• 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 subject matter provides a gas turbine having a system for increasing static pressure recovery downstream of a compressor section (14) of the gas turbine and for reducing gas turbine emission generation. The system includes a bleed diffuser (52) positioned downstream from the compressor section (14) of a gas turbine, and a bleed passage (54) extending from the bleed diffuser (52). The bleed passage (54) is configured to direct air bleed from the bleed diffuser (52) to a secondary combustion system (58). The air flowing into the secondary combustion system (58) is mixed with fuel to form an air / fuel mixture.
    公开号:CH702612B1
    申请号:CH00127/11
    申请日:2011-01-25
    公开日:2016-04-15
    发明作者:Dwight Berry Jonathan；John Hughes Michael
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    Field of the invention
    The present subject matter generally relates to gas turbines, and more particularly to a gas turbine having a system for modifying the static pressure recovery and emission generation of the gas turbine.
    Background of the invention
    To improve the efficiency of combustion of fuel and air within a combustor, gas turbines usually include a diffuser designed to reduce the velocity of the pressurized airflow exiting the compressor section of the gas turbine and the static ones To increase pressure of this pressurized air flow. A diffuser may generally include at least one divergent diffuser wall that allows the pressurized airflow to dissipate or spread across the length of the diffuser. However, when the pressurized air flow passes through the diffuser, the friction along the diffuser wall (s) creates a boundary layer in which the velocity of the air flow is significantly less than the velocity of the main air flow. Thus, the formation of a boundary layer may cause an airflow to enter the combustor section of a gas turbine that exhibits an uneven velocity profile. This can adversely affect combustion in the combustion chambers and reduce the overall efficiency of a gas turbine. In addition, significant flow losses can result when the boundary layer separates from the diffuser wall, which can occur when the divergence angle of the diffuser wall or walls is too wide. Consequently, diffusers usually need to be relatively long in order to obtain the required static pressure recovery without causing boundary layer separation.
    To overcome these problems, bleed diffusers are known, the air that flows adjacent to the diffuser wall or the diffuser walls, tap or remove from the main air flow. In particular, a bleed diffuser may be used to reduce the boundary layer size by diverting the entire boundary layer or a portion thereof away from the main air flow. This can reduce the likelihood of flow losses due to boundary layer separation and also provides a shorter diffuser which can accommodate wide divergence angles to allow a significant increase in static pressure recovery downstream of the compressor section. As a result, however, the improved behavior of a bleed diffuser is often offset by the reduction in overall efficiency and overall performance of the gas turbine caused by bleed of compressed air from the main air flow. Specifically, tapped portions of the air leaving the compressor section reduce the amount of compressed air available for cooling turbine components or increasing the turbine inlet pressure.
    Accordingly, a system in a gas turbine that provides the benefits of a bleed diffuser without the loss of efficiency and performance caused by bleeding compressed air from the main air flow would be welcomed in the art.
    Brief description of the invention
    The invention provides a gas turbine according to claim 1.
    These and other features, aspects, and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present subject matter and, together with the description, serve to explain the principles of the present subject matter.
    Brief description of the drawing
    A complete and an implementation enabling disclosure of the present subject matter, including the best mode thereof, which is directed to a person skilled in the art, is given in the description which refers to the accompanying figures, in which:<Tb> FIG. 1 <SEP> is a cross-sectional view of a part of a gas turbine;<Tb> FIG. 2 <SEP> is a cross-sectional view of one embodiment of a system for increasing static pressure recovery and reducing emissions in a gas turbine in accordance with aspects of the present subject matter;<Tb> FIG. 3 is a greatly enlarged view of an embodiment of a bleed diffuser according to aspects of the present subject matter;<Tb> FIG. FIG. 4 is a sectional view of FIG. 2, particularly illustrating embodiments of a bleed passage, a bracket, and a secondary combustion system in accordance with aspects of the present subject matter; FIG.<Tb> FIG. 5 is a greatly enlarged view of a portion of FIG. 4, particularly illustrating one embodiment of a secondary injector fuel injector in accordance with aspects of the present subject matter;<Tb> FIG. 6 is a cross-sectional view of another embodiment of a system for increasing static pressure recovery and reducing emissions in a gas turbine in accordance with aspects of the present subject matter;<Tb> FIG. FIG. 7 is a sectional view of FIG. 6, particularly illustrating embodiments of a bleed passage, a bracket, and a secondary combustion system in accordance with aspects of the present subject matter. FIG.
    Detailed description of the invention
    [0008] Reference will now be made in detail to embodiments of the present subject matter, one or more examples of which are illustrated in the drawings. Each example is provided for purposes of illustration and not limitation of the present subject matter.
    Referring to FIG. 1, there is illustrated a simplified drawing of a portion of a gas turbine engine 10. The gas turbine engine 10 includes a compressor section 14 that is configured to pressurize air flowing into the turbine 10. The compressed air discharged from the compressor section 14 flows into the combustion chamber section 16 where the air is mixed with a fuel and burned. Hot combustion gases flow from the combustor section 16 to a turbine section 18 to drive the gas turbine 10 and produce power.
    The compressor section 14 is generally characterized by a series of vanes and blades 20 which are used to compress air flowing into the gas turbine engine 10. The pressurized airflow exiting the compressor section 14 passes through a diffuser 22 defined by the compressor exhaust housing 24 before entering the combustor section 16. The diffuser 22 may generally include at least one divergent diffuser wall 26 forming a flow path of increasing cross-section for the airflow. As the cross-sectional area of the flowpath increases, the airflow disperses or disperses, reducing the velocity of the airflow and increasing the static pressure of the airflow. The dispersed airflow leaving the diffuser 22 then passes through an annular plenum chamber 28 to the combustor section 16, generally characterized by a plurality of combustors 30 arranged in an annular array about the axis of the machine (of which FIG 1 illustrates only a single one).
    With continued reference to FIG. 1, each combustor 30 in a gas turbine 10 includes a main combustion system 32 for mixing and combusting an air / fuel mixture and a transition piece 34 for flowing hot combustion gases to a first stage nozzle 36 of the turbine section 18 allow. The main combustion system 32 may include a combustor shell 38, an end cap 40, a plurality of premix fuel nozzle assemblies 42, a flow sleeve 44, and a combustor flame tube 46 disposed in the flow sleeve 44. In operation, dispersed air exiting the diffuser 22 passes through the annular chamber 28 and enters each combustion chamber 30 through the flow sleeve 44 and the impingement sleeve 48 of the transition piece 34 where it is fluidized and mixed with a fuel that is introduced into each Fuel nozzle assembly 42 is injected. The air / fuel mixture exiting each fuel nozzle assembly 42 flows into a reaction zone 50 defined by the combustor flame tube 46 where it is burned. The hot combustion gases then flow through the transition piece 34 to the turbine section 18 to drive the turbine 10 and generate electricity. It should be understood, however, that the main combustion system 32 need not be configured in the manner described above and illustrated herein and generally may have any configuration that allows a combustible mixture to be mixed and combusted. Further, it should be understood that the main combustion system 32 may include other components that are not described or illustrated herein.
    In accordance with one aspect of the present subject matter, FIG. 2 illustrates an embodiment of a gas turbine having a system for increasing the static pressure recovery downstream of a compressor section 14 of the gas turbine engine 10 and for reducing the generation of air polluting emissions. The system includes a bleed diffuser 52 positioned downstream of the compressor section 14 and a bleed passage 54 extending from the bleed diffuser 52. The bleed passage 54 is configured to direct bleed air 56 (Figs. 3, 4, 5, and 7) from the pressurized airflow exiting the compressor section 14 to a secondary combustion system 58 downstream of the main combustion system 32 (Fig. 1) is arranged. The bleed air 56 flowing into the secondary combustion system 58 may be mixed with a fuel to form an air / fuel mixture and combusted in a reaction zone 50 of the combustion chamber 30.
    The bleed diffuser 52 according to the present subject matter may be generally configured to disperse the pressurized airflow flowing out of the compressor section 14. In particular, the bleed diffuser 52 is configured to reduce the velocity of the airflow and thereby increase the static pressure of the airflow passing through the annular chamber 28. In one embodiment, the bleed diffuser 52 includes a vortex controlled diffuser (VCD) defined by the compressor exit housing 24. However, it should be understood that the bleed diffuser 52 need not be a VCD, but generally may have any type of bleed diffuser well known to those skilled in the art. In addition, it should be understood that the VCD employed in the present subject matter need not be configured in the manner described below and illustrated herein, but may have any configuration or arrangement.
    It has been found that a VCD can be used in a gas turbine 10 to give significantly higher static pressure retries on a relatively short length. This increased pressure recovery results in an increased pressure difference between the air flowing through the annular chamber 28 and the hot combustion gases flowing in the combustion chamber 30. As this pressure difference increases, the efficiency and performance of a gas turbine 10 can be significantly improved. In particular, an increased pressure differential may minimize the effect of using a portion of the airflow to cool turbine components and also increase turbine inlet pressure. In addition, the use of a VCD by reducing the length of the rotor can reduce the cost associated with manufacturing a gas turbine 10.
    Referring to FIG. 3, there is illustrated a greatly enlarged view of one embodiment of a bleed diffuser 52. As illustrated, the bleed diffuser 52 includes a hybrid VCD including a swirl chamber 60 and a wide angle post-diffuser 62, both defined by the compressor exhaust housing 24. The vortex chamber 60 may include a vortex wall 64 defining a wall of the vortex chamber 60. It should be understood, however, that a VCD may include more than a single vortex chamber 60. For example, a second vortex chamber (not illustrated) may be formed in the compressor discharge housing 24 on the opposite or to the vortex chamber 60 or upstream of the vortex chamber 60. In addition, the bleed diffuser 52 may further include a forward diffuser 66 to increase the static pressure of the airflow upstream of the swirl chamber 60.
    During operation of the illustrated embodiment, the boundary layer formed when the pressurized airflow passes through the pre-diffuser 66 may be drawn from the main airflow by the arrangement of the vortex wall 64 in the vortex chamber 60. This creates a fluid vortex in the vortex chamber 60 and creates a turbulent shear layer downstream of the chamber 60 which prevents boundary layer separation in the wide angle post diffuser 62. The portion of the airflow drawn from the main airflow may then be bleeded as bleed air 56 from the vortex chamber 60 to cool various components of the gas turbine 10 or to be used in a secondary combustion system 58, as in larger ones Details are described below.
    Referring to FIGS. 2 and 4, a bleed passage 54 extends from the bleed diffuser 52 to a secondary combustion system 58 and may be configured to direct bleed air 56 to the secondary combustion system 58. Thus, in the illustrated embodiment, air drawn into the swirl chamber 60 may be diverted from the main airflow through the bleed passage 54. As illustrated, a lower portion of the bleed passage 54 may be defined by a passage 59 formed in the compressor exit housing 24, while an upper portion of the bleed passage 54 may be defined by a bracket 70 mounted to the compressor exit housing 24 , It should be understood that the bracket 70 may be mounted to the compressor exhaust housing 24 by any suitable means. In addition, seals (not shown) may be included between the compressor exit housing 24 and the retainer 70 to adequately seal the flow of the bleed air 56 as it passes through the bleed passage 54. The bracket 70 may also be configured to be mounted to and support the transition piece 34. For example, as illustrated in FIG. 4, the retainer 70 may be mounted to the impingement sleeve 48 of the transition piece 34. One skilled in the art should understand that the retainer 70 may be secured to the transition piece 34 by any suitable means. It should be further understood that any mounting or securing means used to secure the bracket 70 to the transition piece 34 may be configured to accommodate the thermal expansion of the transition piece 34. Further, seals (not shown) may also be included at the junction between the retainer 70 and the transition piece 34 to sealingly trap the flow of the bleed air 56.
    In the illustrated embodiment, the retainer 70 may also be hollow so as to form at least one flow path for the bleed air 56 flowing through the bleed passage 54. As illustrated in FIG. 4, the retainer 70 may be Y-shaped and may include a first flow path 72 and a second flow path 74 defining an upper portion of the bleed passage 54. In addition, it should be understood that the bleed passage 54 may further include a valve (not illustrated), or generally any other means for shutting off the supply of the bleed air 56 flowing through the bleed passage 54 to the secondary combustion system 58.
    As noted above, the bleed passage 54 may be configured to direct the bleed air 56 to a secondary combustion system 58 located downstream of the main combustion system 32 (FIG. 1) in a combustion chamber 30. The secondary combustion system 58 may generally include, for example, at least one fuel injector 76 configured to receive and dispense fuel from a fuel source 80 (FIGS. 4, 5, and 7). In addition, it should be understood that the secondary combustion system 58 may include other components and may generally have any arrangement or configuration. For example, the secondary combustion system 58 may include a late lean injector or a lean direct injector. Preferably, the secondary combustion system 58 may be configured such that fuel or an air / fuel mixture may be introduced into a reaction zone 50 of the combustor 30 and ignited by the hot combustion gases flowing from the main combustion system 32 (FIG. 1). Thus, the secondary combustion system 58 may allow for higher firing temperatures within a combustor 30 while keeping the rate of formation of harmful emissions, such as NOx, at a minimum. This can be accomplished by exhausting a lean air / fuel mixture from the secondary combustion system 58 to ensure that the combustion reaction temperatures remain below the stoichiometric flame temperature. In addition, short residence times within the transition piece 54 can keep thermal NOx formation rates low.
    An embodiment of the secondary combustion system 58 is illustrated in FIGS. 2, 4 and 5. Referring to FIG. 4, the secondary combustion system 58 may include a pair of fuel injectors 76. For example, a first fuel injector 76 may be disposed in the first flow path 72 of the retainer 70, and a second fuel injector 76 may be disposed in the second flow path 74 of the retainer 70, with the combination of the flow paths 72, 74 defining a portion of the bleed passage 54. In addition, the fuel injectors 76 may be in fluid communication with a fuel source 80 so that the injectors 76 may receive and dispense fuel from the fuel source 80. As particularly illustrated in FIG. 5, the fuel injectors 76 may be mounted in the retainer 70 as a separate component. However, it should be understood that the fuel injectors 76 may be formed as an integral part of the retainer 70. It should be further understood that the illustrated fuel injectors 76 have been simplified for illustrative purposes. Thus, fuel injectors 76 having other components and having different configurations may be used in the system of the present subject matter. For example, the fuel injectors 76 may include swirl vanes for imparting rotational movement to the bleed air 56 flowing past the injectors 76.
    During operation of a gas turbine 10 according to an embodiment of the present subject matter, the secondary combustion system 58 is fed with the bleed air 56 flowing through the bleed passage 54. It should be understood that the pressure differential between the bleed air 56 provided by the bleed passage 54 and the combustion products flowing through the combustion chamber 30 drives the bleed air 56 through the secondary combustion system and into the reaction zone 50 of the combustion chamber 30 , As the bleed air 56 bypasses the fuel injectors 56, it is mixed with the fuel discharged from the injectors 76. The air / fuel mixture then passes through bypass tubes 82, each defining a channel for the air / fuel mixture to pass through the impingement sleeve 48 and the transition piece wall 33, and is in a reaction zone 50 by the Main combustion system 32 ignited combustion products.
    An alternative embodiment of the secondary combustion system 58 is shown in FIGS. 6 and 7. The secondary combustion system 58 includes an air manifold 84 disposed between the bracket 70 and the impingement sleeve 48 of the transition piece 34. The air distributor 84 may be configured to receive bleed air 56 flowing from the bleed passage 54. In particular, the air manifold 84 may be configured to receive the bleed air 56 flowing from the first and second flow paths 72, 74 of the retainer 70, both of which define a portion of the bleed passage 54.
    In addition, the secondary combustion system 58 may include a plurality of fuel injectors 76 disposed in the air manifold 84 at a distance from each other. As illustrated in FIG. 7, the secondary combustion system 58 includes five spaced-apart fuel injectors 76 mounted in the annular air manifold 84, each fuel injector 76 configured to receive and dispense fuel from a fuel source 80. However, it should be understood that any number of fuel injectors 76 may be disposed in the air manifold 84. It should be further understood that the fuel injectors 76 may be formed as an integral part of the air manifold 84 or mounted in the air manifold 84 in the form of a separate component. Further, it should be understood that the air distributor 84 need not be annular, but may have any suitable shape. For example, the air distributor 84 may be semicircular.
    The operation of the secondary combustion system 58 illustrated in Figs. 6 and 7 is similar to the embodiment described above. Specifically, bleed air 56 flowing through the bleed passage 54 enters the air manifold 84 where it is mixed with fuel discharged from each of the fuel injectors 76. The air / fuel mixture then passes through the transfer tubes 82 and is ignited in a reaction zone 50 by the combustion products flowing out of the main combustion system 32 (FIG. 1).
    It should also be understood that the present subject matter may include a combustor section 16 disposed downstream of the compressor section 14. The combustor section 16 may include the combustor 30 configured to receive the pressurized airflow discharged from the compressor section 14. In addition, combustor 30 may include a main combustion system 32 configured to mix and combust an air / fuel mixture. A turbine section 18 may be disposed downstream of the combustor section 16 and configured to receive hot combustion products that flow out of the combustor section 16.
    parts list
    [0026]<Tb> 10 <September> Gas Turbine<Tb> 14 <September> compressor section<Tb> 16 <September> combustor section<Tb> 18 <September> turbine section<tb> 20 <SEP> Guide vanes and blades<Tb> 22 <September> diffuser<Tb> 24 <September> compressor outlet housing<Tb> 26 <September> diffuser wall<tb> 28 <SEP> Annular Plenum Chamber<Tb> 30 <September> combustion chamber<Tb> 32 <September> main combustion system<Tb> 33 <September> transition piece wall<Tb> 34 <September> transition piece<tb> 36 <SEP> First stage nozzle<Tb> 38 <September> combustion chamber housing<Tb> 40 <September> end cover<Tb> 42 <September> Vormischbrennstoffdüsenanordnung<Tb> 44 <September> flow sleeve<Tb> 46 <September> combustor liner<Tb> 48 <September> impingement sleeve<Tb> 50 <September> reaction zone<Tb> 52 <September> bled diffuser<Tb> 54 <September> Disc Channel<Tb> 56 <September> bleed air<tb> 58 <SEP> Secondary combustion system<Tb> 59 <September> Continuity<Tb> 60 <September> cyclone<Tb> 62 <September> Nachdiffusor<Tb> 64 <September> shroud<Tb> 66 <September> Vordiffusor<tb> 70 <SEP> Clamp, bracket<tb> 72, 74 <SEP> Flow paths<Tb> 76 <September> fuel injector<Tb> 80 <September> fuel source<tb> 82 <SEP> Transfer tube, connecting tube<Tb> 84 <September> Air Distribution
    权利要求:
    Claims (9)
    [1]
    A gas turbine (10) having a system for increasing the static pressure recovery downstream of a compressor section (14) of the gas turbine (10) and for reducing the emission generation of the gas turbine (10), the system comprising:a bleed diffuser (52) positioned downstream of the compressor section (14) of the gas turbine (10), the bleed diffuser (52) adapted to disperse a pressurized airflow flowing from the compressor section (14);a secondary combustion system (58) disposed downstream of a main combustion system (32) of the gas turbine (10) in a combustor (30) of the gas turbine (10), the secondary combustion system (58) including at least one fuel injector (76) configured to receive and dispense fuel from a fuel source (80);a bleed passage (54) extending from the bleed diffuser (52) to the secondary combustion system (58), wherein the bleed passage (54) is configured to deliver bleed air (56) from the pressurized air flow to the secondary combustion system (58); to lead; andwherein the bleed air (56) flowing into the secondary combustion system (58) is mixed with the fuel discharged from the at least one fuel injector (76) to form an air / fuel mixture.
    [2]
    2. Gas turbine (10) according to claim 1, wherein the tapping diffuser (52) through a compressor outlet housing (24) of the gas turbine (10) is defined.
    [3]
    3. A gas turbine (10) according to any one of the preceding claims, wherein the bleed diffuser (52) comprises a vortex-controlled diffuser, the vortex-controlled diffuser having at least one swirl chamber (60) adapted to pressurize the bleed air (56) Draw in air flow.
    [4]
    4. Gas turbine (10) according to one of the preceding claims, wherein the bleed passage (54) is defined in part by a passage (59) formed in a compressor outlet housing (24) of the gas turbine (10).
    [5]
    A gas turbine (10) according to any one of the preceding claims, wherein the system includes a retainer (70) configured to support a transition piece (34) of the combustor (30), the bleed passage (54) being partially defined by at least one defined by the holder (70) flow path (72, 74) is defined.
    [6]
    6. The gas turbine (10) according to claim 5, wherein the at least one fuel injector (76) in the at least one flow path (72, 74) of the holder (70) is arranged.
    [7]
    The gas turbine (10) of claim 5, wherein the system includes a first flow path (72) and a second flow path (74) formed in the support (70), the secondary combustion system (58) including a first fuel injector (76 ) and a second fuel injector (76), wherein the first fuel injector (76) is disposed in the first flow path (72), the second fuel injector (76) being disposed in the second flow path (74).
    [8]
    The gas turbine (10) of claim 5, wherein the secondary combustion system (58) includes an air manifold (84) disposed between the support (70) and the transition piece (34), the air manifold (84) configured to to receive the bleed air (56) from the tap channel (54).
    [9]
    The gas turbine (10) of claim 8, wherein the secondary combustion system (58) includes a plurality of fuel injectors (76), wherein the plurality of fuel injectors (76) are disposed in the air manifold (84).
    类似技术:
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    同族专利:
    公开号 | 公开日
    US20110179803A1|2011-07-28|
    JP2011153815A|2011-08-11|
    CN102135034A|2011-07-27|
    DE102011000225B4|2021-05-06|
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    CN102135034B|2014-12-24|
    JP5759185B2|2015-08-05|
    DE102011000225A1|2011-07-28|
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    法律状态:
    2017-03-15| NV| New agent|Representative=s name: GENERAL ELECTRIC TECHNOLOGY GMBH GLOBAL PATENT, CH |
    2021-08-31| PL| Patent ceased|
    优先权:
    申请号 | 申请日 | 专利标题
    US12/694,544|US8381532B2|2010-01-27|2010-01-27|Bled diffuser fed secondary combustion system for gas turbines|
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