Patent Publication Number: US-9423131-B2

Title: Air management arrangement for a late lean injection combustor system and method of routing an airflow

Description:
BACKGROUND OF THE INVENTION 
     The subject matter disclosed herein relates to combustor systems, and more particularly to an air management arrangement for a late lean injection combustor system, as well as a method of routing an airflow within such a late lean injection combustor system. 
     In combustion applications, such as a gas turbine system, for example, a combustor section includes a combustor chamber defined by a combustor liner that is often surrounded by a sleeve, such as a flow sleeve. An airflow typically passes through a passage disposed between the combustor liner and the sleeve for cooling of the combustor liner, as well as routing of the airflow to air-fuel injectors located at a forward end of the combustor liner. The airflow is derived from an air supply that must typically also provide air to other regions for a variety of purposes. Such a region may include late lean injectors that inject air into the combustor chamber in an effort to reduce undesirable emissions into an ambient atmosphere. As late lean injection combustor systems become more prevalent and more of the air supply is employed to provide air to the late lean injectors, efforts to cool the combustor liner are hindered due to the availability of less air from the air supply to be used for cooling purposes within the passage between the sleeve and the combustor liner. 
     Based on the direct supply of airflow to the air-fuel injectors, a combustion system is subject to back pressure when combustion fluctuates and suddenly increases the combustion pressure. The higher pressure inside the combustor chamber will instantaneously “push” a flammable fuel/air mixture into an air supply chamber, such as a compressor discharge casing (CDC). Such flammable mixture may cause damage to the CDC and result in shut down. 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to one aspect of the invention, an air management arrangement for a late lean injection combustor system includes a combustor liner defining a combustor chamber. Also included is a sleeve surrounding at least a portion of the combustor liner, the combustor liner and the sleeve defining a cooling annulus for routing a cooling airflow from proximate an aft end of the combustor liner toward a forward end of the combustor liner. Further included is a cooling airflow divider region configured to split the cooling airflow into a first cooling airflow portion and a second cooling airflow portion, wherein the first cooling airflow portion is directed to at least one primary air-fuel injector, wherein the second cooling airflow portion is directed to at least one lean-direct injector extending through the sleeve and the cooling annulus for injection of the second cooling airflow portion into the combustor chamber. 
     According to another aspect of the invention, a method of routing an airflow for a late lean injector combustor system is provided. The method includes directing a cooling airflow into a cooling annulus defined by a combustor liner and a sleeve surrounding at least a portion of the combustor liner, wherein the cooling airflow is routed through the cooling annulus from proximate an aft end of the combustor liner toward a forward end of the combustor liner. Also included is splitting the cooling airflow into a first cooling airflow portion and a second cooling airflow portion. Further included is routing the first cooling airflow portion to at least one primary air-fuel injector. Yet further included is routing the second cooling airflow portion to at least one lean-direct injector extending through the sleeve and the cooling annulus for injection of the second cooling airflow portion into a combustor chamber. 
     These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
         FIG. 1  is a schematic illustration of a gas turbine system; 
         FIG. 2  is a partial schematic illustration of a combustor section of the gas turbine system; 
         FIG. 3  is a schematic illustration of an air management arrangement for the combustor section; and 
         FIG. 4  is a flow diagram illustrating a method of routing an airflow for the combustor section. 
     
    
    
     The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings. 
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to  FIG. 1 , a gas turbine system is schematically illustrated with reference numeral  10 . The gas turbine system  10  includes a compressor section  12 , a combustor section  14 , a turbine section  16 , a shaft  18  and one or more air-fuel nozzles  20 . It is to be appreciated that one embodiment of the gas turbine system  10  may include a plurality of compressor sections  12 , combustor sections  14 , turbine sections  16 , shafts  18  and one or more air-fuel fuel nozzles  20 . The compressor section  12  and the turbine section  16  are coupled by the shaft  18 . The shaft  18  may be a single shaft or a plurality of shaft segments coupled together to form the shaft  18 . 
     The combustor section  14  uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the gas turbine system  10 . For example, the one or more air-fuel nozzles  20  may be of various types, as will be discussed in detail below, and are in fluid communication with an air supply  22  and a fuel supply  24 . The one or more air-fuel nozzles  20  create an air-fuel mixture, and discharge the air-fuel mixture into the combustor section  14 , thereby causing a combustion that creates a hot pressurized exhaust gas. The combustor section  14  directs the hot pressurized gas through a transition piece into a turbine nozzle (or “stage one nozzle”), and other stages of buckets and nozzles causing rotation of the turbine section  16  within a turbine casing  26 . Rotation of the turbine section  16  causes the shaft  18  to rotate, thereby compressing the air as it flows into the compressor  12 . In an embodiment, hot gas path components are located in and proximate the combustor section  14 , where hot gas flow proximate the components causes creep, oxidation, wear and thermal fatigue of components. As the firing temperature increases, the hot gas path components need to be properly cooled to meet service life and to effectively perform intended functionality. 
     Referring now to  FIG. 2 , the combustor section  14  is schematically illustrated in greater detail. The combustor section  14  includes a transition piece  28  in the form of a duct that is at least partially surrounded by an impingement sleeve  30  disposed radially outwardly of the transition piece  28 . Upstream thereof, proximate a forward region of the impingement sleeve  30  is a combustor liner  32  defining a combustor chamber  34 . The combustor liner  32  is at least partially surrounded by a flow sleeve  36  disposed radially outwardly of the combustor liner  32 . Although the combustor liner  32  and the transition piece  28  have been described as separate components, it is to be appreciated that the combustor liner  32  and the transition piece  28  may be formed as a single, unitary structural component that forms the combustor chamber  34  and a transition zone. Similarly, although the flow sleeve  36  and the impingement sleeve  30  have been described as separate components, it is to be appreciated that the flow sleeve  36  and the impingement sleeve  30  may be formed as a single, unitary sleeve configured to surround at least a portion of the combustor liner  32  and the transition piece  28 , whether separate or integrated components. 
     Irrespective of the precise configuration of the combustor liner  32 , the transition piece  28 , the flow sleeve  36  and the impingement sleeve  30 , a compressor discharge casing  38  is illustrated and includes a compressor discharge exit  40  that is configured to route the air supply  22  that is employed for numerous purposes within the combustor section  14 . The air supply  22  typically originates from the compressor section  12  and enters into the compressor discharge casing  38 . The air supply  22  exits the compressor discharge casing  38  proximate the compressor discharge exit  40  and rushes downstream toward the transition duct  28  and/or the combustor liner  32 . Specifically, rather than routing a portion of the air supply  22  directly to various components, such as air-fuel nozzles, approximately all of the air supply  22  is directed as a cooling airflow  42  to a first cooling annulus  44  defined by the combustor liner  32  and the flow sleeve  36 . The cooling airflow  42  is directed within the first cooling annulus  44  from an aft end  48  of the combustor liner  32  toward a forward end  49  of the combustor liner  32 . As described in detail above, various embodiments relating to the sleeve(s), as well as the combustor liner  32  and transition piece  28  configuration are contemplated, and it is to be understood that the air supply  22  may be directed as the cooling airflow  42  to a second cooling annulus  46  defined by the transition piece  28  and the impingement sleeve  30 . For an embodiment having a single liner or duct defining the combustor chamber  34  surrounded by one or more sleeves, the air supply  22  may be directed as the cooling airflow  42  to such a cooling annulus. For purposes of this description, reference to the first cooling annulus  44  defined by the combustor liner  32  and the flow sleeve  36  is intended to apply to routing of the cooling airflow  42  to any cooling annulus described above. 
     The combustor section  14  is late lean injection (LLI) compatible. An LLI compatible combustor is any combustor with either an exit temperature that exceeds 2500° F. or handles fuels with components that are more reactive than methane with a hot side residence time greater than 10 milliseconds (ms). 
     Irrespective of the embodiment employed in the gas turbine system  10 , at least one, but typically a plurality of lean-direct injectors (“LDIs”)  50 , are each integrated with or structurally supported by a plurality of housings that extend radially into at least one of the transition piece  28  or the combustor liner  32 . The plurality of LDIs  50  extend through the respective component, i.e., the transition piece  28  or the combustor liner  32 , to varying depths. That is, the plurality of LDIs  50  are each configured to supply a second fuel (i.e., LLI fuel) to the combustion zone through fuel injection in a direction that is generally transverse to a predominant flow direction through the transition piece  28  and/or the combustor liner  32 . For each of the above-described embodiments, it is emphasized that the plurality of LDIs  50  may be disposed proximate the transition piece  28  or the combustor liner  32 , in spite of the illustrated embodiments showing disposal of the plurality of LDIs  50  disposed in connection with only one of the transition piece  28  and the combustor liner  32 . Furthermore, the plurality of LDIs  50  may be disposed in connection with both the transition piece  28  and the combustor liner  32 . The plurality of LDIs  50  may be disposed in a single axial circumferential stage that includes multiple currently operating LDIs respectively disposed around a circumference of a single axial location of the transition piece  28  and/or the combustor liner  32 . It is also conceivable that the plurality of LDIs  50  may be situated in a single axial stage, multiple axial stages, or multiple axial circumferential stages. A single axial stage includes a currently operating single LDI. A multiple axial stage includes multiple currently operating LDIs that are respectively disposed at multiple axial locations. A multiple axial circumferential stage includes multiple currently operating LDIs, which are disposed around a circumference of the transition piece  28  and/or the combustor liner  32  at multiple axial locations thereof. 
     Referring now to  FIG. 3 , the cooling airflow  42  is illustrated proximate the forward end  49  of the combustor liner  32 . As shown, the cooling airflow  42  is routed toward the forward end  49  of the combustor liner  32  within the first cooling annulus  44  and around the plurality of LDIs  50 . The cooling airflow  42  provides a convective cooling effect on the combustor liner  32  while flowing toward the forward end  49  of the combustor liner  32 . As noted above, approximately all (i.e., about 100%) of the air supply  22  is directed to the first cooling annulus  44  for cooling purposes. Upon reaching a location proximate the forward end  49  of the combustor liner  32 , a cooling airflow divider region  52 , which as shown in the illustrated embodiment may simply be a walled region of the combustor section  14 , splits the cooling airflow  42  into a first cooling airflow portion  54  and a second cooling airflow portion  56 . 
     The first cooling airflow portion  54  is directed to at least one primary air-fuel injector  58  located at the forward end  49  of the combustor liner  32  for mixing and injection of an air-fuel mixture into the combustor chamber  34 . The at least one primary air-fuel injector  58  is typically aligned relatively parallel to the predominant direction of flow within the combustor chamber  34 . The second cooling airflow portion  56  is directed to the plurality of LDIs  50  for mixing and injection of the LLI fuel, as described above. Although illustrated and described above as being located proximate the forward end  49  of the combustor liner  32 , it is to be appreciated that the cooling airflow divider region  52  may be disposed at any location along the combustor liner  32  and/or the transition piece  28 , as well as any location along the flow sleeve  36  and/or the impingement sleeve  30 . Specifically, the cooling airflow  42  may be split into the first cooling airflow portion  54  and the second cooling airflow portion  56  at any desired location suitable for the particular application of use. Furthermore, the combustor section  14  may include a plurality of cooling airflow divider regions and the cooling airflow  42  may be divided into more than two portions. 
     Routing approximately all of the air supply  22  through the first cooling annulus  44  reduces the likelihood of “flame flash back” pushing out of the combustor chamber  34  upon a sudden increase or fluctuation of combustion pressure within the combustor chamber  34 . In the event of such an increase or fluctuation of combustion pressure, the path that the air-fuel mixture must travel to extend into a sensitive region subject to damage is more tortuous. Specifically, the likelihood of the air-fuel mixture reaching the compressor discharge casing  38  is reduced. Advantageously, in addition to having a longer and more tortuous path, the air-fuel mixture is provided multiple paths to flash back through. In particular, the split of the cooling flow  42  proximate the forward end  49  of the combustor liner  32  allows the air-fuel mixture being pushed back to enter the at least one primary air-fuel injector  58  or one of the plurality of LDIs  50 . For example, if the air-fuel mixture is pushed out of one of the plurality of LDIs  50 , the air-fuel mixture may pass to the at least one primary air-fuel injector  58  for re-entry to the combustor chamber  34 . 
     As illustrated in the flow diagram of  FIG. 4 , and with reference to  FIGS. 1-3 , a method of routing an airflow for a late lean injection combustor system  100  is also provided. The gas turbine system  10  and the combustor section  14  have been previously described and specific structural components need not be described in further detail. The method of routing an airflow for a late lean injection combustor system  100  includes directing a cooling airflow into a cooling annulus  102  defined by the combustor liner  32  and a sleeve surrounding at least a portion of the combustor liner  32 . The cooling airflow is split into a first cooling airflow portion and a second cooling airflow portion  104 . The first cooling airflow portion is routed to at least one primary air-fuel injector  106 , while the second cooling airflow portion is routed to at least one lean-direct injector  108 . 
     Advantageously, approximately all of the air supply  22  is employed to cool various components subjected to extreme thermal conditions, such as the transition piece  28  and/or the combustor liner  32 , for example. By routing the cooling airflow  42  to several air-fuel injectors, including the plurality of LDIs  50 , the air supply  22  serves a dual purpose benefit. Specifically, the cooling air  42  cools various components, then is mixed with a fuel for injection to the combustor chamber  34 . 
     While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.