Abstract:
The present invention provides a mixer for a gas turbine combustor comprising a plurality of generally annular walls interconnected by at least a plurality of first vanes. The vanes are oriented at angles, so as to create a shear layer between the two flows. A fuel is then injected so as to penetrate the shear layer for enhanced mixing. The mixture passes through an extended mixing passage to provide sufficient time and distance for improved mixedness prior to ignition. Multiple embodiments of the present invention are disclosed comprising a plurality of first vanes and a plurality of first and second vanes.

Description:
BACKGROUND OF THE PRIOR ART 
   The present invention relates generally to gas turbine combustors and more specifically to a fuel and air mixing device in a gas turbine combustor. 
   In a gas turbine engine, the combustion section contains a reaction that occurs when fuel and compressed air are mixed together and react after being ignited by an ignition source. Compressed air is directed to one or more combustion chambers from the engine compressor. Fuel injection devices inject a fuel, either liquid or gas, into the compressed air stream and the mixture undergoes a chemical reaction once being exposed to a heat source, such as an igniter. 
   Some examples of prior art mixer devices are shown in  FIGS. 1 and 2 .  FIG. 1  is a cross section of a combustion system disclosed in U.S. Pat. No. 5,515,680, and hereby incorporated by reference. The combustion system utilizes a ring member  31  to inject a fuel transverse to the flow direction of the premixing combustion air, at the outside of a 180 degree bend in the air flow path, in an effort to inject the fuel from a high velocity region towards a lower velocity region for improved mixing. While this technique may improve mixing locally, further improvements can be made such that additional time and distance is provided in the region upstream of the combustor to further enhance premixing.  FIG. 2 , on the other hand, is a cross section of a fuel injector and mixing device disclosed in U.S. Pat. No. 5,165,241 that injects a fuel from the centerbody of the injector, radially outward into the passing air stream, which has previously undergone counter rotating swirl from inner swirler  26  and outer swirler  28 . While this type of mixer attempts to provide improved premixing, it too can be improved by providing a longer time and distance for the fuel and air premixing to be more complete prior to ignition. 
   In order to control emissions levels of oxides of nitrogen (NOx) and carbon monoxide (CO), it is critical that the fuel molecules burn as completely as possible such as to not leave any unburned hydrocarbons to pass into the atmosphere. In order for the fuel to completely burn a number of issues must be addressed, one of which is fuel and air mixedness prior to ignition. Fuel and air mixedness is controlled by factors such as swirl, fuel injection location, and mixing time prior to ignition. Therefore, for the lowest possible emissions, it is most desirable to provide a mixer for a gas turbine combustor that optimizes swirl, fuel injection location, and mixing time such that the combustion process will be as complete as possible. 
   SUMMARY AND OBJECTS OF THE INVENTION 
   The present invention provides a mixer for a gas turbine combustor wherein the mixer comprises a plurality of annular walls containing at least a plurality of first vanes oriented at a first angle in between said annular walls, thereby creating a shear layer. A fuel injector is positioned adjacent the vanes to inject a fuel such that the fuel jet penetrates the shear layer for optimum mixing. Furthermore, the annular walls of the mixer are configured such that sufficient time and distance is provided in order to obtain optimum mixing prior to ignition of the fuel/air mixture. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a cross section of a gas turbine combustor and mixer of the prior art. 
       FIG. 2  is a detailed cross section view of another mixing device of the prior art. 
       FIG. 3  is a cross section of a gas turbine combustor incorporating a mixer in accordance with the preferred embodiment of the present invention. 
       FIG. 4  is a detailed cross section of a portion of the mixer in accordance with the preferred embodiment of the present invention. 
       FIG. 5  is a partial perspective view of a portion of the mixer in accordance with the preferred embodiment of the present invention. 
       FIG. 6  is an additional detailed cross section of a portion of the mixer in accordance with the preferred embodiment of the present invention. 
       FIG. 7  is a plot showing analysis results of percent of fuel that is unmixed at various locations throughout the mixer of the present invention. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
   The preferred embodiment of the present invention will now be described in detail with particular reference made to  FIGS. 3-7 . Referring now to  FIG. 3 , gas turbine combustor  70  is shown in cross section. The present invention pertains to a mixer for combustor  70 . In the preferred embodiment of the present invention, combustor  70  comprises a casing  71 , end cover  72 , combustion liner  73 , and a pilot injector  74 . Another feature of combustor  70  is mixer  75 , which is shown in greater detail in a detailed cross section in  FIG. 4 . 
   Mixer  75 , which serves to provide a region for thorough fuel and air mixing prior to ignition, comprises multiple components depending on the desired level of fuel and air mixedness. For a complete understanding of the invention, all components of mixer  75  are shown in  FIGS. 4-6 . Mixer  75  comprises a first generally annular wall  76  that is located coaxial with a combustor center axis A-A (see  FIG. 3 ). Located radially outward of and coaxial with first generally annular wall  76  is a second generally annular wall  77  having a first portion  77 A and a second portion  77 B having a bend  78  such that a first end  79  of second generally annular wall  77  is located radially inward of first generally annular wall  76  and axially within combustion liner  73 . A third generally annular wall  80  is located radially inward of and coaxial with first generally annular wall  76 . Extending between first generally annular wall  76  and first portion  77 A of second generally annular wall  77  is a plurality of first vanes  81  that are oriented at a first angle relative to centerline A-A. A plurality of second vanes  82  extend between first generally annular wall  76  and third generally annular wall  80 . Second vanes  82  are oriented at a second angle relative to the first angle so as to create a shear layer adjacent first generally annular wall  76 . Depending on the desired swirl level and resulting mixing, the quantity and angles of first vanes  81  and second vanes  82  can vary. For the preferred embodiment, the shear layer resulting from first vanes  81  and second vanes  82  is formed by a difference between the first angle and second angle of between 20 and 60 degrees. 
   An additional feature of mixer  75  is fuel injector  85 , which is located adjacent second generally annular wall  77  for injecting a fuel into the shear layer formed adjacent first generally annular wall  76 . In the preferred embodiment of the present invention, fuel injector  85  comprises an annular manifold  86  having a plurality of injection locations  87  around annular manifold  86 . Furthermore, injection locations  87  are oriented generally perpendicular to center axis A-A. 
   As a result of the radial and axial positions of the generally annular walls  76 ,  77 , and  80  as well as position of combustion liner  73 , a mixing passage  88  is created. Mixing passage  88  is formed between second generally annular wall  77  and combustion liner  73  and serves as a region of extended length for mixing fuel and air, due to bend  78  in second portion  77 B of second generally annular wall  77 . 
   An additional feature of mixer  75  is its ability to compensate for thermal expansion of combustion liner  73 . Combustion liner  73  contains a spring seal  89  that is fixed to the outer surface of combustion liner  73  at a first seal end and is free at a second seal end. The third generally annular wall  80  of mixer  75  engages spring seal  89  proximate its second seal end to provide a means for maintaining the dimensions of mixing passage  88  that is compliant to various thermal changes between combustion liner  73  and mixer  75 . 
   In operation, having provided the aforementioned combustor and mixer geometry, a flow of air is provided to mixer  75 . The airflow is then split with a first portion of air being directed through first vanes  81  and a second portion being directed through second vanes  82 . The airflow portions are swirlered at their respective angles by their respective vanes and form a shear layer, or more specifically, a layer of air in between two rotating flows of different degrees. This shear layer has a thickness, which is attributed to the thickness of first generally annular wall  76  directly upstream of the shear layer. Fuel is then injected into the shear layer to form a premixture in mixing passage  88 . The premixture is directed through bend  78  and into the combustor for ignition. 
   As a result of the swirl vane configuration and orientation, the fuel injection from a manifold configuration into the shear layer, coupled with the mixing passage distance and time, computational analysis has predicted an extremely high rate of mixedness prior to ignition. A plot of this analysis can be seen in  FIG. 7  and shows a cross section of mixer  75  with fluid flowing through the mixer. The dark regions adjacent the swirl vanes represent the air while adjacent the swirlers a jet penetrating the swirling air flow is positioned injecting a fuel generally perpendicular to the center axis. As the premixture travels through mixing passage  88  and towards bend  78 , approximately 14.2% of the fuel molecules have not mixed with air, and if ignition occurred at this location, significant emissions would result. The rate of unmixedness at this location is common to combustors having similar generally axial premixing passages prior to ignition. However, due to bend  78  in mixing passage  88 , and the additional passage length as a result, further mixing occurs. Analysis of unmixed fuel particles at the exit of bend  78 , proximate the entrance to combustion liner  73 , shows only 1.94% of fuel molecules are unmixed. The result of this unmixedness level is even lower emissions. These predictions of unmixedness have been verified by extensive rig testing. Depending on desired performance and emissions, fuel injection hole sizes and position would vary such that the resulting fuel jet penetrates the shear layer as desired. For the present invention, it is preferred to have only one row of fuel injectors  87  circumferentially about annular manifold  86 . 
   In a first alternate embodiment of the present invention, mixer  85  contains only plurality of first vanes  81  between first generally annular wall  76  and second generally annular wall  77 . In this embodiment, the shear layer is formed between the angle of first vanes  81  and the flow passing through a passageway formed by first generally annular wall  76  and combustion liner  73 . While this configuration is simpler to manufacturer and can be manufactured at a lower cost due to the simplified geometry, the mixing benefits associated with the shear layer are not as great given the limited shear generated by the interaction from a single set of vanes and an axial flow. This first alternate embodiment is advantageous if radial space for mixing is limited or sufficient mixing can be achieved with a single set of vanes. 
   A second alternate embodiment maintains the benefits of the preferred embodiment with respect to the shear generated by opposing flow angles from the plurality of first and second vanes, but eliminates seal  89 . Removing seal  89  from the mixer geometry simplifies the manufacturing and reduces the associated cost by eliminating a component that is manufactured from a higher cost alloy having spring capability. However, while removing seal  89  simplifies the manufacturing process and can reduce cost, it does allow for additional thermal movement between combustion liner  73  and mixer  75  than if seal  89  were present, thereby affecting dimensions of mixing passage  88 . Depending on the operating conditions and temperatures of combustor  70 , eliminating seal  89  may not have adverse affects on fuel and air mixing and combustor performance. 
   While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims.