Patent Document

FIELD OF THE DISCLOSURE 
       [0001]    The present application relates generally to gas turbine engines and more particularly relates to a trailing edge of a first stage nozzle or a transition nozzle configured to promote mixing of respective combustion streams downstream thereof before entry into a first stage bucket of a turbine. 
       BACKGROUND 
       [0002]    Annular combustors often are used with gas turbine engines. Generally described, an annular combustor may have a number of individual can combustors that are circumferentially spaced between a compressor and a turbine. Each can combustor separately generates combustion gases that are directed downstream towards the first stage of the turbine. 
         [0003]    The mixing of these separate combustion streams is largely a function of the free stream Mach number at which the mixing is taking place as well as the differences in momentum and energy between the combustion streams. Practically speaking, the axial distance between the exit of the can combustors and the leading edge of a first stage nozzle is relatively small such that little mixing actually may take place before entry into the turbine. 
         [0004]    There is thus a desire to minimize mixing loses. Such reduced mixing loses may reduce overall pressure losses without increasing the axial distance between the combustor and the turbine. Such an improved combustion design thus should improve overall system performance and efficiency. 
       SUMMARY 
       [0005]    The present application and the resultant patent thus provide a disruptive surface on a trailing edge of a stage one nozzle or a transition nozzle to promote mixing of respective combustion streams downstream thereof before entry into a first stage bucket of a turbine. For example, in one embodiment, a gas turbine engine may include a combustor including a combustion flow. The gas turbine engine also may include one or more airfoils forming a first stage nozzle or a transition nozzle disposed downstream of the combustor. Moreover, the gas turbine engine may include a flow disruption surface positioned about a trailing edge of the one or more airfoils to promote mixing of the combustion flow. 
         [0006]    The present application and the resultant patent further provides a method of limiting pressure losses in a gas turbine engine. The method may include positioning a flow disruption surface on a trailing edge of one or more airfoils of a first stage nozzle or a transition nozzle, generating a number of combustion streams in a number of can combustors, substantially mixing the combustion streams with the flow disruption surface, and passing a mixed stream to a stage one bucket. 
         [0007]    The present application and the resultant patent further provides a gas turbine engine. The gas turbine engine may include a number of can combustors forming a number of combustion flows. The gas turbine engine also may include one or more airfoils forming a first stage nozzle or a transition nozzle disposed downstream of the can combustors. Moreover, the gas turbine may include a flow disruption surface positioned about a trailing edge of the one or more airfoils to promote mixing of the combustion flows. 
         [0008]    These and other features and improvements of the present application will become apparent to one of ordinary skill in the art upon review of the following detailed description when taken in conjunction with the several drawings and the appended claims. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]    Reference will now be made to the accompanying drawings, which are not necessarily drawn to scale. 
           [0010]      FIG. 1  is a schematic view of a known gas turbine engine that may be used herein. 
           [0011]      FIG. 2  is a side cross-sectional view of a can combustor that may be used with the gas turbine engine of  FIG. 1 . 
           [0012]      FIG. 3  is a side cross-sectional view of a transition nozzle combustion system that may be used with the gas turbine engine of  FIG. 1 . 
           [0013]      FIG. 4  is a schematic view of a nozzle as may be described herein. 
           [0014]      FIG. 5  is a schematic view of a flow disruption surface as may be described herein. 
           [0015]      FIG. 6  is a schematic view of a flow disruption surface as may be described herein. 
           [0016]      FIG. 7  is a schematic view of a flow disruption surface as may be described herein. 
       
    
    
     DETAILED DESCRIPTION 
       [0017]    Referring now to the drawings, in which like numerals refer to like elements throughout the several views,  FIG. 1  shows a schematic view of gas turbine engine  10  as may be used herein. The gas turbine engine  10  may include a compressor  15 . The compressor  15  compresses an incoming flow of air  20 . The compressor delivers the compressed flow of air  20  to a combustor  25 . The combustor  25  mixes the compressed flow of air  20  with a compressed flow of fuel  30  and ignites the mixture to create a flow of combustion gases  35 . Although only a single combustor  25  is shown, the gas turbine engine  10  may include any number of combustors  25 . In this example, the combustor  25  may be in the form of a number of can combustors as will be described in more detail below. The flow of combustion gases  35  is in turn delivered to a downstream turbine  40 . The flow of combustion gases  35  drives the turbine  40  so as to produce mechanical work. The mechanical work produced in the turbine  40  drives the compressor  15  via a shaft  45  and an external load  50  such as an electrical generator and the like. 
         [0018]    The gas turbine engine  10  may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine  10  may be anyone of a number of different gas turbine engines such as those offered by General Electric Company of Schenectady, New York and the like. The gas turbine engine  10  may have different configurations and may use other types of components. Other types of gas turbine engines also may be used herein. Multiple gas turbine engines, other types of turbines, and other types of power generation equipment also may be used herein together. 
         [0019]      FIG. 2  shows an example of the combustion system  25  that may be used in the gas turbine engine  10 . A typical combustion system  25  may include a head end  60  with a number of fuel nozzles  65 . A liner  68  and a transition piece  70  may extend downstream of the fuel nozzles  65  to an aft end  75  about a number of first stage nozzle vanes  80  of the turbine  40 . An impingement sleeve  85  may surround the liner  68  and the transition piece  70  and provide a cooling flow thereto. Other types of combustors  25  and other types of components and other configurations are also known. 
         [0020]    A cooling flow  90  from the compression system  15  or elsewhere may pass through the impingement sleeve  85 . The cooling flow  90  may be used to cool the liner  68  and the transition piece  70  and then may be used at least in part in charging the flow of combustion gases  35 . A portion of the flow  90  may head towards the aft end  75  and may be used for cooling the first stage nozzle vanes  80  and related components. Other types of cooling flows may be used. The loss of a portion of the cooling flow  90  thus results in a parasitic loss because that portion of the flow  90  is not used for charging the combustion flow  35 . Other components and other configurations also may be used herein. 
         [0021]      FIG. 3  shows an example of a portion of a transition nozzle combustion system  100  as may be described herein. The transition nozzle combustion system  100  may include a transition nozzle  110 . The transition nozzle  110  has an integrated configuration of a liner, a transition piece, and a first stage nozzle vane in a manner similar to that described above. The transition nozzle  110  extends from a head end  120  about the fuel nozzles  65  to a flow region  130  and a transition nozzle aft end  140  about a number of bucket blades in a first turbine stage  150 . The transition nozzle combustion system  100  thus may be considered an integrated combustion system. Other types of combustors in other configurations may be used herein. 
         [0022]    In certain embodiments, as depicted in  FIG. 4 , the trailing edge (i.e., downstream edge) of the first stage nozzle vanes  80  and/or the transition nozzles  110  may include a flow disruption surface to promote mixing of the combustion flows. That is, the trailing edge of the first stage nozzle vanes  80  and/or the transition nozzles  110  may include a chevron mixing component, a lobed mixing component, and/or a fluidics mixing component. In this manner, the trailing edge shape of the first stage nozzle vanes  80  and/or the transition nozzles  110  may be configured to promote mixing of the combustion flow such that the overall pressure loss of the system is reduced. As a result, the trailing edge shape of the first stage nozzle vanes  80  and/or the transition nozzles  110  may reduce the high cycle fatigue load and the heat load on the first stage buckets. The first stage nozzle vanes  80  and/or the transition nozzles  110  may or may not be integral with the transition piece and/or with the combustor. 
         [0023]      FIG. 4  shows an embodiment of an airfoil  400  of a first stage nozzle  80  and/or a transition nozzle  110 . In one example, the airfoil  400  of the first stage nozzle may include a leading edge  402  and a trailing edge  404 . The trailing edge  404  may include a flow disruption surface  406 . The flow disruption surface may be configured to promote mixing of the combustion flows  408 . That is, the flow disruption surface may be configured to promote mixing of the combustion flows  408  downstream thereof before entry into a first stage bucket. The flow disruption surface  406  may include spikes, chevrons, lobes, and/or jets. 
         [0024]    The increased uniformity of the temperature and velocity field created by the enhanced mixing of the trailing edge  404  disruption surface  406  of the airfoils  400  of the first stage nozzle and/or the transition nozzle is beneficial to the rotor blade row mechanical and thermal durability downstream thereof. This is particularly beneficial for a low nozzle count or a transition-nozzle configuration. The enhanced mixing is created by the use of spikes, chevrons, lobes, and/or jets disposed about the trailing edge  404  of the airfoils  400  of the first stage nozzle. This enhanced mixing increases the pressure loss relative to unforced mixing. The addition of the flow disruption surface  406  about the trailing edge  404  of the airfoils  400  of the first stage nozzle and/or the transition nozzle minimizes the amount of mixing that takes place within the bucket domain. The additional pressure loss incurred by the enhanced mixing from the trailing edge  404  of the airfoils  400  of the first stage nozzle and/or the transition nozzle is much lower than the mixing loss incurred should the nozzle wake mix in the downstream bucket. Also, the enhanced mixing reduces the wake strength of the nozzle and the high cycle fatigue loads on the bucket, which allows more economical nozzle configurations to be chosen (such as, but not limited to, lower count and/or closer axial nozzle-bucket spacing). Further, the enhanced mixing makes the incoming velocity and thermal flow distributions more uniform, which reduces the gas loads and thermal loads on the bucket, thereby improving the durability of the bucket. 
         [0025]      FIGS. 5-7  show a number of different embodiments of the flow disruption surface  406  of  FIG. 4  as may be described herein. For example, as depicts in  FIG. 5 , the flow disruption surface  406  of  FIG. 4  may be a chevron mixing joint  500 . In some instances, the chevron mixing joint  500  may include a first set of chevron like spikes  502  and a mating second set of chevron like spikes  504 . As is shown, the depth and angle of the first and second set of chevron like spikes  502 ,  504  may vary. Likewise, the number, size, shape, and configuration of the chevron like spikes  502 ,  504  each may vary. Other components and other configurations may be used herein. 
         [0026]      FIG. 6  shows a further embodiment of the flow disruption surface  406  of  FIG. 4  as may be described herein. In this embodiment, a lobed mixing joint  600  is shown. The lobed mixing joint  600  may include a first set of lobes  602  and a second set of lobes  604 . The first and second set of lobes  602 ,  604  may have a largely sinusoidal wave like shape and may mate therewith. The depth and shape of the first and second set of lobes  602 ,  604  also may vary. The number, size, shape, and configuration of the lobes  602 ,  604  may vary. Other components and configurations may be used herein. 
         [0027]      FIG. 7  shows a further embodiment of the flow disruption surface  406  of  FIG. 4 . In this example, the flow disruption surface  406  of  FIG. 4  may be in the form of a fluidics mixing joint  700  as is shown. The fluidics mixing joint  700  may include a number of jets  702  therein that act as the flow disruption surface  406 . The jets  702  may spray a fluid  704  into the combustion flows. The number, size, shape, and configuration of the jets  702  may vary. Likewise, the nature of the fluid  704  may vary. Other components and configurations may be used herein. 
         [0028]    The embodiments of the flow disruption surface described herein are for purposes of example only. Any other the flow disruption surface geometry or other type of flow disruption surface that encourages mixing of the combustion flows may be used herein. Different types of flow disruption surfaces may be used herein together. Other components and other configurations also may be used herein. 
         [0029]    It should be apparent that the foregoing relates only to certain embodiments of the present application and that numerous changes and modifications may be made herein by one of ordinary skill in the art without departing from the general spirit and scope of the disclosure as defined by the following claims and the equivalents thereof

Technology Category: 2