JP2009109180A - Can annular type dual fuel combustor of multi-annular multistage nozzle flowing in radial direction of lean premix - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To enable stable combustion, since there is a tendency of extremely easily receiving influence of combustion driving vibration (dynamic instability) more than conventional diffusion type combustion to turning stabilizing lean premix combustion of a combustor. <P>SOLUTION: A multi-annular multistage nozzle 120 flowing in the radial direction of a lean premix for generating three independent combustion zones in a can annular type dual fuel gas turbine combustor 100, has a pilot zone for supplying fuel by a gas pilot nozzle 150 and a central cartridge, a flame holder zone for supplying the fuel by inside main gas fuel, a main flame zone for supplying the fuel by outside main gas fuel, a main radial directional swirler 140 for mixing a part of inflow air to the nozzle 120 with the inside main gas fuel and the outside main gas fuel, an end cover assembly, and an outside means for controlling the ratio of the supplied inside main gas fuel and the supplied outside main gas fuel. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates generally to gas turbine combustors, and more specifically to lean premixed radial inflow multi-annular multistage nozzles for can annular dual fuel combustors that dramatically reduce or eliminate combustion dynamics. .

  FIG. 1 shows a high power output including a compressor 12 (partially shown), a plurality of combustors 14 (one shown for convenience and clarity), and a turbine 16 (represented by a single blade). 1 shows a prior art combustor of an industrial gas turbine 10. Although not specifically shown, the turbine 16 is drivably connected to the compressor 12 along a common axis. The compressor 12 pressurizes the intake air and backflows it into the combustor 14, which is used to cool the combustor 14 and provide air for the combustion process. Although only one combustor 14 is shown, the gas turbine 10 includes a plurality of combustors 14 disposed about its outer periphery. The transition duct 18 connects the outlet end of each combustor 14 with the inlet end of the turbine 16 and delivers hot combustion products to the turbine 16.

  Each combustor 14 includes a generally cylindrical combustion casing 24 that is secured to a turbine casing 26 using bolts 28 at the open front end. The rear end of the combustion casing 24 is an end that may include conventional supply pipes, manifolds and associated valves, etc., for supplying gas, liquid fuel and air (and water as needed) to the combustor 14. Closed by the cover assembly 30. The end cover assembly 30 includes a plurality (eg, five) of fuel nozzle assemblies 32 (only one is shown for convenience and clarity) arranged in a circular arrangement around the longitudinal axis of the combustor 14. receive. Each fuel nozzle assembly 32 is substantially cylindrical and includes a rear supply section 52 having an inlet for receiving gas fuel, liquid fuel and air (and water as required) and a forward delivery section 54.

  Mounted in the combustion casing 24 is a substantially cylindrical flow sleeve 34 that is substantially concentric and connected to the outer wall 36 of the transition duct 18 at its forward end. The flow sleeve 34 is connected at its rear end to the combustion casing 24 at the position of the butt joint 37 using a radial flange 35 where the front and rear portions of the combustor casing 24 are joined.

  Within the flow sleeve 34 is a concentrically arranged combustion liner 38 which is connected to the inner wall 40 of the transition duct 18 at its forward end. The rear end of the combustion liner 38 is supported by a combustion liner cap assembly 42 that is supported within the combustion casing 24 by a plurality of struts 39. The outer wall 36 of the transition duct 18 as well as that portion of the flow sleeve 34 extending forward (by bolt 28) where the combustion casing 24 is bolted to the turbine casing 26 form an array of apertures 44 over their respective peripheral surfaces. Air flows back from the compressor 12 through the aperture 44 into the annular space between the flow sleeve 34 and the liner 38 and toward the upstream or rear end of the combustor 14 (shown in FIG. 1). It will be understood that it will be able to flow (as indicated by the flowing arrows).

  The combustion liner cap assembly 42 supports a plurality of premix tubes 46, one for each fuel nozzle assembly 32. More specifically, each premixing tube 46 has its front and rear ends respectively by a front plate 47 and a rear plate 49, each having an opening that is aligned with the premixing tube 46 open at the end. In the combustion liner cap assembly 42. The premix tubes 46 are supported so that the forward delivery section 54 of each fuel nozzle assembly 32 is concentrically arranged.

  The rear plate 49 is mounted with a plurality of rearwardly extending floating collars 48 (one for each premixing tube 46) arranged substantially in alignment with the openings in the rear plate 49. Each floating collar 48 supports an annular air swirler 50 in surrounding relation to a respective fuel nozzle assembly 32. A radial fuel injector 66 is provided downstream of the swirler 50 for discharging gaseous fuel into a premix zone 69 located within the premix tube 46. This arrangement allows the air flowing in the annular space between the liner 38 and the flow sleeve 34 to enter the combustor 14 (end cap) before entering the combustion zone or combustion chamber 70 in the liner 38 downstream of the premix tube 46. It is reversed in the direction of the rear end of the assembly 30 and the sleeve cap assembly 42 to allow the swirler 50 and the premixing tube 46 to flow through. Ignition is accomplished in the multi-combustor 14 using a spark plug 20 with a flame propagation tube 22 (one shown) in a conventional manner.

  In power plant design, reducing the emission of harmful gases such as nitrogen oxides (NOx) into the atmosphere is the biggest concern. To address this problem, low NOx combustors have been developed that utilize lean premixed combustion with multiple burners attached to a single combustion chamber as described in FIG. Each burner includes a flow tube with a centrally located fuel nozzle that includes a cylindrical hub that supports a fuel injector and air swirler and has a flat surface on its downstream end. In addition to the premixed injection stage for low NOx operation, each fuel nozzle may include a diffusion injection stage for start-up and emergency operation and a liquid fuel injection stage for liquid fuel operation.

  Diffusion gas fuel and liquid fuel are typically injected through an orifice located on the flat end face of the fuel nozzle. During low NOx (premix) operation, fuel is injected through the fuel injector and mixes with the swirling air in the flow tube. The diffusion and liquid fuel circuits are typically purged with air during the premix operation to keep the flame gas out of the passage. The combustion flame is stabilized by bluff body recirculation and swirling stop (if there is swirling) behind the fuel nozzle. In premixing systems, strong pressure oscillations are usually generated as a result of combustion instability. Combustion instability is believed to be associated with excitation of span vortices from the bluff end of the combustion nozzle. These pressure oscillations can severely limit device operation and in some cases even cause physical damage to the combustor hardware. Furthermore, the flow of purge air through the diffusion and liquid fuel circuit is injected directly into the recirculation zone. This direct injection reduces the local temperature and strength of recirculation and adversely affects flame stability. Accordingly, there is a need for a low NOx combustor that reduces pressure oscillations and avoids the negative effects associated with direct injection of purge air into the recirculation zone.

  As mentioned above, these modern high power industrial dry low NOx (DLN) can annular gas turbine combustors typically fit a can combustor liner using a flat or angled cap / dome assembly. (Or group) premix nozzles are used. Multiple nozzles require fuel mixing and staging to achieve the intended operability and turndown and performance throughout the design space. However, this approach results in a complex and expensive assembly.

  Similarly, it is difficult to evenly distribute air and fuel to a group of premixed fuel nozzles at the head end, and in general, a sub-ideal non-uniform air flow for all nozzles, Or result in a significant amount of parasitic pressure drop / loss. Swirl-stabilized lean premixed (LP) combustion tends to be very susceptible to combustion-driven vibration (dynamic instability) compared to conventional diffusion combustion.

  Traditionally, in the gas turbine engine industry, flame temperatures (or primary zone temperatures) have been reduced in LP systems to reduce NOx emissions. Since acceptable NOx exhaust emission levels have been reduced to single-digit parts-per-million (ppm) levels (mainly determined by new government regulations), flame temperatures are at least as high as methane content. For fuel, it is very close to the lean blowout (LBO) limit. In such lean mixtures, slight periodic fluctuations in the local fuel-to-air mixture ratio result in local heat dissipation and relatively large periodic fluctuations in the heat dissipation rate, including local periodic flame extinguishing. The discrete oscillation frequency (or tone) can increase in amplitude when the heat release fluctuations are in phase with the sound pressure fluctuations occurring inside the combustion chamber.

  Current LP combustors are becoming leaner and more spatially uniform to further meet low NOx emissions, and there is an increasing need to meet this emission target while operating over a wide range of fuels. In a given system, there is an increased risk of producing unacceptably high levels of combustion dynamics.

  To date, large single nozzle DLN low NOx can annular gas turbine combustion systems have been attempted, but most have not been successful due to operability, durability, and emissions issues. The lack of smart adjustable operating parameters and the lack of multiple independent combustion staging zones led to the introduction of a modular multi-nozzle (group) configuration into the industry. The multi-nozzle design not only facilitates ignition and turndown, but also provides an adjustable operability parameter that avoids the dynamics (or vibration) that occurs while operating in the design working space. Allows staging or grading of fuel distribution to different subgroups.

  The downside of sloping fuel distribution in the combustor is the creation of hot temperature zones that promote NOx production. Thus, if excessive tilting is required to constrain dynamics or instability, regulatory NOx emissions limits may be breached, and in some cases, the unit may become inoperable. The LP combustion dynamics of industrial gas turbines are typically passively reduced in several ways (usually trial and error) that can be expensive and uncertain. This method includes 1) shifting the injection point to change the fuel transfer time from the fuel injection point to the flame front, 2) changing the size of the fuel injection orifice, pressure drop and acoustics across the hole Changing the impedance, and 3) modifying the chamber or nozzle profile (eg, diameter, angle, length) to affect vortex excitation, frequency and amplitude, or flame shape in the chamber .

  These methods attempt to cause any perturbation of heat dissipation to be out of phase (ie, cancel out of phase) with pressure or acoustic perturbations in the combustion chamber. Combustor dynamics have also been reduced or eliminated by adding acoustic attenuation (eg, Helmholtz resonators or quarter wave tubes) to the combustion system. So far, the above methods tend not to be actively designed during the initial design phase of the program, but to be considered and implemented afterwards for the discovery of high combustor dynamics. there were.

Therefore, the possibility to excite or drive discrete combustion oscillations at any load in the design / working space, while having the above average tolerance for the quality of the fuel mixture, is basically in the statistical and absolute sense. There is a need to provide an inexpensive LP combustor that is much less, simple and expandable. If the above solution is found and, as a result, the risk of continually generating discrete dynamics in a given design operating space is greatly reduced, the efficiency and potential of tuning for minimum emissions in a given system Will be extremely increased. In essence, dynamics is no longer such an important and difficult element of the overall combustor design procedure.
US Pat. No. 7,185,494 US Pat. No. 5,203,796 US Pat. No. 6250063 US Pat. No. 5,408,830

  The present invention generates three independent combustion zones in a gas turbine combustor having a lean premixed radial inflow multi-annular multi-stage nozzle, thereby enabling stable combustion with low nitrogen oxide (NOx) emissions. And a method.

  In summary, according to one aspect of the present invention, a lean premixed radial inflow multi-annular multi-stage nozzle is provided that creates three independent combustion zones within a can-annular dual fuel gas turbine combustor. A lean premixed radial inflow multi-annular multistage nozzle (hereinafter referred to as a single large radial nozzle) includes a pilot zone fueled by a central cartridge, a flame holder zone fueled by internal main gas fuel, A main flame zone fueled by external main gas fuel, a main radial swirler that mixes part of the air flowing into the nozzle with the internal main gas fuel supply device and the external main gas fuel supply device, an end cover, and a supply Means for controlling the ratio of the supplied pilot gas fuel, the supplied internal main gas fuel and the supplied external main gas fuel.

  According to another aspect of the present invention, a can-annular dual fuel combustor for a gas turbine engine is provided. The combustor is a lean premixed radial inflow multi-annular multi-stage nozzle (hereinafter referred to as a single large radial nozzle) incorporating an external burner tube mounted on the combustor casing on the end cover and a main radial swirler. Including. The main combustion zone is provided downstream from the outer burner tube of a single large radial nozzle. A source of pressurized air is provided from the compressor source. The air inlet plenum radially surrounds a single large radial nozzle and is radially bounded by the outer wall of the combustor. The pressurized air diffuser receives pressurized air in the reverse flow path from the compressor and discharges the pressurized air to the inlet plenum at the recovery pressure. A fairing mounted on the main radial swirler and surrounding a portion of the outer burner tube is provided to smooth the air flow from the diffuser to the air inlet plenum.

  According to a third aspect of the present invention, there is provided a method utilizing a lean premixed radial inflow multi-annular multistage nozzle (hereinafter referred to as a single large radial nozzle) having an independent combustion zone. The radial nozzle includes a pilot zone, a flame holder zone, and a main zone in a gas turbine combustor that provides stable combustion with low nitrogen oxide (NOx) emissions. The method includes providing a large supply of air to the nozzle, an in-nozzle staging step, dividing the heat release into a number of discrete zones in space, distributing the heat release in time, and downstream center. Venting the recirculation zone.

  These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like numerals represent like elements throughout the drawings.

  The following embodiments of the present invention have several innovative and unique features: (1) multiple (eg, 6) premix nozzles (per can) and a single large combustor chamber cap. Replaced with radial nozzle and liner modifications, thereby enabling significant component savings, cost savings, and dramatic simplification of the combustor head end, (2) Dome diffusion design for the rear To reduce convective cooling of the dome of the liner while reducing static pressure loss by restoring static pressure before premixing fuel and air in a large radial nozzle and using more air for premixing Enabling (3) large amounts of fuel (~ 2 lbm / sec) and air (~ 60 lbm / sec) quickly (eg, with a relatively low pressure drop (eg <4 percent)) , <3 ms) and providing the ability to fully evaporate and mix, and (4) using either gas fuel or liquid fuel, still provide the required system turndown performance and low emissions. However, by dynamically distributing or smearing out both heat dissipation (in time and space) and fuel transport time in the chamber, it is more robust and combustible than current lean premixed (LP) gas turbine combustion systems Has many advantages, including reducing the tendency to cause vibrations.

  The effects of the design are multifaceted: (1) Replacing 5 or more nozzles per can to dramatically reduce costs and parts count, (2) Industrial gas turbines while maintaining the required emission levels Dramatic reduction or complete elimination of combustion dynamics / vibrations at discrete frequencies in the combustion chamber, (3) flexibility of gas and liquid fuels that lead to successful combustion dynamics improvements, (4) water injection and high pressure atomizer air While eliminating the need for DLN with 2 diesel oil-like liquid fuel, and (5) 1 digit (ppmv) harmful emissions.

  Successful transition from a multi-nozzle arrangement to a large single nozzle requires in-nozzle staging. The slanted V-groove flame holder zone used in this design provides an area for fuel staging within the main premix nozzle. For example, if the fuel-air ratio is biased richer near the hub (or center body) in the premixer, the center flameholder zone may be used for handling ignition, mechanical acceleration, low load operation, or sudden load transmission. It is possible to burn at a higher equivalence ratio to the bypass flow that can be advantageous (or necessary). Biasing the fuel air ratio in conjunction with other staging features (such as premixed pilots) allows a single large radial nozzle to replace multiple nozzles per can (eg 6) in a gas turbine combustion system As a result, the combustion system and the engine as a whole are considerably reduced in the number of parts and saved in cost. Reduction of combustion dynamics is achieved while maintaining or reducing harmful emissions (eg, unburned hydrocarbons (UHC), NOx, and carbon monoxide (CO)) relative to the combustion temperatures occurring in the design space. It will be.

  A lean premixed radial inflow multi-annular multi-stage nozzle (hereinafter referred to as a single large radial nozzle), by design, is a well-mixed reactant set fraction (eg, about 33 percent of the nozzle reactant). ) Is reoriented to burn as a tilted array of discrete V-groove zones upstream of the main chamber, it is unlikely to excite the discrete vibration frequency of the combustion drive. The arrangement of axial jets passing through the conical V-groove pack structure reduces discrete dynamics and improves emissions in several ways.

  First, this arrangement divides heat dissipation into a number of discrete reaction zones in space, each reacting on a much smaller spatial scale than the entire combustion chamber. This effectively limits the amount of energy emission that can be emphasizedly coupled at a particular acoustic resonance frequency in the chamber.

  Secondly, the inclined V-groove produces a large number of fuel transfers and distributes (or smears out) the heat release in time. That is, each point along the V-groove length range has a self-relevant transport time, i.e., the time between the fuel injection point and the combustion point. This also effectively limits the amount of heat dissipated energy that can be stressfully coupled at the acoustic resonance frequency of a particular chamber.

  Third, the function of the dewar swirl pack is the “venting” of the downstream central recirculation zone (CRZ) caused by vortex breakdown. The expanded array of jets from the central cone injects non-swirl axial moments directly into the CRZ, reducing the CRZ size and bulk residence time. As a result, this reduces the production of nitrogen oxides (NOx) by reducing the average time that combustion product molecules spend in the primary region (flame) temperature inside the combustor. The concept of “temperature time” for NOx generation becomes increasingly pronounced at a flame temperature of about 2900 F, where the thermal NOx (or Zeldovich) mechanism begins to promote and the impact on the overall system NOx level begins to increase significantly. .

  The nozzle further provides a non-coke design for operation with diesel liquid fuel that does not require water and nebulizer air. The design aspect prevents fuel gallery coking with insulating fuel gallery walls for high reliability. The liquid fuel is rapidly atomized and completely dispersed within the premixed air stream, keeping the fuel away from the hot premixed surface and allowing it to evaporate and immediately mix with air. The liquid injection method does not adversely affect the gas operation. Eliminating the need for water jets and high pressure nebulizer air further provides cost savings.

  Complete fuel-air mixing is fast (about 2 milliseconds) and complete (greater than 97 percent) and requires a low premixer differential pressure (~ 2 percent), which makes shorter and smaller designs And reduces the premixer residence time required to stay below the auto ignition time of diesel in “advanced” gas turbine conditions.

  Several other aspects and advantages of the present invention will become apparent herein. Turndown capability is enhanced through fuel staging (three quasi-independent combustion zones). By using a rear cooling dome, the need for liner cooling air in the flame zone is eliminated.

  Similarly, axisymmetric radial combustion staging does not expose the combustor liner to asymmetric loading, thereby providing improved combustion liner durability.

  Further, as the internal premixer flame holding resistance / margin is improved, flow is accelerated throughout the premixer nozzle and the bulk velocity is maintained above about 300 ft / sec.

  The V-groove lean angle (radial versus axial plane) and the dewarer vane profile were selected as two parameters for optimization. The V-groove lean angle varied between 30 and 60 degrees to maximize the lean angle while still producing a well-defined continuous V-groove wake to support the independent combustion zone. . For unreacted CFD, the 40 degree configuration was the maximum angle that still produced a continuous V-groove wake with other nozzle features held constant. Dewarra vane profile is tuned / optimized by aligning the inlet vane angle with the incoming swirl flow and accelerating the pack flow using a cascade shape, thereby preventing any flow separation in any pack succeeded in.

  FIG. 2 shows an embodiment of the large single radial nozzle of the present invention employed in the combustor of the present invention. The large single radial nozzle combustor 100 includes a generally cylindrical combustor casing 105 that connects to a turbine by plugging a combustion liner into the transition piece so that a transition piece (not shown) is open at the open front end. Can be fixed. The transition piece can then be secured to the turbine casing at the open front end with bolts in the usual manner. The rear end of the combustor casing has an end cover assembly 130 that is compatible with conventional packages for supply tubes, manifolds, valves and fittings for gas fuel, liquid fuel, air, and power (not shown). Closed by. End cover assembly 130 is part of a large radial nozzle assembly 120 that secures it to combustor casing 105.

  In the combustion casing 105, a flow sleeve 106 is mounted in a substantially concentric relationship. Within the flow sleeve 106 is a concentrically arranged combustion liner 110 that is connected to and plugged into the inner wall of a transition liner (not shown) at a forward end 112. The rear end of the combustion liner 110 forms a frustoconical dome 111 on the main combustion chamber 114 that is central to the flow of fuel and combustion products from the large radial nozzle 120. Open and combined with the outer burner tube 113 of the large radial nozzle 120.

  Combustion process air can be drawn from the air compressor to the transition piece (as described above with respect to FIG. 1) and from there through an annular space 115 between the flow sleeve and the outer wall of the combustion liner. At the rear end of the annular space, a concentric mounting diffuser 116 expands air into the inlet plenum 117 of the large radial nozzle 120. The dome 111 serves as an inner concentric wall of the diffuser 116, thereby allowing the dome 111 to be cooled rearward by the air flowing through the diffuser 116. At the same time, the diffuser 116 restores static air pressure before premixing fuel and air at the large radial nozzle 120, resulting in less parasitic pressure loss and more air available for premixing. A fairing 118 around the center of the large radial nozzle 120 facilitates air inflow into the inlet plenum 117 which further reduces parasitic pressure loss.

  The large radial nozzle 120 further includes a main radial swirler 140, a gas pilot nozzle 150, a central flame holder having a V-groove flame holder 160, and an external flame holder 170. The central flame holder 160 and the external flame holder 170 open to the main combustion chamber 114 on their forward ends.

  The end cover 130 may be a generally cylindrical flange designed to support the radial nozzle assembly 120 in the combustor 100 in combination with the combustor casing 105. The rear surface 135 of the end cover 130 allows dual fuel (gas and liquid fuel) and gas pilot nozzle 150 to enter. One of the external main gas supply device 190, the internal main gas supply device 190, and the plurality of liquid gas connections 195 is shown in FIG. 3A. The fuel and air ingress arrangement allows connection of fuel, air and power lines (not shown) to existing combustor configurations.

  FIG. 3A shows an isometric cutaway view showing the internal structure of an embodiment of a large single radial nozzle of the present invention. FIG. 3B shows an axial cross-sectional view showing the internal structure of an embodiment of a large single radial nozzle of the present invention. The nozzle is axisymmetric along the central axis 200.

  End cover assembly 130 includes an end cover plate 205 having a rear section 201, a front section 202 and a central cavity 203. The main radial swirler 140 includes a rear hub 240, a plurality of swirl vanes 250, and a central hub 260 having a central cavity 265 therein. The rear plate 240 is bolted to the end cover plate 205 at the mounting surface 241.

  The central flame holder 160 is mounted on the central hub 260. The central hub 285 of the central flame holder 160 is combined with the central hub 260 of the main radial swirler 140 to support the central flame holder 160 in the radial and axial directions. A radial vane 360 supports the inner burner tube 300 from the central hub 285. A plurality of V-grooves 290 extend between the inner burner tube 300 and the central hub 285. A central cavity 278 is formed in the central hub 285. On the swirl vane 250 of the main radial swirler 140, an external flame holder 170 having a cylindrical outer burner tube 175 is mounted by a base section 180 that extends radially outward and attaches to the top of the swirl vane 250 with bolts 183. The The downstream end 178 of the outer burner tube 175 also extends outwardly and is reinforced to provide support for the combustor conical dome 111 (FIG. 2). Support ledge 179 is combined with combustor conical dome 111. The fuel air mixture from the main radial swirler 140 passes to the flame holder zone through 402 and the main zone through 405 (FIG. 7). Central cavities 203, 265, and 278 allow for insertion of a gas pilot including a liquid pilot and ignition device and a gas pilot nozzle 150 that includes a central cartridge.

  FIG. 4 shows a rear end view of the end cover plate of the large single radial nozzle embodiment of the present invention with respect to an attachable combustor. The end cover plate 205 includes an integrated cylindrical rear section 201 and a small diameter cylindrical front section 202 (FIG. 3A), both layers centered on the central axis 200 of the nozzle. The rear section 201 is sized radially to combine with the rear seating surface of the combustor (not shown) and is axial with respect to the outer circumferential surface 207 of the outer section 201 for attachment to the combustor seating surface. It incorporates a plurality of bolt holes 206 formed close to the direction. The rear section 201 may also include a plurality of pilot holes 208 (also led) for positioning the rear section 201 relative to the combustor seating surface in preparation for bolting. The front seating surface 205 of the front section 202 is also configured to receive a bolt from the mounting surface 241 of the rear plate 240 of the main radial swirler 250 (FIG. 3A) in a circular configuration concentric with the central axis 200 of the nozzle. A plurality of bolt holes 209 may be included.

  Two independent gas fuel supply devices can be connected to the end cover plate 205. The rear section 201 includes an external main gas entry 215 attached to the external main gas supply inlet pipe 216 by an external main gas inlet flange 217 for connection to an external main gas fuel supply device (not shown). The rear section 201 also includes an internal main gas entry 220 with a fitting 219 for connection to an internal main gas supply (not shown). The end cover plate 205 may also include a plurality of liquid fuel supply entry sections 243 that are concentrically located with respect to the central axis 200 of the nozzle.

  FIG. 4 also shows the rear end of the gas pilot nozzle 150. A central cavity 203 is defined in an end cover plate 205 that extends radially from the central axis 200 and extends through a rear section 201 and a front section 202 for insertion of a gas pilot nozzle. The central cavity 203 includes a gas pilot nozzle seating surface 210 (FIG. 3B) with a threaded connection for combination with a gas pilot nozzle rear flange 212 that mounts the gas pilot nozzle 150.

  FIG. 5 shows a fuel gallery of the end cover and back plate of the nozzle embodiment of the present invention. External main gas entry 215 (FIG. 4) is connected to external main gas gallery 310 in end cover plate 205. The outer main gas gallery 310 defines an annular chamber concentric with the central axis 200 of the nozzle. The inner wall 311 and outer wall 312 of the outer main gas gallery 310 are such that the open upper end 315 of the outer main gas gallery 310 communicates with a plurality of corresponding outer main gas channels 665 (FIG. 3B) in the main swirler rear plate 240. It can be located radially with respect to the central axis 200 of the nozzle.

  The internal main gas entry part 220 is connected to the internal main gas gallery 330 in the end cover plate 205. The inner main gas gallery 330 defines an annular chamber concentric with the central axis 200 of the nozzle. The inner wall 317 and the outer wall 318 of the inner main gas gallery 330 may be concentric with the central axis 200 of the nozzle. The inner main gas gallery 330 is disposed radially between the outer main gas channel 310 and the central cavity 203. The inner main gas gallery inner wall 317 and outer wall 318 have internal open end 319 of the internal main gas gallery 330 in communication with the internal main gas channel 680 (FIG. 7) in the main swirler rear plate 240 to swirl the internal main gas. Arranged radially to supply internal main gas injection points 695 on the base surface 242 between 250.

  The liquid fuel supply entry 243 extends axially through the rear section 201 of the end cover 205 and communicates with the main liquid fuel gallery 244. The main liquid fuel gallery 244 defines an annular chamber concentric with the central axis 200 of the nozzle and sealed except for the liquid fuel supply entry 243 and the liquid fuel delivery entry 246. The main liquid fuel gallery 244 is arranged in a radial direction so as to align with the liquid fuel supply entry rear portion 243 and the liquid fuel delivery entry portion 246 front on the end cover plate 205. The liquid fuel delivery entry 246 extends through the front section 202 of the end cover 205 and corresponds to a corresponding liquid in the main radial swirler rear plate 240 that leads to a liquid fuel sprayer 248 in the main swirler rear plate 240. Combined with the fuel delivery entry portion 247. The wall of the main liquid fuel gallery 244, the liquid fuel supply entry portion and the liquid fuel delivery entry portion 246 of the end cover 205, and the fuel delivery entry portion 247 of the rear plate 240 are 290 degrees Fahrenheit where coking of diesel liquid fuel begins. An insulating lining 249 can be provided to maintain a wall temperature below. The attachment 218 is provided outside the liquid fuel supply entry portion in order to connect to the liquid fuel supply device.

  Because the end cover plate 205 and the main swirler rear plate 240 are combined within the intermetallic seating surface 241, it is possible to isolate potential leakage from the fuel cavity along the seating surfaces 204,241. Three annular recesses (FIG. 5) concentric with the nozzle central axis 200 can be provided on the upper seating surface 204 of the end cover plate 205. The first recess 381 is provided outside the external main gas gallery 310. The second recess 382 is provided between the outer main gas gallery 310 and the inner main gas gallery 330. The third recess 383 is provided in the aircraft of the internal main gas gallery 330. The recess can comprise a C-ring or other suitable gasket material to prevent flow along the seating surface.

  6A-6E show views of the main radial swirler of an embodiment of the nozzle of the present invention. FIG. 6A shows an isometric view of the main radial swirler of an embodiment of the nozzle of the present invention. FIG. 6B shows details of the main swirl vane on the main radial swirler of the nozzle embodiment of the present invention. FIG. 6C shows a front view of the main radial swirler of the nozzle embodiment of the present invention. FIG. 6D shows a cross-sectional view through the main swirl vane and central hub of the main radial swirler of the nozzle embodiment of the present invention. FIG. 6E shows a cross-sectional view through the rear plate and central hub of the main radial swirler of the nozzle embodiment of the present invention.

  The main radial swirler 140 includes an integrated central hub 260, a plurality of main swirl vanes 250 mounted on the rear plate 240 and projecting perpendicularly to the rear plate 240 (downstream toward the combustion zone), a gas pilot nozzle 150 and a central cavity 265 that houses a series of internal passages in the rear plate, and a main swirl vane 250 that provides fuel and air flow.

  The rear plate 240 includes a cylindrical flange centered on the central axis 200 of the nozzle. The base surface 241 of the rear plate 240 is sized to be combined radially with the front surface of the end cover plate 205. The mounting surface 242 of the rear plate incorporates a plurality of recesses 371 that receive bolt holes 372 around the rear plate 240. Bolt hole 372 extends through base surface 241 of rear plate 240 and is aligned with bolt hole 209 on front surface 204 of end cover plate 205. The mounting surface 242 of the rear plate 240 allows the mounting of a plurality of main swirl vanes 250 and the storage of fuel injection points in the air flow stream within the main radial swirler 140.

  A plurality of main swirl vanes 250, each including a solid metal airfoil 610, are mounted at right angles to the rear plate 240 and can project axially toward the combustion zone. The main swivel vane 250 can be mounted in the radial machine from the peripheral bolt hole recess 371 and out of the radial machine from the central hub 260. The leading edge 615 of each airfoil protrudes generally radially outward and the trailing edge 620 protrudes generally radially inward. Each airfoil axis 625 may form a pre-designated acute angle α (about 15 degrees) with a radius 635 from the central axis 200 of the nozzle. If the leading edge 615 of the airfoil 610 forms a curved surface, the side surfaces 640, 641 of the airfoil 610 can form a linear taper with respect to the common linear trailing edge 620. The bottom surface 645 and the top surface 650 of the airfoil 610 form a plane. The bottom surface 645 can be attached to the mounting surface 242 of the rear plate 240 by welding or other suitable process.

  A plurality of external main gas fuel injection points 655 are provided along a radius concentric with the central axis 200 of the nozzle on one side 640 of the airfoil 610 just inside the aircraft at the curved leading edge 615. The injection of the external main gas fuel occurs substantially perpendicular to the air flow 660 passing between adjacent swirl vanes. However, the injection points 655 can also be provided on both sides of the airfoil and elsewhere, other than included in embodiments of the present invention. The injection points are spaced approximately evenly in the axial direction along the side surface 640 of the airfoil 610, and external main gas fuel is supplied to the air flow 660 between the airfoils 610 in the circumferential premix space 605. It can be distributed evenly. The airfoil 610 further includes an internal fuel cavity 665 that provides an injection hole 657. The fuel cavity 665 can be a substantially cylindrical hole that rises axially from the base surface 241 into the airfoil 610 and extends to and near the injection hole 657. The fuel cavity 665 guides external main gas fuel from the fuel cavity 310 in the end plate 205. An injection hole 657 in each airfoil 610 extends radially with respect to the cylindrical fuel cavity 665 to supply fuel to the injection point 655. The upper surface 650 of each airfoil 610 can further include a threaded hole 670 for securing the outer burner tube 175 to the main pivot vane 250.

  The internal main gas entry 680 (FIG. 7) in the rear plate 240 extends in a radial direction from the internal main gas gallery 330 in the end cover plate 205 toward the mounting surface 242 of the rear plate 240. An orifice 685 can be provided in each of the internal main gas entry sections 680 to control gas release. The injection point 690 is provided on a mounting surface 242 that is approximately equidistant from the side surfaces 640, 642 of the adjacent airfoil 610 and approximately midway along the side surfaces 640, 642 of the adjacent airfoil 610. In the exemplary 24 airfoils 610, 24 injection points 690 are provided. An injection port 695 at each injection point 690 that extends slightly above the mounting surface 242 of the rear plate 240 allows gas to be injected onto the laminar flow of the mounting surface.

  During gas operation as described above, the gas fuel has a primary radius from a number of injection points 655 that are axially located along the sidewall 640 of the airfoil 610 and from the injection points 695 on the mounting surface 242 of the rear plate 240. It is injected into the air flow of the directional swirler 140. The main gas fuel is supplied from two independent sources as shown in FIG. 4 so as to influence the radial profile of the fuel / air mixture in the annular swirl volume (premixer annulus) 255. That is, the air-fuel mixture near the central hub 260 that eventually passes through the central flame holder device is changed by changing the ratio of the fuel supply from the two sources, so that the air-fuel mixture near the swirl vane 250 It can be rich or sparse compared to (bypass). External means can be supplied to control this ratio between the supplied internal main gas fuel and the supplied external main gas fuel. This can include a throttle controller, pressure controller or other means known in the art that can be operated outside the nozzle.

  A plurality of liquid fuel injection points 245 are also provided on the mounting surface 242 of the rear plate 240 for operation with liquid fuel. The liquid fuel injection point 245 is positioned over the liquid fuel delivery channel 246 in the rear plate 240. The liquid fuel channel 246 in the rear plate 240 can include a thermal insulation layer 249. The liquid fuel injection point 245 is concentric with the central axis 200 and can be positioned to inject liquid fuel in the annular swirl volume 255 approximately at the center of the trailing edge 620 of the airfoil 610. In the exemplary arrangement, six liquid fuel injection points 245 are provided circumferentially equidistantly around the mounting surface 242. Each liquid fuel injection point 245 includes a tip 252 that includes a conical sprayer 248 that attaches to a screw 253 for a liquid fuel delivery channel 247. The sprayer 248 sprays liquid fuel in an axial air flow perpendicular to the mounting surface 242.

  FIG. 7 shows a cross-sectional view of the head end of the combustor of the present invention depicting the air flow and fuel air flow establishing the independent combustion zone of the nozzle embodiment of the present invention. As mentioned above, referring to FIGS. 5-7, the combustor of the present invention incorporating a large single radial nozzle provides three independent combustion zones. The pilot gas nozzle 150 generates a pilot combustion zone Z1. The flame holder combustion zone Z <b> 2 is generated by an axial flow from a dewar 280 that passes beyond the V-shaped groove 290 in the central flame holder 160. The main combustion zone Z <b> 3 is generated by a fuel-air mixture that flows into the main combustion chamber 114 from between the inner burner tube 300 of the central flame holder 160 and the outer burner tube 175 of the outer flame holder 170.

  The air flow from the diffuser 116 enters the inlet plenum 117. The main swirl vane 250 establishes a flow path 660 for incoming air from the inlet plenum 117 for the combustor. About 95 percent of the air entering the nozzle flows between the main swirl vanes 250. Inlet air having external main gas injected from the airfoil 610 and internal main gas injected from the injection point 690 on the mounting surface 242 and / or liquid fuel injected from the atomizer 248 is fed by the airfoil 610. Guided and swiveled counterclockwise (as viewed from the combustion end) through the annular swirl volume 255 (the volume between the swirl vane and the central hub). In the annular swirling volume 255, air and fuel are further mixed by continuous swirling.

  The central hub 260 includes an external truncated cylindrical conical surface centered on the central axis 200 of the nozzle and flows in a circumferential direction from the main swirl vane when standing in the central flame holder 160. Minimize the flow resistance to. The central hub 260 rises from the mounting surface 242 of the rear plate 240 and forms a smooth surface that is concavely inclined inward and forms the radial and axial support portions of the central flame holder 160. Specifically, in its truncated upstream range, central hub 260 provides an outer annular support ledge 273 for central flame holder 160. The inner surface 263 of the central hub 260 defines a cavity 165 that houses the gas pilot nozzle 150 and includes an internal flow path of air to the gas pilot nozzle 150. The central hub inner surface 263 further includes an internal annular mounting ledge 274 for the central flame holder 160.

  The series of internal passages in the rear plate are the internal main gas gallery in the end cover to the external main gas passage from the outer main gas gallery in the end cover to the swirl vane and the gas injector on the mounting surface of the rear plate. A passage for liquid fuel from the liquid fuel delivery entrance in the end cover to the sprayer on the mounting surface of the rear plate, and cooling and pilot pre-cooling from the circumferential outer edge of the rear plate to the radial nozzle center / core. Air passage to the central cavity for the mixed air.

  The internal main gas internal passages 680 to the internal main gas injector tip 695 on the mounting surface 242 of the rear plate 240 may include an orifice 685 in each passage to control the gas flow to the gas injector tip 695. it can. The outer circumferential surface 257 of the rear plate 240 includes a plurality of radial supply holes 275 directed inwardly with respect to the central cavity 265 to supply a flow of cooling air and pilot premixed air to the central cavity 265. The axial passages in the rear plate for the outer main gas 270, the inner main gas 680, and the liquid fuel 247 are located at circumferential positions between the various radial supply holes 275.

  The central flame holder 160 can include a central hub 285, a central cavity 278, a dewar 280, and a plurality of V-grooves 290, an internal burner tube 300 and a support tower 295.

  Since the air flow from between the main swirl vanes 250 is forced into the rotating flow in the annular swirl volume 255, only the exit path is downstream. Approximately 30 percent of the fuel air mixture swirled in the main radial swirler 240 enters the central flame holder 160. The central flame holder 160 includes a support tower 350 located above the central hub 260 of the main radial swirler 240. The support tower 350 provides axial and radial support for the central flame holder 160 in combination with the cylindrical support hub outer support ledge 273 and inner support ledge 274 of the center hub 260. The support arm 355 of the support tower 300 sits on the outer support ledge 273 and the inner support ledge 274. The support tower 295 and the central cavity 280 in the central hub 285 can receive the gas pilot nozzle 150.

  Referring to FIG. 8, on the support tower 295, there is a dewar 280 and a conical V-groove flame holder pack having a concentric hub. The deswirler 280 includes a plurality of annular sections 345 between the central hub 285 and the inner burner tube 295. The annular section 345 is open at the upstream inlet 347 and downstream outlet 348 of the fuel-air mixture. A radial vane 360 is provided between each individual annular section 345 and extends from the central hub 285 to the inner burner tube 295. Each radial vane 360 curves from a somewhat flat slope at the upstream inlet 347 to a steep slope at the downstream outlet 348 of the annular section 345. The flat axial slope at the upstream inlet corresponds to the acceptance of the swirling circumferential flow of fuel-air mixture from the main swirl vane 250 of the main radial swirler 140. About 30 percent of the fuel air mixture from the annular swirl volume 255 of the main radial swirler 140 flows into the annular section 345 of the dewar 280. The changing slope of the radial vane 360 redirects the circumferential flow into an axially oriented flow exiting each individual annular section 360. The reoriented axial flow vents the central recirculation zone (CRZ) as described above.

  With reference to FIGS. 3B and 8, the annular tip 380 of the central hub 285 forms a plane I. The annular tip of the inner burner tube 295 forms a plane II. Plane I is downstream of plane II. The radial vanes 360 of the deswirler 280 at these downstream ends form an inclined edge between the annular tip 380 of the central hub 285 and the tip 385 of the inner burner tube 300 that rises with an inclination of about 30 percent.

  A V-shaped groove 290 is provided at the downstream end of each radial vane 360. V-groove 290 includes a V-shaped element 375 having an open end 376 facing downstream. The apex 377 of the V-shaped element 376 is attached to and extends through the distal end 385 of the inner burner tube 300 along the downstream edge of the radial wall 360 and through the distal end 380 of the central hub.

  The external flame holder 170 includes a generally cylindrical external burner tube 175 that extends at the upstream end and forms an annular seat for combination with the main swirler. The cylindrical tube radially surrounds and extends toward the combustion chamber beyond the central flame holder 160. The downstream end 190 of the external burner tube 175 is strengthened. The ledge 195 provides a seating surface that engages the conical dome 111 (FIG. 2) of the combustor. The annular seating surface 180 of the outer burner tube 175 extends radially outward at its upstream end. The seating surface 180 forms a roof on the main swirler 250 of the main radial swirler 140, thereby restricting the fuel / air mixture exit path from the main radial swirler 140 to the downstream flow paths 402 and 405. The seating surface 180 can be attached to the top of each airfoil of a main swivel vane having multiple bolts, with one bolt in each screw hole on the airfoil. For the remaining 70 percent flow 405 of the fuel-air mixture from the main radial swirler 140 to the combustion space, the annular space is the inner burner tube 295 of the central flame holder 160 and the outer burner tube 175 of the outer flame holder 170. Formed between.

  FIG. 8 shows the central flame holder, gas pilot annulus, and central cartridge of an embodiment of the large single radial nozzle of the present invention. The remaining 5 percent air flow path of the incoming air from the inlet air plenum passes from the circumferential edge 257 of the rear plate 240 through a plurality of radial entries 275 (12 in the embodiment) to the central cavity 260 of the nozzle 140. To supply.

  As a means of achieving ignition, combustor turndown and improving stability, a central gas pilot nozzle 150 is positioned inside the conical flame holder volume at the upstream end of the smallest diameter. The gas pilot nozzle 150 provides a central cartridge 155 that can include an igniter / flame detector and a liquid gas pilot.

  Approximately 5 percent of the air flow to the radial nozzles entering through the radial flow holes 275 in the circumferential surface 257 of the rear plate 240 relative to the main radial swirler 140 is divided internally. About 80 percent of this air is directed to the annular axially swirling gas pilot premixer 855 through the air supply annulus between the inner wall of the central cavity 265 of the central hub and the outer surface 812 of the annular shell 810 of the gas pilot nozzle 150. And flow forward. The remaining portion of the air flows through a plurality of radial feed holes 875 in the annular shell 810 to the central cartridge used for liquid pilot spray and cooling and purging of the central cartridge tip.

  9A and 9B show the nozzle body of the gas pilot nozzle of the large single radial nozzle embodiment of the present invention.

  The gas pilot nozzle 150 includes a body 805 with an annular shell 810 that can be backloaded through the end cover plate 205 into the central cavity 203 of the nozzle 140. The annular shell 810 includes a rear flange 815 having a plurality of bolt holes 816 at its rear end for attaching its front surface 817 to a seating ledge 210 in the central cavity 203 of the end cover 205. The rear flange 815 also includes a central hole 818 for insertion of the central cartridge 155 and includes a raised surface 819 on the rear surface 820 around the central cavity for bolting the central cartridge 155 to the gas pilot rear flange 815. The screw hole 821 is incorporated. The rear flange 815 also includes an entry 230 that connects to a pilot gas fuel supply for gas pilot operation.

  The gas pilot nozzle body 805 extends through the central cavities 203, 265, 278 of the nozzle 120 to the cylindrical hub 370 of the central flame holder 160. The gas pilot annular shell 810 tapers gradually from the rear end to the front end. The annular shell 810 includes a lower shell 835, a tapered shell 840, a central shell 845 and a tapered head 850.

  An annular gas pilot airflow space 864 is also defined between the inner wall 368 of the cylindrical hub 370 and the inner wall 296 of the support tower 295 having the outer surfaces 842, 847 of the tapered shell 840 and the central shell 845. Air from the inner radial end 277 of the central radial feed hole 275 in the rear plate 240 enters the gas pilot air flow space 864 and flows to the axial swirling gas pilot premixer 855 in the forward axial direction.

  An entry 230 in the rear flange 815 for pilot gas fuel provides an internal pilot gas fuel cavity 862 in the annular shell 810. An internal pilot gas fuel cavity 842 in the lower shell 835 supplies pilot gas fuel to the annular pilot gas space 866 between the inner wall of the annular shell 810 and the outer surface 872 of the central cartridge 155. The tapered head 850 extends in close proximity to the cylindrical hub 370 to form a gas pilot annulus 825 between the outer surface 830 of the tapered annular head 850 and the inner surface 368 of the cylindrical hub 370. A plurality of pilot gas fuel holes 860 extend radially through the annular shell at an upstream inlet between adjacent axial mixing vanes 857 that provide pilot gas fuel injection points. The forward portion of the central shell 845 houses a plurality of axial mixing vanes 857 in the overall shape of the airfoil on the outer surface 847 to mix the gas pilot fuel and air moving downstream of the airflow space 864, This constitutes an annular axial swirling gas pilot premixer 855.

  Center cartridge 155 includes a cylindrical body 405 mounted on rear flange 224. The central cartridge 155 is inserted into the central cavity 203 of the gas pilot nozzle body 805 and bolted onto the rear surface 820 raised through the rear flange 224. The rear flange 224 includes an axial entry for connecting on its circumferential surface to the igniter and flame detector 236, a radial entry 232 for liquid pilot fuel, and a radial entry for air. Part 234. The center cartridge 155 is aligned with the center axis 200 of the nozzle.

  FIG. 10 shows an axial cross-sectional view of the central cartridge 155 of the nozzle embodiment of the present invention. The central cartridge 155 is radially enclosed within the cartridge wall 872 and at the downstream end near the end tip 885. The igniter 875 extends axially from the center cartridge flange 224 to the end tip 885. A liquid fuel pilot 880 extends from the central cartridge flange 224 to the end tip 885. Air cavity 873 receives air for use in the central cartridge. Air to the central cartridge enters through the radial feed holes 275 in the back plate 240 and exits through the holes 277 to the space 864 between the pilot nozzle 150 and the inner surface 368 of the support tower 270. A portion of the air entering the space 864 enters the central cartridge 155 through a cartridge supply hole 870 that fills the air cavity around the igniter 875 and the liquid fuel pilot 880 and extends forward relative to the tip impact shield 865. The tip collision shield 865 seals the upper end portion of the cavity and includes a plurality of tip holes 867 (18 holes in the embodiment of the present invention). Air from the tip impact shield 865 is directed to the annular air channel 876 at the downstream end of the igniter 875 that supports ignition. Air from the tip impact shield 865 is also supplied to the conical annulus 881 around the heat shield 882 on the liquid fuel pilot 880. A plurality of offset blast holes 883 are provided through the air blast sprayer shroud 884. Liquid pilot fuel is provided through a cylindrical cavity 890 in the liquid pilot body 891. The frustoconical annulus of the spin chamber wall 892 at the tip defines a spin chamber 893 therein for liquid fuel. The annular air gap 894 is provided around the liquid pilot main body 891 for heat insulation. Around the annular air gap 894 is provided an air assist annulus 895 connected to the air assist supply device in the flange 224 of the central cartridge 155. Arranged within the air assist annulus 895 is an air assist swirler 896 (in the embodiment, including eight swirl vanes 897). The swirl vane 897 gives a swirl motion to the auxiliary air introduced into the spin chamber 893.

  FIG. 11 shows another embodiment of a central flame holder for a single large radial nozzle 900. Here, the flame holder 905 includes a perforated cup 910. The perforated cup 910 consists of a plurality of holes 920 around a central axis 915 through which 30 percent (about) of the air fuel mixture 950 originating from the annular region passes, partly pilot air fuel originating from the pilot nozzle 960 Mixing with the air-fuel mixture 955, some burns at the exit of these holes 920. The hole 920 includes a fillet 930 to minimize corner separation. A shroud 940 is provided around the cup 910 to direct the flow into the cup. The lower convex end 945 of the cup 920 is open and conformal to receive the pilot air / fuel mixture 955 from the gas pilot nozzle 960. Therefore, in this case, heat dissipation occurs in three stages. The first stage is a pilot zone. The second stage is an air-fuel mixture that burns at the outlets of these holes 920 and the third stage is a flow that bypasses the perforated cup 920.

  The foregoing describes a single large radial nozzle for a gas turbine combustor that provides a major improvement in operation over a multi-nozzle design. First, in combination with a controllable external main gas fuel injection path and an internal main gas fuel injection path, the nozzle premixer design, particularly the in-nozzle combustion staging provided by the conical de-swar V-groove flame holder, This is an inherent aspect of the design. This aspect allows multiple nozzles (per combustor) to be replaced with a single nozzle, resulting in a major cost and component reduction. Secondly, combustion dynamics / vibration reduction is brought about by smearing out the number of fuel transfers, and heat dissipation in the chamber is a novel method. This unique property can also allow a wide range of fuels to be combusted without the need for hardware modifications or replacement. Finally, a combustor headend design and an annular dome diffuser that recovers pressure, while nozzles are used to convectively cool the back of the liner dome without the need for a separate cooling air source. The method integrated into the combustor dome can improve simplicity as well as functionality.

  At present, the nozzle of the present invention is sized for a GE9FB high power industrial engine, but works with almost any combustor ring design (eg, 7H, 9H, 7FB, 7FA, 9FA, 6C, etc.). The size can be enlarged or reduced. This design can be incorporated into existing packages or introduced as a new product offering.

  While only certain features of the invention have been illustrated and described herein, many modifications and changes will occur to those skilled in the art. Therefore, it is to be understood that the appended claims are intended to cover all modifications and changes that fall within the true spirit of the invention.

1 shows a prior art combustor having multiple nozzles. FIG. 1 shows an embodiment of a combustor of the present invention including a large single radial nozzle combustor of the present invention. FIG. FIG. 3 is an isometric cutaway view showing the internal structure of an embodiment of the large single radial nozzle structure of the present invention. FIG. 3 is an axial sectional view showing the internal structure of an embodiment of the large single radial nozzle structure of the present invention. FIG. 3 is a feed end view of an end cover plate of an embodiment of a large single radial nozzle of the present invention. FIG. 4 shows the fuel gallery and fuel entry portion of the end cover and rear plate of an embodiment of the large single radial nozzle of the present invention. 1 is an isometric view of a main radial swirler of an embodiment of a large single radial nozzle of the present invention. FIG. FIG. 4 is a detailed view of the main swirl vanes on the main radial swirler of the large single radial nozzle embodiment of the present invention. 1 is a front view of a main radial swirler of an embodiment of a large single radial nozzle of the present invention. FIG. FIG. 3 is a cross-sectional view through the main swirl vane and central hub of the main radial swirler of an embodiment of the large single radial nozzle of the present invention. FIG. 3 is a cross-sectional view through the rear plate and central hub of the main radial swirler of an embodiment of the large single radial nozzle of the present invention. 1 is a cross-sectional view of the head end of a combustor of the present invention depicting the air flow and independent combustion zones of an embodiment of a large single radial nozzle of the present invention. FIG. 3 shows a central flame holder, gas pilot annulus, and central cartridge of an embodiment of a large single radial nozzle of the present invention. The figure which shows the nozzle body of the gas pilot nozzle of embodiment of the large sized single radial nozzle of this invention. The figure which shows the nozzle body of the gas pilot nozzle of embodiment of the large sized single radial nozzle of this invention. FIG. 3 is an axial cross-sectional view of a dual fuel center cartridge of an embodiment of the large single radial nozzle of the present invention. FIG. 4 shows another embodiment of a large single radial nozzle flame holder cup of the present invention.

Explanation of symbols

DESCRIPTION OF SYMBOLS 10 Gas turbine 12 Compressor 14 Combustor 16 Turbine blade 18 Transition duct 20 Spark plug 22 Flame propagation tube 24 Combustion casing 26 Turbine casing 28 Bolt 30 End cover assembly 32 Fuel nozzle assembly 34 Flow sleeve 36 Transition duct outer wall 38 Liner 39 Struts 42 Combustion liner cap assembly 44 Aperture 46 Premix tube 47 Front plate 48 Floating collar 49 Rear plate 50 Fuel swirler 52 Rear supply section 54 Front delivery section 66 Radial fuel injector 69 Premix zone 70 Combustion chamber 100 Can annular Type dual fuel combustor 105 combustion casing 106 flow sleeve 110 combustor liner 111 conical dome 112 extending portion 113 external burner tube 114 Combustion chamber 115 annular space 116 diffuser 117 inlet plenum 118 fairing 120 lean premix radially flowing Maruchianyura multistage nozzle (large single radial nozzle)
130 End cover assembly 135 Rear surface 138 Front surface 140 Main radial swirler 150 Gas pilot nozzle 155 Central cartridge 160 Central flame holder 170 External flame holder 175 External burner pipe 178 Downstream end of burner pipe 179 Support ledge 180 Base section 183 Bolt 185 Main external gas supply device 190 Main internal gas supply device 195 Liquid gas supply device 200 Central shaft 201 Rear section 202 Front section 203 Central cavity 204 Front surface 205 End cover plate 206 Combustor bolt hole 207 External circumferential surface 208 Pilot hole 209 Bolt hole 210 Gas pilot nozzle seating surface 212 Gas pilot nozzle flange 214 Gas pilot nozzle rear flange bolt 215 Main external gas Inlet 220 Main internal gas inlet 223 Central cartridge 224 Central cartridge flange 225 Central cartridge flange bolt 230 Gas pilot connection 232 Liquid fuel pilot connection 234 Air auxiliary connection 236 Ignition device / flame detector connection 240 Rear plate 241 Base Surface 242 Mounting surface 243 Liquid fuel supply entry rear portion 244 Main liquid fuel gallery 245 Liquid fuel injection point 246 Liquid fuel delivery entry portion end cover 247 Liquid fuel delivery entry portion rear plate 248 Sprayer 249 Insulation lining 250 Swivel vane 252 Fuel tip 255 Annular swirl volume 257 Circumferential edge of rear plate 260 Central hub 263 Inner surface 265 Central cavity 270 Cylindrical support hub 273 External support ledge 274 Internal mounting Wedge 275 radial air passage 277 outlet 278 central cavity 280 dewarer 285 central hub 290 V-groove pack 295 support tower 296 support tower inner wall 300 inner burner tube 310 main outer gas gallery 311 inner wall 312 outer wall 315 upper end 317 inner wall 318 outer wall 319 Upper end 330 Main internal gas gallery 345 Annular compartment 347 Annular compartment inlet 348 Annular compartment downstream outlet 350 Cylindrical support tower 355 Support arm 360 Radial vane 365 Internal burner tube wall 366 Inner wall 368 Cylindrical hub inner wall 370 Cylindrical hub 371 Recess 372 Bolt hole 375 V-shaped element 376 Open end 377 Apex 380 Annular tip 381 First recess 382 Second recess 383 Third recess 385 Tip of tip internal burner tube End 390 Inclined edge 400 Air flow from diffuser 401 Air flow from inlet plenum to main radial swirler 402 Flow to central flame holder zone 405 Flow to main zone 605 Circumferential premixing space 610 Airfoil 615 Front Edge 620 Trailing edge 625 Axis 630 Acute angle 635 Radius 640 Side face with injection point 642 Side face without injection point 645 Bottom face 650 Top face 655 Main external gas fuel injection point 657 Main external gas fuel injection hole 660 Air flow 665 Fuel cavity 670 Screw hole 680 Main internal gas injection entrance 685 Orifice 690 Injection point 695 Main internal gas injection point 805 Main body 810 Annular shell 815 Rear flange 816 Bolt hole 817 Front seat surface 818 Center hole 819 Raised surface 820 Rear surface 821 Screw hole 825 Gas pilot animation 830 outer surface 835 lower shell 836 outer surface of lower shell 840 tapered shell 842 outer surface of tapered shell 845 central shell 847 outer axial surface 850 tapered head 852 outer surface of tapered head 855 gas pilot premixer 857 mixing vane 860 pilot gas hole 862 Internal pilot gas fuel cavity 864 Gas pilot air flow space 865 Tip impact shield 866 Tip hole 870 Cartridge supply hole 872 Outer surface of central cartridge 873 Air cavity 875 Ignition device 880 Liquid fuel pilot 881 Conical annulus 882 Heat shield 883 Offset blast hole 884 Air Blast sprayer shroud 885 End tip 890 Cylindrical fuel cavity 891 Liquid pilot body 892 Pin chamber wall 893 Spin chamber 894 Annular air gap 895 Air assist annulus 896 Air assist swirler 900 Large single radial nozzle 905 Center flame holder 910 Perforated cup 915 Center shaft 920 Hole 930 Fillet 940 Shroud 945 Inner burner tube wall 950 Center support 30 percent fuel-air mixture entering the flamearm 955 Pilot-air fuel-air mixture 960 Pilot nozzle

Claims (10)

A lean premixed radial inflow multi-annular multi-stage nozzle 120 that produces three independent combustion zones within a can-annular dual fuel gas turbine combustor 100 comprising:
A pilot zone Z1 fueled by a central cartridge 223 during liquid operation and by a central gas pilot nozzle 150 during gas operation;
A central flame holder zone Z2 fueled by an internal main gas fuel supply device;
A main flame zone Z3 fueled by an external main gas fuel supply device;
A rear plate 240 that is axially aligned and mechanically fastened to the end cover 130; and a plurality of swirl vanes 250 that are arranged in a circular arrangement around the central axis of the nozzle and spaced at approximately equal distances. A premixing volume 205 in the circumferential space between the individual swirl vanes 250, a central hub 260 including a central cavity 278 therein, and an annular swirl volume between the plurality of swirl vanes 250 and the central hub 260 A main radial swirler 140 that mixes a portion of the air flowing into the nozzle 120 with the internal main gas fuel supply device and the external main gas fuel supply device,
Means for controlling the ratio between the internal main gas fuel supply 220 and the external main gas fuel supply 215;
An end cover 130;
A nozzle 120 comprising: The main radial swirler 140 is
A cylindrical rear plate 240;
A central hub 260 projecting axially downstream from the downstream surface 242 of the rear plate 240, including a smooth conical surface 270 truncated at the downstream end;
A cavity 665 connecting the external main gas fuel supply device from the end cover 205 to the plurality of swirl vanes 250;
A plurality of injector nozzles 690 mounted on the downstream surface 242 of the rear plate 240;
A cavity 680 connecting an internal main gas fuel supply device from the end cover 205 to the plurality of injector nozzles 690;
A plurality of liquid fuel sprayers 248 mounted on the downstream surface 242 of the rear plate 240;
A cavity 247 connecting a liquid fuel supply device from the end cover 205 to the plurality of liquid fuel sprayers;
A central cavity 278 along the central axis of the nozzle 120;
A plurality of radially oriented cavities 250 connecting an outer circumferential surface 275 of the rear plate with the central cavity 278 to supply air to the central cartridge 223 and the gas pilot nozzle 150;
The nozzle 120 according to claim 1, comprising: Each of the plurality of swirl vanes 250 is
Projecting axially downstream from the downstream surface 242 of the rear plate 240 toward the combustion end of the nozzle 120, the center line 225 forms a predetermined angle 630 with a radius 635 from the central axis 200 of the nozzle 120. An airfoil portion 610 defining a circumferential premixing space 605 for airflow from the outside of the main swirler 250 to the annular swirling volume 255 between the main swirling vane 250 and the central hub 270;
An internal cavity 665 in each airfoil 610 for the external main gas fuel supply in the rear plate;
A plurality of gas fuel injection holes 657 for distributing an external main gas fuel supply from the internal cavity 665 to the premix space 605;
The nozzle 120 according to claim 2, comprising: The end cover is
A cylindrical plate 205 including an outer radial mounting surface 207 for mechanical attachment to the combustor 100 and an inner radial mounting surface 204 for attachment to the rear plate 240;
A cavity connecting an external main gas fuel supply 185 to the rear plate 240;
A cavity connecting an internal main gas fuel supply device 190 to the rear plate 240;
A plurality of cavities 246 connecting a liquid fuel supply device 195 to the rear plate 240;
A central cavity 203 including a mounting flange 210 for receiving and mounting the gas pilot nozzle 150;
The nozzle 120 according to claim 1, comprising: The central flame holder zone Z2 is
A central hub 285;
An internal burner tube wall 365;
Adjacent compartments including a plurality of segmented radial compartments 345 each formed as an annular segment bounded on the outer radius by the inner burner tube wall 365 and on the inner radius by the outer wall of the central hub 385 345 are separated by a radially inclined radial wall radial vane 360, the slope upstream of the compartment 345 for swirling a portion of the fuel-air mixture in the annular swirl volume 255. Diverting the circumferential flow of the fuel-air mixture in the annular swirling volume 255 of the main radial swirler 250, gradually increasing from the side inlet 347 to the downstream outlet 348 from the compartment 345 and downstream A de-swarler 280 that redirects the airflow in the axial direction;
A V-groove flame holder pack 290 including a plurality of radially oriented arms 360 spaced approximately equidistantly around the circumferential circumference of the inner burner tube wall 365;
With
The arm 360 is attached to and extends from the downstream end of the central hub 285 to the downstream axial end of the inner burner tube 300, and the attachment position on the inner burner wall 365 is attached to the hub extension. Located downstream from the position, thereby a predetermined radial-to-axial angle 630 of the radial orientation arm 360, a convex recess in the radial orientation arm 360, and the convex recess 375 facing upstream. Form a vertex 377 of
The nozzle 120 according to claim 1. The central hub 285 is
Formed irregularly on the inner surface 296 of the cylindrical tube 296 to accommodate the central gas pilot nozzle 150 and adapted at the upstream end to combine with the central hub 260 of the main radial swirler 140 Further comprising a hub extension 380 on the downstream axial end of the cylindrical tube 295, comprising a cylindrical tube 295 having a central cavity 278 and surrounded by a V-groove 290 at equidistant intervals.
The nozzle according to claim 5. The central gas pilot nozzle 150 is
A central cavity at the upstream end and a radially enlarged bolt flange 815 to fit the rear plate and the central hub 285 of the flame holder zone within the central cavities 203, 270, 278 of the end cover 205; A generally cylindrical nozzle body 805 adapted to mate,
A plurality of radial air supply holes 870 that are axially aligned on the body 805 to receive airflow from the rear plate 240;
A central cartridge 155 in the central cavities 203, 270, 278;
An annulus 864 adapted to supply pilot gas fuel from a gas pilot nozzle 150 to the gas pilot at a downstream end of the central cartridge 155;
A plurality of pilot mixing vanes 857 adapted to mix pilot gas fuel and air;
A plurality of radially extending holes 860 passing through wall 872 of said central cartridge 155 upstream between adjacent pilot mixing vanes 857;
An annulus 825 located downstream of the pilot mixing vane 857, adapted to supply a mixed pilot gas air mixture to the pilot zone Z1;
The nozzle 120 according to claim 1, comprising: The central cartridge is
A liquid fuel pilot 880;
An air auxiliary supply device 873 for the liquid fuel pilot 880;
An ignition device 875;
The nozzle according to claim 7, comprising: The main flame zone Z3 is further
A cylindrical inner burner tube wall 365 centered on the central axis 200 of the nozzle 120;
A cylindrical outer burner tube 175 centered on the central axis 200 of the nozzle 120, protruding axially downstream from the main radial swirler 140 and having a diameter larger than the inner burner tube wall 365;
The outer burner tube at an upstream end extending radially outward and having a roof formed on a plurality of main swirl vanes 250 for delivering fuel and air to the annular mixing zone 175 base portion 180;
The nozzle 120 of claim 1 further comprising: A can annular dual fuel combustor 100 for a gas turbine engine comprising:
A lean premixed radial inflow multi-annular multi-stage nozzle 120 including an inner burner tube 300, an outer burner tube 113, and a main radial swirler 240 mounted on the combustor casing 105 on the end cover 130;
A main combustion zone Z1 downstream from the external burner tube 113 of the nozzle;
A source of pressurized air from the combustor;
An air inlet plenum 117 that radially surrounds the nozzle 120 and is radially bounded by the casing wall 105 of the combustor;
A pressurized air diffuser 116 that receives the pressurized air in a reverse flow path from the compressor and discharges the pressurized air to the inlet plenum at a recovery pressure;
With
The diffuser 116 includes an inner wall coinciding with the rear side of the dome 111 on the main combustion zone, thereby providing back cooling to the dome 111 from the pressurized air passing through the diffuser 116;
In order to smooth the air flow from the diffuser 116 to the air inlet plenum 117, a fairing 118 mounted on the main radial swirler and surrounding a portion of the outer burner tube 113 is provided.
A can-annular dual-fuel combustor 100 characterized by that.

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