Abstract:
A fuel system for a gas turbine engine includes a plurality of duplex nozzles arranged on each side of top dead center and a plurality of simplex nozzles. A primary manifold is operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles and a secondary manifold is operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. An equalizer valve that is in communication with both the primary manifold and the secondary manifold distributes fuel at various pressures to both the primary and secondary manifolds.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation in part of U.S. application Ser. No. 13/301,856 filed on Nov. 22, 2011 and claims priority to U.S. Provisional Application No. 61/706,908 filed on Sep. 28, 2012. 
     
    
     BACKGROUND 
       [0002]    Gas turbine engines, such as those which power modern commercial and military aircraft, generally include a compressor for pressurizing an airflow, 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 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 nozzles project into a forward section of the combustion chamber through a respective nozzle guide to supply the fuel to be mixed with the pressurized air. 
         [0003]    It remains desirable for gas turbine engine manufacturers to develop combustor configurations that reduce emissions and noise with improved operational efficiencies. 
       SUMMARY 
       [0004]    A fuel system for a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a plurality of duplex nozzles arranged on each side of top dead center, a plurality of simplex nozzles, a primary manifold operable to communicate fuel to a primary flow jet in each of the plurality of duplex nozzles, a secondary manifold operable to communicate fuel to a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles, and an equalizer valve in communication with the primary manifold and the secondary manifold. The equalizer valve is movable between an open position and a closed position, the closed position is operable to permit a supply of fuel pressure to the primary manifold that is greater than fuel pressure to the secondary manifold, and the open position is operable to permit supply of fuel pressure to the primary manifold with essentially an equal fuel pressure as the secondary manifold. 
         [0005]    In a further embodiment of the foregoing fuel system, the plurality of duplex nozzles are arranged with respect to a fuel igniter. 
         [0006]    In a further embodiment of any of the foregoing fuel systems, at least one of the plurality of duplex nozzles are arranged adjacent to a fuel igniter. 
         [0007]    In a further embodiment of any of the foregoing fuel systems, at least one of the plurality of duplex nozzles are arranged opposite the fuel igniter. 
         [0008]    In a further embodiment of any of the foregoing fuel systems, the primary flow jet in each of the plurality of duplex nozzles is surrounded by the secondary flow jet. 
         [0009]    In a further embodiment of any of the foregoing fuel systems, the plurality of duplex nozzles include ten (10) duplex nozzles and the plurality of simplex nozzles includes six (6) simplex nozzles. 
         [0010]    In a further embodiment of any of the foregoing fuel systems, the plurality of simplex nozzles includes simplex nozzles arranged on either side of a bottom dead center position. 
         [0011]    A method of noise control from a combustor of a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor including forming a high fuel-air ratio circumferential zone at a top dead center position and forming a low fuel-air ratio circumferential zone at a bottom dead center position. 
         [0012]    In a further embodiment of the foregoing method, further including alternating the local circumferential zones with varied fuel-air ratios. 
         [0013]    In a further embodiment of any of the foregoing methods, further including forming the local circumferential zones as a high-low-high-low local fuel-air ratios. 
         [0014]    In a further embodiment of any of the foregoing methods, further including locating at least one of a plurality of duplex nozzles adjacent to a fuel igniter to form at least one high local fuel-air ratio within at least one of the plurality of circumferential zones. 
         [0015]    In a further embodiment of any of the foregoing methods, further including locating at least one of the plurality of duplex nozzles opposite the fuel igniter to form at least one high local fuel-air ratio within at least one of the plurality of circumferential zones. 
         [0016]    In a further embodiment of any of the foregoing methods, further includes alternating a plurality of duplex nozzles and a plurality of simplex nozzles to define the plurality of circumferential zones wherein the plurality of duplex nozzles includes ten (10) duplex nozzles and the plurality of simplex nozzles includes six (6) simplex nozzles. 
         [0017]    In a further embodiment of any of the foregoing methods, further including selectively equalizing a fuel pressure between a primary manifold and a secondary manifold, the primary manifold in communication with a primary flow jet in each of the plurality of duplex nozzles and the secondary manifold in communication with a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. 
         [0018]    In a further embodiment of any of the foregoing methods, further including selectively opening a valve between the primary manifold and the secondary manifold. 
         [0019]    In a further embodiment of any of the foregoing methods, further including selectively equalizing the fuel pressure in response to a non-low power conditions. 
         [0020]    In a further embodiment of any of the foregoing methods, further including selectively dividing a fuel pressure between a primary manifold and a secondary manifold, the primary manifold in communication with a primary flow jet in each of the plurality of duplex nozzles and the secondary manifold in communication with a secondary flow jet in each of the plurality of duplex nozzles and a secondary flow jet in each of the plurality of simplex nozzles. 
         [0021]    In a further embodiment of any of the foregoing methods, further including selectively dividing the fuel pressure in response to a low power conditions. 
         [0022]    In a further embodiment of any of the foregoing methods, the low power condition includes the power required for approach conditions and a margin above at least partially into a cruise condition. 
         [0023]    In a further embodiment of any of the foregoing methods, further including selectively forming the plurality of local circumferential zones with varied fuel-air ratios within the combustor in response to a low power condition. 
         [0024]    Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
         [0025]    These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]    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: 
           [0027]      FIG. 1  is a schematic cross-section of a gas turbine engine; 
           [0028]      FIG. 2  is a partial sectional view of an exemplary annular combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
           [0029]      FIG. 3  is a schematic view of the exemplary combustor; 
           [0030]      FIG. 4  is a schematic view of a duplex nozzle; 
           [0031]      FIG. 5  is a schematic view of a simplex nozzle; 
           [0032]      FIG. 6  is a schematic view of a fuel system for the exemplary combustor; and 
           [0033]      FIG. 7  is a schematic view of a method of operation of the fuel system. 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      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. 
         [0035]    The engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
         [0036]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  and high pressure turbine  54 . 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. 
         [0037]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel within 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 speed spool  30  and high speed spool  32  in response to the expansion. 
         [0038]    With reference to  FIG. 2 , the combustor  56  generally includes an outer liner  60  and an inner liner  62  disposed within a combustor case  64 . An annular combustion chamber  66  is defined between the outer liner  60  and the inner liner  62 . It should be understood that although a particular combustor is illustrated, other combustor types with various liner panel arrangements will also benefit herefrom. 
         [0039]    The outer liner  60  and the combustor case  64  define an outer annular plenum  76  and the inner liner  62  and the combustor case  64  define an inner annular plenum  78 . The liners  60 ,  62  contain the flame for direction toward the turbine section  28 . Each liner  60 ,  62  generally includes a support shell  68 ,  70  which supports one or more liner panels  72 ,  74  mounted to a hot side of the respective support shell  68 ,  70 . The liner panels  72 ,  74  define a liner panel array which may be generally annular in shape. Each of the liner panels  72 ,  74  may be generally rectilinear and manufactured of, for example, a nickel based super alloy or ceramic material. 
         [0040]    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 nozzles  86  (one shown) and a multiple of nozzle guides  90  (one shown) that defines a central opening  92 . 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  94  that accommodate the respective nozzle  86  and introduce air into the forward end of the combustion chamber  66 . Each nozzle  86  projects through one of the hood ports  94  and through the central opening  92  within the respective nozzle guide  90  along a nozzle axis F. 
         [0041]    With reference to  FIG. 3 , the multiple of nozzles  86  (sixteen shown) are arranged to supply fuel to the combustor  56 . The multiple of nozzles  86  and surrounding structure generate a swirling, intimately blended fuel-air mixture that supports combustion in the forward section of the combustion chamber  66 . 
         [0042]    Airflow in the combustor  56  is generally uniform and fuel-air variation in the disclosed non-limiting embodiment is selectively generated by the multiple of nozzles  86 . More specifically, the variation is selectively generated through a plurality of duplex nozzles  86 D (one shown in  FIG. 4 ) and a plurality of simplex nozzles  86 S (one shown in  FIG. 5 ). 
         [0043]    In the example embodiment, ten (10) duplex nozzles  86 D and six (6) simplex nozzles  86 S are provided to provide the desired uniformity at high power settings. Moreover, duplex nozzles  86 D provided at the top dead center TDC location, schematically shown at  16 -D and  1 -D provide enhanced stability and to facilitate starting. The example disclosed configuration including the duplex nozzles  86 D at the  16 -D and  1 -D positions provides improvements in noise and tones generated by the combustor  56  during operation. 
         [0044]    With reference to  FIG. 4 , each of the duplex nozzles  86 D include a primary flow jet  100  and secondary flow jets  102 . The primary flow jet  100  is defined generally along axis F and the secondary flow jets  102  are generally transverse to axis F. Each of the simplex nozzles  86 S include only the secondary flow jets  102  ( FIG. 5 ). It should be appreciated that various jet arrangements may alternatively or additionally be provided. 
         [0045]    With reference to  FIG. 6 , a fuel system  106  communicates fuel to the multiple of nozzles  86 . The fuel system generally includes a fuel metering unit (FMU)  108 , a flow divider valve (FDV)  110 , a primary manifold  112 , a secondary manifold  114 , and an equalizing valve  116  which selectively permits fuel communication between the primary manifold  112  and the secondary manifold  114 . The fuel metering unit (FMU)  108  is a hydromechanical unit that controls fuel flow and the flow divider valve (FDV)  110  proportions the fuel flow to the primary manifold  112  and the secondary manifold  114 . The primary manifold  112  communicates fuel to the primary flow jet  100  in each of the duplex nozzles  86 D and the secondary flow manifold  114  communicates fuel to the secondary flow jets  102  in each of the duplex nozzles  86 D and each of the simplex nozzles  86 S. 
         [0046]    A module  120  executes a fuel control algorithm  122  ( FIG. 7 ). The functions of the algorithm  122  are disclosed in terms of functional block diagrams, and it should be understood by those skilled in the art with the benefit of this disclosure that these functions may be enacted in either dedicated hardware circuitry or programmed software routines capable of execution in a microprocessor based electronics control embodiment. In one non-limiting embodiment, the module  120  may be a portion of a flight control computer, a portion of a Full Authority Digital Engine Control (FDEC) or other system. 
         [0047]    The module  120  typically includes a processor, a memory, and an interface. The processor may be any type of known microprocessor having desired performance characteristics. The memory may, for example only, include computer readable medium which stores the data and control algorithms described herein. The interface facilitates communication with the fuel metering unit (FMU)  108 , the equalizing valve  116 , as well as other avionics and systems. 
         [0048]    Reduction in tone amplitude has been demonstrated through the formation of local circumferential zones with different fuel-air ratio mixtures. The different fuel-air ratio mixtures in the alternating circumferential zones as defined by the nozzles  86 D,  86 S ( FIG. 3 ) vary the delay time of heat release, and consequently provide a differential coupling to the associated naturally occurring acoustic frequencies. That is, high-low-high-low local fuel-air ratios are defined about the circumference of the combustor  56 . 
         [0049]    The fuel system  106  locates the duplex nozzles  86 D adjacent to a set of fuel igniters  124  and opposite the set of fuel igniters  124  ( FIG. 3 ). It should be appreciated that the fuel igniters  124  may be located in other circumferential positions and the duplex nozzles  86 D would be adjusted in accordance therewith. That is, if the fuel igniters  124  were mounted at bottom dead center (BDC) at positions  8  and  9 , for example, the duplex nozzles  86 D would be located in positions  7 - 10  and  15 - 2 . In the disclosed, non-limiting embodiment of sixteen (16) nozzles  86  where each of the nozzles  86  are separated by 22.5 degrees, the duplex nozzles  86 D are located in positions  1 ,  4 - 7  and  12 - 16  while positions  8 - 11 ,  2  and  3  utilize simplex nozzles  86 S. The example duplex nozzles  86 D are therefore located on either side of the top dead center (TDC) at positions  16  and  1 . Moreover, in the disclosed example, simplex nozzles  86 S are arranged on either side of BDC at positions  8  and  9 . 
         [0050]    At low power conditions, the equalizing valve  116  is closed such that the primary manifold  112  is provided with greater fuel pressure than the secondary manifold  114  to drive a fuel flow distortion. The increased fuel pressure drop (and fuel flow) increases the overall (primary plus secondary) fuel flow to the duplex nozzles  86 D in relation to the simplex nozzles  86 S. That is, when the equalizing valve  116  is closed, the duplex nozzles  86 D generate the relatively high fuel-air ratio mixtures and the simplex nozzles  86 S provide the relatively low fuel-air ratio mixtures. The varied fuel-air ratio mixtures dampen tangential and axial pressure waves within the combustor  56  to control combustor tones and enhance combustor stability. In one example in which the FMU  108  supplies fuel at approximately 100 psi at a low power condition and approximately 1200 psi at a high power condition, an approximate 50-150 psi difference is provided between the duplex nozzles  86 D and the simplex nozzles  86 S when the equalizing valve  116  is closed. It should be appreciate that “low power” as defined herein may include moderate power such as that required for approach conditions and a margin above at least partially into a cruise condition. 
         [0051]    At non-low power conditions, the equalizing valve  116  is open such that the primary manifold  112  and the secondary flow manifold  114  receive equalized flow such that the duplex nozzles  86 D and the simplex nozzles  86 S generate a symmetric uniform fuel-air ratio throughout the combustor  56 . It should be appreciated that “non-low power” as defined herein may include cruise power conditions and above such as a take-off flight condition. 
         [0052]    The equalizing valve  116  allows the selection of fuel asymmetry at low power conditions where combustor tones predominate and reversion to symmetric operation at high power conditions where uniform fuel-air ratio mixture distribution is desired. The local fuel-air mixture ratio control further facilitates enhanced stability for snap transient decelerations. Selective operation of the equalizing valve  116  thereby reduces tone amplitudes over, for example, idle and approach power settings so that community and aircraft cabin noise are minimized yet selectively permits symmetric high power operation during, for example, cruise power settings to enhance downstream turbine durability. Additionally, fuel flow through the primary manifold may be increased to further mitigate specific tones and noise frequencies that may be generated during operation. 
         [0053]    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. 
         [0054]    Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
         [0055]    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. 
         [0056]    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. 
         [0057]    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.