• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The present invention relates to a gas turbine with a one-stage fuel injection, wherein the gas turbine (10) has a plurality of combustion chambers. Each of the combustion chambers has a plurality of outer fuel nozzles disposed about a longitudinal axis of the combustion chamber and operable to inject fuel for ignition in a primary reaction zone of the combustion chamber. A central fuel nozzle is disposed substantially along the longitudinal axis of the combustion chamber and is configured to inject fuel that flows downstream from the primary reaction zone into another reaction zone before it ignites. The central nozzle is of a physical nature such that the flame is stabilized at an equivalence ratio greater than 0.65 and is destabilized at an equivalence ratio of up to 0.65. The physical nature of the central fuel nozzle differs from the outer fuel nozzles, and the central fuel nozzle is designed to minimize turbulence and flow recirculation of the fuel delivered by the central fuel nozzle as compared to the outer fuel nozzles.
    公开号:CH703230B1
    申请号:CH01263/11
    申请日:2011-07-28
    公开日:2016-11-30
    发明作者:Lee Vandervort Christian；Meier Haynes Joel
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    Background to the invention
    The disclosed embodiments of the invention relate generally to gas and liquid fuel turbines that include tube and ring combustion systems.
    The DLN technology (so-called "dry-low-NOx technology") is routinely used to influence emissions in industrial gas turbines with gaseous fuel combustion and with tube-ring combustion systems by premixing fuel and air. The main advantage of premixing is that it provides a uniform rate of combustion, which results in relatively constant reaction zone temperatures. With careful air regulation, these temperatures can be optimized to produce very low emissions of nitrogen oxides (NOx), carbon monoxides (CO) and unburned hydrocarbons (UHC). Modulation of a central premix fuel nozzle can enable the operating range to be expanded by keeping the fuel / air ratio and corresponding reaction rates of the outer nozzles relatively constant while the fuel supply to the turbine is varied.
    The multi-stage fuel injection, hereinafter also referred to as fuel staging, is known to those skilled in the art as a means of achieving higher turbine inlet temperatures with uniform heat output. Axially staged systems involve multiple levels of fuel injection along the combustor flow path. The use of axial fuel staging requires special design considerations to inject the fuel into the high temperature combustion products. The high temperature and pressure in the vicinity of the latter stages of an axially staged combustion chamber have prevented the development of stable structures suitable for commercial use.
    The object on which the present invention is based is to provide a new gas turbine which, with a single-stage arrangement, achieves a fuel staging in order to achieve lower maximum combustion temperatures. Such gas turbines have the advantage that they have correspondingly low NOx and CO emissions.
    Brief description of the invention
    [0005] This object is achieved by the subject matter of independent claim 1. Advantageous developments of the present invention are the subject matter of the dependent claims.
    Brief description of the drawings
    These and other features, aspects and advantages of the present invention will be better understood from reading the following detailed description with reference to the accompanying drawings, in which like reference numbers characterize like parts throughout, wherein: FIG. 1 is an illustration a combustor operability or flame stability for a gas turbine combustion system; 2 is a graph of the stoichiometric ratio of fuel and air (X-axis) versus NOx levels at 15% O2 (Y-axis) showing the benefit of lean burn late; Fig. 3 shows the ranges of flame stability for a premixed combustion system, the range "1" being the range in which conventional fuel nozzles are unable to stabilize a flame (conventional lean output), the range "2" being the Area is where the improved fuel nozzle is unable to stabilize a flame (extended lean output) and area "3" is an area where all fuel nozzles can stabilize a flame; 4 is a schematic cross-sectional view of a can and annulus combustor of a turbine in accordance with an embodiment; FIG. 5 is a schematic front view of an end cover and fuel nozzle assembly in accordance with an embodiment; 6 is a schematic cross-sectional view of an outer fuel nozzle in accordance with some embodiments; 7 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 8 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 9 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 10 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 11 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 12 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 13 is a schematic cross-sectional view of a central fuel nozzle in accordance with an embodiment; 14A is an illustration of flame shapes in a conventional tubular and ring combustor; 14B is an illustration of flame shapes for a tubular ring combustor in accordance with an embodiment; and FIG. 14C is an illustration of flame shapes for a tubular and ring combustor in accordance with one embodiment.
    Detailed description of the invention
    Before beginning to explain some embodiments of the invention in detail, it should be noted in this context that the various embodiments of the invention are not limited to their application in terms of the details of the construction and the arrangement of the components, as explained in the following description or shown in the drawing.
    The terms "first", "second", and the like, as used herein, are not intended to imply any order, quantity, or relevance, but rather are used to distinguish one element from another. The indefinite articles (a, an, an, ...) are not intended to imply a restriction on the quantity, but rather to indicate the presence of at least one item to which the indefinite article relates. The term "about" when used with a quantity includes the specified value and has the meaning determined by the context (for example, it includes the error range when measuring a certain quantity).
    Reference to "an embodiment" means throughout the description that a particular feature, structure, or condition that is described in connection with an embodiment is included in at least one embodiment. Therefore, the appearances of the expression “in one exemplary embodiment” in various places in the description do not necessarily mean the reference to the same exemplary embodiment. Furthermore, certain features, structures or textures can be combined in any suitable manner in one or more exemplary embodiments.
    A gas turbine is provided which uses a combustion staging according to the invention with single-stage fuel injection in order to achieve very low emissions during the combustion of gaseous fuel.
    Here, a fuel nozzle assembly has a physical nature so that stabilization of the flame is partially avoided, so the gas turbine can operate without the use of downstream fuel injection. The desired low emissions are thereby provided.
    Figure 1 is a graphical representation of flame stability for a conventional gas turbine combustion system. As shown, flame stability is a function of the fuel-to-air ratio and air flow. There is an area of stable combustion, the size of which can potentially be affected by several variables, including the type of fuel. The fuel nozzle arrangement of the gas turbine as proposed herein is physically such that for some of the fuel nozzles the range of stable combustion is reduced and the range of flame stability is shifted to higher values.
    Avoidance of flame stabilization in turn allows unburned fuel to propagate downstream, beyond the primary reaction zone of the adjacent fuel nozzles. This means that a flame supported by the present nozzle will not burn on the spot, but within the combustion chamber zone of the arrangement and / or the combustion chamber. The result is comparable to that achieved with axial fuel staging without the conventional requirement of injecting fuel downstream.
    The advantages of axial fuel staging or late lean injection for NOx emissions by a premixed flame are graphically illustrated in FIG. The solid line shows the conventional case of the ratio of NOx relative to the fuel-air-stoichiometric ratio (hereinafter also the equivalence ratio), while the ratio of NOx relative to the equivalence ratio in nozzles, arrangements and combustion chambers with axial fuel staging is shown in dotted lines (by experts this is sometimes referred to as lean fuel injection). As shown, there is an extended operating range for fuel-air ratios, which is already able to work with the required NOx emissions. The late or delayed introduction of some of the fuel enables the entire flame zone to be expanded, which in turn results in a lowering of the peak temperatures and a reduction in NOx emissions. However, it has not yet been possible to provide an expedient means of arranging the downstream fuel nozzles or the late injection fuel nozzles in the path of the high temperature combustion gas. In the gas turbine according to the invention, this extended operating range is made available in that the late injection can be avoided by the physical nature of the central nozzle, while the same effect is achieved. However, while the present nozzle is able to provide the benefits of late lean injection without requiring such a configuration, the nozzle is also able to use high fuel-air ratios, i.e., greater than 0.65 for operating modes with low power.
    Figure 3 is a graph of NOx emissions versus fuel to air ratio. The right portion of the graph shows the normal range of lean output for a premix fuel nozzle. The central area depicts an area of expanded lean output that can be achieved through the use of embodiments of the present central fuel nozzle. The left area shows an area in which flame stabilization at the central fuel nozzle cannot occur due to insufficient fuel flow or a fuel-air ratio that is far too low.
    Therefore, a nozzle is provided herein which, according to the requirements, is part of a combustion chamber arrangement which is present in a tube-ring configuration in an industrial gas turbine. The present nozzles, assemblies, and combustion chambers are advantageously set for low to moderate fuel-air equivalence ratios, for example, fuel-air equivalence ratios less than 0.65, such as are typically used in high power modes.
    Figure 4 is a schematic cross-sectional illustration through one of the combustors of a turbine, including a tubular and ring combustor configuration. The gas turbine 10 includes a compressor 12 (partially shown), a plurality of combustors 14 (one shown), and a turbine, here illustrated by a single vane 16. Although not explicitly shown, the turbine is in driving connection with the compressor 12 along a common axis. The compressor 12 pressurizes inlet air, which then flows in the reverse direction to the combustion chamber 14, where it is used to cool the combustion chamber 14 and provide air for the combustion process.
    As indicated above, the gas turbine contains a plurality of combustion chambers 14 which are arranged in the circumferential region of the gas turbine. A double wall transition duct 18 connects the outlet end of each combustor to the inlet end of the turbine to deliver the hot products of combustion to the turbine. The ignition in the various combustion chambers 14 is achieved by a spark plug 20 in connection with flashover tubes 22 in a known manner.
    Each combustor 14 includes a substantially cylindrical combustor housing 24 that is attached to the turbine housing 26 via bolts 28 at an open front end. The end of the combustor housing remote from the turbine (also referred to as the rear) is closed by an end cover assembly 30 which includes supply pipes, manifolds and associated valves for supplying gaseous fuel, liquid fuel, air and water to the combustor 14, as will be described in more detail below . The end cover assembly 30 receives a plurality (e.g., 3 to 6) of "outer" fuel nozzles 32 (one is shown) and a central nozzle 33, with the "outer" fuel nozzles 32 being arranged in an annular arrangement about the longitudinal axis of the combustion chamber 14 .
    In the combustion chamber housing 24, a substantially cylindrical flow sleeve 34 is arranged essentially concentrically thereto, which is connected with its front end to the outer wall 36 of the double-walled transition channel 18. The rear end of the flow sleeve 34 is connected to the combustion chamber housing 24 by means of a radial flange 35 at a butt joint 37, at which the front and rear sections of the combustion chamber housing 24 are connected.
    A combustion chamber wall 38 is arranged concentrically in the flow sleeve 34 and is connected with its front end to the inner wall 40 of the transition channel 18. The rear end of the combustion chamber wall 38 is supported by a combustion chamber wall cap arrangement 42, which in turn is supported in the combustion chamber housing by a plurality of struts 39 and an associated fastening arrangement. The outer wall 36 of the transition duct 18 and this part of the flow sleeve 34, which extend forward from the point at which the combustion chamber housing 24 is connected to the turbine housing by bolts, are provided with an arrangement of openings 44 on their respective peripheral surfaces to allow the air to reverse flow from the compressor 12 through the openings 44 into the annulus between the flow sleeve 34 and the combustion chamber wall 38 toward the upstream or rear end of the combustion chamber (as shown by the flow arrows in FIG. 4).
    The combustor wall cap assembly 42 carries a plurality of premix tubes 46, one for each "outer" fuel nozzle 32 and another for the central nozzle 33. More specifically, each premix tube 46 is within the combustor cap assembly 42 at its front and rear ends by front and rear plates 47 and 49, respectively, which are each provided with openings which are aligned with the premixing tubes 46 having open ends. The front plate 47 (a baffle plate which is provided with a series of cooling openings) can be shielded from the thermal radiation of the combustion chamber flame by shielding plates (not shown).
    The rear plate 49 is mounted on a plurality of rearwardly extending slide collars 48 (one for each premix tube 46, arranged substantially in alignment with the openings in the rear plate), each carrying an air swirler 50 which has a radially outermost wall of the surrounding nozzles concerned. The arrangement is such that the air flowing in the annular space between the combustion chamber wall 38 and the flow sleeve 34 is forced to reverse the direction again in the rear end of the combustion chamber (between the end cover arrangement 30 and the sleeve opening 44) and through the swirl generator 50 and the premixing tubes 46 to stream. The structural details of the combustor wall cap assembly 42, the manner in which the combustor wall cap assembly is retained in the combustor housing, and the manner in which the premix tubes 46 are retained in the combustor wall cap assembly is the subject of U.S. Patent 5,259,184, incorporated herein by reference and its content is fully incorporated by reference.
    FIG. 5 schematically shows a front-end representation and a fuel nozzle arrangement of an exemplary embodiment of an end cover arrangement of a tubular ring combustion chamber, as is shown in FIG. As discussed above, outer fuel nozzles 32 and a central nozzle 33 are attached to end cover 30. The end cover 30 includes internal passages that supply gaseous and / or liquid fuel, water, and air to be atomized to the nozzles, as detailed below. The lines and pipes for supplying the various fluids are in turn connected to the outer surface of the end cover assembly.
    The outer fuel nozzles 32 may be configured to deliver premixed gaseous fuel, liquid fuel, injected water, atomized air, and / or distributed fuel. In some embodiments, the outer fuel nozzle assemblies 32 and the central nozzle 33 are configured to supply premixed gaseous fuel.
    Referring to Fig. 6, each outer fuel nozzle 32 includes a rear supply section 72 with inlets for receiving liquid fuel, injected water, air to be atomized, and gaseous fuel to be premixed, and with suitable connecting passages to supply any of the above fluids. As mentioned above, the outer fuel nozzles 32 are each designed to receive gaseous fuel and to supply it to a relevant passage in a front or remote supply section 74 of the fuel nozzle arrangement. The outer fuel nozzles 32 can be set up to be oriented essentially parallel to the longitudinal axis (axis of symmetry) of the central fuel nozzle 33, or they can run outwardly inclined relative to this axis so that their flames are inclined with respect to the combustion chamber wall. Such a configuration allows the central nozzle fuel to travel further downstream prior to ignition. Although the exact angle is not critical as long as the above-described goal is achieved, the angle of inclination can be limited by the combustion chamber wall. Appropriate angles of inclination with respect to the longitudinal axis of the central fuel nozzle 33 are to be expected in the range from approximately 1 degree to approximately 7 degrees.
    In the illustrated embodiment, the front supply section of the outer fuel nozzle 32 contains a number of concentric tubes. The tubes 76 and 78 describe at least one gas passage 80 which receives premixed gas-fuel from the premixed gas-fuel inlet 82 in the rear supply section 72 via line 84. The premix gas passages 80 communicate with a plurality of radial fuel injectors 86 (FIG. 5), each of which are provided with a plurality of injection ports or holes 88 for delivering premix gas fuel into the premix zone within the premix tube 46. The injected fuel mixes with air flowing in the reverse direction from the compressor 12.
    A second passage 90 is formed between the concentric tubes 78 and 92 and is used to bring atomizing air from the atomizing air inlet 94 via the orifice 96 into the combustion zone 70 of the combustion chamber 14. A third passage 98 is formed between concentric tubes 92 and 100 and is used to bring water from water inlet 102 into combustion zone 70 to effect NOx reductions in a manner known to those skilled in the art.
    The tube 100 itself, which is the innermost of a series of concentric tubes for forming the outer fuel nozzle 32, forms a central passage via the liquid fuel inlet 106. The liquid fuel exits the nozzle via an outlet opening 108 in the center of the outer fuel nozzle 32 .
    Thus, all of the outer fuel nozzles 32 and the central fuel nozzle 33 provide premixed gaseous fuel. The central fuel nozzle 33, but not the outer fuel nozzles 32, provide passive air purging, and each of the outer fuel nozzles 32, but not the central fuel nozzle 33, is configured to provide liquid fuel, emission control water, and atomized air. A number of quaternary pins (not shown) are circumferentially positioned around the front combustor housing 24 to distribute fuel through eight holes per pin.
    The central fuel nozzle 33 has a physical constitution that minimizes turbulence and flow recirculation, so that the flame stability is insufficient. The central fuel nozzle 33 is therefore able to achieve flame destabilization at equivalence ratios below 0.65. Non-limiting examples of physical properties that enable such capability of the central fuel nozzle 33 include one or more aerodynamic features such as a streamlined nozzle tip, a nozzle tip air purge, a streamlined swirler, dual swirlers, dual counter-rotation swirlers, swirl-nozzle nozzles. Combinations, inlet flow conditioners, premix tube exit nozzle funnels and / or diverging premix tube walls.
    For example, in some embodiments, the central fuel nozzle 33 may be provided with a streamlined tip, alone or in combination with a nozzle tip air purge, which both cools the rear portion of the nozzle and removes the remaining recirculation zones. As a result, the flame has difficulties in staying in the region of the nozzle tip, that is to say that the central fuel nozzle 33 has a reduced flame stability compared to a conventional central fuel nozzle. And thus, premixed fuel discharged from the central fuel nozzle 33 flows downstream prior to ignition. The result is comparable to the effect of the axial fuel staging, but advantageously does not require any downstream fuel injection.
    In some embodiments, the central fuel nozzle 33 can have any number of swirlers in any configuration. For example, the central fuel nozzle can be provided with a streamlined swirler, dual swirlers, dual counter-rotating swirlers, a swirl generator combined with a nozzle or fuel injection pin 124, and so on. Any of these swirl generators can be provided for the rotational or counter-rotational flow of fluids discharged there and can serve to destabilize the flame present at the tip of the central fuel nozzle 33. The central fuel nozzle 33 can also be provided with a premixing tube with a “bell-shaped” outlet. Inlet flow conditioners can also be used to achieve the desired flame destabilization. Any of these measures can be used alone or in any combination. Several embodiments of such configurations of the central fuel nozzle 33 are shown in FIGS. 7-13.
    An embodiment of the central fuel nozzle 33 is shown in FIG. As shown, the central fuel nozzle assembly 33 includes a proximal end or rear supply section 52 having a passage 56 extending through the central nozzle assembly 33 for receiving a passive air purge. In operation, the inlet 54 is configured to receive air from the compressor discharge section 114 via an extraction connection 112, both of which are shown in FIG. 4. The central passage 56 passively supplies air to the combustion zone 70 of the combustion chamber 14 (FIG. 4) via the nozzle tip air purge orifices 58 formed at the foremost end 60 of the central fuel nozzle 33. With respect to a turbine 10, the distal or forward discharge end 60 of the central fuel nozzle 33 is located in the premix tube 46 and close to the combustion zone 70.
    Inlets 62 are formed in the rear supply section 52 of the central fuel nozzle 33 to premix gas-fuel. The premix gas passage 64 or passages 64 are in communication with a plurality of radial fuel injectors 66 (FIG. 5) each provided with a plurality of fuel injection ports or holes 68 for dispensing premix gas fuel into a premix zone , which is arranged in the premix pipe 46.
    8 and 9 show two additional exemplary embodiments of the central fuel nozzle 33. More precisely, the central fuel nozzle 33 in the exemplary embodiments shown in FIGS cool and prevent a flame from adhering to it. The exemplary embodiments shown in FIGS. 8 and 9 also use swirl generators in order to destabilize the flame, the exemplary embodiment according to FIG. 8 showing a single streamlined, annular swirl generator 118 and the exemplary embodiment according to FIG. 9 using a dual annular swirl generator 118.
    10 shows an additional exemplary embodiment of a central fuel nozzle 33, the premixing tube 120 being provided with a bell-shaped outlet 122. It is believed that providing a premix tube 120 with such an outlet can reduce turbulence and flow recirculations, which in turn can improve flame stability. 11 shows an embodiment of a central fuel nozzle 33 in which a swirl generator 118 and fuel injection pins 124 are combined in order to form a “swirl nozzle” 126. This exemplary embodiment therefore advantageously provides a more aerodynamic configuration with fewer opportunities for the formation of turbulence, for vortex formation or for recirculation. 11 also shows a bell-shaped outlet of a premix tube 120, although, as noted above, this is not necessarily the case and any particular configuration that allows the central fuel nozzle 33 to have a destabilized flame at less than zero fuel to air ratios , 65 can be used alone or in combination with one or more of other such textures.
    Fig. 12 shows a further embodiment of the central fuel nozzle 33, in which an inlet flow conditioner 128 is present in the vicinity of the combined swirl generator 118 and the fuel pins 124 or the swirl nozzle 126. The inlet flow conditioner 128 can be viewed as a flow straightener and serves to provide a uniformly directed inlet flow to the swirl generator or to the swirl nozzle. The advantage is that there is less turbulence, eddy formation or recirculation.
    13 shows an embodiment of a central fuel nozzle 33 in which the premixing tube 120 is provided with a bell mouth 122, the bell mouth 122 diverging conically with respect to the longitudinal axis of the central fuel nozzle 33.
    Flame shapes for conventional fuel nozzle arrangements in comparison to that according to the invention are shown in FIGS. 14A to 14C. 14A shows a conventional single stage combustor system with later lean burn and a stabilized flame on all fuel injectors. In contrast to this, FIG. 14B shows a fuel nozzle arrangement according to the invention which comprises exemplary embodiments of the present central fuel nozzle and shows a destabilized flame at the central fuel nozzle, which only ignites further downstream from the central fuel nozzle. 14C shows another embodiment in which the outer fuel nozzles are inclined outwards, which before ignition causes convection of unburned fuel even further downstream.
    The turbine operates on gaseous or liquid fuel in a number of modes. In the first mode of operation, premixed gas fuel is provided to two of the outer fuel nozzles 32 and the central fuel nozzle 33 in order to accelerate the turbine. From the ignition and the completion of the ignition rollover and up to about 95% of the speed, the flow of premixed fuel to the central fuel nozzle 33 is switched off, and this portion of the fuel is diverted to two of the outer fuel nozzles 32. From about 95% of the speed and During operation with a low load, the flow of the premixed fuel to the outer fuel nozzles 32 is switched off, and this portion of the fuel for premixing the gaseous fuel is delivered to the central fuel nozzle 33. As the load on the unit is increased further, premixed gaseous fuel is provided to two of the outer fuel nozzles 32 and to the central fuel nozzle 33. From about 20% of the load, the flow from two of the outer fuel nozzles is distributed to three of the outer fuel nozzles 32, while the flow through the central fuel nozzle 33 is maintained. From about 30% of the load, the flow of premixed gaseous fuel through the central fuel nozzle 33 is switched off and this portion of the premixed gas fuel is released through the two of the outer fuel nozzles so that all the outer fuel nozzles 32 emit premixed gaseous fuel. For a short period of time, the fuel is delivered exclusively to the outer, so-called quaternary premixing nozzles. The central fuel nozzle 33 is then operated again to dispense premixed gaseous fuel through the premixed gas-fuel passages 64. This operating mode is used with controlled fuel proportions for the premix gas nozzles up to 100% of the load ratio. In order to work in the first and the last two operating modes, the central fuel nozzle must generally be able to stabilize the flame at equivalence ratios greater than 0.65.
    A gas turbine is provided with an improved central fuel nozzle. The central fuel nozzle has destabilizing flame holding characteristics, that is, the nozzle is unable to stabilize a flame up to an equivalence ratio of about 0.65. It follows that the heat given off by the flame is delayed, as a result of which lower maximum flame temperatures and correspondingly lower NOx levels are achieved. The flame stabilization capability is retained for higher equivalence ratios to support the operation of the combustion chamber in other sections of the load range.
    List of reference symbols
    10 gas turbine 12 compressor 14 combustion chambers 16 blades 18 transition duct 20 spark plug 22 flashover tubes 24 combustion chamber housing 26 turbine housing 28 bolt 30 end cover assembly 32 outer fuel nozzle 33 central fuel nozzle 34 flow sleeve 35 flange connection 36 outer wall 37 butt joint 38 combustion chamber wall 39 struts 40 inner wall 42 combustion chamber wall cap arrangement 44 Openings 46 Premix Tube 47 Back Plate 48 Movable Collar 49 Back Plate 50 Swirler 52 Back Supply Section 54 Passive Purge Air Inlet 56 Passage 60 Front End 62 Inlets 64 Premix Gas Passages 68 Fuel Injection Port 70 Combustion Zone 72 Rear Supply Section 74 Remote Discharge Section 76 Pipe 78 Pipe 80 Premix Gas Passage 82 Premix Gas Passage 82 Fuel 84 Line 86 Radial Fuel Injectors 88 Fuel Injector Orifices 90 Passage 92 Tube 94 Atomizing Air Inlet 96 Orifice 98 Passage 100 Tube 102 Water inlet 106 liquid fuel inlet 108 discharge port 112 extraction port 113 compressor discharge section 114 nozzle tip air purge port 116 nozzle tip 118 swirl generator 120 premix tube 122 bell-shaped outlet 124 fuel injection pin 126 swirl nozzle 128 inlet flow conditioner
    权利要求:
    Claims (5)
    [1]
    1. A gas turbine with a one-stage fuel injection, wherein the gas turbine (10) comprises:a plurality of combustors (14) disposed about a circumference of the gas turbine (10) to form a fuel injection stage, each of the combustors (14) comprising:a plurality of outer fuel nozzles (32) disposed about a longitudinal axis of the combustion chamber (14) and operable to inject fuel for ignition in a primary reaction zone of the combustion chamber (14);a central fuel nozzle (33) disposed substantially along the longitudinal axis of the combustion chamber (14) and configured to inject fuel that flows downstream from the primary reaction zone into another reaction zone before it ignites;the central nozzle (33) is of a physical nature such that the flame is stabilized at an equivalence ratio greater than 0.65 and destabilized at an equivalence ratio of up to 0.65,wherein the physical nature of the central fuel nozzle (33) is different from the outer fuel nozzles (32) and the central fuel nozzle is configured to reduce turbulence and flow recirculation of the fuel delivered from the central fuel nozzle (33) compared to the outer fuel nozzles (32 ) minimized.
    [2]
    2. Gas turbine according to claim 1, wherein the central fuel nozzle (33) comprises:a rearward supply section (52), a forward end (60), and a central passageway (56) interconnecting the rearward supply section (52) and forward end (60), the forward end (60) having nozzle tip air purge ports (114) to passively supply air from the central passage (56) to cool the tip (116) of the central fuel nozzle (33) and to prevent the adhesion of a flame to the tip (116).
    [3]
    3. The gas turbine of claim 2, wherein the central fuel nozzle (33) further comprises a premix tube (46) at least partially surrounding the central fuel nozzle (33), and wherein the rear supply section (52) includes inlets (62), premix gas passages (64). for receiving premix fuel through the inlets (62) and radial fuel injectors (66) to deliver premixed gas fuel into a premix zone in the premix tube (46).
    [4]
    4. Gas turbine according to claim 1, wherein the central fuel nozzle (33) comprises:at least one streamlined annular swirler (118), an inlet flow conditioner (128), and / or a premix tube (46) having a bell-shaped diverging exit (122).
    [5]
    5. A gas turbine according to any one of claims 1 to 4, wherein each of the outer fuel nozzles (32) with respect to the longitudinal axis of the combustion chamber (14) outwardly in the direction of a combustion chamber wall (38) is inclined.
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    同族专利:
    公开号 | 公开日
    US20120023952A1|2012-02-02|
    CH703230A2|2012-01-31|
    CN102345879A|2012-02-08|
    JP5925442B2|2016-05-25|
    US9557050B2|2017-01-31|
    JP2012032144A|2012-02-16|
    DE102011052159A1|2012-02-02|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2021-02-26| PL| Patent ceased|
    优先权:
    申请号 | 申请日 | 专利标题
    US12/847,688|US9557050B2|2010-07-30|2010-07-30|Fuel nozzle and assembly and gas turbine comprising the same|
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