Patent Publication Number: US-9404657-B2

Title: Combuster with radial fuel injection

Description:
Applicant hereby claims priority to U.S. Patent Application No. 61/707,033 filed Sep. 28, 2012, the disclosure of which is herein incorporated by reference. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a fuel nozzle arrangement therefor. 
     Gas turbine engines, such as those which power modern commercial and military aircraft, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases. The combustor generally includes radially spaced apart inner and outer liners that define an annular combustion chamber therebetween. Arrays of circumferentially distributed combustion air holes penetrate multiple axial locations along each liner to radially admit the pressurized air into the combustion chamber. A plurality of circumferentially distributed fuel injectors axially project into a forward section of the combustion chamber to supply the fuel for mixing with the pressurized air. 
     Combustion of the hydrocarbon fuel in the presence of pressurized air may produce nitrogen oxide (NO x ) emissions that are subject to excessively stringent controls by regulatory authorities, and thus may be sought to be minimized as much as possible. 
     At least one known strategy for minimizing NO x  emissions is referred to as rich burn, quick quench, lean burn (RQL) combustion. The RQL strategy recognizes that the conditions for NO x  formation are most favorable at elevated combustion flame temperatures, such as when a fuel-air ratio is at or near stoichiometric. A combustor configured for RQL combustion includes three serially arranged combustion zones: a Rich burn zone at the forward end of the combustor, a Quench or dilution zone axially aft of the rich burn zone, and a Lean burn zone axially aft of the quench zone. 
     During engine operation, a portion of the pressurized air discharged from the compressor enters the rich burn zone of the combustion chamber. Concurrently, the fuel injectors introduce a stoichiometrically excessive quantity of fuel into the rich burn zone. Although the resulting stoichiometrically fuel rich fuel-air mixture is ignited and burned to partially release the energy content of the fuel, NO x  formation may still occur. 
     The fuel rich combustion products then enter the quench zone where jets of pressurized air radially enter through combustion air holes from the compressor and into the quench zone of the combustion chamber. The pressurized air mixes with the combustion products to support further combustion of the fuel with air by progressively deriching the fuel rich combustion products as they flow axially through the quench zone and mix with the air. Initially, the fuel-air ratio of the combustion products changes from fuel rich to stoichiometric, causing an attendant rise in the combustion flame temperature. Since the quantity of NO x  produced in a given time interval is known to increase exponentially with flame temperature, quantities of NO x  may be produced during the initial quench process. As the quenching continues, the fuel-air ratio of the combustion products changes from stoichiometric to fuel lean, causing an attendant reduction in the flame temperature. However, until the mixture is diluted to a fuel-air ratio substantially lower than stoichiometric, the flame temperature remains high enough to generate NO x . 
     Finally, the deriched combustion products from the quench zone flow axially into the lean burn zone. Additional pressurized air in this zone supports ongoing combustion to release energy from the fuel. The additional pressurized air in this zone also regulates the peak temperature and spatial temperature profile of the combustion products to reduce turbine exposure to excessive temperatures and excessive temperature gradients. 
     SUMMARY 
     A combustor for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes an forward fuel injection system in communication with a combustion chamber, and a downstream fuel injection system that communicates with said combustion chamber downstream of said forward fuel injection system. 
     In a further embodiment of the foregoing embodiment, the downstream fuel injection system at least partially surrounds the combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially inboard of the combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially outboard of the combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially outboard and radially inboard of the combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system includes a multiple of fuel nozzle assemblies axially upstream of a necked region of the combustor. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system includes a multiple of fuel nozzle assemblies within a first two-thirds of the combustor. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially inboard of the combustion chamber, a main supply line of a radially inner fuel injection manifold extends through a forward assembly. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially inboard of the combustion chamber, a main supply line of a radially inner fuel injection manifold extends through a downstream vane 
     A gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes an forward fuel injection system in communication with a combustion chamber and a downstream fuel injection system around said combustion chamber, said downstream fuel injection system communicates with said combustion chamber downstream of said forward fuel injection system. 
     In a further embodiment of the foregoing embodiment, the downstream fuel injection system is radially inboard of said combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially outboard of said combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system is radially outboard and radially inboard of said combustion chamber. 
     In a further embodiment of any of the foregoing embodiments, the downstream fuel injection system includes a multiple of fuel nozzle assemblies axially upstream of a necked region of said combustor. In the alternative or additionally thereto, in the foregoing embodiment the downstream fuel injection system includes a multiple of fuel nozzle assemblies within a first two-thirds of said combustor. In the alternative or additionally thereto, in the foregoing embodiment the downstream fuel injection system is radially inboard of said combustion chamber, a main supply line of a radially inner fuel injection manifold extends through a forward assembly. In the alternative or additionally thereto, in the foregoing embodiment the downstream fuel injection system is radially inboard of said combustion chamber, a main supply line of a radially inner fuel injection manifold extends through a downstream vane. 
     A method of communicating fuel to a combustor of a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure includes communicating fuel axially into a combustion chamber and communicating fuel radially into the combustion chamber. 
     In a further embodiment of the foregoing embodiment, the method includes communicating fuel radially inward into the combustion chamber. 
     In a further embodiment of the foregoing embodiment, the method includes communicating fuel radially outward into the combustion chamber. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of a gas turbine engine; 
         FIG. 2  is a partial longitudinal schematic sectional view of an exemplary annular combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a partial lateral schematic sectional view of an exemplary annular combustor of  FIG. 2 ; and 
         FIG. 4  is a partial longitudinal schematic sectional view of an exemplary annular combustor according to another non-limiting embodiment, that may be used with the gas turbine engine shown in  FIG. 1 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT as well as aero-derivative/electrical power engine applications. 
     The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing structures  38 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  (“LPC”) and a low pressure turbine  46  (“LPT”). The inner shaft  40  drives the fan  42  directly or through a geared architecture  48  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
     The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  (“HPC”) and high pressure turbine  54  (“HPT”). A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed with the fuel and burned in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines  54 ,  46  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. 
     The main engine shafts  40 ,  50  are supported at a plurality of points by bearing structures  38  within the static structure  36 . It should be understood that various bearing structures  38  at various locations may alternatively or additionally be provided. 
     In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the low pressure compressor  44  and low pressure turbine  46  and render increased pressure in a fewer number of stages. 
     A pressure ratio associated with the low pressure turbine  46  is pressure measured prior to the inlet of the low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about 5 (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     In one embodiment, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of “T”/518.70.5. in which “T” represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     With reference to  FIG. 2 , the combustor  56  generally includes a combustor outer liner  60  and a combustor inner liner  62 . The outer liner  60  and the inner liner  62  are spaced inward from a diffuser case  64  such that a combustion chamber  66  is defined therebetween. The combustion chamber  66  is generally annular in shape and is defined between combustor liners  60 ,  62 . 
     The outer liner  60  and the diffuser case  64  define an outer annular plenum  76  and the inner liner  62  and the case  64  define an inner annular plenum  78 . It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner panel arrangements will also benefit herefrom. 
     Each liner  60 ,  62  generally includes a respective support shell  68 ,  70  that supports one or more respective liner panels  72 ,  74  mounted to a hot side of the respective support shell  68 ,  70 . Each of the liner panels  72 ,  74  may be generally rectilinear and manufactured of, for example, a nickel based super alloy or ceramic material. 
     The combustor  56  further includes a forward assembly  80  immediately downstream of the compressor section  24  to receive compressed airflow therefrom. The forward assembly  80  generally includes an annular hood  82 , a bulkhead assembly  84 , a multiple of axial fuel nozzles  86  (one shown; illustrated schematically) and a multiple of swirler assemblies  90  (one shown; illustrated schematically) that define a central opening. The annular hood  82  extends radially between, and is secured to, the forwardmost ends of the liners  60 ,  62 . The annular hood  82  includes a multiple of circumferentially distributed hood ports  82 P that accommodate the respective fuel nozzle  86  and introduces air into the forward end of the combustion chamber  66 . The centerline of the fuel nozzle  86  is concurrent with the centerline F of the respective swirler assembly  90 . Each fuel nozzle  86  may be secured to the diffuser case  64  to project through one of the hood ports  82 P and through the central opening  90 A within the respective swirler assembly  90 . It should be understood that some combustors, such as lean or front-end staged combustors, may have more complex front end geometries in which fuel nozzles may be oriented other than in a circumferential pattern. 
     Each swirler assembly  90  is circumferentially aligned with, and/or concentric to, one of the hood ports  82 P to project through the bulkhead assembly  84 . Each bulkhead assembly  84  includes a bulkhead support shell  84 S secured to the liners  60 ,  62 , and a multiple of circumferentially distributed bulkhead heatshields segments  98  secured to the bulkhead support shell  84 S around the central opening  90 A. 
     The forward assembly  80  directs a portion of the core airflow into the forward end of the combustion chamber  66  while the remainder enters the outer annular plenum  76  and the inner annular plenum  78 . The multiple of axial fuel nozzles  86 , swirler assemblies  90  and associated fuel communication structure defines a forward fuel injection system  92  that supports combustion in the combustion chamber  66 . 
     A downstream fuel injection system  94  communicates with the combustion chamber  66  downstream of the forward fuel injection system  92 . The downstream fuel injection system  94  introduces a portion of the fuel required for desired combustion performance, e.g., emissions, operability, durability as well as to lean-out the fuel contribution provided by the multiple of axial fuel nozzles  86  generally parallel to axis F. 
     The downstream fuel injection system  94  generally includes a radially outer fuel injection manifold  96  located in the outer annular plenum  76  and/or a radially inner fuel injection manifold  98  located in the inner annular plenum  78 . It should be appreciated that the downstream fuel injection system  94  may include only the radially outer fuel injection manifold  96 ; only the radially inner fuel injection manifold  98  or both (shown). 
     The radially outer fuel injection manifold  96  may be mounted to the diffuser case  64 . Alternatively, the radially outer fuel injection manifold  96  may be mounted to the shell  68 . The radially inner fuel injection manifold  98  may be mounted to the diffuser case or shell  70 . It should be appreciated that various mount arrangements may alternatively or additionally provided such as location of the outer fuel injection manifold  96  mounted inside or outside the diffuser case  64 . 
     The radially outer fuel injection manifold  96  and the radially inner fuel injection manifold  98  may be manufactured of a series of straight tube sections  96 T,  98 T that may be connected together by a series of joints or fittings via braze or weld methods ( FIG. 3 ). It should be appreciated that various assembly methods and component structures may be alternatively or additionally be provided. 
     The radially outer fuel injection manifold  96  includes a multiple of radially extending supply lines  100  which terminate in an outer fuel nozzle assembly  102  that project predominantly radially toward the centerline F of the combustor chamber  66 . The multiple of radially extending supply lines  100  may include, for example, compliant fuel lines or pigtails that accommodate relative growth and part movement. In one disclosed non-limiting embodiment, the outer fuel nozzle assembly  102  includes fuel injector ports  104 A encased by an air swirler  106 A that promote mixing of the fuel spray with air from within the diffuser case  64  to facilitate generation of the fuel-air distribution required for combustion. 
     The radially inner fuel injection manifold  98  likewise includes a multiple of radially extending supply lines  108  which terminate in an inner fuel nozzle assembly  110  that project predominantly radially toward the centerline F of the combustor chamber  66 . The multiple of radially extending supply lines  108  may include, for example, compliant fuel lines or pigtails that accommodate relative growth and part movement. In one disclosed non-limiting embodiment, the inner fuel nozzle assembly  110  include fuel injector ports  104 B encased by an air swirler  106 B that promote mixing of the fuel spray with air from within the diffuser case  64  to facilitate generation of the fuel-air distribution required for combustion. 
     The radially inner fuel injection manifold  98  includes a main supply line  112  which may be arranged to pass through the relatively cooler forward assembly  80  to provide communication with the multiple of radially extending supply lines  108 . Alternatively, the main supply line  112  may pass through a downstream vane  114  such as a Nozzle Guide Vane ( FIG. 4 ). It should be appreciated that the main supply line  112  may be a secondary or intermediary fuel line to, for example, facilitate assembly. 
     Given operational temperatures from the HPC  52 , the radially outer fuel injection manifold  96  and the radially inner fuel injection manifold  98  may be subject to soaking temperatures that may promote coking. The radially outer fuel injection manifold  96  and the radially inner fuel injection manifold  98  and other associated lines may be configured with a protective, low-conductivity sheath, a coating, a cooled tube-in-tube construction, be relatively oversized compared to fuel flow or other insulation that provides thermal resistance between the relatively hot air temperatures in the diffuser case  64  and the relatively cold fuel temperatures in the fuel lines, manifolds and nozzles. Alternatively, or in addition, the downstream fuel injection system  94  may communicate through or with the bypass stream of the engine and may include a thermal management or heat exchange system to further maintain low fuel temperatures. 
     The outer and inner fuel nozzle assemblies  102 ,  110  project through openings in the combustor  56  to supply fuel to the combustor between the bulkhead assembly  84  and a combustor exit  66   x . In one disclosed non-limiting embodiment, the outer and inner fuel nozzle assemblies  102 ,  110  project through openings in the combustor  56  located within the first two-thirds of the combustor chamber  66 . In another disclosed non-limiting embodiment, the outer and inner fuel nozzle assemblies  102 ,  110  project through openings in the combustor  66  between 20-70% of the axial length. In another disclosed non-limiting embodiment, the outer and inner fuel nozzle assemblies  102 ,  110  project through openings in the combustor  66  upstream of a necked region  56 N of the combustor  56 . That is, an internal height of the bulkhead assembly  84  is greater than the combustor exit  66   x.    
     Spark energy may be provided to the combustor  56  through a frequency-pulsed igniter arrangement  116  (illustrated schematically) which provides a continuous spark or other ignition source. The frequency-pulsed igniter arrangement  116  may be located in conventional as well as other locations within the combustor  56 . 
     The fuel required for combustion is, thus, provided by the both the axial fuel nozzles  86  and the fuel nozzles  102 ,  110  associated with the radially outer fuel injection manifold  96  and the radially inner fuel injection manifold  98 . The distributed fuel injection and fuel-air mixing provided thereby may be tailored to optimize emissions, e.g., NOx, COx, smoke, particulates, etc., as well as control of combustor thermals, durability, profile and pattern factors that impact the downstream turbine section. 
     It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.