Patent Publication Number: US-11022313-B2

Title: Combustor assembly for a turbine engine

Description:
FIELD OF THE INVENTION 
     The present subject matter relates generally to a gas turbine engine, or more particularly to a combustor assembly for a gas turbine engine. 
     BACKGROUND OF THE INVENTION 
     A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere. 
     Traditionally, the combustion section includes a combustor for receiving compressed air and fuel and combusting the combination to provide the turbine section with the combustion gasses. The fuel and air is typically provided with an equivalence ratio of about 1:1 such that substantially stoichiometric combustion takes place. However, such may lead to relatively high peak temperatures, and further to undesirable amounts of NOx formation. 
     Accordingly, a combustion section for a gas turbine engine capable of avoiding these issues would be useful. More specifically, a combustion section capable of generating combustion gases having a reduced amount of NOx, while efficiently combusting all of the fuel would be particularly beneficial. 
     BRIEF DESCRIPTION OF THE INVENTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     In one exemplary embodiment of the present disclosure, a rich-quench-lean combustor assembly is provided for a gas turbine engine defining an axial direction. The combustor assembly includes a liner extending between a forward end and an aft end, the liner including a plurality of quench air jets positioned between the forward end and the aft end. The combustor assembly additionally includes a dome attached to or formed integrally with the liner, the dome and the liner together defining at least in part a combustion chamber. The combustor assembly additionally includes a fuel nozzle attached to the dome, the fuel nozzle configured as a premix fuel nozzle for providing a substantially homogenous mixture of fuel and air to the combustion chamber, the mixture of fuel and air having an equivalence ratio of at least 1.5. 
     In another exemplary embodiment of the present disclosure, a gas turbine engine is provided. The gas turbine engine defines an axial direction and includes a compressor section and a turbine section arranged in serial flow order. The gas turbine engine also includes a combustor assembly. The combustor assembly includes a liner extending between a forward end and an aft end, the liner including a plurality of quench air jets positioned between the forward end and the aft end. The combustor assembly additionally includes a dome attached to or formed integrally with the liner, the dome and the liner together defining at least in part a combustion chamber. The combustor assembly additionally includes a fuel nozzle attached to the dome, the fuel nozzle configured as a premix fuel nozzle for providing a substantially homogenous mixture of fuel and air to the combustion chamber, the mixture of fuel and air having an equivalence ratio of at least 1.5. 
     In an exemplary aspect of the present disclosure, a method is provided for operating a combustor assembly of a gas turbine engine. The combustor assembly includes a liner attached to or formed integrally with a dome, the liner including a plurality of quench air jets and the dome including a fuel nozzle attached thereto. The method includes providing a substantially homogenous mixture of fuel and air having an equivalence ratio of at least about 1.5 through the fuel nozzle to a combustion chamber defined at least in part by the dome and the liner. The method also includes providing the combustion chamber with a quench airflow through the plurality of quench air jets of the liner at a location downstream from the fuel nozzle. 
     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 invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which: 
         FIG. 1  is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter. 
         FIG. 2  is a close-up, cross-sectional view of a combustor assembly in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 3  is a close-up, cutaway view of a fuel nozzle in accordance with an exemplary embodiment of the present disclosure. 
         FIG. 4  is a close-up, cross-sectional view of a portion of the exemplary combustor assembly of  FIG. 2 . 
         FIG. 5  is a perspective view of a section of the exemplary combustor assembly of  FIG. 2 , with a flow sleeve removed for clarity. 
         FIG. 6  is a close-up, side, cross-sectional view of a primary stage air jet of the exemplary combustor assembly of  FIG. 2  in accordance with an exemplary aspect of the present disclosure. 
         FIG. 7  is a close-up, outside view of the exemplary primary stage air jet of  FIG. 6 . 
         FIG. 8  is a forward-looking-aft view of a combustion chamber, dome, and outer liner of the exemplary combustor assembly of  FIG. 2 , in accordance with an exemplary aspect of the present disclosure. 
         FIG. 9  is a forward-looking-aft view of a combustion chamber, dome, and outer liner of a combustor assembly in accordance with another exemplary embodiment of the present disclosure. 
         FIG. 10  is a flow diagram of a method for operating a combustor assembly in accordance with an exemplary aspect of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components. The terms “forward” and “aft” refer to relative positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust. The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows. 
     Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,  FIG. 1  is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of  FIG. 1 , the gas turbine engine is a high-bypass turbofan jet engine  10 , referred to herein as “turbofan engine  10 .” As shown in  FIG. 1 , the turbofan engine  10  defines an axial direction A (extending parallel to a longitudinal axis  12  provided for reference), a radial direction R, and a circumferential direction C (i.e., a direction extending about the axial direction A; not depicted). In general, the turbofan  10  includes a fan section  14  and a core turbine engine  16  disposed downstream from the fan section  14 . 
     The exemplary core turbine engine  16  depicted generally includes a substantially tubular outer casing  18  that defines an annular inlet  20 . The outer casing  18  encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor  22  and a high pressure (HP) compressor  24 ; a combustion section  26 ; a turbine section including a high pressure (HP) turbine  28  and a low pressure (LP) turbine  30 ; and a jet exhaust nozzle section  32 . A high pressure (HP) shaft or spool  34  drivingly connects the HP turbine  28  to the HP compressor  24 . A low pressure (LP) shaft or spool  36  drivingly connects the LP turbine  30  to the LP compressor  22 . 
     For the embodiment depicted, the fan section  14  includes a variable pitch fan  38  having a plurality of fan blades  40  coupled to a disk  42  in a spaced apart manner. As depicted, the fan blades  40  extend outwardly from disk  42  generally along the radial direction R. Each fan blade  40  is rotatable relative to the disk  42  about a pitch axis P by virtue of the fan blades  40  being operatively coupled to a suitable actuation member  44  configured to collectively vary the pitch of the fan blades  40  in unison. The fan blades  40 , disk  42 , and actuation member  44  are together rotatable about the longitudinal axis  12  by LP shaft  36  across a power gear box  46 . The power gear box  46  includes a plurality of gears for stepping down the rotational speed of the LP shaft  36  to a more efficient rotational fan speed. 
     Referring still to the exemplary embodiment of  FIG. 1 , the disk  42  is covered by rotatable front nacelle  48  aerodynamically contoured to promote an airflow through the plurality of fan blades  40 . Additionally, the exemplary fan section  14  includes an annular fan casing or outer nacelle  50  that circumferentially surrounds the fan  38  and/or at least a portion of the core turbine engine  16 . For the embodiment depicted, the nacelle  50  is supported relative to the core turbine engine  16  by a plurality of circumferentially-spaced outlet guide vanes  52 , and a downstream section  54  of the nacelle  50  extends over an outer portion of the core turbine engine  16  so as to define a bypass airflow passage  56  therebetween. 
     During operation of the turbofan engine  10 , a volume of air  58  enters the turbofan  10  through an associated inlet  60  of the nacelle  50  and/or fan section  14 . As the volume of air  58  passes across the fan blades  40 , a first portion of the air  58  as indicated by arrows  62  is directed or routed into the bypass airflow passage  56  and a second portion of the air  58  as indicated by arrow  64  is directed or routed into the LP compressor  22 . The ratio between the first portion of air  62  and the second portion of air  64  is commonly known as a bypass ratio. The pressure of the second portion of air  64  is then increased as it is routed through the high pressure (HP) compressor  24  and into the combustion section  26 , where it is mixed with fuel and burned to provide combustion gases  66 . 
     The combustion gases  66  are routed through the HP turbine  28  where a portion of thermal and/or kinetic energy from the combustion gases  66  is extracted via sequential stages of HP turbine stator vanes  68  that are coupled to the outer casing  18  and HP turbine rotor blades  70  that are coupled to the HP shaft or spool  34 , thus causing the HP shaft or spool  34  to rotate, thereby supporting operation of the HP compressor  24 . The combustion gases  66  are then routed through the LP turbine  30  where a second portion of thermal and kinetic energy is extracted from the combustion gases  66  via sequential stages of LP turbine stator vanes  72  that are coupled to the outer casing  18  and LP turbine rotor blades  74  that are coupled to the LP shaft or spool  36 , thus causing the LP shaft or spool  36  to rotate, thereby supporting operation of the LP compressor  22  and/or rotation of the fan  38 . 
     The combustion gases  66  are subsequently routed through the jet exhaust nozzle section  32  of the core turbine engine  16  to provide propulsive thrust. Simultaneously, the pressure of the first portion of air  62  is substantially increased as the first portion of air  62  is routed through the bypass airflow passage  56  before it is exhausted from a fan nozzle exhaust section  76  of the turbofan  10 , also providing propulsive thrust. The HP turbine  28 , the LP turbine  30 , and the jet exhaust nozzle section  32  at least partially define a hot gas path  78  for routing the combustion gases  66  through the core turbine engine  16 . 
     It should be appreciated, however, that the exemplary turbofan engine  10  depicted in  FIG. 1  is by way of example only, and that in other exemplary embodiments, the turbofan engine  10  may have any other suitable configuration. Additionally, it will be appreciated that in other embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine, such as a suitable aeronautical gas turbine engine (e.g., turboshaft, turboprop, turbojet, etc.), land-based gas turbine engine (e.g., power generation gas turbine engine), aero-derivative gas turbine engine (e.g., marine applications), etc. 
     Referring now to  FIG. 2 , a close-up cross-sectional view is provided of a rich-quench-lean (“RQL”) combustor assembly (“combustor assembly  100 ”) in accordance with an exemplary embodiment of the present disclosure. In certain embodiments, for example, the combustor assembly  100  of  FIG. 2  may be positioned in the combustion section  26  of the exemplary turbofan engine  10  of  FIG. 1 . Alternatively, however, it may be positioned in any other suitable gas turbine engine. For example, in other embodiments, the combustor assembly  100  may be incorporated into one or more of a turboshaft engine, a turboprop engine, a turbojet engine, a land-based gas turbine engine for power generation, an aero-derivative or marine gas turbine engine, etc. 
     As shown, the combustor assembly  100  generally includes an inner liner  102  extending between an aft end  104  and a forward end  106  generally along the axial direction A, as well as an outer liner  108  also extending between an aft end  110  and a forward end  112  generally along the axial direction A. The inner and outer liners  102 ,  108  are each attached to or formed integrally with a dome  114 . More particularly, for the embodiment depicted, the inner and outer liners  102 ,  108  are each formed integrally with the dome  114 , such that the inner liner  102 , outer liner  108 , and dome  114  form a one-piece combustor liner extending continuously from the aft  104  end of the inner liner  102 , to the forward end  106  of the inner liner  102 , across the dome  114 , to the forward end  110  of the outer liner  108 , and to the aft end  112  of the outer liner  108 . Although not depicted in  FIG. 2 , the one-piece combustor liner may additionally extend along the circumferential direction C. In certain embodiments, the one-piece combustor liner may extend continuously along the circumferential direction C, or alternatively, the combustor assembly  100  may include a plurality of one-piece combustor liners arranged along the circumferential direction C. Furthermore, as will be discussed in greater detail below, the inner and outer liners  102 ,  108  and dome  114  together at least partially define a combustion chamber  116 , the combustion chamber  116  having a centerline  118  extending therethrough. 
     For the embodiment depicted, the inner liner  102 , the outer liner  108 , and dome  114  are each formed of a ceramic matrix composite (CMC) material, which is a non-metallic material having high temperature capability. Exemplary CMC materials utilized for such liners  102 ,  108  and the dome  114  may include silicon carbide, silicon, silica or alumina matrix materials and combinations thereof. Ceramic fibers may be embedded within the matrix, such as oxidation stable reinforcing fibers including monofilaments like sapphire and silicon carbide (e.g., Textron&#39;s SCS-6), as well as rovings and yarn including silicon carbide (e.g., Nippon Carbon&#39;s NICALON®, Ube Industries&#39; TYRANNO®, and Dow Corning&#39;s SYLRAMIC®), alumina silicates (e.g., Nextel&#39;s 440 and 480), and chopped whiskers and fibers (e.g., Nextel&#39;s 440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite and montmorillonite). It should be appreciated, however, that in other embodiments, one or more of the inner liner  102 , outer liner  108 , and dome  114  may be formed of any other suitable material, such as a traditional metal alloy. 
     Referring still to  FIG. 2 , the combustor assembly  100  further includes a flow sleeve  120  enclosing the inner liner  102 , outer liner  108 , and dome  114 . The flow sleeve  120  generally includes an outer portion  122 , an inner portion  124 , and a forward portion  126 . For the embodiment depicted, the combustor assembly  100  is mounted within an outer casing of the gas turbine engine, and more particularly, the flow sleeve  120  is mounted to an outer combustor casing  128  of the gas turbine engine using one or more mounting features  130 . The one or more mounting features  130 , for the embodiment depicted, are attached to the outer portion  122  of the flow sleeve  120 . Additionally, although not depicted, the combustor assembly  100  may include a plurality of such mounting features  130  attached to the outer portion  122  of the flow sleeve  120  and spaced along the circumferential direction C. 
     For the exemplary flow sleeve  120  depicted, the outer portion  122 , forward portion  126 , and inner portion  124  are formed integrally as a one-piece flow sleeve, extending continuously. Additionally, although not depicted, the flow sleeve  120  may additionally extend continuously along the circumferential direction C. However, in other embodiments, one or more of the outer portion  122 , forward portion  126 , and inner portion  124  may be formed separately and attached in any suitable manner. Additionally, in other embodiments, the flow sleeve  120  may be formed of a plurality of individual flow sleeves  120  spaced along the circumferential direction C. 
     As is also depicted in  FIG. 2 , the combustor assembly  100  is positioned downstream of a diffuser  132  of the gas turbine engine. The diffuser  132  is positioned at an aft end of a combustion section of the gas turbine engine for providing compressed air  134  from the combustion section to the combustor assembly  100 . The combustor assembly  100  further includes a fuel nozzle  135  for receiving a portion of the compressed air  134 , mixing the received portion of the compressed air  134  with fuel, and providing such fuel-air mixture to the combustion chamber  116 . As will be shown and discussed in greater detail below, the fuel nozzle  135  depicted is configured as a pre-mix fuel nozzle and the combustor assembly  100  further includes a plurality of fuel nozzles  135  spaced substantially evenly along the circumferential direction C (see, e.g.,  FIGS. 5, 8 ). Additionally, it should be appreciated that in certain embodiments, the fuel may be a natural gas (such as methane), LNG, propane, lean methane, high H2 content fuel, a petroleum distillate (such as No. 2 distillate fuel), kerosene, heavy fuel oil, marine diesel fuel, or any other suitable fuel. 
     Referring now briefly to  FIG. 3 , a close-up, perspective, cross-sectional view is provided of a pre-mix fuel nozzle  135  in accordance with one or more embodiments of the present disclosure as may be incorporated in the exemplary combustor assembly  100  of  FIG. 2 . 
     The fuel nozzle  135  of  FIG. 3  generally includes a centerbody  136  and an outer sleeve  138  that generally surrounds the centerbody  136 . The outer sleeve  138  includes one or more radial vanes  140  forming a radial swirler  142 . An inner sleeve  144  is disposed between the centerbody  136  and the outer sleeve  138 . The inner sleeve  144 , at least a part of which is generally disposed upstream of the radial swirler  142 , includes a contoured shroud  146  at an aft end. Generally, the contoured shroud  146  is aerodynamically contoured to promote mixing of a liquid or gaseous fuel and air. For example, the contoured shroud  146  includes a plurality of lobes  148 . 
     Moreover, the radial swirler  142  is disposed radially outward of the contoured shroud  146  and a fuel injection port  150 . The fuel injection port  150  is defined between the inner sleeve  144  and a portion of the outer sleeve  138 , and provides for a flow of fuel from a fuel circuit (not depicted) defined by the fuel nozzle  135 . As shown, at least one fuel injection port  150  is disposed upstream, or forward, of the contoured shroud  146 . 
     The positioning of the radial swirler  142  to the contoured shroud  146  and fuel injection port  150  is such that compressed air, such as a portion of the compressed air  134  from the diffuser  132  (see  FIG. 2 ), entering through the radial swirler  142  converges and mixes with a liquid or gaseous fuel exiting a fuel injection port  150 . The contoured shroud  146  may aid in positioning the fuel exiting the fuel injection port  150  such that the convergence of air  126  through the radial swirler  142  may deliver high levels of fuel-air mixing. 
     Generally upstream in the fuel nozzle  135  from the contoured shroud  146 , an outer surface  152  of the centerbody  136  and an inner surface  154  of the inner sleeve  144  includes a plurality axially oriented vanes  156  forming an axial swirler  158 . The axial swirler  158  may have any geometry between at least one outer surface  152  of the centerbody  136  and at least one inner surface  154  of the inner sleeve  144 , and is not limited to any particular geometry, unless otherwise specified. Neither the centerbody  136  nor the inner sleeve  144  is bound to one diametric value for its entire structure. Furthermore, the centerbody  136  and subsequent surrounding features may have other radial cross-sectional forms, such as an elliptical or polygonal radial cross section. 
     Notably, a relationship of the outer sleeve  138  to the centerbody  136  creates an annular circuit  160  substantially along a length of the centerbody  136 . In the embodiment shown in  FIG. 3 , the radial cross-sectional area of the annular circuit  160  at a first location is greater than at a second location downstream of the first location. Thus, the annular circuit  160  may converge as it extends downstream along a center axis  162 . More specifically, the inner surface  132  of the outer sleeve  138  is converging toward the center axis  162  as the inner surface  132  extends downstream, while the outer surface  152  of the centerbody  136  is diverging from the center axis  162 . In other embodiments, however, the inner surface  132  of the outer sleeve  138  may have any other suitable shape relative to the outer surface  152  of the centerbody  136 . 
     Referring still to the exemplary fuel nozzle  135  of  FIG. 3 , the centerbody  136  includes a first centerbody circuit  164 , a second centerbody circuit  166 , and a third centerbody circuit  168  leading to at least one circuit outlet  170  to egress a fluid (e.g. liquid or gaseous fuel, air, inert gas, or combination thereof). The second centerbody circuit  166  and the third centerbody circuit  168  are positioned generally co-axial to the first centerbody circuit  164 . In another embodiment, the second centerbody circuit  166  or the third centerbody circuit  168  may be tunnels within the centerbody (i.e., not annular cavities), radially outward from a first centerbody circuit  164 . Any combination of centerbody circuits  164 ,  166 ,  168  may be fluidly connected toward the downstream end of the centerbody  136  before egressing through the centerbody outlet  170 . In another embodiment, any centerbody circuit  164 ,  166 ,  168 , or a combination thereof, may egress independently to a circuit outlet  170  without fluid interconnection. The exemplary fuel nozzle  135  may be configured for independent variable flow rates within each centerbody circuit  164 ,  166 ,  168 . It should be apparent to one skilled in the art that additional centerbody circuits (fourth, fifth . . . Nth) may be installed and arranged in substantially similar manner as the first, second, and third circuits  164 ,  166 ,  168  described herein. 
     The exemplary fuel nozzle  135  depicted in  FIG. 3 , and described herein, may allow for the fuel nozzle  135  to provide a substantially homogenous mixture of fuel and air to the combustion chamber  116  during operation of the combustor assembly  100 . As used herein, “substantially homogenous” means more mixed than not, i.e., having a mixedness greater than fifty percent (50%). For example, in certain embodiments, the fuel nozzle  135  may be configured to provide a mixture of fuel and air to the combustion chamber  116  having at least about a seventy percent (70%) mixedness. More specifically, in certain embodiments, the fuel nozzle  135  may be configured to provide a mixture of fuel and air to the combustion chamber  116  having at least about an eighty percent (80%) mixedness. It should be appreciated, that as used herein, terms of approximation, such as “about” or “approximately,” refer to being within a ten percent (10%) margin of error. 
     Moreover, as used herein, the term “mixedness” with respect to a mixture of fuel and air, refers to a calculation for determining how a fuel species varies over a surface. Mixedness may be calculated generally using the following formula: 
               Mixedness   =     1   -       σ   f       f   _           ,         
where σ f  is a standard deviation of mixture fraction and  f  is a mass weighted average of mixture fraction. Each may be calculated in the manner described below.
 
     The mixture fraction, f, is defined in terms of the atomic mass fraction, which may be expressed as follows: 
               f   =         x   i     -     x     i   ,   ox             x     i   ,   fuel       -     x     i   ,   ox             ,         
where x i  is the elemental mass fraction for the element, x i,ox  denotes the oxidizer, and x i,fuel  denotes the value at the fuel stream.
 
     Further, the standard deviation, σ f , of the mixture fraction on a surface is computed using the following formula: 
     
       
         
           
             
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     Further, still, the mass weighted average of the mixture fraction,  f , is calculated by dividing the summation of the value mixture fraction multiplied by the absolute value of the dot product of the facet area and momentum vectors by the summation of the absolute value of the dot product of the facet area and momentum vectors, as is indicated in the following formula: 
     
       
         
           
             
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     Mixedness calculated in accordance with the above method may provide a mixedness at an outlet of a fuel nozzle. 
     Further, it should be appreciated, that in other exemplary embodiments, the combustor assembly  100  may have any other suitable fuel nozzle, and that the present application is not limited to the exemplary fuel nozzle  135  depicted in  FIG. 3  and described above, unless specifically so limited by the claims. 
     Referring now back to  FIG. 2 , as stated, the combustor assembly  100  is configured as an RQL combustor assembly  100  (i.e., a combustor assembly providing for a rich combustion, a quench air, and subsequently a lean combustion). In order to achieve an initial rich combustion, the fuel nozzle  135  is further configured to provide a mixture of fuel and air to the combustion chamber  116  having an equivalence ratio of at least 1.5. More specifically, for the embodiment depicted the fuel nozzle  135  is configured to provide the combustion chamber  116  with a mixture of fuel and air having an equivalence ratio at least about two (2). Notably, as used herein, “equivalence ratio” refers to a ratio of fuel to air. 
     As depicted, the combustion chamber  116  includes a primary combustion zone  172  and a secondary combustion zone  174 . The substantially homogenous mixture of fuel and air, having an equivalence ratio of at least 1.5, is combusted in the primary combustion zone  172 . However, given the relatively high equivalence ratio of such a mixture of fuel and air, incomplete combustion occurs (i.e., less than stoichiometric combustion). In order to complete the combustion process, the combustor assembly  100  is configured to introduce an additional amount of air downstream of the primary combustion zone  172  forming the secondary combustion zone  174 . More specifically, the inner liner  102  of the combustor assembly  100  includes a plurality of quench air jets  176  positioned between the forward end  106  and the aft end  104 , and similarly the outer liner  108  of the combustor assembly  100  includes a plurality of quench air jets  176  positioned between the forward  112  end and the aft end  110 . 
     As will be appreciated, a majority of the compressed air  134  from the diffuser  132  is received by the combustor assembly  100  for combustion (i.e., “compressed air for combustion  134 A”), while in at least certain embodiments, a portion of the compressed air  134  from the diffuser  132  is diverted downstream for cooling operations (not shown). In order to achieve a desired equivalence ratio, along with the desired mixedness of the fuel and air mixture provided to the combustion chamber  116 , a majority of the compressed air for combustion  134 A is introduced into the combustion chamber  116  through the plurality of quench air jets  176  of the inner liner  102  and outer liner  108 . More specifically, in at least certain embodiments, at least about sixty percent (60%) of the compressed air for combustion  134 A is introduced to the combustion chamber  116  through the quench air jets  176  of the inner liner  102  and the outer liner  108  is a quench airflow  134 B. For example, in certain embodiments, at least about seventy percent (70%) of the compressed air for combustion  134 A is introduced to the combustion chamber  116  through the quench air jets  176  of the inner liner  102  and the outer liner  108  as quench airflow  134 B. 
     The quench airflow  134 B introduced through the quench air jets  176  of the inner liner  102  and outer liner  108  may mix with the combustion gases from the primary combustion zone  172  in the secondary combustion zone  174 . The mixture of quench airflow  134 B through the quench air jets  176  and combustion gases from the primary combustion zone  172  may result in a lean combustion mixture within the secondary combustion zone  174 . For example, such a mixture of fuel and air may have an equivalence ratio of less than about 0.75. More specifically, in certain embodiments at least, the secondary combustion zone  174  may define an equivalence ratio of less than about 0.65. 
     For the embodiment depicted in  FIG. 2 , the quench air jets  176  of the inner liner  102  and the outer liner  108  are each positioned approximately halfway along a length of the inner liner  102  and the outer liner  108 , respectively (i.e., about halfway between the respective forward ends  106 ,  112  and aft ends  104 ,  110 ). Notably, the exemplary combustion chamber  116  depicted converges aft of the primary combustion zone  172 , and the quench air jets  176  of the inner liner  102  and the outer liner  108  are positioned at the convergence. More specifically, the exemplary combustor assembly  100  depicted defines a forward height  178  within the combustion chamber  116  between the outer liner  108  and the inner liner  102  at a location forward of the plurality of quench air jets  176  of the outer liner  108  and the inner liner  102 . Additionally, the combustor assembly  100  defines an aft height  180  within the combustion chamber  116  between the outer liner  108  and the inner liner  102  at a location aft of the plurality of quench air jets  176  of the outer liner  108  and the inner liner  102 . A ratio of the forward height  178  to the aft height  180  is at least about 1.75:1. For example, in certain embodiments, the ratio of the forward height  178  to the aft height  180  may be at least about 2:1. 
     As used herein, the forward height  178  is defined in a direction that extends perpendicular to the centerline  118  of the combustion chamber  116  and intersects with a longitudinal axis of the gas turbine engine (e.g., axis  12  of  FIG. 1 ). Similarly, the aft height  180  is defined in a direction that extends perpendicular to the centerline  118  of the combustion chamber  116  and intersects with the longitudinal axis of the gas turbine engine. Further, each of the forward height  178  and aft height  180  refer to the greatest heights in the respective directions, and more particularly, for the embodiment depicted, refer to heights immediately forward and immediately aft, respectively, of the quench air jets  176  of the inner liner  102  and outer liner  108 . 
     Briefly, it should also be appreciated, that for the embodiment depicted, the plurality of quench air jets  176  of the inner liner  102  and of the outer liner  108  are each positioned at approximate the same position along the centerline  118  of the combustion chamber  116 , i.e., such that they are aligned. However, in other embodiments, the quench air jets  176  of the inner liner  102  may be offset from the quench air jets  176  of the outer liner  108  along the centerline  118 . For example, in certain embodiments, the quench air jets  176  of the inner liner  102  may be positioned forward of the quench air jets  176  of the outer liner  108 , or alternatively, the quench air jets  176  of the inner liner  102  may be positioned aft of the quench air jets  176  of the outer liner  108 . 
     Reference will now be made also to  FIGS. 4 and 5 .  FIG. 4  provides a close-up, cross-sectional view of a section of the exemplary combustor assembly  100  of  FIG. 2 , and  FIG. 5  provides a perspective view of a section of the exemplary combustor assembly  100  of  FIG. 2 , with the flow sleeve  120  removed for clarity. 
     As previously discussed, the mixture of fuel and air provided to the primary combustion zone  172  of the combustion chamber  116  by the fuel nozzles  135  defines a relatively high equivalence ratio. In order to maintain the desired equivalence ratio, the combustor assembly  100  is configured such that substantially no cooling air enters the combustion chamber  116  forward of the quench air jets  176  of the inner liner  102  and the outer liner  108 . More particularly, for the embodiment depicted, the inner liner  102  defines a forward section  182  extending between the quench air jets  176  and the dome  114  and the outer liner  108  similarly defines a forward section  184  extending between the quench air jets  176  and the dome  114 . For the embodiment depicted, the forward section  182  of the inner liner  102 , the forward section  184  of the outer liner  108 , and the dome  114  are each configured to prevent a flow of cooling air from entering the combustion chamber  116 , and are further configured to be cooled substantially by one or both of impingement cooling or convective cooling. More particularly, for the embodiment depicted, each of the forward section  182  of the inner liner  102 , the forward section  184  of the outer liner  108 , and the dome  114  are free from any cooling holes and are configured to be cooled through a flow of impingement air onto an outer surface  186  of the respective components. 
     Further, for the embodiment depicted, the impingement air for cooling the forward section  184  of the outer liner  108 , the forward section  182  of the inner liner  102 , and the dome  114  is provided through impingement cooling holes  188  defined by the flow sleeve  120 . The impingement cooling holes  188  through the flow sleeve  120  are all positioned forward of the quench air jets  176  of the inner and outer liners  102 ,  108 . During operation, a portion of the compressed air for combustion  134 A not provided to the combustion chamber  116  through the fuel nozzles  135  flows through the impingement cooling holes  188  defined by the flow sleeve  120 , through a chamber  190  (defined between the flow sleeve  120  and the outer liner  108 , dome  114 , and inner liner  102 ) directly onto the outer surfaces  186  of the forward section  194  of the outer liner  108 , the forward section  182  of the inner liner  102 , and the dome  114  for cooling such components. Notably, while such a cooling method may not provide a level of cooling attainable through inclusion of cooling holes through the inner and outer liners  102 ,  108  and/or dome  114 , as discussed above, the inner and outer liners  102 ,  108  and dome  114  may be formed of a CMC material. Additionally, as the combustor assembly  100  is configured for incomplete combustion in the primary combustion zone, a temperature within the combustion chamber  116  proximate the forward sections  182 ,  184  of the inner and outer liners  102 ,  108  may be reduced. Accordingly, with such an embodiment the components may be capable of withstanding the temperatures necessary for operation without inclusion of cooling holes. 
     Moreover, referring still to  FIGS. 4 and 5 , for the embodiment depicted, the quench air jets  176  of the outer liner  108  and the quench air jets  176  of the inner liner  102  are each configured in a two-stage configuration. More particularly, referring first to the quench air jets  176  of the outer liner  108 , the quench air jets  176  include a plurality of primary stage air jets  192  and a plurality of secondary stage air jets  194 . The plurality of primary stage air jets  192  are each spaced from the plurality of secondary stage air jets  194  along the axial direction A and along the centerline  118 . 
     Further, for the embodiment depicted the plurality of primary stage air jets  192  are each configured as relatively large air jets, while the plurality of secondary stage air jets  194  are each configured as relatively small air jets. For example, each of the plurality of primary stage air jets  192  define a cross-sectional area  196  (see cross-section close up in  FIG. 4 ) and each of the plurality of secondary stage air jets  194  also defines a cross-sectional area  198  (see cross-section close up in  FIG. 5 ). For the embodiment depicted, the cross-sectional area  196  of the primary stage air jets  192  is greater than the cross-sectional area  198  of the secondary stage jets. For example, in at least certain embodiments, a ratio of the cross-sectional area  196  of the primary stage air jets  192  to the cross-sectional area  198  of the secondary stage air jets  194  may be at least about 1.75:1, such as at least about 2:1, or at least about 2.25:1. It should be appreciated, that for embodiments where a cross-section of the primary stage air jets  192  and/or secondary stage air jets  194  vary, e.g., along the circumferential direction C (see, e.g.,  FIG. 9 ), the cross-sectional area  196  and cross-sectional area  198  refer to an average cross-sectional area. 
     In addition to differing in size, the exemplary combustor assembly  100  depicted includes a greater number of secondary stage air jets  194  as compared to the primary stage air jets  192 . More particularly, for the embodiment depicted, the combustor assembly  100  defines a ratio of a number of secondary stage air jets  194  to a number of primary stage air jets  192  of at least about 1.5:1. For example, in certain embodiments, the combustor assembly  100  may define a ratio of a number of secondary stage air jets  194  to a number of primary stage air jets  192  of at least about 1.75:1, of at least about 2:1, or of at least about 2.25:1. 
     Further, at least for the embodiment depicted, the number of primary stage air jets  192  correlates to a number of fuel nozzles  135 . For example, in the embodiment depicted, the combustor assembly  100  defines a ratio of a number of primary stage air jets  192  to a number of fuel nozzles  135  of at least about 1.5:1. More particularly, in certain embodiments, the combustor assembly  100  may define a ratio of a number of primary stage air jets  192  to a number of fuel nozzles  135  of at least about 1.75:1, of at least about 2:1, or of at least about 2.25:1. 
     Further still, as briefly mentioned above, the quench air jets  176  of the inner liner  102  and the outer liner  108  are positioned at the convergence of the combustion chamber  116 , about halfway along a length of the inner liner  102  and outer liner  108 . For the embodiment depicted, the plurality of primary stage air jets  192  are positioned at a forward end of the convergence, and the plurality of secondary stage air jets  194  are positioned at an aft end of the convergence. Additionally, the primary stage air jets  192  and secondary stage air jets  194  may be separated a distance (i.e., a distance from an edge of one opening to an edge of the other opening along the centerline  118 ) of less than about ten percent (10%) of a total length of the inner liner  102  or outer liner  108  along the centerline  118 . Moreover, as may be seen most clearly in  FIG. 4 , the plurality of primary stage air jets  192  are oriented substantially perpendicularly to the centerline  118  of the combustion chamber  116  (i.e., a length-wise centerline  200  of each primary air jet  194  is substantially perpendicular to a local centerline  118  of the combustion chamber  116 ). By contrast, the plurality of secondary stage air jets  194  are, for the embodiment, oriented oblique relative to the centerline  118  of the combustion chamber  116  (i.e., a length-wise centerline  202  of each of the plurality of secondary stage air jets  194  define an angle relative to a local centerline  118  of the combustion chamber  116 ). For example, the centerline  118  of each of the plurality of secondary stage air jets  194  may define an angle relative to a local centerline  118  of the combustion chamber  116  less than about seventy-five (75) degrees, such as less than about sixty (60) degrees. It should be appreciated, however, that in other exemplary embodiments, the plurality of primary air jets may instead be oriented oblique to the centerline  118  of the combustion chamber  116 , and further, in certain embodiments, the plurality of secondary stage air jets  194  may be oriented perpendicularly to the centerline  118  of the combustion chamber  116 . 
     Notably, aft of the quench air jets  176 , the exemplary inner liner  102  and outer liner  108  depicted each include a plurality of film cooling holes  206 . The film cooling holes  206  provide for cooling of an inside surface (i.e., a surface adjacent to the combustion chamber  116 ) of the inner liner  102  and outer liner  108 . It should be appreciated, that the film cooling holes  206  may be at least about 1/10 of a size of the secondary stage air jets  194 . 
     Referring now briefly to  FIG. 6 , providing a close-up, cross-sectional view of the exemplary primary stage air jet  192  depicted in  FIG. 4 , the outer liner  108  includes an inlet transition  208  immediately forward of each of the plurality of primary stage air jets  192 . The inlet transition  208  may reduce an amount of separation of the compressed air for combustion  134 A from the outer liner  108  as the compressed air for combustion  134 A flows over the outer liner  108  and into the primary stage air jets  192 . For the embodiment depicted, the inlet transition  208  defines a radius of curvature at least about 0.65 inches. For example, in certain embodiments, the inlet transition  208  may define a radius of curvature of at least about 0.75 inches. As used herein, the term “radius of curvature” refers to a radius of a circle that touches a curve at a given point and has the same tangent and curvature at that point. 
     Further, referring now briefly to  FIG. 7 , providing an outer view of the primary stage air jet  192  of the outer liner  108  depicted in  FIG. 4 , the exemplary primary stage air jet  192  defines an inlet  210  having a generally elliptical shape. For the embodiment depicted, the elliptical shape of the inlet includes a minor radius of curvature  212  of at least about 0.25 inches and a major radius of curvature  214  of at least about 0.4 inches. However, in other embodiments, the elliptical shape of the inlet  210  may instead define any other suitable major and/or minor radius of curvature  212 ,  214 . For example, in other embodiments, the elliptical shape of the inlet  210  may define a minor radius of curvature  212  of at least about 0.3 inches and a major radius of curvature  214  of at least about 0.5 inches. Additionally, it should be appreciated that in other embodiments, the inlet of the primary stage air jets  192  may have any other suitable shape, and similarly, a cross-section of the primary stage air jets  192  may have any other suitable shape. For example, although for the embodiment depicted, the cross-sections of the primary stage air jets  182  are substantially elliptical, in other embodiments they may be substantially circular, squared, triangular, polygonal, etc. 
     By contrast, referring now back to  FIG. 5 , for the embodiment depicted, each of the plurality of secondary stage air jets  194  are configured as elongated slots, elongated along a widthwise direction  216 . Furthermore, for the embodiment depicted, the elongated slots define an oblique angle  218  relative to the axial direction A and relative to the centerline  118  of the combustion chamber  116  (see particularly, the close up of elongated slot in  FIG. 5 ). For example, the elongated slots (i.e., the widthwise direction  216 ) may define an angle  218  of at least about thirty degrees (30°) relative to the axial direction A and centerline  118 , such as at least about forty-five degrees (45°) relative to the axial direction A and centerline  118 . However, in other embodiments, the widthwise direction  216  of the elongated slots may define any other suitable angle  218  with the axial direction A and centerline  118 . Additionally, in still other embodiments, it should be appreciated, that the secondary stage air jets  194  may instead be circular in shape, elliptical in shape, or have any other suitable shape. 
     As stated, in certain embodiments, at least about sixty percent (60%) of the compressed air for combustion  134 A is introduced to the combustion chamber  116  through the quench air jets  176  as a quench airflow  134 B. Inclusion of the plurality of primary stage air jets  192  and the plurality secondary stage air jets  194  in accordance with the present disclosure may allow for substantially even distribution of the quench airflow  134 B between the plurality of primary stage air jets  192  and the plurality of secondary stage air jets  194 . For example, in certain embodiments, the combustor assembly  100  may be configured to provide between about forty percent (40%) and about sixty percent (60%) of the quench airflow  134 B through the plurality of primary stage air jets  192  and between about forty percent (40%) and about sixty percent (60%) of the quench airflow  134 B through the plurality of secondary stage air jets  194 . More specifically, in certain embodiments, the combustor assembly  100  may be configured to provide about fifty-five percent (55%) of the quench airflow  134 B through the plurality of primary stage air jets  192  and about forty-five (45%) of the quench airflow  134 B through the plurality of secondary stage air jets  194 . 
     The plurality of quench air jets  176  configured in such a manner may allow for the quench airflow  134 B to reach the combustion gasses from the primary combustion zone  172  across an entire height of the secondary combustion zone  174 . More particularly, with such a configuration, the primary stage air jets  192  may be configured to provide deep penetration of the quench airflow  134 B, while the secondary stage air jets  194  may be configured to provide relatively shallow penetration of the quench airflow  134 B. 
     It should be appreciated, that although the description above with reference to  FIGS. 4 through 6  is directed to the outer liner  108  and the quench air jets  176  of the outer liner  108 , in certain embodiments, the quench air jets  176  of the inner liner  102  may be configured in substantially the same manner. 
     Reference will now be made to  FIG. 8 .  FIG. 8  provides a view along a centerline  118  of the combustion chamber  116  depicted in  FIG. 4 , looking forward from an aft end. More particularly,  FIG. 8  provides a view of the outer liner  108 , along with the primary stage air jets  192  and secondary stage air jets  194 , the combustion chamber  116 , and an aft end of two of the plurality of fuel nozzles  135 . As is depicted, and as may be seen in other of the Figures, the plurality of primary stage air jets  192  are each spaced along the circumferential direction C, and similarly, the plurality of secondary stage air jets  194  are also spaced along the circumferential direction C. Particularly for the embodiment depicted, each of the plurality of primary stage air jets  192  and the plurality of secondary stage air jets  194  are substantially evenly spaced along the circumferential direction C. As may also be seen in  FIG. 8 , each of the plurality of primary stage air jets  192  are substantially the same size (i.e., each defines substantially the same cross-sectional area) and further, each of the plurality of secondary stage air jets  194  are also substantially the same size (i.e., each defines substantially the same cross-sectional area). 
     It should be appreciated, however, that in other exemplary embodiments of the present disclosure, one or both of the plurality of primary stage air jets  192  or the plurality of secondary stage air jets  194  may define a variable spacing along the circumferential direction C and/or a variable size. For example, referring now to  FIG. 9 , a view is provided along a centerline  118  of a combustion chamber  116  of a combustor assembly  100 , looking forward from an aft end, in accordance with another exemplary embodiment of the present disclosure is provided. 
     For the embodiment of  FIG. 9 , at least one of the plurality of primary stage air jets  192  or plurality of secondary stage air jets  194  are unevenly spaced along the circumferential direction C. More particularly, for the embodiment of  FIG. 9 , the plurality of secondary stage air jets  194  are unevenly spaced along the circumferential direction C. For the embodiment depicted, the uneven spacing along the circumferential direction C of the plurality of secondary stage air jets  194  correlates to a position of the plurality of fuel nozzles  135 . Specifically, for the embodiment depicted, the plurality of secondary stage air jets  194  are unevenly spaced along the circumferential direction C such that they are more closely spaced immediately downstream of the respective fuel nozzles  135 . 
     Moreover, for the embodiment depicted, at least one of the plurality primary stage air jets  192  or plurality of secondary stage air jets  194  include a variable size along the circumferential direction C, which for the embodiment depicted also correlates to a position of the plurality of fuel nozzles  135 . Specifically, for the embodiment depicted, the plurality primary stage air jets  192  include a variable sizing along the circumferential direction C, such that the primary stage air jets  192  located immediately downstream of a respective fuel nozzle  135  are larger (i.e. define a larger cross-sectional area) as compared to other primary stage air jets  192  not located immediately downstream of a respective fuel nozzle  135 . 
     It should be appreciated, however, that in other exemplary embodiments, the combustor assembly  100  may have any other suitable configuration, and that the above description is not meant to be limiting unless the claims specifically provide for such limitations. 
     A combustor assembly  100  in accordance with one or more embodiments of the present disclosure may allow for a more efficient combustor assembly  100 , with reduced emissions. More particularly, a combustor assembly  100  in accordance with one or more embodiments of the present disclosure may allow for reduced NOx emissions. For example, by including a primary combustion zone  172  with a relatively high equivalence ratio, and immediately downstream including a secondary combustion zone  174  with a relatively low equivalence ratio (facilitated by the two-stage quench air jets  176 , lack of forward cooling holes, and/or pre-mixing nozzles), the relatively high combustion temperatures which may generate a maximum amount of NOx may be minimized. 
     Further, inclusion of the two-stage quench air jets  176  may allow for a plurality primary stage air jets  192  to achieve relatively deep penetration of a quench airflow into the combustion chamber  116 , while the plurality of secondary stage air jets  194  may provide for relatively shallow penetration of a quench airflow into the combustion chamber  116 , such that a more uniform provision of quench airflow to the secondary combustion zone  174  may be provided 
     Referring now to  FIG. 10 , a flow diagram is provided of an exemplary method ( 300 ) for operating a combustor assembly of a gas turbine engine. In certain exemplary aspects, the method ( 300 ) may operate one or more of the exemplary combustor assemblies described above with reference to  FIGS. 2 through 8 . Accordingly, the exemplary combustor assembly may include a liner attached to or formed integrally with a dome, with the liner including a plurality of quench air jets and the dome including a fuel nozzle attached thereto. 
     The exemplary method ( 300 ) includes at ( 302 ) providing a substantially homogenous mixture of fuel and air having an equivalence ratio of at least about 1.5 through the fuel nozzle to a combustion chamber defined at least in part by the dome and the liner. Providing the substantially homogenous mixture of fuel and air at ( 302 ) may include providing a mixture of fuel and air to the combustion chamber having a mixedness of at least about seventy percent (70%), such as at least about eighty percent (80%). 
     Additionally, the exemplary method ( 300 ) includes at ( 304 ) providing the combustion chamber with a quench airflow through the plurality of quench air jets of the liner at a location downstream from the fuel nozzle. For the exemplary aspect depicted, the exemplary method is operable with a combustor assembly wherein the liner is an outer liner and wherein the combustor assembly further includes an inner liner. The outer liner and inner liner each include a plurality of quench air jets positioned between a respective forward end and aft end of the liners, with each set of quench air jets including a plurality of primary stage air jets spaced from a plurality of secondary stage air jets along an axial direction and along a centerline of the combustion chamber. Accordingly, for the exemplary aspect depicted, providing the combustion chamber with a quench airflow at ( 304 ) includes at ( 306 ) providing between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the plurality of primary stage air jets and at ( 308 ) providing between about forty percent (40%) and about sixty percent (60%) of the quench airflow through the plurality of secondary stage air jets. 
     A combustor assembly operated in accordance with exemplary method described herein may allow for more efficient operation of the combustor assembly with reduced emissions. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.