuel mixture in the secondary nozzle  102 , i.e. control of the secondary fuel  82 , provides control of the flame speed and prevents the flame from crossing the venturi  46  into the primary combustion chamber  40 . 
     Referring to  FIGS. 10 and 11 , secondary nozzle  124  comprises an inlet flow conditioner (IFC)  126 , an air swirler assembly  132  with natural gas fuel injection, and a diffusion gas tip  146 . A shroud extension  134  extends from the air swirler assembly  132 . 
     Air enters the secondary nozzle  124  from a high pressure plenum  90 , which surrounds the entire secondary nozzle  124  except the discharge end, which enters the combustor reaction zone  94 . Most of the air for combustion enters the premixer via the IFC  126 . The IFC  126  includes a perforated cylindrical outer wall  128  at the outside diameter, and a perforated end cap  130  at the upstream end. Premixer air enters the IFC  126  via the perforations in the end cap  130  and the cylindrical outer wall  128 . 
     The function of the IFC  126  is to prepare the air flow velocity distribution for entry into the premixer. The principle of the IFC  126  is based on the concept of backpressuring the premix air before it enters the premixer. This allows for better angular distribution of premix air flow. The perforated wall and endcap  128 ,  130  perform the function of backpressuring the system and evenly distributing the flow circumferentially around the IFC annulus. Depending on the desired flow distribution within the premixer, appropriate hole patterns for the perforated wall and endcap  128 ,  130  are selected. 
     Referring to  FIG. 11 , the air swirler assembly of the secondary nozzle  124  comprises a plurality of swirling vanes  140  and a plurality of spokes, or pegs,  142  provided between the swirling vanes  140 . Each spoke  142  comprises a plurality of fuel injection holes  144  for injecting fuel into the air swirled by the vanes  140 . Natural gas inlet ports  136  allow natural gas to be introduced into fuel passages  138  that are in communication with the spokes  142 . A nozzle extension  148  is provided between the air swirler assembly and the diffusion gas tip  146 . A bellows  150  may be provided to compensate for differences in thermal expansions. 
     Although the various embodiments described above include diffusion nozzles as the primary nozzles, it should be appreciated that the primary nozzles may be premixed nozzles, for example having the same or similar configuration as the secondary nozzles. 
     The flame tolerant nozzle enhances the fuel flexibility of the combustion system. The flame tolerant nozzle as the secondary nozzle in the combustor makes the combustor capable of burning full syngas as well as natural gas. The flame tolerant nozzle may be used as a secondary nozzle in the combustor and thus make the combustor capable of burning full syngas or high hydrogen, as well as natural gas. The flame tolerant nozzle, combined with a primary dual fuel nozzle, will make the combustor capable of burning both natural gas and full syngas fuels. It expands the combustor&#39;s fuel flexibility envelope to cover a wide range of Wobbe number and reactivity, and can be applied to oil and gas industrial programs. 
     The cooling features of the flame tolerant nozzle, including for example, the fuel cooled center body, the tip of the center body, the swirling vanes of the pre-mixer, and the air cooled burner tube, enable the nozzle to withstand prolonged flame holding events. During such a flame holding event, the cooling features protect the nozzle from any hardware damage and allows time for detection and correction measures that blow the flame out of the pre-mixer and reestablish pre-mixed flame under normal mode operation. 
     While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.