• 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:
    A combustor (10) for a gas turbine includes a plurality of radially outer nozzles (26) disposed in an annular array, each of the radially outer nozzles having an outlet end arranged to direct fuel and / or air to a first combustion chamber (32 ). A central nozzle (28) has an outlet end disposed axially upstream of the outlet ends of the radially outer nozzles and configured and arranged to supply fuel and air to a second combustion chamber (36) axially upstream of the first combustion chamber. The second combustion chamber (36) opens into the first combustion chamber (32) and has a sufficient length to keep a flame of the central nozzle limited to the second combustion chamber.
    公开号:CH702737B1
    申请号:CH00257/11
    申请日:2011-02-14
    公开日:2016-02-15
    发明作者:Almaz Valeev；Sergey Anatolievich Meshkov；Mark Hadley；Stanley Widener；Geoffrey Myers；Valery Alexandrovich Mitrofanov
    申请人:Gen Electric；
    IPC主号:
    专利说明:

    This invention relates to gas turbine technology, and more particularly to an axially stepped nozzle configuration of a gas turbine combustor that promotes improved CO burnout.
    Background of the invention
    Currently, there is a limit to an otherwise desirable reduction of the exit temperature of combustion gases due to the CO content contained in the combustion gases. In other words, the combustor exit temperature must be kept relatively high to ensure that the CO burnout meets the required CO emission levels. In order to maintain the combustion chamber exit temperature sufficiently high to maintain low CO levels under low load or no load conditions, the customer must either shut down the turbine or keep the turbine "online", even during the periods with low power requirements, so that in this way the fuel consumption amount is increased.
    Thus, there is a need for a mechanism by which the amount of CO generated by combustion in the gas turbine can be reduced, so that the part load capability for the customer can be improved. In particular, if the CO levels in the combustor could be reduced under lower load or no load conditions, customers would consume less fuel during periods of reduced power consumption. This, in turn, would lead to direct fuel economy, but without shutting down the turbine and then having to restart when demand reestablished, thus providing improvements in reliability.
    Brief description of the invention
    The invention provides a combustion chamber according to claim 1 and a method according to claim 8. Further variants of the invention are presented in the dependent claims.
    The invention will now be explained in greater detail in conjunction with the drawings below.
    Brief description of the drawings
    [0006]<Tb> FIG. 1 shows a cross-section through a gas turbine combustor according to a first exemplary, but non-limiting embodiment of the invention;<Tb> FIG. FIG. 2 shows an enlarged partial perspective view of the combustor illustrated in FIG. 1; FIG.<Tb> FIG. Fig. 3 <SEP> is a partial cutaway perspective view of the combustor illustrated in Fig. 2; and<Tb> FIG. 4 shows a schematic representation of a combustion chamber configuration according to another exemplary but not limiting embodiment.
    Detailed description of the invention
    Referring now to FIGS. 1-3, there is illustrated a gas turbine combustor 10 according to an exemplary, but non-limiting embodiment of the invention. It is to be understood that combustor 10 is usually combined with various other similar combustors disposed in an annular array around the gas turbine housing, with each combustor providing combustion gases to the first stage of the turbine. Each combustor 10 is supplied with air from a compressor (not shown). The compressor air is flowed (as indicated by the flow arrows) in the reverse direction into an annular channel 12, which is arranged between a transition piece 14 and a combustion chamber flame tube 16, which are arranged radially inwardly and axially aligned, on the one hand and the radially outer, axially aligned flow sleeves 18th and 20, on the other hand. The compressor air flows into the channel 12 through impingement cooling holes 22, 24 in the respective flow sleeves 18 and 20 so as to achieve cooling at the transition piece and the combustion chamber flame tube before reversing the flow direction at the inlet end of the combustion chamber. Generally, and under certain operating conditions, the air flows into air injectors associated with each of a number of six radially outer nozzles 26 and one middle nozzle 28 (the number of nozzles in the combustion chamber usually varies between 6 and 8) where pre-mixes with fuel supplied to the nozzles via the combustor end cover 30. The air / fuel mixture from the radially outer nozzles 28 is injected into the combustion zone or main combustion chamber 32. Ignition is achieved by spark plugs (not shown) in conjunction with sparkover tubes (also not illustrated) connecting adjacent combustion chambers. Hot combustion gases flow from the combustion chamber 32 into the transition piece 14 and then to the first stage of the gas turbine, which is represented by a single vane 34. Up to this point, the combustor as described is generally well known, the invention being the arrangement of the central nozzle 28 with respect to the radially outer nozzles 26 and 30 and the establishment of a second or (primary) combustion chamber 36 upstream of the latter first combustion chamber (or main combustion chamber) 32 is concerned.
    More specifically, and continuing to refer particularly to Figs. 2 and 3, the center nozzle 28 is reset in an upstream direction (relative to a flow direction of the combustion gases from left to right in the various figures). In other words, the middle nozzle 28 is disposed axially behind the outlets of the radially outer surrounding nozzles 26. A combustor cap 38 supports the outlet ends of the outer nozzles, but is configured and mounted to allow compressor air to flow between the cap and the housing wall 40 (Figure 1). A substantially cylindrical tubular member 42 extends rearward from the cap 38 to the outlet end of the central nozzle 28 and thus forms the primary combustion chamber 36 which opens at the foremost plate 44 of the cap 38 into the main combustion chamber 32. The length of the space 36 is determined to be sufficient to allow complete combustion of CO while protecting the flame of the center nozzle from the cold ambient air entering the main space 32 via the radially outer nozzles 26.
    Fuel is supplied to the radially outer nozzle tubes (two of which are illustrated at 46 (FIG. 1) and the central nozzle tube 48 through the end cap 30, as noted above, while air is conventionally configured at the radially outer nozzles 26 at inlets 50 Premix swirlers (two of which are illustrated in FIG. 3) and the center nozzle 28 are fed through a premix swirl inlet via openings 52 in the radial vane 54.
    In low load ranges up to full speed without load (FSNL, full-speed no-load), the fuel is fed only to the central nozzle 28, while air flows through the radially outer nozzles 26. By limiting the flame of the central nozzle to the primary combustion chamber 36, this is protected from the cold air supplied by the radially outer nozzles 26 and is thus not subjected to any undesirable temperature drop. As a result, by holding the flame of the middle nozzle at a high temperature and with sufficient fuel volume for the middle nozzle 28, the flame of the middle nozzle will burn out the resident CO. The reduction in CO levels, in turn, allows the turbine operator to lower the gas turbine even further in terms of load, with associated reduced fuel consumption when power requirements are low.
    When the load is increased, a point comes at which the amount of fuel required for combustion is greater than can be absorbed by the central nozzle 28. Then, the radially outer nozzles 26 are used, wherein the fuel supplied to the radially outer nozzles is mixed with the combustion material supplied by the compressor, as described above. The combustion flames connected to the outer nozzles 26 are anchored downstream of the primary combustion chamber 36 within the main combustion chamber 32. The radially outer nozzles 26 may be "fired" or ignited simultaneously or in any predetermined order (or simultaneously, for example, in groups of two or three), as may be dictated by optimization of combustion for particular combustor applications.
    In any event, the flame of the central nozzle at FSFL remains anchored in the primary combustion chamber 36, while the flames of the outer nozzles remain anchored in the main combustion chamber 32 downstream of the primary combustion chamber 36. Because the tubular member 42 defining the primary combustion chamber 36 is directly exposed to the flame of the central nozzle, it must be removed by any suitable means, such as applying a thermal barrier coating, impact testing, adding turbulators, or a combination of the foregoing. be cooled.
    In an optimized application of the invention to a specific turbine model, one-third (1/3) of the combustion air flows through the central nozzle while two thirds (2/3) flow through the array of outer nozzles, with a phi ratio of about 0.6 (where Phi is an equivalence ratio defined by the ratio of the actual fuel / air ratio to the stoichiometric value). Typical Phi values are in the range of 0.50 to 0.65.
    In an alternative mode of operation at FSFL, the flame in central nozzle 28 may be extinguished for a relatively short time and subsequently re-fueled so that the flame is reignited (and held) downstream of primary combustion chamber 36. By re-igniting the flame of the central nozzle in the main combustion chamber 40 and keeping it away from the primary combustion chamber 36, the temperature of the tubular element 42 will be cooler, and the mixing zone for the fuel and air supplied to the central nozzle 28 will become extended, resulting in better mixing and lower NOx emissions. In this alternate FSFL mode of operation, it may be advantageous to taper the wall of the tubular member 42 in the downstream direction inwardly. The higher velocity of the fuel / air mixture flowing through the reduced cross-section would prevent the flame of the middle nozzle from moving upstream in the primary combustion chamber. It should be noted that in the event that it is decided to reignite the flame in the primary combustion chamber, it is necessary to provide a spark plug or other ignition device in the space.
    In yet another exemplary, but non-limiting embodiment, more than a single nozzle may be protected from the cold air flowing through the surrounding or adjacent nozzles at FSNL. For example, a central nozzle and one or two further nozzles in the outer assembly could be reset in the same manner as described above in connection with the middle nozzle 28. In addition, the one or two additional nozzles could be arranged a single oblong, oval or other shape having combustion chamber, that is, the shape of the space would be determined by the number and location of the reset nozzles. Such an arrangement is illustrated in FIG. 4 in which a central nozzle 128 and a radially outer nozzle 126 surrounded by a surrounding assembly are recessed within a second combustion chamber 136 defined by an elongate tubular member 142.
    This developed multi-stage combustor is thus capable of isolating the reacting flames of the fuel-supplied nozzles (eg, the middle nozzle 28) from the extremely cold ambient air coming from adjacent non-fueled nozzles (For example, the radially outer nozzles 26) in the part-load or zero-load range by creating a combustion zone in a recessed combustion chamber (the primary combustion chamber 36) for complete burnout of CO at the end of this space.
    LIST OF REFERENCE NUMBERS
    [0017]<Tb> 10 <September> combustion chamber<tb> 12 <SEP> Annular channel<Tb> 14 <September> transition piece<Tb> 16 <September> combustor liner<tb> 18, 20 <SEP> Flow sleeves<tb> 20, 24 <SEP> Cooling holes<tb> 26, 126 <SEP> Outer nozzles<tb> 28, 128 <SEP> Medium Nozzle<tb> 30, 50 <SEP> Facility<tb> 32 <SEP> First combustion chamber<tb> 34 <SEP> Single vane<tb> 36, 136 <SEP> Second combustion chamber<Tb> 38 <September> combustor cap<Tb> 40 <September> housing wall<tb> 42 <SEP> Tubular element<tb> 44 <SEP> Annular plate<tb> 46 <SEP> Outer Nozzle Tubes<tb> 48 <SEP> Middle nozzle tube<Tb> 50 <September> Verwirblereinlässe<Tb> 52 <September> openings
    权利要求:
    Claims (12)
    [1]
    A combustor (10) for a gas turbine, comprising:a plurality of radially outer nozzles (26) disposed in a substantially annular array, each of the radially outer nozzles having an outlet end arranged to supply fuel and / or air to a first combustion chamber (32);at least one central nozzle (28) having an outlet end located axially upstream of the outlet ends of the radially outer nozzles is configured and arranged to direct fuel and air to a second combustion chamber (36) axially upstream of the first combustion chamber (32). supply, wherein the second combustion chamber (36) opens into the first combustion chamber (32) and has a sufficient length to keep a flame of the central nozzle limited to the second combustion chamber (36).
    [2]
    The combustor of claim 1, wherein the outlet ends of the radially outer nozzles (26) are retained in an annular plate (44), and wherein a tubular member (42) defining the second combustion chamber extends from the annular plate (44) extends in an upstream direction.
    [3]
    3. The combustor of claim 1, wherein in addition to the central nozzle (128), one or more of the plurality of radially outer nozzles (126) have outlet ends upstream of the remaining ones of the plurality of radially outer nozzles.
    [4]
    4. The combustor of claim 3, wherein the one or more radially outer nozzles (126) having outlet ends are configured and arranged upstream of the remaining ones of the plurality of radially outer nozzles to supply fuel and air to the second combustion chamber (136).
    [5]
    The combustor of claim 4, wherein the outlet ends of the remaining ones of the plurality of radially outer nozzles (126) are retained in an annular plate, and wherein a tubular member (142) defining the second combustion chamber extends from the annular plate in an upstream direction extends.
    [6]
    A combustion chamber according to claim 1 or 4, wherein means (30, 50) are provided to supply either air alone or air and fuel to the plurality of radially outer nozzles.
    [7]
    A combustor according to claim 6, further comprising means (30, 52) for supplying fuel and air to the central nozzle.
    [8]
    8. A method of operating a gas turbine having at least one combustion chamber (10) according to claim 1, the method comprising:a) at zero or low load conditions, supplying fuel and air to the central nozzle (28) and only air to the outer array of nozzles (26), while a flame generated by the central nozzle (28) from one through the outer Arrangement of nozzles (26) air flowing is isolated; andb) under higher load conditions, supplying a fuel / air mixture through both the outer array of nozzles (26) and the central nozzle (28) so that flames generated by the outer array of nozzles are maintained in a first combustion chamber (32) and a flame generated by the central nozzle is maintained in a second combustion chamber (36) upstream of the first combustion chamber.
    [9]
    9. The method of claim 8, further comprising:c. Extinguishing the flame generated by the central nozzle (28); andd. Re-igniting a new flame generated by the central nozzle (28) in the first combustion chamber (32).
    [10]
    10. The method of claim 8, wherein the first combustion chamber (32) has a sufficient length to burn out CO at low or zero load levels.
    [11]
    The method of claim 8, including cooling a tubular member (42) defining the second combustion chamber (36).
    [12]
    12. The method of claim 11, wherein the cooling is performed by impingement cooling, thermal protection coating, turbulators or any combination of these.
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    同族专利:
    公开号 | 公开日
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    优先权:
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