US20090077972A1

US20090077972A1 - Toroidal ring manifold for secondary fuel nozzle of a dln gas turbine

Info

Publication number: US20090077972A1
Application number: US11/859,087
Authority: US
Inventors: Arjun Singh
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2007-09-21
Filing date: 2007-09-21
Publication date: 2009-03-26
Also published as: CH697908A2; JP2009074792A; DE102008037374A1; CN101392917A

Abstract

A toroidal ring manifold for effectively dispersing premixing fuel with air in the secondary fuel nozzle of a combustor for a Dry Low NO_x(DLN) gas turbine, thereby providing stable combustion with low nitrogen oxide (NO_x) emissions. The toroidal ring manifold is centered around a centerbody hub of a secondary fuel nozzle assembly in a premixing volume between the nozzle centerbody hub body and a centerbody cap. The ring manifold receives fuel from the nozzle body and dispenses premix fuel from a plurality of rows of individual holes on its downstream surface into an axial airstream. The number and location of the rows, the number, size and spacing of holes in each row, and the radial position of the toroidal ring manifold within the premixing volume are optimized to promote premixing.

Description

BACKGROUND OF THE INVENTION

The invention relates generally to a secondary fuel nozzle for a gas turbine combustor and more specifically to a toroidal ring manifold for fuel premixing in the secondary fuel nozzle of a Dry Low NOx (DLN) gas turbine.
FIG. 1 illustrates a prior art combustor for a gas turbine 12, which includes a compressor 14 (partially shown), a plurality of combustors 16 (one shown for convenience and clarity), and a turbine represented by a single blade 18. Although not specifically shown, the turbine 16 is drivingly connected to the compressor 14 along a common axis. The compressor 14 pressurizes inlet air, which is then reverse flowed to the combustor 14 where it is used to cool the combustor 16 and to provide air to the combustion process. Although only one combustor 14 is shown, the gas turbine 10 includes a plurality of combustors 16 located about the periphery thereof. A transition duct 20 connects the outlet end of each combustor 16 with the inlet end of the turbine 18 to deliver the hot products of combustion to the turbine 18.
Each combustor 16 comprises a primary or upstream combustion chamber 24 and a secondary of downstream combustion chamber 26 separated by a venturi throat region 28. The combustor 16 is surrounded by a combustor flow sleeve 30, which channels compressor discharge air flow to the combustor. The combustor is further surrounded by an outer casing 31, which is bolted to the turbine casing 32.
Primary nozzles 36 provide fuel delivery to the upstream combustion chamber 24 and are arranged in an annular array around a central secondary nozzle 38. Each of the primary nozzles 36 protrudes into the primary combustion chamber 24 through a rear wall 40. Secondary nozzle 38 extends from a rear wall 40 to the throat region 28 in order to introduce fuel into the secondary combustion chamber 26. Fuel is delivered to the primary nozzles 36 through fuel lines (not shown) in a manner well known in the art.
Combustion air is introduced into the fuel stage through air swirlers 42 positioned adjacent the outlet ends of nozzles 36. The swirlers 42 introduce swirling combustion air, which mixes with the fuel from nozzles 36 and provides an ignitable mixture for combustion on startup, in chamber 24. Combustion air for the swirlers 42 is derived from the compressor 14 and the routing of air between the combustion flow sleeve 30 and the wall 44 of the combustion chamber. The cylindrical wall 44 of the combustor is provided with slots or louvers 46 in the primary combustion chamber 24, and similar slots or louvers 48 downstream of the secondary combustion chamber 26 for cooling purposes, and for introducing dilution air into the combustion zones to prevent substantial rises in flame temperature. The secondary nozzle 38 is located within a centerbody 50 and extends through a liner 52 provided with a swirler 54 through which combustion air is introduced for mixing with fuel from the secondary nozzle.
Referring now to FIG. 2, a gas-only secondary fuel nozzle assembly 56 is illustrated. Fuel is supplied to sustain a flame by diffusion pipe P₁and to sustain a premixed flame by pipe P₂which, at the inlet to the secondary fuel nozzle assembly 56, are arranged concentrically relative to each other.
The following will primarily describe the premix fuel secondary nozzle assembly 56. A rearward component, or gas body, 58 includes an outer sleeve portion 60 and an inner hollow core portion 62 provided with a central bore 94 forming a premix fuel passage 64. A plurality of axial air passages 68 are formed in a forward half of the rearward component 58 in surrounding relationship to the premix passage 64. A like number of radial wall portions (e.g., four) are arranged about the end of sleeve portion 60 and each includes an inclined, radial aperture 70 for permitting air within the liner 52 to enter a corresponding air passage 68. The rearward end of component 58 is adapted to receive the fuel pipes P₁, P₂, respectively, as shown in FIG. 2, within a mounting flange 77.
A plurality of radial holes 78 are provided about the circumference of the forward portion of component 58, permitting a like number of radial gas injector tubes (pegs) 80 to be received therein to thereby establish communication with the premix passage 64. Each peg 80 is provided with a plurality of apertures or orifices 82 so that fuel from the premix passage 64 may be discharged into a premixing area 90 between the secondary nozzle assembly 56 and liner 52 for mixing with combustion air within the liner. The pegs 80 are designed to distribute fuel into the airflow. Good mixing of fuel and air in the premixing area 90 is necessary to minimize nitrogen-oxide (NO_x) emissions. A flame holding swirler 116 which may or may not be integral with the nozzle is located at the forward end of the secondary nozzle, extending radially between the reduced diameter forward end 108 and the liner 52 for swirling the premixed fuel/air flowing within the liner. Combustion air will enter the secondary nozzle assembly 56 via holes 70 and will flow through a premix passage defined by passage 64, pilot bore 98 and pilot orifice 100. This fuel, along with air from swirler slots 96, provides a diffusion flame sub-pilot. At the same time, a majority of the fuel supplied to the premix passage will flow into the gas injectors 80 for discharge from orifices 82 toward the liner 52 where it is mixed with air.
As illustrated in FIG. 3A-3B, premixing of fuel with air as performed in prior art secondary fuel nozzles may include the plurality of pegs 80, equally spaced around the periphery of the secondary nozzle body 75 in the premixing volume 40. Each peg 80 may include a central cavity 85 running the length of the peg. The inner end of each peg may be attached to the nozzle body at the location of the radial fuel holes, thereby establishing communication between the fuel cavity in the nozzle body and the central cavity of the peg, as previously described with respect to FIG. 2. Along a downstream surface of the peg 80, a plurality of the fuel discharge holes 82 are provided from the central internal cavity, thereby providing for discharge of premix fuel into the airflow between the secondary nozzle body 75 and the liner 52. Three radially-located fuel discharge holes 82 are provided along the downstream side of the peg 80. Positioning of the hole location along the row of holes was varied. In this prior art secondary nozzle, six pegs are evenly distributed around the circumference of the secondary nozzle body 75, with three orifices for fuel dispersal along the downstream side of the peg. However, the effective mixing of fuel and air is not complete. More complete mixing of the fuel and air can lead to lower NO_xemissions and more stable combustion.
The above described nozzle construction provides for the premixed mode of operation via a diffusion flame and, once in the premixed mode, operates to turn off the diffusion flame and start the premixing flame, for sustained operation. However, elevated emissions from a gas turbine is the result of insufficient mixing of air and fuel prior to burning in the combustion chamber. The existing peg design, described above, is not able to mix fuel and air properly to obtain the requisite degree of mixing for low emissions. Attempts to change the location of holes in the pegs have not been able to achieve satisfactory fuel and air mixing.
FIG. 4 illustrates a fuel distribution device 150 for a secondary fuel nozzle as described in U.S. Pat. No. 6,446,439 and U.S. Pat. No. 6,282,904 by Kraft et al. An annular fuel manifold 155 is mounted to a support sleeve 160 through support cylinders 165. The manifold 155 presents a rectangular cross-section. The support sleeve 160 is affixed to the body of a secondary fuel nozzle (not shown) by welding. Fuel in the body of the secondary nozzle, passes through holes 170 in the support sleeve and through the support cylinders 165 into the hollow annular fuel manifold 155. The annular fuel manifold 155 is positioned in an airstream 175 around secondary nozzle body (not shown). Fuel is distributed from the downstream face 180 of the annular fuel manifold through an array of apertures 185. The apertures 185 may be at a first radial distance 186 or a second radial distance 187 within the airstream from a central axis. The direction of the apertures 185 with respect to the airflow may be collinear or at an angle. However, the rectangular-shaped annulus limits the angles of the apertures may make with respect to the direction of the airstream.
The cylindrical-shaped annular cylindrical fuel manifold 155 for fuel premix distribution may provide for radial and circumferential fuel distribution over the peg arrangement. However the annular manifold has limitations on mixing, stemming from the limited flow angles that may be created within respect to the airflow, and particularly with respect to the radial distribution of fuel into the airstream.
Accordingly, there is a need to provide an alternate structure to improve the fuel-air premixing in the secondary nozzle to promote lower emissions and improved combustion dynamics.

BRIEF DESCRIPTION OF THE INVENTION

The present invention relates to a toroidal ring manifold for effectively dispersing premixing fuel with air in the secondary nozzle of a combustor for a DLN gas turbine, thereby providing stable combustion with low nitrogen oxide (NO_x) emissions.
Briefly in accordance with one aspect of the present invention, a toroidal ring manifold for dispersing fuel into a premixing zone of a secondary fuel nozzle for a DLN gas turbine combustor is provided. The toroidal ring manifold includes a toroidal ring manifold shell of a generally toroidal shape, with a cavity within the shell. A plurality of radial penetrations on an inner toroidal surface of the toroidal ring manifold shell extend into the shell cavity, each of the plurality of radial penetrations being located in a predetermined arrangement on the inner toroidal surface.
A plurality of support arms extend inward radially from the inner toroidal surface of the toroidal ring manifold shell. One support arm of the plurality of support arms may be attached to the toroidal ring manifold shell at each of the plurality of radial penetrations and extend inward radially from the inner surface of the toroidal ring manifold shell. The support arm further includes an axial internal cavity for transporting premix fuel from an inner radial end to an outer radial end at the toroidal ring manifold shell. A plurality of penetrations through one poloidal surface of the toroidal ring manifold shell for dispersing premix fuel are also provided and located according to a predetermined arrangement.
In accordance with another aspect of the present invention, a secondary fuel nozzle assembly for a gas turbine combustor is provided. The secondary fuel nozzle assembly includes a premix fuel connection at the rearward end of the secondary fuel nozzle assembly and a nozzle body connected at a rearward end to the premix fuel connection. A toroidal ring manifold is centered radially around the nozzle body for distributing fuel from the nozzle body into an axial airflow path around the nozzle body. A support structure is provided for supporting the toroidal ring manifold. Further, a communication path is provided for fuel from the nozzle body to the toroidal ring manifold.
In accordance with a third aspect of the present invention, a combustor for a DLN gas turbine is provided. The combustor includes a secondary fuel nozzle assembly with a fuel connection at the rearward end and a nozzle body connected at a rearward end to the fuel connection. A toroidal ring manifold may be centered radially around the nozzle body for distributing fuel from the nozzle body into an axial airflow path around the nozzle body. A support structure is provided for the toroidal ring manifold. A communication path is provided for fuel from the nozzle body to the toroidal ring manifold. A liner circumferentially surrounds the secondary nozzle assembly, including a penetration at the rearward end for admitting airflow to a premixing volume. The premixing volume encompasses a generally annular shape defined between the secondary nozzle assembly and an inner wall of the liner. A swirler may be mounted at the forward end of the secondary nozzle assembly.

BRIEF DESCRIPTION OF THE DRAWING

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 characters represent like pails throughout the drawings, wherein:

FIG. 1 is a partial cross-sectional view of a prior art known dry low NO_xcombustor;

FIG. 2A is a partial cross sectional view of a prior art secondary premixed/diffusion fuel nozzle;

FIG. 2B is an exploded cross-sectional view of the prior art secondary fuel nozzle illustrated in FIG. 2;

FIG. 3A illustrates a peg arrangement for the prior art secondary fuel nozzle;

FIG. 3B illustrates the arrangement of fuel discharge holes in the peg of the prior art secondary nozzle;

FIG. 4 illustrates a prior art manifold for fuel premix.

FIG. 5A illustrates a cross-section of an inventive toroidal shell;

FIG. 5B illustrates a side view of a toroidal ring manifold around a nozzle body;

FIG. 5C illustrates a toroidal ring manifold cross section illustrating row angles for fuel mixing holes on the downstream surface;

FIG. 6A illustrates an external view of a downstream side for a preferred embodiment of the inventive toroidal ring manifold;

FIG. 6B illustrates an internal view of the downstream side for a preferred embodiment of the inventive toroidal ring manifold;

FIG. 7A illustrates an isometric view of assembly arrangements for the inventive toroidal ring manifold to a secondary fuel nozzle body.

FIG. 7B illustrates a sectional view of the inventive toroidal ring manifold through a secondary fuel nozzle, support arm and manifold shell;

FIG. 8 illustrates a secondary fuel assembly for a gas turbine incorporating an embodiment of the inventive toroidal ring manifold;

FIG. 9 illustrates a combustor for a gas turbine incorporating the inventive toroidal ring manifold.

FIG. 10A illustrates a radial profile of calculated unmixedness for the prior art peg secondary fuel nozzle; and

FIG. 10B illustrates a radial profile of calculated unmixedness for the secondary fuel nozzle with an embodiment of the inventive toroidal ring manifold.

DETAILED DESCRIPTION OF THE INVENTION

The following embodiments of the present invention have many advantages, including effectively dispersing premixing fuel with air in the secondary nozzle of a combustor for a Dry Low NO_x(DLN) gas turbine, thereby providing stable combustion with low nitrogen oxide (NO_x) emissions.
According to one aspect of the present invention a toroidal ring manifold is provided to disperse fuel into the premixing volume between a secondary nozzle body and the finer. FIG. 5A illustrates a cross-section of a toroidal shell. A toroidal ring manifold 200 may comprise a generally toroidal-shaped shell 210, with a circular ring cross section 215. The toroidal-shaped shell 210 is centered on a poloidal axis 220. When the toroidal ring manifold 200 is united with a nozzle body 400 (partially shown) as part of a secondary fuel assembly, the poloidal axis may be coincident with a longitudinal axis 305 of the nozzle body 400. A toroidal axis 225 runs through the center 230 of circular ring cross-section 215. The toroidal ring manifold 200 includes a ring height H 235, defined between the center 240 of the toroidal ring manifold on the poloidal axis 220 and the center 230 of the circular ring cross-section 215 on toroidal axis 225. Setting of the ring height H 235 and a cross-section diameter 236 adjusts the positioning of the ring outer surface 237 relative to a premixing space (not shown) around the body of the secondary nozzle. The outer surface 237 of the toroidal ring manifold 200 may be divided into an upper surface 238 and a lower surface 239 by an equator of the ring.
A locus of points forming a row circumferentially around the toroidal ring manifold (for example first point 250 on the surface of the circular cross-sections 215) may be defined by row angle θ ₁ 260 formed between a line running parallel to the poloidal axis 220 through the center 230 of the ring cross-section 215 and a line 270 between center 230 of the ring-cross section 215 and the point on the surface 250. For convention, angle θ₁(toward the poloidal axis) is defined as positive. A second row angle θ ₂ 275 describes point 255 and an associated locus of centers for a second row, also on the outer surface of the toroidal ring manifold 200.
FIG. 5B illustrates a side view of a toroidal ring manifold 200 around a nozzle body 400. Toroidal ring manifold 200 includes fuel mixing holes 320 on a downstream surface 290. Height H 235 of the toroidal ring manifold establishes positioning of the manifold and hence of the fuel mixing holes 320 relative to the nozzle body 400. FIG. 5C illustrates a toroidal ring manifold cross section illustrating row angles for fuel mixing holes on the downstream surface. The row angles θ ₁ 260 and θ ₂ 275 may define the locus of centers for fuel mixing holes 320 (described in greater detail below).
FIG. 6A illustrates a downstream elevation view of the toroidal ring manifold. FIG. 6B illustrates an internal elevation view of the downstream surface of the toroidal ring manifold. The toroidal ring manifold 200, centered on the poloidal axis 230 includes a central cavity 290 within the manifold shell 210. The central cavity 290 may also be toroidal-shaped. A plurality of radial penetrations 295 may be provided through the outer surface 237 of the toroidal manifold shell 210, where the penetrations 295 extend into the central cavity 290, according to a predetermined arrangement. The toroidal ring manifold 210 may also include a plurality of support arms 300 extending inward radially from the lower surface 239 of the manifold shell 210. In one aspect of the inventive manifold, four support arms 300 and the corresponding radial penetrations 295 may be evenly distributed about the lower surface 239 of manifold shell 210.
The plurality of support arms 300 may each include an axial internal cavity 325 for fuel delivery, where the internal cavity 325 extends from an inner radial end 326 to an outer radial end 327 of the support arm 300. The outer radial end 327 may be connected to the toroidal ring manifold 200 at the location of one of the corresponding radial penetrations 295 of manifold shell 210, thereby providing a communication path for fuel through the support arms 300 and into the central cavity 290 of the toroidal ring manifold 200.
FIG. 7A illustrates an isometric view of assembly arrangements for the inventive toroidal ring manifold to a secondary fuel nozzle body. FIG. 7B illustrates a sectional view through a secondary fuel nozzle body 350, support arm 300 and manifold shell 210. The manifold shell 210 may be comprised of four shells section 330. Each shell section 330 includes a radial penetration 295 for accepting the outer radial end 327 of the support arm 300. The support arm 300 includes threaded connections 328 on inner radial end 326 and threaded connections 329 on the outer radial end 327. The radial penetration 295 on the manifold shell 210 includes threads 297 to mate with threads 329 to join the outer radial end 327 of support arm 300 to the shell section 330. Further radial holes 361 may penetrate wall 351 of nozzle body 350 extending to the premix fuel passage 360, thereunder. Radial holes 361 may include threads 365 to permit mating with threads 328 on the inner end 326 of support arm 300 for attaching the support arm 300 to the secondary fuel nozzle body 350. A flow path 355 for premix fuel is thereby provided from premix fuel passage 360 in the secondary fuel nozzle body 350, through radial holes 361, through cavity 325 in the support arms, and radial penetrations 295 in the shell section 330. Shell sections 330 may be connected to form toroidal ring manifold 200 by welding or other suitable connection means.
Referring again to FIG. 6A-6B, a plurality of penetrations, through the downstream poloidal surface of the toroidal ring manifold shell 210 and into the shell cavity 290 may be located according to a predetermined arrangement. The predetermined arrangement of penetrations are optimized to provide from mixing of fuel from the cavity 290 of the shell 210 with air flow surrounding the manifold shell 210. The predetermined arrangement may include at least one row of holes. Loci for the centers of individual holes in a row (ring) 280, 285 of holes may be set at an individual predetermined angle with respect to the ring cross-section. It is preferred that the individual holes 310, 315 within a row of holes 280, 285 be evenly spaced circumferentially around the locus of centers for a specific row. Further it is preferred that the individual holes within each row be of the same diameter. Holes for separate rows of holes may be of the same diameter or a different diameter. It is also preferred that the toroidal ring manifold shell include two rows of holes.
A preferred embodiment may include two rows or rings (a first row 280 and a second row 285) of holes in a downstream poloidal surface of the toroidal ring manifold shell. It is further preferred that the locus of centers for individual holes 310 for the first row 280 be at a predetermined positive angle θ₁ 260 (FIG. 5C), with respect to the ring cross-section 215 and the locus of centers for individual holes 315 of the second row 285 be at a predetermined negative angle θ₂ 275 (FIG. 5C) with respect to the ring cross-section 215. Further, it is preferred that the diameter size of the individual holes 310 of the first row 280 be smaller than the diameter size of the individual holes 315 of the second row 285. In a preferred embodiment of the inventive manifold, the individual holes 310 of the first row 280 may include diameters of about 0.082 inch and the individual holes 315 of the second row 285 may include diameters of about 0.116 inch.
The number of individual holes of a first row of holes may equal or not equal the number of individual holes of other rows. In a preferred embodiment of the inventive manifold with two rows of holes, the first row 280 may include 16 individual holes 310 and the second row 285 may include 8 individual holes 315. Further, it is preferred that the individual holes 310 of the first row 280 and the individual holes 315 of the second row 285 be evenly spaced relative to each other so as to promote an evenly distributed circumferential fuel-air (F/A ratio) in the premix volume.
Another aspect of the present invention provides a secondary fuel nozzle assembly 500 for a gas turbine combustor including an embodiment of the inventive toroidal ring manifold 200 (of FIGS. 5-7) is illustrated in FIG. 8. FIG. 8 illustrates the inventive toroidal ring manifold 200 mounted to the nozzle centerbody 350. The toroidal ring manifold 200 is positioned in the premixing volume 380 for distributing secondary premix fuel around the centerbody hub 395 and within the centerbody cap 390. A fuel oil connection 435, a secondary gas connection 436, a tertiary gas connection 437 and a water connection 438 may be provided at the rearward end of the secondary fuel nozzle assembly 500. The toroidal ring manifold 200 is centered radially around the nozzle centerbody 350 for distributing premix fuel from the nozzle centerbody 350 into an axial airflow path 450 around the nozzle centerbody. A support structure is provided for supporting the manifold ling shell from the centerbody hub 395.
Within the nozzle centerbody hub 395 of the secondary fuel nozzle assembly 500, a plurality of internal fuel cavities (passages) are provided for supplying different fuel types to the premixing volume 380 and to the tip 410 of the nozzle. Fuel oil may be provided to the tip 410 through fuel oil cavity 415. Secondary gas fuel cavity 426 connects the secondary gas fuel for premixing to a plurality of radial holes 365 through the outer wall 351 of centerbody hub 395. The plurality of radial holes 365 are arranged in a predetermined distribution around the circumference of the centerbody hub 395. The predetermined distribution of radial holes 365 are arranged to match and align with the support arms (not shown) of the toroidal ring manifold to provide a communication path for the premix fuel through the cavities of the support arm and into the shell cavity of the toroidal ring manifold 200, as previously described. The secondary fuel may further be distributed through secondary fuel passages 422 and 423 to pilot holes 440 at the nozzle tip 410 for a gas pilot flame. Tertiary gas fuel may be supplied through tertiary gas passages 425 to the nozzle tip 410. An injection water passage 424 for water injection at the nozzle tip 410 may further be provided.
Secondary gas fuel, released through the fuel mixing holes 320 into airstream 450, mixes with the air in premixing volume 380. The secondary gas fuel is further mixed with the airstream 450 by swirler 430 at the tip 410 of the nozzle.
The structure of the inventive toroidal ring manifold, its support arms, and the attachment to the centerbody hub have previously been described with respect to FIGS. 4-8. The secondary fuel nozzle assembly 500 includes these described elements of the toroidal ring manifold. Further, wlile the inventive toroidal ring manifold 200 is shown with one secondary fuel nozzle assembly 500, it should be understood that various embodiments of the inventive toroidal ring manifold may be combined with secondary fuel nozzles that include different arrangements and combinations of fuel sources and fuel passage arrangements.
A combustor 600 for a DLN gas turbine is also provided under another aspect of the present invention. The combustor 600 may include the secondary fuel nozzle assembly 500, including the toroidal ring manifold 200, as previously described. The combustor 600 may also a plurality of primary fuel nozzle assemblies 510 surrounding the secondary fuel nozzle assembly. Fuel connections 540 may be provided at the rear end of the combustor.
FIG. 9 illustrates an embodiment of the inventive combustor. A compressor 514 (partially shown) is mounted to the combustor 600 and provides compressed air to the combustor. A flow sleeve 530 surrounding the primary fuel nozzle assemblies 510 and secondary fuel nozzle assembly 500 accepts air from the compressor 514 through flow holes 532. Air flow between the flow sleeve 530 and combustion liner 535 is directed to the rear of the primary fuel nozzle assemblies 510 and secondary fuel assembly 500. Airflow 545 enters the rear end of the secondary fuel nozzle assembly 500 and flows around the toroidal ring manifold 200 between centerbody hub 565 and centerbody cap (liner) 570. The toroidal ring manifold 200 disperses fuel into the airstream from the fuel mixing holes 320 (FIG. 8) on its downstream surface. Air and secondary gas fuel mix 550 in premix volume 555 and may flow through swirler 560 situated between the forward end of nozzle tip 580 and the center body cap 570.
The structure of the inventive toroidal ring manifold, its support arms, and the attachment to the centerbody hub have previously been described with respect to FIGS. 4-8. The combustor 600 includes these described elements of the toroidal ring manifold 200. Further, while the inventive toroidal ring manifold 200 is shown with one combustor, it should be understood that various embodiments of the invention may be combined with secondary fuel nozzles that include different arrangements and combinations of airflow paths, fuel sources and fuel passage arrangements.
As previously described, the toroidal ring manifold shell further includes a radius being set to a predetermined height to align the at least one row of holes radially within the premixing volume between the secondary fuel nozzle assembly and the liner. The predetermined arrangement for the plurality of penetrations through a downstream poloidal surface of the toroidal ring manifold shell preferentially includes two rows of holes, with a locus for the centers of the holes for each row being set at an individual predetermined angle with respect to the ring cross-section, including a first row of holes with a locus for the centers of the holes for the first row being set at a predetermined positive angle with respect to the ring cross-section and a second row of holes, with a locus for the centers of the holes for the row being set at a predetermined negative angle with respect to the ring cross-section. The diameter for the individual holes within first row of holes may be smaller than the diameter for the individual holes within the second row of holes, and the centers for the individual holes within first row of holes may be staggered circumferentially with respect to centers for the individual holes within the second row of holes.
While the inventive toroidal ring manifold is shown with one combustor arrangement, it should be understood that various embodiments of the invention may be combined with other combustor embodiments that include different arrangements and combinations of fuel, air and water sources, flow paths and discharge.
The objective of the above-described design is to establish a constant fuel-air (F/A) ratio in the premix volume downstream of the toroidal ring manifold. It is desired to reduce both the circumferential F/A ratio variation and the radial F/A ratio variation in the premix volume and particularly at the exit of the swirler. The number of individual holes in a row of holes may be used to reduce the circumferential F/A ratio variation. The ring height and the row angles for the locus of centers of the rows of holes may be used to reduce the radial variation in F/A ratio and control the exit radial F/A profile downstream of the swirler.
A parameter representing unmixedness has been defined for comparing the relative performance of different cases with varied values for embodiments of the inventive design and the prior art. Unmixedness=ΣA_i(Φ_I−Φ_global)²/ΣA_i, where Ai=area of a cell, Φ_I=cell equivalence ratio and Φ_global=global equivalence ratio (0.4828). Unmixedness was computed for an existing peg design for the secondary fuel nozzle to be 0.06642. A design space was employed to evaluate the effect of individual hole locations and ring height on premixedness at the exit of the swirler. A calculation of unmixedness was performed for ranges of the following parameters, as follows: ring diameter (0.35-0.45 inch); ring height (1.2-1.7 inch); angle θ₁(0-120 degrees); angle θ₂(0-120 degrees).
Due to strong linearity in the responses, a reliable transfer function was difficult to obtain. Optimization of parameters was done with sampling and meta-model methods. Unnixedness was calculated at the exit of the swirler for the above-described cases. Ring height and row angle parameters were optimized to minimize the unmixedness, yielding: ring height approximately 1.35 inch, angle θ₁approximately 58.7 degrees and angle θ₂approximately 1.7 degrees. With the parameters optimized, unmixedness at the swirler exit was calculated as 0.01, considerably below the unmixedness of 0.06642 of the existing peg design.
Further, the toroidal ring manifold cross-section provides presents an aerodynamic design with respect to the airstream flowing between the liner and the nozzle body, superior to other less aerodynamic arrangements in prior art. For example, the inventive toroidal ring manifold presents a pressure drop roughly equivalent to the prior art peg design, thereby continuing to allow adequate fuel flow and air flow through the premixing volume and swirler while providing superior mixing.
FIG. 10A illustrates a radial profile of calculated unmixedness for the prior art peg secondary fuel nozzle. FIG. 10B illustrates a radial profile of calculated unmixedness for the secondary fuel nozzle with an embodiment of the inventive toroidal ring manifold. Together FIGS. 10A and 10B illustrate the relative unmixedness of fuel and air on a plane at the discharge of the swirler utilizing the prior art pegs and inventive toroidal ring manifold, respectively. The vertical axis represents distance within the premixing volume from the outside of the secondary fuel nozzle to the finer. The horizontal axis represents the equivalence ratio at specific radial distances from center points. For the peg design, the calculated equivalence ratio 1010 varies from 0.22 to 0.8 over the radial distance. For the inventive toroidal ring manifold design, the calculated equivalence 1020 ratio varies from 0.35 to 0.63 over the radial distance, indicating a much more even radial F/A ratio, thereby promoting more even downstream combustion, with the resulting reduction in combustion dynamics and in NO_xemissions.
While various embodiments are described herein, it will be appreciated from the specification that various combinations of elements, variations or improvements therein may be made, and are within the scope of the invention.

Claims

1. A toroidal ring manifold for dispensing fuel into a premixing zone of a secondary fuel nozzle for a Dry Low NO_x(DLN) gas turbine combustor, the toroidal ring manifold comprising:

a toroidal ring manifold shell of a generally toroidal shape, including a central cavity within the shell;

a plurality of radial penetrations on an inner surface of the toroidal ring manifold shell, each penetration of the plurality of penetrations extending into the central cavity and being located in a predetermined arrangement on the inner surface;

a plurality of support arms extending inward radially from the inner surface of the toroidal ring manifold shell, one support arm of the plurality of support arms being attached to the toroidal ring manifold shell at each of the plurality of radial penetrations and extending inward radially from the inner surface of the toroidal ring manifold shell, wherein the support arm further includes an axial internal cavity for transporting fuel from an inner radial end to an outer radial end at the toroidal ring manifold shell; and

a plurality of penetrations through one poloidal surface of the toroidal ring manifold shell, the plurality of penetrations being located according to a predetermined arrangement.

2. The toroidal ring manifold according to claim 1, each support arm of the plurality of support arms each comprising:

a means at the inner radial end for attaching the support arm to a fuel source and a means at the outer radial end for attaching the support arm to the toroidal ring manifold, wherein the means may include a threaded connection.

3. The toroidal ring manifold according to claim 1, wherein the predetermined arrangement for the plurality of penetrations through one poloidal surface of the toroidal ring manifold shell comprises:

at least one row of holes, with a locus for centers of holes for each row being set at an individual predetermined angle with respect to a ring cross-section.

4. The toroidal ring manifold according to claim 3 wherein individual holes within each row of the at least one row of holes are evenly spaced circumferentially along the locus for the centers of the holes for each row.

5. The toroidal ring manifold according to claim 4, wherein:

individual holes within each row of the at least one row of holes are of the same diameter;

the holes within separate rows of the at least one row of holes may be at least one of a same diameter and a different diameter; and

a toroidal radius of the toroidal ring manifold shell is set to a predetermined size for positioning the at least one row of holes with respect to the poloidal axis.

6. The toroidal ring manifold according to claim 3, wherein the predetermined arrangement for the plurality of penetrations through one poloidal surface of the toroidal ring manifold shell comprises:

two rows of holes, with a locus for the centers of the holes for each row being set at an individual predetermined angle with respect to the ring cross-section.

7. The toroidal ring manifold according to claim 6, wherein the two rows of holes, with a locus for the centers of the holes for each row being set at an individual predetermined angle with respect to the ring cross-section; comprises:

a first row of holes, with a locus for the centers of the holes for the first row being set at a predetermined positive angle with respect to the ring cross-section; and

a second row of holes, with a locus for the centers of the holes for the row being set at a predetermined negative angle with respect to the ring cross-section.

8. The toroidal ring manifold according to claim 7, wherein a diameter for the individual holes within first row of holes are smaller than a diameter for the individual holes within the second row of holes.

9. The toroidal ring manifold according to claim 7, wherein centers for the individual holes within first row of holes are staggered circumferentially with respect to centers for the individual holes within the second row of holes.

10. A secondary fuel nozzle assembly for a Dry Low NO_x(DLN) gas turbine combustor, the secondary fuel nozzle assembly comprising:

a fuel supply at a rearward end of the secondary fuel nozzle assembly;

a nozzle body connected at the rearward end to the fuel supply;

a liner axially surrounding the nozzle body;

a toroidal ring manifold centered radially around the nozzle body for distributing fuel from the nozzle body into an axial airflow path around the nozzle body;

a premixing volume between the nozzle body and the liner;

a support structure for the toroidal ring manifold; and

a communication path for fuel from the nozzle body to the toroidal ring manifold.

11. The secondary fuel nozzle assembly for a gas turbine combustor; the secondary fuel nozzle assembly further comprising:

an internal fuel cavity within the nozzle body for delivering fuel from the fuel supply;

a plurality of radial passages from the internal fuel cavity arranged in a predetermined distribution around a circumference of the nozzle body;

a plurality of penetrations on an inner toroidal surface of the toroidal ring manifold, the plurality of radial penetrations being equal to the plurality of radial passages from the internal fuel cavity wherein the plurality of radial penetrations are arranged axially and circumferentially to align with the predetermined distribution of radial passages from the internal fuel cavity;

a plurality of radial support arms equal in number to the plurality of radial passages from the internal fuel cavity, each support arm including an internal passage between an inner radial end and an outer radial end, the inner radial end being attached to the nozzle body at the location of one of the plurality of radial passages from the internal fuel cavity and the outer radial end being attached to the toroidal ring manifold at the location of the penetrations of the inner toroidal surface; and

a predetermined arrangement for a plurality of penetrations through a downstream poloidal surface of the toroidal ring manifold shell including at least one row of holes, with a locus for centers of holes for each row being set at an individual predetermined angle with respect to a ring cross-section of the toroidal ring manifold.

12. The secondary fuel nozzle assembly according to claim 11, wherein the predetermined arrangement for the plurality of penetrations through one poloidal surface of the toroidal ring manifold shell comprises:

individual holes, within each row of the at least one row of holes, being evenly spaced circumferentially around the locus for centers.

13. The secondary fuel nozzle assembly according to claim 12, wherein:

the individual holes within separate rows of the at least one row of holes may include at least one of a same diameter and a different diameter; and,

a toroidal radius of the toroidal ring manifold shell may include a predetermined size for positioning the at least one row of holes with respect to the poloidal axis.

14. The secondary fuel nozzle assembly according to claim 13, wherein the predetermined arrangement for the plurality of penetrations through a downstream poloidal surface of the toroidal ring manifold shell comprises:

two rows of holes, with a locus for the centers of the holes for each row being set at an individual predetermined angle with respect to the ring cross-section, including a first row of holes with a locus for the centers of the holes for the first row being set at a predetermined positive angle with respect to the ring cross-section and a second row of holes, with a locus for the centers of the holes for the row being set at a predetermined negative angle with respect to the ring cross-section.

15. The secondary fuel nozzle assembly according to claim 14, wherein the diameter for the individual holes within first row of holes are larger than the diameter for the individual holes within the second row of holes.

16. The secondary fuel nozzle assembly according to claim 15, wherein centers for the individual holes within first row of holes are staggered circumferentially with respect to centers for the individual holes within the second row of holes.

17. A combustor for a Dry Low NO_x(DLN) gas turbine, the combustor comprising:

a secondary fuel nozzle assembly including a premix fuel connection at the rearward end of the secondary fuel nozzle assembly; a nozzle body connected at a rearward end to the fuel connection; a toroidal ring manifold centered radially around the nozzle body for distributing fuel from the nozzle body into an axial airflow path around the nozzle body; a support structure for the toroidal ring manifold; and a communication path for fuel from the nozzle body to the toroidal ring manifold;

a liner circumferentially surrounding the secondary fuel nozzle assembly, including a penetration at the rearward end for admitting airflow to a premixing volume;

the premixing volume of a generally annular shape defined between the secondary fuel nozzle assembly and an inner wall of the liner; and

a swirler mounted at the forward end of the secondary fuel nozzle assembly.

18. The combustor for a DLN gas turbine according to claim 17, the secondary fuel nozzle assembly further comprising:

an internal fuel cavity within the nozzle body for delivering premix fuel from the fuel connection;

a plurality of radial passages from the internal fuel cavity arranged in a predetermined distribution around a circumference of the nozzle body,

a plurality of radial support arms equal in number to the plurality of radial passages from the internal fuel cavity, each support arm including an internal passage between an inner radial end and an outer radial end, the inner radial end being attached to the nozzle body at the location of one of the plurality of radial passages from the internal fuel cavity and the outer radial end being attached to the toroidal ring manifold at the location of the penetrations of the inner toroidal surface;

a predetermined arrangement for a plurality of penetrations through a downstream poloidal surface of the toroidal ring manifold shell including at least one row of holes, with a locus for centers of holes for each row being set at an individual predetermined angle with respect to a ring cross-section of the toroidal ring manifold; and

a toroidal radius of the toroidal ring manifold shell being set to a predetermined value to align the at least one row of holes radially within the premixing volume between the secondary fuel nozzle assembly and the liner.

19. The combustor for a DLN gas turbine according to claim 18, wherein the predetermined arrangement for the plurality of penetrations through a downstream poloidal surface of the toroidal ring manifold shell comprises:

20. The combustor for a DLN gas turbine according to claim 19, wherein the diameter for the individual holes within first row of holes are larger than the diameter for the individual holes within the second row of holes, and the centers for the individual holes within first row of holes are staggered circumferentially with respect to centers for the individual holes within the second row of holes.