Abstract:
A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Description:
BACKGROUND OF THE INVENTION  
         [0001]    The present invention relates to a miniature gas turbine engine, and more particularly to a combustor system in which each combustion region approximates a well-stirred reactor.  
           [0002]    Miniature turbojet engines (100 lbf thrust and smaller) are often utilized in single usage applications such as reconnaissance drones, cruise missiles, decoy and other weapon applications, including air-launched and ground-launched weapon systems. The use of such an engine greatly extends the range of the weapon in comparison to the more conventional solid fuel rocket engine. Miniature gas turbine engines are difficult to fabricate economically for general expendable usage in large numbers.  
           [0003]    To achieve economically feasible extended range expendable propulsion sources for such weapon system, it is necessary that the gas turbine engines be manufactured relatively inexpensively yet provide a high degree of reliability and efficiency. One component that greatly affects performance yet is rather complicated to manufacture is the combustor system.  
           [0004]    Miniature gas turbine engines typically utilize annular combustor shapes that wrap around other engine features such as an exhaust tailpipe or a turbine wheel to minimize frontal area in order to maximize the thrust per unit drag. If the engine frontal area is minimized, the combustor internal volume must be utilized optimally.  
           [0005]    Miniature gas turbine engine combustor systems may not have room for conventional fuel injection systems and require high-density, high-viscosity fuels to maximize thrust. The combustor system must accommodate these fuels and provide reliable ignition and stable operation. These requirements are a challenge given the size and cost limitations for an expendable system.  
           [0006]    Accordingly, it is desirable to provide an inexpensive and reliable combustor system having a minimal frontal area for a miniature gas turbine engine which achieves stability throughout the flight envelope and combustion efficiency at cruise conditions.  
         SUMMARY OF THE INVENTION  
         [0007]    The combustor system according to the present invention includes an annular combustor liner with a minimized number of fuel injection tubes, no hot-side cooling, and reduced hole count. Fuel injection is via orifices that direct fuel into fuel-air mixture tubes, which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region.  
           [0008]    Fuel system pressures are kept relatively low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air mixture tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.  
           [0009]    In the primary burning region, the fuel-air stoichiometry varies from fuel-lean to fuel-rich and the reaction residence time is maximized by keeping the percentage of total air flow into this region as low as possible. In the second burning region, the fuel-air mixture is tailored to maintain the maximum possible combustion time for best efficiency throughout the engine envelope. The secondary region results in near-stoichiometric fuel/air ratios and consequently, maximum flame temperatures.  
           [0010]    Non-reacting excess air is dumped into the engine flow stream just ahead of the turbine nozzle. Some air is dumped at the outer wall and some is passed around the combustor to provide external convective wall cooling. The combustor air flows and fuel injection are tailored to meet operating requirements.  
           [0011]    The present invention therefore provides an inexpensive and reliable combustor system having a minimal frontal area for a miniature gas turbine engine which achieves stability throughout the flight envelope and combustion efficiency at cruise conditions. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0012]    The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows:  
         [0013]    [0013]FIG. 1 is a general perspective view an exemplary vehicle embodiment for use with the present invention;  
         [0014]    [0014]FIG. 2 is a schematic view of a gas turbine engine having a start system according to the present invention;  
         [0015]    [0015]FIG. 3 is a longitudinal sectional view of a combustor system of FIG. 2;  
         [0016]    [0016]FIG. 4 is a radial sectional view of a primary burning region of the combustor system of FIG. 3;  
         [0017]    [0017]FIG. 5 is a perspective view of a combustor liner the combustor system of FIG. 2; and  
         [0018]    [0018]FIG. 6 is a longitudinal sectional view of another combustor system of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0019]    [0019]FIG. 1 illustrates a general schematic view of a vehicle  100  including an expendable miniature gas turbine engine  10  according to the present invention. The vehicle  100  includes a body  102  and one or more aerodynamic surfaces  104 . The engine  10  is coupled to, or within, the body  102 . An intake  106  provides air to the engine  10 , and an exhaust pipe  108  exhausts the thrust therefrom. The engine  10  of the invention may also be used in other single usage and reusable applications such as reconnaissance drones, cruise missiles, decoys and other weapon and non-weapon applications.  
         [0020]    Referring to FIG. 2, the miniature gas turbine engine  10  generally includes a housing  14 , a rotor shaft  16  rotationally mounted to a forward bearing  18 , a combustion system  20  and an exhaust pipe (nozzle)  22 . The rotor shaft  16  rotates about a longitudinal axis X although other forms of rotors, such as a monorotor configuration, would also benefit from the present invention. In the illustrated rotor configuration, a rotor  24  includes compressor blades  26  facing forward toward an inlet  28  and turbine blades  30  facing rearward toward the exhaust  22  to define a turbine wheel. The forwardly extending shaft  16  is received in the bearings  18  and is preferably coupled to a fuel pump (illustrated schematically at  32 ) to provide fuel to an annular combustor liner  34  through a fuel manifold  36 .  
         [0021]    Referring to FIG. 3, a cross-section of the combustion system  20  is illustrated. The combustion system  20  generally includes the annular combustor liner  34 , the fuel manifold  36  and an igniter  38 . The combustor liner  34  is a reverse flow annular combustor, and thus has a leading end  40  generally disposed toward the rear of turbine engine  10 , and a trailing end  42  generally disposed toward the front of the turbine engine  10 . The combustor liner  34  includes an outer wall  44  in the form of a metal tube having an outer surface  46  and an opposing inner surface  48 . The combustor liner  34  further includes an inner wall  50 , and a dome  52  generally connected to, and joining, the inner and outer walls  44 ,  50  at respective annular lines of intersection  54  and  56 .  
         [0022]    The exhaust pipe  22  extends rearwardly of the engine  10  from throat  60 , and interfaces with rear housing wall  62 , whereby the combustor liner  34  is enclosed on its outer and rear surfaces by housing  14  and on its inner surface by the exhaust pipe  58 . The combustor liner  34  interfaces with the exhaust pipe  22  through a combustor exit  64  such that exhaust gases from the combustor liner  34  are directed through the exhaust pipe  22  generating a high velocity thrust (illustrated schematically by arrow T).  
         [0023]    A compressor discharge plenum  66  is located between the outer wall  44  of the combustor liner  34  and the housing  14 . The discharge plenum  66  distributes air from the compressor (FIG. 2) into the combustor liner  34  through fuel-air tubes  68  which feed a fuel-rich mixture of fuel and air into the leading end  40  of the combustor liner  34  to form a primary burning region P. It should be understood that the term “tubes” is to be construed to broadly include openings, holes, apertures, bent metal deflectors and the like in addition to separate cylindrical member. Moreover, any hole shape, including elliptical, rectangular, triangular and any hole condition including plain sharp-edged, plunged and the like will benefit from the present invention.  
         [0024]    Fuel is introduced into the combustor liner  34  through a fuel passageway  70  which communicates fuel from the fuel manifold  36  into each of the fuel-air tubes  68  through a fuel orifice  71 . Fuel orifices  71  are preferably drilled holes which direct fuel into the tubes  68  at a prescribed location. The fuel orifices  71  control fuel system pressure through proper predetermined sizing and quantity. The fuel orifices  71  preferably produce a predetermined allowable level of fuel pressure drop at the maximum required fuel flow rate such as 150 psid. The fuel orifices  71  essentially just pour fuel into tubes  68 . That is, fuel may just trickle from the fuel orifice  71  at low speed turbine engine  10  operation or a stream from the orifice  71  at high speed operation. At either extreme, a fine fuel spray is not necessary since a great deal of fuel-air mixing occurs within the fuel-air feed tubes  68  such that the fuel manifold  36  need not require precision machining.  
         [0025]    Optimum fuel-air mixing is required to obtain optimum combustor performance. The fuel-air mixing tubes  68  are preferably designed with enough length and air momentum/velocity to break up and evaporate as much fuel as possible. It should be understood that as the present invention is directed toward expendable gas turbine engines longevity concerns relating to extending the fuel-air mixing tubes  68  relatively deep into the combustor  34  (FIG. 4) without complicated cooling systems is of minimal concern.  
         [0026]    Since the fuel orifices  71  are relatively simple holes, and since the tubes  68  are directly attached to the combustor  34 , fueling is inexpensive and requires minimal hardware on the engine case. Control of the fuel flow rate into the fuel manifold is performed by any variable system. The size of the fuel manifold jet holes is preferably set to maximize fuel jet velocity and maintain fuel flow uniformity from hole-to-hole.  
         [0027]    The fuel-air mixture within the injection tubes  68  is most preferably of a fuel-rich quality and the air velocities through these tubes  68  are of relatively high velocity, e.g., Mach 0.3 and greater. The fuel injection of the present invention makes the combustor relatively independent of the type of fuel burned. A wide range of fuels ranging from gasesous (methane, propane, natural gas) to pure, light hydrocarbons (hexane, octane, butane) to aviation fuels (Jet-A, Jet-A1, JP-4, JP-5, JP-10, JP-8) to heavy diesel fuels (DF1, DF2, marine diesel) can be burned with fuel manifold system and combustor air apportionment readily available to one of ordinary skill in the art.  
         [0028]    Since the primary zone stoichiometry is variable by design and since the reaction times in the primary zone are short by design, primary zone flame temperatures may be “tuned” such that they are low for certain chosen engine operating conditions. These predetermined designed low flame temperatures and short reaction times result in a combustion system that is capable of achieving extremely low levels of NO x  with a wide variety of liquid fuels.  
         [0029]    Initial ignition of the combustion process is performed by a spark-gap or pyrotechnic flare-type igniter  38  preferably located through the dome  52 . The igniter  38  is placed in a position down-swirl of one of the fuel-air injection tubes  68  to ensure contact with fuel as it enters the combustor  34 . Under extreme cold conditions and at low engine speeds fuel break-up is preferably assisted by a jet of high-velocity air, oxygen or air/oxygen mixture directed into the fuel-air tube  68  just upstream of the igniter  38 . This oxygen jet is used to improve ignition only and is not needed during normal combustor operation. Once ignition is initiated, the igniter is no longer needed since the combustor  34  is a continuous ignition device.  
         [0030]    The air flow through fuel-air mixing tubes  68  breaks the fuel into small droplets and mixes the fuel with air before the fuel-air mixture enters the combustor liner  34 . Fuel is further mixed with the combustion air by strong aerodynamic forces within the combustor. Fuel break-up occurs through air-blast atomization, tube-wall impingement and vaporization. The discharge direction of the fuel-air mixture is generally circumferential and axial aft as the fuel-air tubes  68  preferably extend into the combustor liner  34  as a circumferential row which directs the mixture generally toward the dome  52  and igniter  38  (FIG. 5).  
         [0031]    The fuel air mixture is mixed with additional air injected through a row of secondary air-fed tubes  70  downstream of the fuel-air mixing tubes  68 . The secondary air-feed tubes  70  are located approximately midway between the combustor dome  52  and the exit  64  to form a secondary burning region S. It should be understood that the term “tubes” is to be construed to broadly include openings, holes, apertures, bent metal deflectors and the like in addition to separate cylindrical member. Moreover, any hole shape, including elliptical, rectangular, triangular and any hole condition including plain sharp-edged, plunged and the like will benefit from the present invention.  
         [0032]    A row of dilution air-feed tubes  74  are located just upstream of the combustor exit  64  to form a final dilution mixing region D. The sets of tubes  68 ,  70  and  72  produce a generally circumferential air velocity into the combustor liner  34 . It should be understood that the high degree of swirl produced via this air direction provides for high mixing and maximizes the path length experienced by the fuel entering the combustor.  
         [0033]    The combustor liner  34  is maintained at acceptable temperature levels by designing the combustor  34  for high air velocities. The high air velocity through the compressor discharge plenum  66  and over the external portion of the combustor  34  provides for convective cooling. It should be understood that other combustor cooling techniques, e.g., splash cooling, film cooling, effusion cooling or the like which require air injection into the combustor may also be used, but is preferably designed to avoid interference with the primary and secondary burning. Such additional cooling techniques will necessarily require slightly larger combustor volumes.  
         [0034]    The air flow into the combustor  34  is apportioned to provide the two burning regions P, S and the dilution-mixing region D. The two burning regions P, S allows the combustor  34  to operate at minimum overall burning time. In the primary burning region P, the fuel/air stoichiometry is preferably designed to be rich at full power engine operating conditions. Combustion occurs in this region at temperatures that maintain high enough flame speeds for adequate stability, but all the fuel cannot react. The fuel that is unable to react in the primary region P is mixed with air in the secondary region S and then burned. The secondary region S results in near-stoichiometric fuel/air ratios and consequently, maximum flame temperatures. It should be understood that one of skill in the art utilizing the teaching of the present invention is readily able to design such a near-stoichiometric fuel/air ratio.  
         [0035]    The fuel/air ratios in the two burning regions P, S vary with engine operating condition, so the air apportionment within the combustor  34  is preferably designed to accommodate the full engine flight envelope. Flame temperatures within the primary burning region P are critical and must be maintained at all times in order to maintain stable, efficient combustion. If the primary region is too lean or too rich, the flame temperature drops and burning rates fall to levels too low to maintain combustion. Each combustion region approximates a “well-stirred reactor” and the combination of the two regions P, S results in a “best possible” use of combustion volume.  
         [0036]    Downstream of the secondary burning region S, air is injected to mix out the hot flame gases at the dilution-mixing region D. The dilution-mixing region D is designed to preferably provides cool enough temperatures to avoid damage to the downstream turbine  30 . It should be understood that the mixing air may alternatively or additionally be introduced through tubes, drilled holes, or plunged holes and may be fed through the inner or outer combustor wall.  
         [0037]    Referring to FIG. 6, the igniter  38  is positioned between the two burning regions P, S and the fuel-air injection tubes  68  are located within the dome  52  of the combustor liner  34 . That is, the fuel-air injection tubes  68  are disposed toward the rear of turbine engine  10  through a leading end  40  of the combustor liner  34  generally between the inner and outer walls. The igniter  38  is placed in circumferential position about the outer wall of the combustor  34 . The FIG. 6 arrangement provides advantages of the above-described design with different packaging constraints.  
         [0038]    The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention.