Abstract:
The present application and the resultant patent provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface; passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region; substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation-in-part of copending U.S. patent application Ser. No. 13/036,084, filed on Feb. 28, 2011, which is hereby incorporated by reference in its entirety. 
     
    
     TECHNICAL FIELD 
       [0002]    The present application relates generally to gas turbine engines and more particularly relates to a joint between adjacent can combustors to promote mixing of the respective combustion streams downstream thereof before entry into a first stage of a turbine, and to related methods of improving durability of a first stage bucket. 
       BACKGROUND OF THE INVENTION 
       [0003]    Can-annular combustors often are used with gas turbine engines. Generally described, a can-annular combustor may have a number of individual can combustors that are circumferentially spaced in an annular arrangement between a compressor and a turbine. Each can combustor separately generates combustion gases that are directed downstream towards the first stage of the turbine. 
         [0004]    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. Moreover, a stagnant flow region or wake in a low flow velocity region may exist downstream of a joint between adjacent can combustors due to the bluntness of the joint. As such, the non-uniform combustor flows may have a Mach number of only about 0.1 when leaving the can combustors. 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. 
         [0005]    The combustor flows then may be strongly accelerated in the first stage nozzle to a Mach number of about 1.0. This acceleration may exaggerate the non-uniformities in the flow fields and hence create high mixing losses downstream thereof. As the now strongly nonuniform flow field enters the first stage bucket, the majority of mixing losses may take place therein as the wakes from the can combustor joints may be mixed by an unsteady flow process. Due to the nonuniform flow and unsteady mixing, the first stage bucket may be subjected to high cycle fatigue loads and thermal loads that significantly reduce durability of the first stage bucket. 
         [0006]    There is thus a desire for an improved combustor design that may minimize mixing loses. Such reduced mixing loses may reduce overall pressure losses without increasing the axial distance between the combustor and the turbine, which may improve overall system performance and efficiency. Such an improved combustor design also may reduce high cycle fatigue loads and thermal loads on the first stage bucket, which may improve durability of the first stage bucket. 
       SUMMARY OF THE INVENTION 
       [0007]    The present application and the resultant patent thus provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint including a flow disruption surface; passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region; substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine. 
         [0008]    The present application and the resultant patent further provide a gas turbine engine. The gas turbine engine may include a first can combustor generating a first combustion flow; a second can combustor generating a second combustion flow, wherein the first can combustor and the second can combustor meet at a joint including a flow disruption surface; and a turbine comprising a first stage bucket; wherein the flow disruption surface promotes mixing of the first combustion flow and the second combustion flow to form a mixed combustion flow in a mixing region upstream of the first stage bucket to improve durability of the first stage bucket. 
         [0009]    The present application and the resultant patent further provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a number of combustion flows in a number of can combustors positioned in a circumferential array, wherein each pair of adjacent can combustors meets at a joint including a flow disruption surface; passing the number of combustion flows over the flow disruption surfaces and to a mixing region; substantially mixing the number of combustion flows in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine. 
         [0010]    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 
         [0011]      FIG. 1  is a schematic view of a known gas turbine engine that may be used herein. 
           [0012]      FIG. 2  is a side cross-sectional view of a can combustor that may be used with the gas turbine engine of  FIG. 1 . 
           [0013]      FIG. 3  is an end plan view of a number of adjacent can combustors. 
           [0014]      FIG. 4  is a schematic view of a number of adjacent can combustors and the first two rows of turbine airfoils with a wake downstream of the can combustors. 
           [0015]      FIG. 5  is a schematic view of a number of adjacent can combustors and the first two rows of turbine airfoils illustrating the use of the can combustor mixing joints as may be described herein. 
           [0016]      FIG. 6  is a perspective view of a can combustor mixing joint as may be described herein. 
           [0017]      FIG. 7  is an end plan view of an alternative embodiment of a can combustor mixing joint as may be described herein. 
           [0018]      FIG. 8  is an end plan view of an alternative embodiment of a can combustor mixing joint as may be described herein. 
       
    
    
     DETAILED DESCRIPTION 
       [0019]    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. 
         [0020]    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 any one 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. 
         [0021]      FIG. 2  shows one example of the can combustor  25 . Generally described, the can combustor  25  may include a head end  55 . The head end  55  generally includes the various manifolds that supply the necessary flows of air  20  and fuel  30 . The can combustor  25  also includes an end cover  60 . A number of fuel nozzles  65  may be positioned within the end cover  60 . A combustion zone  70  may extend downstream of the fuel nozzles  65 . The combustion zone  70  may be enclosed within a liner  75 . A transition piece  80  may extend downstream of the combustion zone  70 . The can combustor  25  described herein is for the purpose of example only. Many other types of combustor designs may be used herein. Other components and other configurations also may be used herein. 
         [0022]    As is shown in  FIG. 3 , a number of the can combustors  25  may be positioned adjacent one another in a circumferential array. Likewise, as is shown in FIG.  4 , each pair of adjacent can combustors  25  may meet at a joint  85 . As was described above, the flows of combustion gases  35  through the pair of adjacent can combustors  25  may create a wake  90  downstream of the joint  85 . Specifically, the flows of combustion gases  35  may create the wake  90  immediately downstream of the joint  85 , as is shown. The wake  90  may be a stagnant flow in a low velocity flow region  92 . The wakes  90  of the flows of combustion gases  35  through the number of can combustors  25  extend into the airfoils  95  of the turbine  40 . Specifically, the wakes  90  extend into the airfoils  95  of a first stage nozzle  96 , wherein the flows of combustion gases  35  are accelerated so as to exaggerate the non-uniformities therein. The flows of combustion gases  35  then exit the first stage nozzle  96  and enter a first stage bucket  97 . The wakes  90  of the flows of combustion gases  35  generally mix in the first stage bucket  97  but incur significant mixing and pressure losses. Other components and other configurations may be used herein. 
         [0023]      FIG. 5  shows as portion of a gas turbine engine  100  as may be described herein. The gas turbine engine  100  includes a number of adjacent can combustors  110  positioned in a circumferential array. In this example, three (3) can combustors  110  are shown: a first can combustor  120  with a first combustion flow  125 , a second can combustor  130  with a second combustion flow  135 , and a third can combustor  140  with a third combustion flow  145 . Any number of adjacent can combustors  110  may be used herein. Each pair of adjacent can combustors  110  meets at a mixing joint  150 . Each mixing joint  150  may have a flow disruption surface  155  defined thereon so as to promote mixing of the combustion flows  125 ,  135 ,  145 . The gas turbine engine  100  further includes a turbine  160  positioned downstream of the can combustors  110 . The turbine  160  includes a number of airfoils  170 . In this example, the airfoils  170  may be arranged as a first stage nozzle  180  and a first stage bucket  190  of the turbine  160 . Any number of nozzles and buckets may be used herein. Other components and other configurations may be used herein. 
         [0024]      FIGS. 6-8  show a number of different embodiments of the mixing joint  150  between adjacent can combustors  110  as may be described herein.  FIG. 6  shows a chevron mixing joint  200 . The chevron mixing joint  200  may include a first set of chevron like spikes  210  defined by the first can combustor  120  and a corresponding second set of chevron like spikes  220  defined by the second can combustor  130 , which define the flow disruption surfaces  155 . Specifically, the first set of chevron like spikes  210  may be defined by a downstream edge of a first wall  230  of the first can combustor  120 , and the second set of chevron like spikes  220  may be defined by a downstream edge of a second wall  240  of the second can combustor  130 . In this manner, the first and second can combustors  120 ,  130  meet at the chevron mixing joint  200  between the first wall  230  and the second wall  240 . As is shown, the flow disruption surfaces  155  may face downstream from the first and second can combustors  120 ,  130  and toward the turbine  160 . Further, as is shown, the depth and angle of the first and second sets of chevron like spikes  210 ,  220  may vary from the first can combustor  120  to the second can combustor  130 . Likewise, the number, size, shape, and configuration of the chevron like spikes  210 ,  220  each may vary. Other components and other configurations may be used herein. 
         [0025]      FIG. 7  shows a further embodiment of the mixing joint  150  as may be described herein. In this embodiment, a lobed mixing joint  250  is shown. The lobed mixing joint  250  may include a first set of lobes  260  defined by the first can combustor  120  and a second set  270  of lobes defined by the second can combustor  130 , which define the flow disruption surfaces  155 . Specifically, the first set of lobes  260  may be defined by the downstream edge of the first wall  230  of the first can combustor  120 , and the second set of lobes  270  may be defined by the downstream edge of the second wall  240  of the second can combustor  130 . In this manner, the first and second can combustors  120 ,  130  meet at the lobed mixing joint  250  between the first wall  230  and the second wall  240 . As is shown, the flow disruption surfaces  155  may face downstream from the first and second can combustors  120 ,  130  and toward the turbine  160 . The first and second sets of lobes  260 ,  270  may have a largely sinusoidal wave like shape and may mate therewith. The depth and shape of the first and second set of lobes  260 ,  270  also may vary. The number, size, shape, and configuration of the lobes  260 ,  270  may vary. Other components and configurations may be used herein. 
         [0026]      FIG. 8  shows a further embodiment of the mixing joint  150  as may be described herein. In this embodiment, the mixing joint  150  may be in the form of a fluidics mixing joint  280 , as is shown. The fluidics mixing joint  280  may include a number of jets  290  therein that act as a flow disruption surface  155 . Specifically, as is shown, the jets may be positioned between the first wall  230  of the first can combustor  120  and the second wall  240  of the second can combustor  130 . The jets  290  may spray a fluid  300  into the flows of combustion gases  125 ,  135  as they exit the first can combustor  120  and the second can combustor  130 . The number, size, shape, and configuration of the jets  290  may vary. Likewise, the nature of the fluid  300  may vary. Other components and configurations may be used herein. 
         [0027]    Referring again to  FIG. 5 , the use of the mixing joints  150  described herein may enhance the mixing of the combustion flows  125 ,  135 ,  145  from adjacent can combustors  120 ,  130 ,  140 . Specifically, the various geometries of the flow disruption surfaces  155  of the mixing joints  150  may enhance the mixing of the combustion flows  125 ,  135 ,  145  in a mixing region  305  positioned downstream of the mixing joints  150 . As is shown, the mixing region  305  may be positioned immediately downstream of the mixing joints  150 . As a result of the enhanced mixing, a wake  310  formed by the combustion flows  125 ,  135 ,  145  may be much smaller than the wake  90  described above. Because the enhanced mixing of the combustion flows  125 ,  135 ,  145  occurs in the mixing region  305  before entry into the first stage nozzle  180 , the mixing may result in significantly less mixing losses as compared to mixing downstream in the first stage nozzle  180 , the first stage bucket  190 , or elsewhere. The enhanced mixing thus may reduce the overall pressure losses in the gas turbine engine  100  as a whole without increasing the axial distance between the can combustors  110  and the turbine  160 . 
         [0028]    Use of the gas turbine engine  100  described herein may include generating the combustion flows  125 ,  135 ,  145  in the adjacent can combustors  120 ,  130 ,  140  and then passing the combustion flows  125 ,  135 ,  145  over the flow disruption surfaces  155  of the mixing joints  150  and to the mixing region  305 . The combustion flows  125 ,  135 ,  145  may be passed over the flow disruption surfaces  155  and to the mixing region  305  at a first velocity. In this manner, the combustion flows  125 ,  135 ,  145  from the adjacent can combustors  120 ,  130 ,  140  may substantially mix in the mixing region  305  to form a mixed combustion flow  315  upstream of and before entry into the turbine  160 . In other words, the mixed combustion flow  315  may be a substantially homogenous mixture of the combustion flows  125 ,  135 ,  145  from the adjacent can combustors  120 ,  130 ,  140 . The mixed combustion flow  315  then may be passed to the first stage nozzle  180  of the turbine  160 , in which the mixed combustion flow  315  may be accelerated to a second velocity greater than the first velocity. The mixed combustion flow  315  then may be passed to the first stage bucket  190  of the turbine  160  at the second velocity. Because the mixed combustion flow  315  is formed upstream of and before entry into the turbine  160 , the passing of the mixed combustion flow  315  to the first stage bucket  190  may generate a substantially uniform velocity field in the first stage bucket  190 . Moreover, because the mixed combustion flow  315  is formed upstream of and before entry into the turbine  160 , the passing of the mixed combustion flow  315  to the first stage bucket  190  may generate a substantially uniform temperature field in the first stage bucket  190 . In this manner, the first stage turbine bucket  190  may be subjected to reduced high cycle fatigue loads as well as reduced thermal loads. The mixed combustion flow  315  formed during use of the gas turbine engine  100  thus may improve the durability of the first stage bucket  190 . 
         [0029]    The embodiments of the mixing joint  150  described herein are for purposes of example only. Other mixing joint geometries or other types of flow disruption surfaces  155  that enhance mixing of the combustion flows  125 ,  135 ,  145  from adjacent can combustors  120 ,  130 ,  140  before entry into the turbine  160  may be used herein. Different types of flow disruption surfaces  155  may be used herein together. Other components and other configurations also may be used herein. 
         [0030]    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 invention as defined by the following claims and the equivalents thereof