Abstract:
The present invention discloses a novel apparatus and method for operating a gas turbine combustor having a structural configuration proximate a pilot region of the combustor which seeks to minimize the onset of thermo acoustic dynamics. The pilot region of the combustor includes a generally cylindrical extension having an outlet end with an irregular profile which incorporates asymmetries into the system so as to destroy any coherent structures.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of U.S. Provisional Patent Application Ser. No. 61/708,323 filed on Oct. 1, 2012. 
     
    
     TECHNICAL FIELD 
       [0002]    The present invention relates generally to a system and method for improving combustion stability and reducing emissions in a gas turbine combustor. More specifically, the improvements in a combustor premixer address acoustic dynamic instabilities and can also reduce thermal stresses, thus improving structural integrity and component life. 
       BACKGROUND OF THE INVENTION 
       [0003]    In an effort to reduce the amount of pollution emissions from gas-powered turbines, governmental agencies have enacted numerous regulations requiring reductions in the amount of oxides of nitrogen (NOx) and carbon monoxide (CO). Lower combustion emissions can often be attributed to a more efficient combustion process, with specific regard to fuel injector location and mixing effectiveness. 
         [0004]    Early combustion systems utilized diffusion type nozzles, where fuel is mixed with air external to the fuel nozzle by diffusion, proximate the flame zone. Diffusion type nozzles have been known to produce high emissions due to the fact that the fuel and air burn stoichiometrically at high temperature to maintain adequate combustor stability and low combustion dynamics. 
         [0005]    An enhancement in combustion technology is the utilization of premixing, such that the fuel and air mix prior to combustion to form a homogeneous mixture that burns at a lower temperature than a diffusion type flame and produces lower NOx emissions. Premixing fuel and air together before combustion allows for the fuel and air to form a more homogeneous mixture, which will burn more completely, resulting in lower emissions. However, in this configuration the fuel is injected in relatively the same plane of the combustor, and prevents any possibility of improvement through altering the mixing length. 
         [0006]    Premixing can occur either internal to the fuel nozzle or external thereto, as long as it is upstream of the combustion zone. An example of a premixing combustor  100  of the prior art is shown in  FIG. 1 . The combustor  100  is a type of reverse flow premixing combustor utilizing a pilot nozzle  102 , a radial inflow mixer  104 , and a plurality of main stage mixers  106  and  108 . The pilot portion of the combustor  100  is separated from the main stage combustion area by a center divider portion  110 . The center divider portion  110  separates the fuel injected by the pilot nozzle  102  from the fuel injected by the main stage mixers  106  and  108 . While the combustor  100  of the prior art has improved emissions levels and ability to operate at reduced load settings, analysis and testing has demonstrated the onset of thermo acoustic dynamics due to symmetries generated in the burner as a result of the burner geometry, such as the center divider portion. 
         [0007]    As one skilled in the art understands, mechanisms that cause thermo-acoustic instabilities are coherent structures generated by the burner. One type of combustor known to exhibit such instabilities is a combustor having a cylindrical shape. What is needed is a system that can provide flame stability and low emissions benefits at a part load condition while also reducing thermo-acoustic instabilities generated by coherent flame structures. 
       SUMMARY 
       [0008]    The present invention discloses a gas turbine combustor having a structural configuration proximate a pilot region of the combustor which seeks to minimize the onset of thermo acoustic dynamics. The pilot region, or center region of the combustor, is configured to incorporate asymmetries into the system so as to destroy any coherent structures in the resulting flame. 
         [0009]    In an embodiment of the present invention, a combustor is disclosed having a combustion liner located within a flow sleeve with a dome located at a forward end of the flow sleeve and encompassing at least a forward portion of the combustion liner. The combustor also comprises a generally cylindrical extension projecting into the combustion liner from the dome, where the outlet end of the extension has an irregular profile. 
         [0010]    In an alternate embodiment of the present invention, an extension for a dome of a gas turbine combustor is disclosed. The extension comprises a generally cylindrical member extending along an axis of the combustor where the generally cylindrical member has an outlet end configured to not be located in a single plane perpendicular to the axis of the combustor. 
         [0011]    In yet another embodiment of the present invention, a method is provided for isolating a main stage of fuel injectors from a pilot fuel nozzle in order to reduce acoustic dynamics in the combustor. The method comprises providing a combustion liner having a dome and extension component where air is injected into the combustion liner and a first stream of fuel is injected into the extension piece to mix with a portion of the air to form a pilot flame. A second stream of fuel is injected into another portion of the air located outside of the combustion liner. This mixture is then directed into the combustion liner in a way such that the second stream of fuel is separated from the first stream of fuel by the extension piece. 
         [0012]    Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention. The instant invention will now be described with particular reference to the accompanying drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         [0013]    The present invention is described in detail below with reference to the attached drawing figures, wherein: 
           [0014]      FIG. 1  is a cross section view of a gas turbine combustion system of the prior art. 
           [0015]      FIG. 2  is a cross section view of a gas turbine combustion system in accordance with an embodiment of the present invention. 
           [0016]      FIG. 3  is a perspective view of a portion of the gas turbine combustion system of  FIG. 2  in accordance with an embodiment of the present invention. 
           [0017]      FIG. 4A  is a detailed cross section view of a portion of the gas turbine combustion system of  FIG. 2  in accordance with an embodiment of the present invention. 
           [0018]      FIG. 4B  is an alternate detailed cross section view of a portion of the gas turbine combustion system of  FIG. 2  in accordance with an embodiment of the present invention. 
           [0019]      FIG. 5  is an alternate perspective view of a portion of a gas turbine combustion system in accordance with an alternate embodiment of the present invention. 
           [0020]      FIG. 6  is a cross section of the portion of a gas turbine combustor of  FIG. 5  in accordance with an alternate embodiment of the present invention. 
           [0021]      FIG. 7  is a perspective view of a portion of a gas turbine combustion system in accordance with yet another alternate embodiment of the present invention. 
           [0022]      FIG. 8  is a cross section of the portion of a gas turbine combustor of  FIG. 7  in accordance with an alternate embodiment of the present invention. 
           [0023]      FIG. 9  is a perspective view of a portion of a gas turbine combustion system in accordance with an additional embodiment of the present invention. 
           [0024]      FIG. 10  is a perspective view of a portion of a gas turbine combustion system in accordance with yet another embodiment of the present invention. 
           [0025]      FIG. 11  is a perspective view of a portion of a gas turbine combustion system in accordance with a further embodiment of the present invention. 
           [0026]      FIG. 12  depicts the process of isolating a main stage of fuel injectors from a pilot stage in accordance with an embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0027]    By way of reference, this application incorporates the subject matter of U.S. Pat. Nos. 6,935,116, 6,986,254, 7,137,256, 7,237,384, 7,513,115, 7,677,025, and 7,308,793. 
         [0028]    The preferred embodiment of the present invention will now be described in detail with specific reference to  FIGS. 2-12 . The combustion system of the present invention utilizes premixing fuel and air prior to combustion in combination with precise staging of fuel flow to the combustor to achieve reduced emissions at multiple operating load conditions. Reconfigured combustor geometry is provided to target a reduction of combustion acoustic pressure fluctuations, to reduce thermal stresses, cracking and detrimental thermo-acoustic coherent structures. 
         [0029]    Referring now to  FIG. 2 , a gas turbine combustion system  200  is provided comprising a generally cylindrical flow sleeve  202  and a generally cylindrical combustion liner  204  located at least partially within the flow sleeve  202 . The combustion system  200  also comprises a dome  206  located axially forward of the flow sleeve  202 . The dome  206  is positioned such that it encompasses at least a forward portion  208  of the combustion liner  204 . The dome  206  also has a hemispherical head end  210  and an opening  212  that is coaxial with a center axis A-A of the combustion system  200 . The gas turbine combustion system  200  also comprises a pilot nozzle  214  extending generally along the center axis A-A of the combustion system  200  and a radial inflow mixer  216 , each for directing a supply of fuel to pass into the combustion liner  204  along or near the center axis A-A. 
         [0030]    Referring also to  FIGS. 3 ,  4 A and  4 B, the gas turbine combustion system  200  also comprises a generally cylindrical extension  218  projecting into the combustion liner  204  from the dome  206 . The precise length of generally cylindrical extension  218  can vary and is chosen based upon the operating parameters defining turndown as the main fuel stage is isolated from the pilot stage by separating the flame regions and avoiding flame quenching at lower operating temperatures. The extension  218  has an inlet end  220  positioned at the opening  212  of the dome  206  and an opposing outlet end  222 , which is positioned at a distance within the combustion liner  204 . As discussed above, combustors that have a cylindrical structure with uniform exit planes are subject to cracking due to thermal gradients causing circumferential stresses within the cylindrical structure. Furthermore, these combustors also have tendencies to produce thermo-acoustic dynamics having a coherent structure. That is, the acoustic waves formed within the combustor have a uniform structure due to the symmetric structure within the combustor. The combustor of the prior art depicted in  FIG. 1  has been known to exhibit circumferential stress-induced cracking and to produce acoustic waves in the center divider portion  110 , due to its symmetric structure. 
         [0031]    As one skilled in the art will understand, acoustic waves are a by-product of the combustion process due to vortices being shed at a cylindrical burner outlet. When these vortices are convected into the flame, a fluctuation in the heat release occurs. When the acoustic fluctuations amplify the shedding of vortices, a constructive interference with the heat release can occur causing high amplitude dynamics. These high dynamics can cause cracking in the combustor. 
         [0032]    The present invention provides reconfigured combustor geometry to help reduce fluctuations in heat release. In the prior art combustor of  FIG. 1 , the combustor  100  included a center divider portion  110  for separating the flow of fuel in the pilot nozzle  102  from the fuel from main stage injectors  106  and  108 . The center divider portion  110  has a cylindrical cross section and a uniform exit plane perpendicular to the flow of fuel and air. As such, vortices shed at the exit plane of the center divider portion  110  are convected into the surrounding main stage flame, which is produced by injection of fuel from injectors  106  and  108 . Because of the uniform exit plane of the center divider portion  110 , these vortices have been known to cause a fluctuation in heat release and cause high amplitude dynamics. Further, the large temperature gradient experienced by the center divider portion  110  creates circumferential stresses causing cracking of the divider portion. 
         [0033]    To improve the prior art combustor design while maintaining the benefit of separate fuel injection circuits required for a combustor having the specified design and staging configuration, the outlet end  222  of the generally cylindrical extension  218  in combustion system  200  is configured to have an irregular profile or shape. An irregular profile or shape has been shown to reduce the temperature gradient and dynamics levels. A variety of irregular shapes can be used for the outlet end  222  of the generally cylindrical extension  218 .  FIGS. 3-6  depict some of the alternate embodiments of the generally cylindrical extension component having an irregular profile or shape to the outlet end. 
         [0034]    Referring to  FIGS. 3-4B , the irregular profile or shape of the outlet end  222  comprises a planar edge  224  extending generally perpendicular to the center axis A-A where the planar edge  224  is interrupted by a series of semi-circular cutouts  226 . The semi-circular cutouts  226  provide a non-uniform exit plane from the generally cylindrical extension  218 . That is, as the flow exits the generally cylindrical extension  218 , it will exit into the surrounding flow at slightly different axial locations due to the cutouts  226 . As a result, asymmetries are introduced into the exit flow from the generally cylindrical extension  218 , which disrupts any coherent structures being formed that could otherwise amplify if injected in a symmetrical pattern. In addition, the semi-circular cutouts  226  tend to reduce the cracking in the generally cylindrical extension  218  by relieving circumferential stresses induced by the thermal gradients in the generally cylindrical extension  218 . The exact size, quantity and spacing of the semi-circular cutouts  226  about the outlet end  222  can vary depending on a variety of factors such as frequency of combustion dynamics that should be damped, the flow velocity, flame position, and delay times. For the embodiment of the present invention depicted in  FIGS. 3-4B , twelve semi-circular cutouts  226  are equally spaced about the outlet end  222  of the generally cylindrical extension  218 . Depending on the combustor design and operating conditions, the cutouts  226  can also be positioned about the outlet end  222  in a non-equal or irregular pattern 
         [0035]    The irregular profile or shape is not limited to semi-circular cutouts. Alternatively, the irregular profile or shape of the outlet end of the extension  218  can take on other shapes, including but not limited to, a saw tooth pattern, a plurality of rectangular cutouts, and elliptical or sinusoidal cutouts. 
         [0036]    An alternate embodiment of the present invention is depicted in  FIGS. 5 and 6 . The alternate embodiment discloses a generally cylindrical extension  600  having a different geometry than that of the cylindrical extension  218  discussed above. The generally cylindrical extension  600  has an inlet end  602  and an opposing outlet end  604 . The cylindrical extension  600  is coupled to the dome and functions similar to the prior configuration discussed above and pictured in  FIGS. 2-4B . The main difference with the alternate generally cylindrical extension  600  is with respect to the irregular shape of the outlet end  604 . For the embodiment depicted in  FIGS. 5-6 , the outlet end  604  forms a plane taken at an angle a relative to the center axis A-A, such that the outlet end  604  is not in a single plane perpendicular to the center axis A-A of the combustion system. As with the semi-circular cutouts in the outlet end of the cylindrical extension  218 , the angular planar cut at outlet end  604  of cylindrical extension  600  provides an alternate way of introducing asymmetries into the flow of the combustion liner. 
         [0037]    Yet another embodiment of the present invention is depicted with respect to  FIGS. 7 and 8 . This alternate embodiment discloses a generally cylindrical extension  700  having a different geometry than the embodiments discussed above. The generally cylindrical extension  700  has an inlet end  702  and an opposing outlet end  704 . The cylindrical extension  700  is coupled to the dome and functions similar to the prior configuration discussed above and pictured in  FIGS. 2-6 . In the configuration depicted in  FIGS. 7 and 8 , it is possible to obtain the acoustic benefits driven primarily by the configuration of  FIGS. 5 and 6 , with the thermal stress reductions that can be obtained through the cutouts in the outlet end of the extension, as depicted in  FIGS. 3-4B . That is, the main difference with this alternate generally cylindrical extension  700  is with respect to the irregular shape of the outlet end  704 . For the embodiment depicted in  FIGS. 7 and 8 , the outlet end  704  forms a plane taken at an angle a relative to the center axis A-A, such that the outlet end  704  is not in a single plane perpendicular to the center axis A-A of the combustion system. As discussed above, the angular planar cut at outlet end  704  of cylindrical extension  700  provides a way of introducing asymmetries into the flow of the combustion liner Furthermore, and as discussed above, including a plurality of cutouts  706  in the outlet end  704  helps reduce the thermal stresses within the generally cylindrical extension  700 . Although generally semi-circular cutouts  706  are shown in  FIGS. 7 and 8 , the size and shape of these cutouts can vary to include other shapes, such as, but not limited to rectangular, elliptical, sinusoidal or saw-tooth shape. 
         [0038]    A series of alternate embodiments of the present invention are depicted in  FIGS. 9-11 , where the outlet end of the dome extension portion of the present invention can take on a variety of shapes in order to target certain frequencies of combustion acoustic pressure fluctuations. These alternative shapes to the outlet end may also aid in reducing thermal stresses in the dome extension. For example, the irregular profile of outlet end may consist of a variety of geometries, such as planar edges, continuous peaks and valleys or a combination of non-uniform exit plane geometries. The spacing of the features generating these profiles may be equal about the circumference of the outlet end or unequally spaced, depending on the frequency range of combustion acoustic pressure fluctuations being targeted. 
         [0039]    Referring first to  FIG. 9 , this alternate embodiment discloses a generally cylindrical extension  900  having a different geometry than the embodiments discussed above. The generally cylindrical extension  900  has an inlet end  902  (not shown) that is coupled to the dome and functions similar to the prior configuration discussed above and pictured in  FIGS. 2-8 . The generally cylindrical extension  900  also has an opposing outlet end  904 . In the configuration depicted in  FIG. 9 , it is possible to obtain the acoustic benefits driven primarily by the configuration of  FIGS. 5 and 6 , with the thermal stress reductions that can be obtained through the cutouts in the outlet end of the extension, as depicted in  FIGS. 3-4B . That is, similar to the configuration discussed above with respect to  FIGS. 7 and 8 , the main difference with this alternate generally cylindrical extension  900  is with respect to the irregular shape of the outlet end  904 .  FIG. 9  depicts an outlet end  904  having a wave-like profile formed by a series of axial exit planes where the effective outlet end  904  varies axially along a length of the extension  900 . These waves have a series of peaks  906  and troughs  908 , which are essentially formed by connecting a series of axially-spaced planar cuts. The peaks  906  and troughs  908  can be uniformly spaced or non-uniformly spaced. As a result of this outlet end profile, fuel flow from the pilot nozzle mixes with the surrounding fuel-air mixture in a non-uniform and axially spaced fashion, thereby introducing asymmetries into the exit flow, which disrupts any coherent structures being formed that could otherwise amplify if injected in a symmetrical pattern. 
         [0040]      FIG. 10  provides yet another alternative embodiment of an outlet end geometry for the extension. In this embodiment, a generally cylindrical extension  1000  has an inlet end  1002  (not shown) that is coupled to the dome and functions similar to the prior configuration discussed above and pictured in  FIGS. 2-8 . The generally cylindrical extension  1000  also has an opposing outlet end  1004 . As discussed above, a profile of the outlet end  1004  can be non-uniform. This is shown in  FIG. 10 , which depicts a generally cylindrical extension  1000 , where the outlet end  1004  exhibits a non-uniform profile along the axial distance forming the outlet end  1004  extends. As with the embodiment depicted in  FIG. 9 , fuel flow from the pilot nozzle, which extends along a center axis, can mix with the surrounding fuel-air mixture in a non-uniform and axially spaced fashion, thereby providing a way of targeting a reduction of certain frequencies of combustion acoustic pressure fluctuations. 
         [0041]    Referring now to  FIG. 11 , a portion of the gas turbine combustion system is shown including a generally cylindrical extension  1100  having an inlet end (not shown) that is coupled to the dome and functions similar to the prior configurations discussed above and pictured in  FIGS. 2-8 . The generally cylindrical extension  1100  also has an opposing outlet end  1104 . As discussed above, a profile of outlet end  1104  can be non-uniform. More specifically, the outlet end  1104  can have an outlet edge formed by multiple axially-spaced exit planes, as discussed above, but these multiple axially-spaced planes are taken at varying radii relative to the center axis of the combustor, thereby defining radial peaks  1106  and valleys  1108  in the generally cylindrical extension  1100 . That is, the generally cylindrical extension  1100  can flare radially inward or outward relative to the center axis of the combustor, as represented by arc-shaped portion  1110  of generally cylindrical extension  1100 . 
         [0042]    The present invention also provides a way of isolating a main stage of fuel injectors from a pilot fuel nozzle such that acoustic dynamics in the combustion system are reduced. Referring now to  FIG. 12 , the process  1200  for isolating the main stage of fuel injectors is depicted. In a step  1202 , a combustion liner is provided for a combustion system with the combustion liner having a hemispherical dome with an opening located therein and a generally cylindrical extension positioned at the opening and extending into the combustion liner. As discussed above, the generally cylindrical extension piece has an irregular profile or shape to the outlet end. Next, in a step  1204 , a flow of compressed air is injected into the combustion liner and around the hemispherical dome. In a step  1206 , a first stream of fuel is injected into the generally cylindrical extension piece in order to mix with a portion of the compressed air injected in step  1204  for providing a pilot flame. A second stream of fuel is injected in a step  1208  from a position radially outward of the combustion liner such that the second stream of fuel mixes with compressed air from step  1204  and the fuel-air mixture reverses flow direction upon contact with the hemispherical dome and enters the combustion liner to form a main injection flame. The extension piece serves to separate the stream of fuel for the pilot flame from the stream of fuel for the main injection flame. The irregular shape or profile of the extension piece creates asymmetries in the fuel injection location and thereby destroys any coherent structures between the pilot flame and main injection flame. 
         [0043]    While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims. The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive. Alternative embodiments and required operations will become apparent to those of ordinary skill in the art to which the present invention pertains without departing from its scope. 
         [0044]    From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.