Patent Abstract:
A system and method for tuning a gas turbine combustion system having a plurality of seals positioned between the combustion system and the turbine inlet is disclosed. The system and method provide ways of permitting a predetermined amount of compressed air to bypass the combustion system and enter the turbine so as to control emissions and dynamics of the combustion system. The seals contain a plurality of holes to meter airflow passing therethrough and are positioned such that they can be removed from the engine and modified to increase or decrease the amount of air passing therethrough.

Full Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     Not applicable. 
     TECHNICAL FIELD 
     The present invention generally relates to gas turbine engines. More particularly, embodiments of the present invention relate to a combustion system and a method of operation of the combustion system in order to provide an additional way of controlling engine emissions and combustion dynamics. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engines operate to produce mechanical work or thrust. For land-based gas turbine engines, a generator is typically coupled to the shaft, such that the mechanical work produced is harnessed to generate electricity. A typical gas turbine engine comprises a compressor, at least one combustor, and a turbine, with the compressor and turbine coupled together through an axial shaft. In operation, air passes through the compressor, where the pressure of the air increases and then passes to a combustion section, where fuel is mixed with the compressed air in one or more combustion chambers. The hot combustion gases then pass into the turbine and drive the turbine. As the turbine rotates, the compressor turns, since they are coupled together along a common shaft. The turning of the shaft also drives the generator for electrical applications. The gas turbine engine also must operate within the confines of the environmental regulations for the area in which the engine is located. As a result, more advanced combustion systems have been developed to more efficiently mix fuel and air so as to provide more complete combustion, which results in lower emissions. 
     Low emissions combustion systems require the fuel and air being mixed to be properly proportioned in order to obtain optimal results. Fuel flows are usually tightly controlled through carefully sized orifices in the fuel nozzles and controlled fuel valves. Airflows may actually vary due to distributions driven by the compressor exit profile and the amount of air required to cool the turbine section. Because the amount of air introduced into the combustion system significantly affects reaction zone temperature and performance of the combustion system, an adjustable air mass is advantageous for regulating the combustion process. 
     A general issue with gas turbines, and especially industrial gas turbines, is the need to be able to tune the combustors to avoid issues such as lean blow out (LBO), where the combustor is operating too lean and is not receiving enough fuel, for a given amount of air, causing the flame to be extinguished. Another known problem of tuning a gas turbine combustor include excessive combustion dynamics caused by rapid changes in pressures within the combustor. 
     To compensate and control these combustion instabilities, prior gas turbine combustors incorporated additional dilution holes in the combustion liner or a transition piece in order to control the amount of air being used in the combustion process. However, these forms of “air control” have been known to adversely effect emissions of the combustion system, at least with respect to carbon monoxide. 
     SUMMARY 
     Embodiments of the present invention are directed towards a system and method for, among other things, tuning a gas turbine engine to avoid operational and emissions issues found in prior art designs. 
     In one embodiment of the present invention, a gas turbine combustion system comprises a combustion liner, a flow sleeve encompassing the combustion liner, an end cap positioned near an end of the combustion liner and the flow sleeve. A plurality of fuel nozzles extend through the cap and towards the combustion liner. A transition duct couples the aft end of the combustion liner to an inlet of the turbine in order to direct the flow of hot combustion gases from the combustor to the turbine. A plurality of tunable side seals are positioned between adjacent transition ducts and the inlet of the turbine. The plurality of side seals each have one or more openings located therein that permit a controlled amount of air to pass therethrough and bypass the combustion system. 
     In an alternate embodiment, a method of tuning a combustion system of a gas turbine engine is disclosed. A portion of an airflow source to be supplied to the combustion system is determined and then, a size and quantity of openings for a plurality of seals is determined in which the size and quantity will result in the portion of an airflow source being supplied to the combustion system by permitting the remainder of the airflow source to bypass the combustion system. Once the size and quantity of openings are determined, the openings are placed in the plurality of seals and the seals are then placed in the gas turbine engine to regulate the amount of airflow permitted to bypass the combustion system. 
     In yet another alternate embodiment, a tunable side seal for use in a gas turbine combustor is disclosed wherein the seal comprises one or more sheets of material secured together having one or more holes located through the one or more sheets. The seal is sized and configured to be positioned between sidewalls of adjacent transition ducts and a turbine inlet. Furthermore, the seals are oriented in a manner so as to be accessible from outside of a gas turbine engine such that the seal can be removed and the one or more holes altered to adjust the amount of air permitted to pass therethrough. 
     Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       The present invention is described in detail below with reference to the attached drawing figures, wherein: 
         FIG. 1  depicts a perspective view of a portion of a gas turbine engine of the prior art; 
         FIG. 2  depicts a perspective view of a portion of a gas turbine engine in accordance with an embodiment of the present invention; 
         FIG. 3  depicts a cross section of a gas turbine engine in accordance with an embodiment of the present invention; 
         FIG. 4  depicts an elevation view of a seal used in an embodiment of the present invention; 
         FIG. 5  depicts an elevation view of an alternate seal in an embodiment of the present invention; 
         FIG. 6  depicts an elevation view of yet another seal in an embodiment of the present invention; and, 
         FIG. 7  is a chart identifying a method of tuning a combustion system of a gas turbine engine in accordance with an embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION 
     The subject matter of the present invention is described with specificity herein to meet statutory requirements. However, the description itself is not intended to limit the scope of this patent. Rather, the inventors have contemplated that the claimed subject matter might also be embodied in other ways, to include different components, combinations of components, steps, or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies. 
     Referring initially to  FIG. 1 , a view of a portion of a combustion system  100  of the prior art is disclosed. The combustion system  100  includes a plurality of combustion liners (not shown) with each liner coupled to a transition duct  102  and the transition duct  102  is in turn coupled to the turbine inlet  104 . Transition ducts  102  direct the flow of hot combustion gases from a combustion liner to the turbine inlet  104 . Prior art combustors attempted to direct all of the air from the compressor (except for that used for turbine cooling) to the combustion system  100  for maximum efficiency by placing solid seals  106  between adjacent transition ducts  102  and the turbine inlet  106 . As previously disclosed, a gas turbine operator or manufacturer could place or adjust size and location of dilution holes in the combustion liner or transition duct  102  in an effort to tailor the airflow to the combustion system. However, such efforts affected the combustion system emissions as well as the temperature profile entering the turbine. Furthermore, the use of solid seals  106  has also resulted in too much air being provided to the combustion system, resulting in an overly lean fuel-air mixture. 
     Referring to  FIGS. 2-7 , multiple embodiments of the present invention are shown.  FIG. 2  depicts a portion of a gas turbine combustion system  200  having a tunable side seal  202 , where the seal  202  is shown in greater detail in  FIGS. 4-6 . Referring to  FIG. 3 , a tunable gas turbine combustion system  200  comprises a combustion liner  204 , a flow sleeve  206  encompassing the combustion liner  204  and an end cap  208  positioned proximate a forward end of the combustion liner  204  and flow sleeve  206 . A plurality of fuel nozzles  210  extend through openings in the end cap  208  with the fuel nozzles  210  extending towards the combustion liner  204 . Coupled to the aft end of the combustion liner  204  is a transition duct  212  that directs the hot combustion gases from the combustion liner  204  into a turbine inlet  214 . In the embodiment shown in  FIG. 3 , a double-walled transition duct is utilized. Referring to  FIGS. 3 and 4 , a plurality of tunable side seals  202  are located adjacent to the transition duct  204  and have one or more openings  218  located therein. The openings  218 A aid in tuning the combustion system  200  by permitting a predetermined amount of air to pass therethrough. As a result of the openings  218 A, a controlled portion of air bypasses the combustion system  200 , including the combustion liner  204  and transition duct  212 . Directing a predetermined amount of air through the side seals  202  provides the operator with a way of tuning the combustion system  200  by setting a quantity and size of openings  218 A which will regulate the amount of air directed to the combustion system  200 . 
     The combustion system  200  is generally a can-annular system where there are a plurality of individual combustion systems arranged about a centerline or longitudinal axis of a gas turbine engine as shown in  FIG. 3 . Each combustion liner  204  and transition duct  212  feed hot combustion gases into a portion of the turbine inlet  214 . As a result of the combustion system orientation, the plurality of side seals  202  are oriented generally radially outward relative to the centerline A-A, as shown in  FIG. 3 . An additional advantage provided by this seal orientation is the ability to remove the plurality of side seals  202  from the combustion system  200 . This allows for the one or more openings  218 A to be altered in size and/or quantity if an operator determines the amount of air passing therethrough, and bypassing the combustion system  200 , is either too much or too little. Openings  218 A can be welded closed should there be too much air passing therethrough, or the size of the openings can be increased if the air flow is too little. For example, a plurality of side seals  202  can be used to regulate the amount of air permitted to bypass the combustion system compatible with a General Electric Frame 7FA gas turbine engine. The seal arrangement for this type of combustion system generally permits up to approximately 2% of air from the compressor to bypass the combustion system and pass directly into the turbine. The present invention is not limited to this engine, but instead can be used on a variety of engine types and the total amount of air permitted to pass therethrough can vary. 
     The plurality of side seals  202  can be fabricated from a variety of materials and sizes depending upon the size and shape of slots between the transition duct  212  and turbine inlet  214  and the operating conditions. Because of the elevated operating temperatures, the plurality of seals  202  are generally fabricated from a high temperature cobalt-based alloy such as Haynes  188 . In an embodiment of the invention, the plurality of seals  202  are each generally fabricated from sheet metal, including an embodiment in which a plurality of sheets of metal are fixed together by brazing or a series of spot welds, such that the seal is flexible along the seal axis (S-A), as shown in  FIG. 4 . Due to the seal construction, the openings should be placed in areas absent of a weld or braze material so as to not initiate cracks in the joints between sheets of metal forming the seal. 
     In an embodiment of the present invention, a tunable side seal  202  in a gas turbine combustion system is disclosed. The tunable side seal  202  is fabricated from one or more sheets of material  220  having one or more openings or holes located through the one or more sheets. As an example, the side seal  202  can be fabricated from a cobalt-based alloy. The tunable side seal  202  is sized to be positioned between sidewalls (e.g.  232  and  234  of  FIG. 2 ) of adjacent transition ducts  212  and the turbine inlet  214 , as shown in  FIG. 4 . The exact size of the seals and their thickness depends on the configuration of the slot. However, slightly undersizing the thickness of the seal  202  compared to the slot will aid in permitting the seal  202  to be removed. 
     Where a seal  202  is fabricated from a plurality of sheets of metal that are fixed together along a seal centerline SC, the seal is flexible about its centerline. This flexibility also aids in the installation and removal of the seals  202  when the openings are to be adjusted. 
     As previously discussed, the plurality of seals  202  each has a plurality of openings or holes. The openings can be a variety of shapes and sizes depending upon the amount of air desired to pass through the seal. However, in order to avoid creating non-uniform cooling or “hot-spots” at the turbine inlet  214 , it is preferred that the same amount of air pass through each seal around the combustion system. Such a cooling scheme can be created by a uniform set of elliptically-shaped holes  218 A as shown in  FIG. 4 , a set of circular holes  218 B as shown in  FIG. 5 , or a varying pattern of holes  218 C across the seal as shown in  FIG. 6  as long as the total flow permitted to pass through each seal is generally equal around the turbine inlet  214 . 
     An additional alternate embodiment of the present invention discloses a method  700  of tuning a combustion system of a gas turbine engine, and is shown in  FIG. 7 . The method  700  comprises a step  702  of determining a portion of an airflow source that is to be supplied to the combustion system. Then, in a step  704 , the size and quantity of openings for the plurality of seals that will result in the desired portion of the airflow source to be supplied to the combustion system is determined. Then, in a step  706  the holes are placed in the plurality of seals, and then in a step  708 , the plurality of seals having the holes are placed into the gas turbine engine in a region between adjacent transition ducts and an inlet of the turbine. Once the seals are installed in the gas turbine engine and the engine runs, measurements and operational data can be recorded such that, in a step  710 , a determination can be made as to whether the combustion system is operating outside of its pre-determined limits. If the combustion system is not operating outside of its limits, then the process ends in a step  712 . However, if the determination is made that the combustion system is operating outside of the limits, and a change in air flow is desired, then in a step  714 , the seals are removed from the engine, and in a step  716 , the quantity and/or size of the openings are adjusted such that the flow of air bypassing the combustion system can be changed. If the airflow is too great, the hole size can be reduced or quantity of holes reduced. If the air flow is too little, the hole size can be increased or quantity of holes can be increased. 
     The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive. Alternative embodiments will become apparent to those of ordinary skill in the art to which the present invention pertains without departing from its scope. 
     From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.

Technology Classification (CPC): 5