Abstract:
A fuel injection system for a gas turbine engine includes a vane in an airflow path within the gas turbine engine, the vane includes an air channel with an outlet in communication with the airflow path; and a fuel nozzle within the vane operable to inject fuel into the air channel to at least partial premix and prevaporize the fuel with a secondary airflow from within the vane in the air channel prior to entry into the airflow path through the outlet. A method of injecting fuel within a gas turbine engine includes at least partially premixing and prevaporizing fuel with a secondary airflow from within a vane in an air channel within the vane, the vane within an airflow path of the gas turbine engine.

Description:
[0001]    The present disclosure claims priority to U.S. Provisional Patent Disclosure Ser. No. 61/754,365, filed Jan. 18, 2013. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
       [0002]    This disclosure was made with Government support under FA8650-11-M-2157 awarded by The United States Air Force. The Government has certain rights in this invention. 
     
    
     BACKGROUND 
       [0003]    The present disclosure relates to gas turbine engines, and more particularly to a fuel injection system therefor. 
         [0004]    Gas turbine engines, such as those which power modern aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust. On military engines, downstream of the turbine section, an augmentor section, or “afterburner”, is operable to selectively increase the thrust. The increase in thrust is produced when fuel is injected into the core exhaust gases downstream of the turbine section and burned with the oxygen contained therein with the aid of flameholders to generate a second combustion. 
         [0005]    Typically, the injected fuel is controlled to penetrate relatively deep into the core exhaust gases to provide good mixing and increase augmentor efficiency as well as the magnitude of the supplemental engine thrust. Such deep fuel penetration, however, is dependent on the fuel flow rate which may negatively impact flame stability and increase augmentor instabilities commonly called “screech” as the fuel penetrates away for the flameholder device. 
       SUMMARY 
       [0006]    A fuel injection system for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a vane in an airflow path within the gas turbine engine, the vane includes an air channel with an outlet in communication with the airflow path; and a fuel nozzle within the vane operable to inject fuel into the air channel to at least partial premix and prevaporize the fuel with a secondary airflow from within the vane in the air channel prior to entry into the airflow path through the outlet. 
         [0007]    A further embodiment of the present disclosure includes, wherein the fuel nozzle is directed downstream with respect to airflow through the air channel. 
         [0008]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the fuel nozzle is directed upstream with respect to airflow through the air channel. 
         [0009]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is a core airflow path within the gas turbine engine. 
         [0010]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is a primary combustion gas exhaust airflow. 
         [0011]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is within a combustor section of the gas turbine engine. 
         [0012]    A further embodiment of any of the foregoing embodiments of the present disclosure includes a second fuel injector within the vane, the second fuel injector is positioned downstream with respect to the outlet from the air channel. 
         [0013]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the second fuel injector directly injects fuel into the primary combustion gas exhaust airflow. 
         [0014]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein an interior of the vane receives the secondary airflow such that the interior is at a higher pressure than an airflow within the airflow path. 
         [0015]    A gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a vane in an airflow path within the gas turbine engine, the vane includes an air channel with an outlet in communication with the airflow path; a first fuel nozzle within the vane operable to inject fuel into the air channel to at least partial premix and prevaporize the fuel with a secondary airflow from within the vane in the air channel prior to entry into the airflow path through the outlet; and a second fuel nozzle within the vane operable to directly inject fuel into the airflow path. 
         [0016]    A further embodiment of any of the foregoing embodiments of the present disclosure includes a spraybar within the vane, the spraybar in communication with the first fuel nozzle and the second fuel nozzle. 
         [0017]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein an interior of the vane receives the secondary airflow such that the interior is at a higher pressure than the primary combustion gas exhaust airflow of the airflow path. 
         [0018]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is a core airflow path within the gas turbine engine. 
         [0019]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is a primary combustion gas exhaust airflow. 
         [0020]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the airflow path is within a combustor section of the gas turbine engine. 
         [0021]    A method of injecting fuel within a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes at least partially premixing and prevaporizing fuel with a secondary airflow from within a vane in an air channel within the vane, the vane within an airflow path of the gas turbine engine. 
         [0022]    A further embodiment of any of the foregoing embodiments of the present disclosure includes directly injecting the fuel into the airflow path; and selectively activating the at least partial premixing and prevaporizing independent of the directly injecting. 
         [0023]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the selectively activating occurs during a high power operating condition. 
         [0024]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the vane is within an augmentor section. 
         [0025]    A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the vane is within a combustor section. 
         [0026]    The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]    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: 
           [0028]      FIG. 1  is a general schematic view of an exemplary gas turbine engine embodiment for use with the present disclosure; 
           [0029]      FIG. 2  is an expanded sectional view of an vane within an augmentor section of the gas turbine engine according to one disclosed non-limiting embodiment; 
           [0030]      FIG. 3  is a sectional view of the secondary fuel injector; 
           [0031]      FIG. 4  is a side view of the vane illustrating a multiple of outlets from Carbureted Fuel Injection System (CFIS) and a Jet-in-Cross Flow (JCF) Fuel Injection System; 
           [0032]      FIG. 5  is an expanded view of a CFIS fuel injector according to another disclosed non-limiting embedment; 
           [0033]      FIG. 6  is an expanded view of a CFIS fuel injector according to another disclosed non-limiting embedment; and 
           [0034]      FIG. 7  is an expanded sectional view of a vane within a combustor section of the gas turbine engine according to another disclosed non-limiting embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0035]      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 , a turbine section  28 , an augmenter section  30  and a nozzle section  32 . The sections are defined along a central longitudinal engine axis A. Although depicted as an augmented low bypass turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are applicable to other gas turbine engine architectures to include non-augmented engines, geared architecture engines, direct drive turbofans, turbojet, turboshaft, multi-stream variable cycle adaptive engines and other engine architectures. Variable cycle gas turbine engines power aircraft over a range of operating conditions and essentially alters a bypass ratio during flight to achieve countervailing objectives such as high specific thrust for high-energy maneuvers yet optimizes fuel efficiency for cruise and loiter operational modes. 
         [0036]    The compressor section  24 , the combustor section  26  and the turbine section  28  are generally referred to as the engine core. The fan section  22  and a low pressure turbine  34  of the turbine section  28  are coupled by a first shaft  36  to define a low spool. The compressor section  24  and a high pressure turbine  38  of the turbine section  28  are coupled by a second shaft  40  to define a high spool. 
         [0037]    An outer engine structure  42  and an inner engine structure  44  define a generally annular secondary airflow path  46  around a primary airflow path  48  of the engine core. It should be understood that various structure may define the outer engine structure  42  and the inner engine structure  44  to essentially define an exoskeleton. 
         [0038]    Air that enters the fan section  22  is divided between a core flow through the primary airflow path  48  and a secondary airflow through the secondary airflow path  46 . The core flow passes through the combustor section  26 , the turbine section  28 , then the augmentor section  30  where fuel may be selectively injected and burned to generate additional thrust through the nozzle section  32 . The secondary airflow may be utilized for a multiple of purposes to include, for example, cooling and pressurization. The secondary airflow as defined herein is any flow different than the primary combustion gas exhaust airflow. The secondary airflow passes through an annulus defined by the outer engine case structure  42  and the inner engine structure  44  then may be at least partially injected into the primary airflow path  48  adjacent the augmentor section  30  and the nozzle section  32 . 
         [0039]    With reference to  FIG. 2 , the augmenter section  30  generally includes a turbine exhaust case (TEC)  50  and a center body  52  with a tail cone  54 . The TEC  50  generally includes an outer case  51  of the outer engine structure  42  and a concentrically spaced inner liner  53  that operates as a heat shield to protect the outer case  51  from the core exhaust gas flow. Air discharged from, for example, the fan section  22  is communicated through the secondary airflow path  46  defined in part by the outer case  51  and the inner liner  53 . 
         [0040]    Circumferentially arrayed vanes  56  extend generally radially between the center body  52  and the TEC  50 . Each of the vanes  56  have circumferentially opposite first and second walls  58 ,  60  through which secondary airflow to cool and pressurize the vanes  56 . The multiple of vanes  56  operate, in one disclosed non-limiting embodiment, as bluffbody flameholders by providing a rear-facing flame holder surface  59  to hold the flame. Combustion instability is a phenomenon that can occur in high-output combustion systems and may limit operation, which, if allowed to occur for prolonged periods, may damage hardware as a result of cyclic stresses. Such high-frequency combustion instability in thrust augmenters is commonly referred to as “screech”. 
         [0041]    Each of the particular vanes  56  contain a spraybar  62  that extends from a fuel manifold  64  (illustrated schematically) of an augmentor fuel injection system  66 . The fuel manifold  64  may be located radially inboard such that the spraybars  62  extend radially outward, as shown, or the fuel manifold may alternatively or additionally be located radially outboard such that the spraybars extend radially inward. The spraybars  62  spray fuel through the circumferentially opposite first and second walls  58 ,  60  of the vanes  56  at generally right angles directly into the core exhaust gas stream downstream of the turbine section  28 . The rear-facing flame holder surface  59  provides a low velocity region in the core exhaust gas stream to facilitate flame stability in the augmentor section  30 . An igniter or pilot system is operated to ignite and maintain ignition of the fuel sprayed into the augmentor section  30 . 
         [0042]    With reference to  FIG. 3 , at least one of the multiple of vanes  56  includes a Carbureted Fuel Injection System (CFIS)  70  in addition to a Jet-in-Cross Flow (JCF) fuel injection system  72  that can be positioning upstream, downstream, or adjacent to the JCF injection system  72 . In this disclosed non-limiting embodiment, the CFIS.  70  and the JCF fuel injection system  72  receive fuel from the spraybars  62  to spray fuel into the core exhaust gas stream downstream of the turbine section  24  to mix with oxygen and ignite to generate a second combustion and increased thrust. Although only a single CFIS fuel injector  74  and a single JCF fuel injector  76  are schematically illustrated in each of the respective first and second walls  58 ,  60 , it should be appreciated that any number may be included in each vane  56 . Although a single CFIS fuel injector  74  is shown upstream of a single JCF fuel injector  76  are schematically illustrated in each of the respective first and second walls  58 ,  60 , it should be appreciated that several orientations of the injectors are possible in each vanes  56 . In one disclosed non-limiting embodiment, a multiple of CFIS fuel injectors  74  are radially distributed along a span of the vane  56 . Furthermore, it should be appreciated that all or only a subset of the vanes  56  may include the CFIS system  70  and the JCF fuel injection system  72 . 
         [0043]    Each CFIS fuel injector  74  generally includes an air channel  78  and a fuel nozzle  80  within the air channel  78  to provide carbureted fuel injection. “Carbureted” as defined herein includes the at least partial premixing of fuel within the air channel  78 . Each air channel  78  defines an inlet  82  which may include a bell-mouth  84  within the vane  56  and an outlet  86  through the respective first and second walls  58 ,  60  ( FIG. 4 ) that may be circular or of other shapes such as a slot and oriented perpendicular to the first and second walls  58 ,  60  or at angles. 
         [0044]    An interior  88  of the vane  56  receives secondary airflow from the secondary airflow path  46  such that the interior  88  is at a higher pressure than the primary combustion gas exhaust airflow. Secondary airflow is thereby communicated through the air channel  78  into the primary combustion gas exhaust airflow. 
         [0045]    The fuel nozzle  80  injects fuel into the air channel  78 . In one disclosed non-limiting embodiment, the fuel nozzle  80  is directed downstream with the airflow through the air channel  78 . In other disclosed non-limiting embodiments, the fuel nozzle  80  is directed upstream ( FIG. 5 ) or transverse ( FIG. 6 ) to the airflow within the air channel  78 . It should be appreciated that various fuel injection geometries into the air channel  78  may be provided to premix and pre-vaporize fuel with air in the air channel  78  prior to ejection through the outlet  86 . 
         [0046]    In one disclosed non-limiting embodiment, the CFIS system  70  pre-vaporizes and premixes about 5-20% of the total fuel sprayed by the augmenter section  30 . The CFIS system  70  thereby produces a well-prepared fuel-air mixture which then flows out and along the first and second walls  58 ,  60  of the vane  56  to feed the flow located in the wake of the rear-facing flame holder surface  59  behind the vane  56  which facilitates combustion stability. With this arrangement, the JCF fuel injection system  72  may still inject a portion (e.g. 80-95%) of the total fuel injected into the augmenter section  30  with high penetration to facilitate overall combustion efficiency. Through adjustment of the JCF/CFIS fuel split, optimal fueling of the wake can be realized over all engine and flight conditions which results in a robustly stabilized system that is insensitive to flight and operating conditions to provide stable, screech-free operation. Although the portion of fuel to the CFIS fuel injector  74  is discussed as 5-20%, it should be appreciated that any proportion of fueling between the CFIS and JCF injectors in applicable. 
         [0047]    In one disclosed non-limiting embodiment, the CFIS system  70  and the JCF fuel injection system  72  may inject fuel in accords with a predetermined fuel split. That is, the CFIS system  70  may, for example only, inject about 5-20% of the total fuel sprayed into the augmenter section  30  and the JCF fuel injection system  72  injects the remainder. In another disclosed non-limiting embodiment, the CFIS system  70  is selectively activated at particular engine operational conditions such as at high power. 
         [0048]    The CFIS.  70  allows the fuel distribution to be optimally tuned for different operational conditions. “Screech” generally occurs at high flight speeds where the pressure and combustion rate in the augmenter is greatest. Research has shown that combustion instability is linked to the static stability of a flameholder and fuel which is supplied into the flameholder wake as well as control of fuel to the augmenter section  30  alters the combustion process and may be used to avoid screech. 
         [0049]    As the CFIS.  70  is located within the vanes  56 , the CFIS.  70  has minimal—if any—influence on external geometry or cooling yet increases overall system capabilities and reduces life cycle costs. The vanes  56  may also be readily retrofit to the engine  20 . 
         [0050]    With reference to  FIG. 7 , in another disclosed non-limiting embodiment, the vanes  56 ′ may alternatively or additionally be located in other engine sections such as the combustor section  26 . That is, the vanes  56 ′ may be, for example, located within a combustor  90  generally between an outer combustor wall  92  and an inner combustor wall  94 . The outer combustor wall assembly  92  and the inner combustor wall  94  are spaced apart such that a combustion chamber  96  is defined therebetween to receive carbureted fuel from the outlet  86 ′ and direct fuel injection from the fuel injectors  76 ′ of the vanes  56 ′ as discussed above. It should be appreciated that various engine sections will alternatively or additionally benefit herefrom. 
         [0051]    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. 
         [0052]    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. 
         [0053]    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. 
         [0054]    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.