Abstract:
A fuel nozzle assembly for a combustor in a gas turbine including: a gaseous fuel passage and a fuel nozzle at a distal end of the passage; an air tube concentric with the fuel nozzle and defining an air passage between the air tube and the fuel nozzle, wherein the air tube includes a distal section extending axially beyond the fuel injection nozzle; a first fuel-air mixing zone inside the distal section of the air tube, wherein the first fuel-air mixing zone is downstream of the fuel injection nozzle; a flame holder including a porous structure with thermal barrier coating and micro swirlers and defining a downstream end of the first fuel-air mixing zone, wherein fuel and air from the first fuel-air mixing zone pass through the porous structure of the flame holder and into a combustion zone of the combustor.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates to the mixing region of the fuel nozzle assembly for a combustor in a gas turbine burning on Syngas or hydrogen fuels. Less forgiving properties of hydrogen (H2) and syngas fuels such as higher flame speeds, lower ignition times makes it impossible to use prior art designs applicable only for burning natural gas fuels. 
     Industrial gas turbines have a combustion section typically formed by an array of can-annular combustors. Each combustor includes a fuel nozzle mixing region that provides specified amounts of fuel-air mixture to a combustion zone within the combustor. The fuel-air mixture is allowed to burn inside the combustion zone to generate hot, pressurized combustion gases that drive a turbine. 
     Natural gas, e.g., primarily methane, is a common fuel for industrial gas turbines. Rapid depletion of hydrocarbon resources has led to an increased focus on using coal derived H 2  and/or syngas for industrial gas turbines. The flame speed of hydrogen and syngas is significantly higher, e.g., six to seven times faster, than the flame speed of natural gas. The burner needs to be designed to operate for greater flame speed of hydrogen and syngas which could increase the propensity for flame flashing back in to the mixing region of the fuel nozzle assembly. Flame holding in the unburnt mixing region has the potential to damage the components of the nozzle assembly. There is a strong need to design and develop devices and methods to prevent propagation of flame into the fuel nozzle assembly. 
     Syngas refers to a gas mixture available in varying amounts of carbon monoxide and hydrogen generated by the gasification of a carbon containing fuel to a gaseous product. Syngas examples include steam reforming of natural gas or liquid hydrocarbons to produce hydrogen, the gasification of coal and in some types of waste-to-energy gasification facilities. Syngas is combustible and often used as a fuel source. Syngas may be produced, for example, by gasification of coal or municipal waste. 
     Existing combustor operating on natural gas may need major modifications to accommodate additional burning of hydrogen and syngas fuels. For example, the higher flame speed of hydrogen and syngas (as compared to natural gas) may require combustor adjustments to ensure that the flame is stabilized in the combustion zone and does not propagate upstream into the mixing region of the fuel nozzle assembly. There is a strong need to develop methods and devices to modify the existing natural gas combustor designs to allow burning of hydrogen and syngas fuels. 
     BRIEF DESCRIPTION OF THE INVENTION 
     A fuel nozzle arrangement is disclosed for a combustor in a gas turbine, the assembly including: a gaseous fuel nozzle having a center axis and extending along the center axis, the fuel injection nozzle including a gaseous fuel passage and a fuel nozzle at a distal end of the passage; an air tube concentric with the fuel nozzle and defining an air passage between the air tube and the fuel nozzle, wherein the air tube includes a distal section extending axially beyond the fuel injection nozzle; a first fuel-air mixing zone defined by and inside the distal section of the air tube, wherein said first fuel-air mixing zone is downstream of the fuel injection nozzle; a flame holder comprising a porous structure and defining a downstream end of the first fuel-air mixing zone, wherein fuel and air from the first fuel-air mixing zone pass through the porous structure of the flame holder and into a combustion zone of the combustor. 
     A method is disclosed for combusting a gaseous fuel in a combustor of a gas turbine, the method comprising: injecting a gaseous fuel into an air tube of a fuel injection assembly; mixing air and gaseous fuel in the air tube, wherein in the air passes through the air tube and the gaseous fuel is discharged from a nozzle into the air tube; passing the mixture of air and gaseous fuel through a porous medium at a distal end of the air tube, and combusting the mixture of air and gaseous fuel downstream of the porous medium in a combustion zone of the combustor. 
     A method is disclosed for modifying the mixing region of a natural gas nozzle assembly to a syngas or hydrogen nozzle assembly, the method includes: placing a porous flame stabilizer a distal end of an air tube of the nozzle assembly, and allowing fuel and air from the nozzle assembly through the flame stabilizer before the mixture is burnt in a combustion zone. The method may further include selecting the porous flame stabilizer in such a way that it creates the necessary pressure drop between a combustion zone immediately downstream of the flame stabilizer and an air fuel mixture in the air tube and immediately upstream of the flame stabilizer, wherein the required pressure drop is sufficient to prevent propagation of flame through the flame stabilizer. Higher pressure drops in the porous structure results in increased fuel-air mixture velocities for stabilizing the high flame speeds of H2/syngas fuels. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a side, cross-sectional view of a combustor in an industrial gas turbine. 
         FIG. 2  is side, cross-section of the mixing region of a fuel-air nozzle assembly and a partial cross-section of a combustor. 
         FIG. 3  is a side view of a porous flame stabilizer. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     A porous flame stabilizer has been developed for insertion into a mixing region of a fuel nozzle assembly of a combustor for an industrial gas turbine. The flame stabilizer has a high porosity to allow sufficient amount of fuel and air mixture to flow through the media at a higher velocity and design pressure drops. The porous structure prevents the propagation of flame upstream in to the structure and the mixing region. The propagation of flame is prevented by allowing higher mixture velocities in the porous structure and the structure can itself act like an arrestor to the flame. The porous structure may include a thermal barrier coating (TBC) on a downstream region of the structure. The TBC shields the porous structure from being exposed to flame residing downstream of the structure. 
       FIG. 1  shows a combustor  10 , in partial cross-section, for a gas turbine  12  having a compressor  14  (partially shown), a plurality of combustors  10  (one shown), and a turbine represented here by a single turbine blade  16 . The turbine is drivingly connected to the compressor along a shaft  17 . Compressor air (C) reverse flows to the combustor  10  where it is used to cool the combustor and to provide air to the combustion process. 
     The gas turbine includes a plurality of combustors  10  arranged in an annular array about the periphery of the gas turbine casing  18 . High pressure air from the compressor  14  flows (see flow arrow C) to the combustor through a compressed air inlet  20  near the hot gas outlet  22  of the combustor. The compressed air flows (C—in a counter-direction to the combustion gases within the combustor) through an annular passage defined by the combustor flow sleeve  24  and the combustor liner  26  to a combustor inlet  28 . 
     Each combustor  10  includes a substantially cylindrical combustion casing  42  which is secured to the gas turbine casing  18 . The inlet end  28  of the combustion casing is closed by an end cover assembly  44  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  44  receives a plurality (for example, five) outer fuel nozzle assemblies  30 ,  32  arranged in an annular array about a longitudinal axis of the combustor. The array of outer fuel nozzle assemblies  32  is arranged around a center fuel nozzle assembly  30  that may be small (in terms of size and fuel flow) relative to the outer nozzle assemblies  32 . 
     Fuel, e.g., syngas, hydrogen, natural gas or a mixture of two or more of these gases, is supplied to the inlet of each fuel nozzle assemblies  30 ,  32  by fuel piping and manifolds  34  connected to the end cover assembly  44 . Gaseous fuel enters an inlet to a fuel nozzle assembly  35  having a gas passage cylinder extending along an axis of the nozzle assembly  30 ,  32 . Gaseous fuel is discharged from a distal end of the fuel nozzle assembly  35  and into an air tube gas passage(s)  48 . The air tube is concentric with the nozzle assembly, which is housed in the air tube. Compressor air (C) enters the inlet  28 , flows through the air tube and mixes with gaseous fuel discharged from the nozzle assembly  35 . The mixture of fuel and air flows into a combustion zone  46  downstream of the nozzle assemblies  30 ,  32 . 
     Each fuel nozzle assembly  30 ,  32  provide controlled amounts of fuel-air mixture to the combustion zone. The air and fuel are initially mixed in a distal end of the air tube  48  and the mixture flows into the combustion zone  46  generally defined by an air-cooled flame tube  36 . Ignition of the fuel-air mixture is achieved in the combustion zone by spark plug(s) in conjunction with cross fire tubes (not shown) between combustors  10 . At the downstream end of the combustion zone  46 , hot combustion gases flow through a double-walled transition duct  40  that connects the outlet end  22  of each combustor with the inlet end of the turbine (see blade  16 ) to deliver the hot combustion gases to the turbine. 
       FIG. 2  is a side, cross-sectional view of a fuel nozzle assembly  30 ,  32  in a combustor  10 . The fuel nozzle assembly includes a gaseous fuel nozzle assembly  35  extending along an axis of the assembly  30 ,  32 . The nozzle extends through an air tube  48 . Fuel and air manifolds at the end cover assembly  44  provide gaseous fuel and air in a controlled ratio or amount to the nozzle and air tube, respectively. The fuel nozzle  35  and air tube  48  may be conventional components of a fuel nozzle for a combustor of a natural gas turbine. For example, U.S. Published Patent Applications 2003-0121269 A1 and 2006-0288706 A1 show exemplary fuel nozzle assemblies for an industrial gas turbine capable of operating on a natural gas fuel. 
     The air tube  48  may be a cylindrical gas passage formed of a thin metal tube. The air tube is concentric with the fuel nozzle  35  which is contained within the tube. The fuel discharge nozzle  50  at the end of the fuel nozzle  35  is within the air tube  48 . The distal portion  52  of the air tube extends beyond the fuel discharge nozzle  50 . Gaseous fuel discharges from the nozzle  50  into the distal portion  52  of the air tube. Compressor air flowing through the air tube begins to mix with the gaseous in the distal portion of the air tube. 
     Swirl vanes  54 , e.g., a thin metal disc with radial vanes, may be in the air tube upstream of the nozzle  50 . The swirl vanes impart a rotation to the air flow that promotes mixing with fuel and the expansion of the mixture into the larger volume of the combustion zone  46 . Swirl vanes are conventional components often included in the air tube of natural gas air fuel nozzles  30 ,  32 . The swirl vanes may be retained when the air fuel nozzles are modified to operate on hydrogen gas or syngas. Alternatively, the swirl vanes may be removed when the nozzles are modified to operate on hydrogen gas or syngas. If the swirl vanes are removed, a new swirl component is preferably added to the nozzles  30 ,  32  to swirl the fuel-air mixture and to promote mixing of the fuel and air to enhance combustion and flame stabilization. The modified air fuel nozzles may be capable of operating on natural gas, hydrogen, syngas or a combination of these gases. The fuel-air mixture discharging from the porous structure with micro swirlers results in formation of multiple micro flames producing lower NOx, CO and higher flame stability. 
     A high porosity flame stabilizer  56  may be positioned at the outlet of the air tube  48 . The flame stabilizer helps in increasing fuel-air velocities through the air tube and into the combustion zone  46 . In addition, the flame stabilizer may impart a swirl to the fuel-air mixture. Microswirlers, e.g., small vanes or cork-screw shaped flow passages, may be embedded in the stabilizer. The flame stabilizer arrests flame and prevents the propagation of flame upstream of the stabilizer into the air tube. The flame stabilizer also behaves like a passive control device for mitigating high frequency thermo acoustic oscillations. 
     The flame speed of hydrogen and syngas may be significantly faster, e.g., six to seven times as fast, as the flame speed of natural gas, e.g., methane. The flame speed may exceed the flow velocity of the air fuel mixture passing through the air tube. But for case with no flame stabilizer, the syngas or hydrogen flame may propagate upstream into the air tube and fuel discharge nozzle. To avoid such propagation of the flame, the flame stabilizer increases the fuel-air mixture velocities and arrests the propagation of the flame at the downstream face of the flame stabilizer. 
     The high porosity of the flame stabilizer  56  allows the air and fuel mixture to flow through the porous media of the stabilizer at a sufficient rate to provide effective combustion and generate sufficient volumes of hot combustion gases in the combustion zone  46  to drive the turbine  16 . Sufficiently high pressure drop across the flame stabilizer (represented by the right pointing arrow  58 ) is sufficient to prevent a fast moving flame (represented by the left pointing arrow  60 ) from entering and/or passing through the porous media of the stabilizer. An optimum pressure drop is chosen depending on the flame speed of the gaseous fuel and the flow rate of the air fuel mixture through the air tube. The porosity and thickness of the flame stabilizer is selected to achieve the desired pressure drop. Assuming that the pressure drop is properly selected, the upstream extend of combustion should be adjacent to the downstream face of the porous media of the flame stabilizer  56 . Accordingly, the porous flame stabilizer preferably anchors the flame slightly off the downstream face of the porous media of the flame stabilizer  56 . 
     The downstream face of the flame stabilizer may be coated with the thermal barrier coating (TBC)  62 , e.g., a high temperature ceramic. The TBC shields the stabilizer from the heat, e.g., radiant and conductive, of the combustion flame. The TBC is preferably applied to the surfaces of the stabilizer exposed to the flame. 
       FIG. 3  is a perspective view of an exemplary flame stabilizer  56 . A honeycomb structure  64  is one example of a porous flame stabilizer. An array of multiple passages is illustrated by dotted lines showing a single passage  66 . The flow passages  66  are formed by the honeycomb structure and may be constricted at the outlet ends  68 . The constrictions may, for example, be formed by coating the ends  68  so as to form bulbous or anvil shaped side walls between the passages. The coating applied to constrict the outlet of the passages may be a thermal barrier coating (TBC). The build-up of the TBC may form the flow constrictions in the passages. The constriction of outlet ends of the passages  66  may be used to determine the desired pressure drop across the stabilizer  56 . Further the blunt ends of the sidewalls may form eddy flows that enhance air fuel mixing and contribute to flame stabilization at the downstream face of the flame stabilizer. 
     Further, the passages  66  may spiral or cork-screw through the stabilizer. The spiral or cork-screw passages impart swirl to the fuel-air mixture that can supplement swirl vanes upstream of the stabilizer or replace the swirl vanes. In addition to a honeycomb structure, the flame stabilizer may be formed of structures such as: a matrix of interconnected fibers, a mesh and a sponge. These are exemplary structures. Further, the flame stabilizers may be a disc that fits onto the end of each air tube, a plug that fits into the end of each air tube or some other structure through which flows the fuel-air mixture. It is preferred that the flame stabilizers be added to the combustor with minimal modification needed to the combustor. 
     The flame stabilizer  56  may provide a relatively low cost and easy to install device for converting a natural gas combustor in a gas turbine to a combustor capable of burning hydrogen or syngas. To convert a natural gas burning gas turbine to hydrogen or syngas, a flame stabilizer may be positioned in the discharge end or adjacent the discharge end of a flame tube in each fuel nozzle  30 ,  32  of each combustor of the gas turbine. Optionally, the swirl vanes  52  may be removed and replaced by the flame stabilizer. Further, the fuel manifold and fuel supply lines may be modified to accept hydrogen or syngas. 
     The flame stabilizer  56  promotes stable combustion in the combustion zone  46 , even for fuels having fast flame speeds. A potential benefit of enhanced stable combustion is an decrease in the fuel-air ratio to achieve stable combustion. The fuel-air ratio is the proportions of gaseous fuel and air that are mixed in the Increasing the range of fuel-air ratios fuel nozzles  30 ,  32 . Increasing the range of fuel-air ratios that provide stable combustion may allow for fuel-air ratios that result in low nitric-oxide emissions, increased fuel economy, lower combustion temperatures and acceptable thermo acoustic pulsations. 
     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.