Patent Publication Number: US-2013232978-A1

Title: Fuel air premixer for gas turbine engine

Description:
BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a fuel-air premixer therefor. 
     Gas turbine engines, such as those powering modern commercial and military aircraft, 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 fuel air premixer project into a forward section of the combustion chamber to supply fuel mixed with pressurized air. 
     Future gas turbine combustors may be required to meet aggressive emission requirements, particularly NOx. Combustor schemes under development to achieve these goals will require a high degree of fuel/air mixing prior to combustion. The high combustor inlet temperatures may thereby lead to short autoignition time scales such that the fuel/air mixing must occur in a very short time. 
     SUMMARY 
     A fuel-air premixer for a combustor of a turbine engine according to one aspect of the present disclosure includes a central passage along an axis. The central passage is operable to communicate an unswirled airflow. An outer annular passage is located around the axis. The outer annular passage includes a multiple of first swirl vanes that are operable to communicate a first swirled airflow in a first direction. An inner annular passage is located around the axis between the central passage and the outer annular passage. The inner annular passage includes a multiple of second swirl vanes that are operable to communicate a second swirled airflow in a second direction different than the first direction. 
     In a further embodiment of the above, the multiple of first swirl vanes and the multiple of second swirl vanes define respective chord directions C1 and C2 that are transverse to the axis and to each other. 
     In a further embodiment of any of the above, the fuel-air premixer includes a swirler body extending around the outer annular passage, the inner annular passage, and the central passage. 
     In a further embodiment of any of the above, the multiple of first swirl vanes extend between the swirler body and a splitter plate, the multiple of second swirl vanes extend between the splitter plate and a central nozzle body that defines the central passage. 
     In a further embodiment of any of the above, the swirler body radially contracts downstream of the central nozzle body. 
     In a further embodiment of any of the above, the central nozzle body includes a multiple of fuel passages disposed radially outboard and parallel to the central passage. 
     In a further embodiment of any of the above, the multiple of fuel passages include respective fuel orifices that each extend in a radial direction. 
     In a further embodiment of any of the above, the fuel orifices are located downstream of the splitter plate. 
     In a further embodiment of any of the above, the fuel orifices are located downstream of the splitter plate at an axial distance greater than approximately five (5) diameters of the fuel orifices. 
     In a further embodiment of any of the above, the swirler body radially contracts downstream of the central nozzle body. 
     In a further embodiment of any of the above, the swirler body is axially spaced from the central nozzle body. 
     A gas turbine engine according to one aspect of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. The combustor includes a fuel-air premixer having a central passage along an axis. The central passage is operable to communicate an unswirled airflow. An outer annular passage is located around the axis. The outer annular passage includes a multiple of first swirl vanes that are operable to communicate a first swirled airflow in a first direction. An inner annular passage is located around the axis between the central passage and the outer annular passage. The inner annular passage includes a multiple of second swirl vanes that are operable to communicate a second swirled airflow in a second direction different than the first direction. 
     A method of communicating fuel and air to a combustor of a turbine engine according to one aspect of the present disclosure includes communicating an unswirled airflow along an axis, communicating a first swirled airflow in a first direction around the unswirled airflow, communicating a second swirled airflow in a second direction different than the first direction forming a turbulent region and injecting fuel into the turbulent region. 
     A further embodiment of the above includes radially injecting the fuel outward relative to the axis into the turbulent region at a location approximately equivalent to a one-quarter auto-ignition time relative to an end of a swirler body. 
     A further embodiment of any of the above includes choking the unswirled airflow, the first swirled airflow and the second swirled airflow downstream of the turbulent region. 
     A further embodiment of any of the above includes providing essentially zero net swirl at an end of a swirler body that receives the unswirled airflow, the first swirled airflow and the second swirled airflow. 
    
    
     
       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 perspective partial sectional view of an exemplary annular combustor that may be used with the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a sectional view of a fuel air premixer; and 
         FIG. 4  is a schematic view of a premixer length with respect to an auto-ignition time. 
     
    
    
     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 and land-based engines. 
     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. 
     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. 
     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. 
     With reference to  FIG. 2 , the combustor  56  generally includes an outer combustor liner  60  and an inner combustor liner  62 . The outer combustor liner  60  and the inner combustor liner  62  are spaced inward from a combustor case  64  such that a combustion chamber  66  is defined there between. The combustion chamber  66  is generally annular in shape and is defined between combustor liners  60 ,  62 . 
     The outer combustor liner  60  and the combustor case  64  define an outer annular plenum  76  and the inner combustor liner  62  and the combustor 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. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto. 
     The combustor liners  60 ,  62  contain the flame for direction toward the turbine section  28 . Each combustor 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. 
     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 premixers  86  (one shown) and a multiple of fuel nozzle guides  90  (one shown) along an opening  92 . In this example, the premixer  86  is shown at the forward assembly  80 . It is to be understood, however, that the location and size of the premixers  86  can vary depending on the particular design of the combustor. Put another way, the size and number of premixers  86  can be varied from the illustrated example and one or more premixers  86  can additionally or alternatively be located through either or both of the support shells  68 ,  70 . 
     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 fuel-air premixer  86  and introduce air into the forward end of the combustion chamber  66 . Each fuel-air premixer  86  may be secured to the outer case  64  and projects through one of the hood ports  94  and through the opening  92  of the respective fuel nozzle guide  90 . It should be understood that various additional or alternative structure may also be utilized. 
     Each of the fuel nozzle guides  90  is circumferentially aligned with one of the hood ports  94  to project through the bulkhead assembly  84 . Each bulkhead assembly  84  includes a bulkhead support shell  96  secured to the liners  60 ,  62 , and a multiple of circumferentially distributed bulkhead heat shield segments  98  secured to the bulkhead support shell  96  around the central opening  92 . 
     The forward assembly  80  introduces primary core combustion air 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 premixers  86  and surrounding structure generate a swirling, intimately blended fuel-air mixture that supports combustion in the combustion chamber  66 . 
     With reference to  FIG. 3 , each of the multiple of fuel-air premixers  86  include a swirler body  100 , a central nozzle body  102 , a main nozzle body  104 , a multiple of fuel conduits  106 , a splitter plate  108 , a multiple of first swirl vanes  110  and a multiple of second swirl vanes  112 . The multiple of fuel conduits  106  extend between the central nozzle body  102  and the main nozzle body  104  to permit airflow into the central nozzle body  102 . That is, the multiple of fuel conduits  106  are essentially tubes which are of minimal cross-sectional area so as to not block airflow thereby. The swirler body  100  includes a convergent section  100 C that generally corresponds with a frustro-conical end  102 A of the nozzle body  102  to maintain relatively high airflow velocities to prevent flashback. It should be appreciated that additional or alternative components may also be utilized. 
     A central passage  114  is defined along a premixer axis P within the central nozzle body  102 . The central passage  114  facilitates the communication of unswirled airflow into the swirler body  100 . An outer annular passage  116  is defined around the axis P between the swirler body  100  and the splitter plate  108 . The multiple of first swirl vanes  110  extend between the swirler body  100  and the splitter plate  108  to facilitate the communication of a first swirled airflow into the swirler body  100 . An inner annular passage  118  is defined around the axis P between the splitter plate  108  and the central nozzle body  102 . The multiple of second swirl vanes  112  extend between the splitter plate  108  and the central nozzle body  102  to facilitate the communication of a second swirled airflow into the swirler body  100 . 
     The multiple of first swirl vanes  110  and the multiple of second swirl vanes  112  define respective chord directions C1 and C2 (shown schematically) that are transverse to the axis P and to each other. The transverse orientation generates annular swirled airflows. In this example, the airflows generated by the first swirl vanes  110  and the second swirl vanes  112  have component vectors in opposite directions to generate a highly turbulent region in the vicinity of an orifice  120  from each of a multiple of fuel passages  122  within the fuel conduits  106  and central nozzle body  102 . For example, the multiple of first swirl vanes  110  generate annular swirled flow in a clockwise direction and the multiple of second swirl vanes  112  generate annular swirled flow in a counter-clockwise direction, or vice versa. The highly turbulent region facilitates the mixing of the radially injected fuel but then the counter-swirling essentially cancel each other out to provide an essentially zero net swirl at an end  100 A of the swirler body  100 . 
     The fuel orifices  120  from each of the multiple of fuel passages  122  extend in a radial direction toward the highly turbulent region. The orifices  120  are axially located at a location L ( FIG. 4 ) approximately equivalent to a one-quarter (¼) auto-ignition time relative to an end  100 A of the swirler body  100  or, described another way, the orifices  120  are located downstream of the splitter plate  108  at an axial distance greater than approximately five (5) diameters of one of the orifices  120 , assuming that the orifices  120  are of equivalent diameters. If the orifices  120  are not of equivalent diameters, the five (5) diameter axial distance can be with regard to an average diameter of the orifices  120 , smallest diameter of the orifices or largest diameter of the orifices, for example. 
     The given location L limits or essentially eliminates the potential for the fuel/air mixture to auto-ignite within the premixer  86 . High-efficiency engines may operate at high pressure ratios, and may have high combustor inlet temperatures. These high temperatures can lead to reduced auto-ignition times, which represent a practical limit to the residence time of the fuel/air mixture in the premixer  86 . However, the given location L that corresponds to a one-quarter (¼) auto-ignition ensures that the residence time is well below the auto-ignition time. 
     The fuel-air premixer  86  achieves effective fuel/air mixing through injection of the fuel jets F into the high-shear region R where the counter-swirling air flows meet. This mixing layer is characterized by high levels of turbulence, which operate to further atomize the fuel into small droplets and to disperse those droplets through the swirler body  100 . Small droplets evaporate quickly, and once the fuel has been vaporized, the turbulent air flow acts to mix the fuel vapor with the air. 
     The fuel-air premixer  86  efficiently mixes liquid fuel with air by the end of the swirler body  100  to enable low pollutant emissions through minimization of lean or rich excursions from the design fuel/air ratio that would lead to higher emissions. The liquid fuel is mostly vaporized by the end  100 A of the premixer  86 , especially at high power operating conditions to improve mixing and reduce pollutant. The mixing of the fuel and air generated by the swirl vanes  110  and  112  also limits or eliminates flame “flash back” into the premixer  86 . That is, the highly turbulent region in the vicinity of the orifice  120  from the fuel passages  122  within the fuel conduits  106  and central nozzle body  102  and cancellation of the counter-swirling to provide an essentially zero net swirl at an end  100 A of the swirler body  100  limit or eliminate regions of low or negative velocity that would otherwise allow the flame to propagate from the combustor upstream into the premixer, where significant damage may be the result. 
     The premixer  86  is also relatively uncomplicated to manufacture, simple and scaleable as many low-emissions combustor designs use a large number of premixers in a staged array. 
     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. 
     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. 
     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.