Patent Publication Number: US-11377970-B2

Title: System and method for providing compressed air to a gas turbine combustor

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     Not applicable. 
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     Not applicable. 
     TECHNICAL FIELD 
     This present disclosure relates generally to a system for improving airflow supply and distribution to a gas turbine combustor. More specifically, embodiments of the present disclosure relate to a reconfigured air flow inlet region between a transition duct and a flow sleeve of the gas turbine combustor. 
     BACKGROUND OF THE DISCLOSURE 
     A gas turbine engine typically comprises a multi-stage compressor coupled to a multi-stage turbine via an axial shaft. Air enters the gas turbine engine through the compressor where its temperature and pressure increase as it passes through subsequent stages of the compressor. The compressed air is then directed to one or more combustors where it mixes with a fuel source to create a combustible mixture. This mixture is ignited in the one or more combustors to create a flow of hot combustion gases. These gases are directed into the turbine causing the turbine to rotate, thereby driving the compressor. The output of the gas turbine engine can be mechanical thrust via exhaust from the turbine or shaft power from the rotation of an axial shaft, where the axial shaft can drive a generator to produce electricity. 
     In a typical industrial gas turbine engine, the combustor section comprises a plurality of can-annular combustors. In this arrangement, a plurality of individual combustors is arranged about the axis of the gas turbine engine. Each of the combustors typically comprises a combustion liner positioned within a flow sleeve and one or more fuel nozzles located at an inlet of the combustion liner. Compressed air passes between the flow sleeve and the combustion liner and along an exterior surface of the combustion liner prior to being mixed with fuel in the combustion liner. By directing compressed air over the combustion liner, the air cools the combustion liner and is pre-heated prior to combustion, resulting in a more efficient combustion process. The air from the engine compressor can also be used to cool a transition duct, which, as one skilled in the art understands, is used to direct hot combustion gases from the combustion liner to the turbine inlet. 
     In prior art combustion systems, a portion of the compressed air was injected into the passage between the combustion liner and flow sleeve through a series of injection ports in the flow sleeve. This prior art configuration is shown in  FIG. 1  and includes a flow sleeve  100  having a plurality of openings  102  and injection ports or tubes  104 . Positioned within the flow sleeve  100  is a combustion liner  106  which is coupled to a transition duct  108 . The transition duct  108  includes an outer cooling sleeve  110 . Air from the engine compressor enters a channel  112  formed between the transition duct  108  and the outer cooling sleeve  110  and flows along an outer wall of the transition duct  108  and an outer wall of the combustion liner  106 , as indicated by the arrows in  FIG. 1 . The openings  102  and injection ports  104  of the flow sleeve  100  provide jets of cooling air aimed towards the combustion liner  106 . This arrangement creates a cross flow of cooling air resulting in an adverse interaction between air entering through the openings  102  and injection ports  104  and air in the channel  112 . As such, cooling of an aft end of the combustion liner is not as effective as desired. 
     BRIEF SUMMARY OF THE DISCLOSURE 
     The following presents a simplified summary of the disclosure to provide a basic understanding of some aspects thereof. This summary is not an extensive overview of the application. It is not intended to identify critical elements of the disclosure or to delineate the scope of the disclosure. Its sole purpose is to present some concepts of the disclosure in a simplified form as a prelude to the more detailed description that is presented elsewhere herein. 
     The present disclosure provides systems and methods for improving a flow of cooling air to a gas turbine combustion system, thereby providing a more uniform distribution of cooling air along a combustion liner. 
     In an embodiment of the disclosure, a transition duct for a gas turbine engine is provided and comprises an inlet ring, a duct body connected to the inlet ring, and an aft frame connected to the duct body. A bellmouth is positioned radially outward of the inlet ring and encompasses the inlet ring. A plurality of struts extends between the bellmouth and the inlet ring, where the struts have a leading edge, an opposing trailing edge, and a thickness. In this configuration, air for combustion in the gas turbine engine passes through the bellmouth, between the plurality of struts and is directed to a combustion system coupled to the transition duct. 
     In an alternate embodiment of the disclosure, a flow inlet device for a gas turbine combustor is provided. The flow inlet device comprises an inlet ring, a bellmouth positioned radially outward of and encompassing the inlet ring, and a plurality of struts extending between the inlet ring and the bellmouth. The inlet ring and the bellmouth direct air for use in the gas turbine combustor between the plurality of struts. 
     In yet another embodiment of the disclosure, a method of increasing airflow to a gas turbine combustor is provided. The method provides a transition duct for a gas turbine engine having an inlet ring, a duct body connected to the inlet ring, an aft frame connected to the duct body, a bellmouth positioned radially outward and encompassing the inlet ring, and a plurality of struts positioned between the bellmouth and the inlet ring, where the struts have a leading edge, an opposing trailing edge, and a thickness. A flow sleeve is coupled to the transition duct and a flow of air is directed through the bellmouth, between the plurality of struts, and towards an inlet of the gas turbine combustor. 
     The present disclosure is aimed at providing an improved way of directing cooling air into and along a gas turbine combustion system including improvements to various combustor hardware, such that overall cooling air distribution is improved. 
     These and other features of the present disclosure can be best understood from the following description and claims. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       The present disclosure is described in detail below with reference to the attached drawing figures, wherein: 
         FIG. 1  is a cross section view of a portion of a gas turbine combustor in accordance with the prior art. 
         FIG. 2  is a perspective view of a transition duct of a gas turbine combustor in accordance with an embodiment of the present disclosure. 
         FIG. 3  is an alternate perspective view of the transition duct of  FIG. 2  in accordance with an embodiment of the present disclosure. 
         FIG. 4  is a detailed perspective view of a portion of the transition duct of  FIG. 3  in accordance with an embodiment of the present disclosure. 
         FIG. 5  is an elevation view of the transition duct of  FIG. 2  in accordance with an embodiment of the present disclosure. 
         FIG. 6  is a partial cross section view of the transition duct of  FIG. 5  in accordance with an embodiment of the present disclosure. 
         FIG. 7  is an elevation view of a portion of the transition duct of  FIG. 2  in accordance with an embodiment of the present disclosure. 
         FIG. 8  is a partial cross section view of the transition duct of  FIG. 7  in accordance with an embodiment of the present disclosure. 
         FIG. 9  is a partial cross section view of a transition duct, flow sleeve, and combustion liner in accordance with an embodiment of the present disclosure. 
         FIG. 10  is an alternate perspective view of a transition duct in accordance with an embodiment of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     The present disclosure is intended for use in a gas turbine engine, such as a gas turbine used for aircraft engines and/or power generation. As such, the present disclosure is capable of being used in a variety of turbine operating environments, regardless of the manufacturer. 
     As those skilled in the art will readily appreciate, a gas turbine engine is circumferentially disposed about an engine centerline, or axial centerline axis. The engine includes a compressor, a combustion section and a turbine with the turbine coupled to the compressor via an engine shaft. As is well known in the art, air compressed in the compressor is mixed with fuel which is burned in the combustion section and expanded in turbine. For certain gas turbine engines, such as industrial gas turbines used in power generation, the combustion system comprises a plurality of interconnected can-annular combustion chambers with each combustion chamber directing hot combustion gases to a turbine inlet via a transition duct. The transition duct typically has a varying geometric profile in order to connect a cylindrical combustor to a portion of an annular turbine inlet. 
     Various embodiments of the present disclosure are depicted in  FIGS. 2-10 . Referring initially to  FIG. 2 , a transition duct  200  capable of connecting a combustion liner to the turbine is provided. The transition duct  200  comprises an inlet ring  202  connected to a duct body  204 , which together form a gas path profile for directing hot combustion gases to the turbine. The duct body  204  is typically actively cooled due to the operating temperatures of the transition duct  200 . For the embodiment depicted in  FIG. 2 , a plurality of cooling holes  206  are placed in the duct body  204 . The cooling holes  206  can vary in size, shape, orientation, and spacing in order to provide the required cooling flow to the duct body  204 , as various surfaces of the duct body  204  will require different amounts of cooling air. 
     Connected to an opposite end of the duct body  204  is an aft frame  208 . The aft frame  208  is formed in a shape corresponding to a portion of an inlet of the turbine section (not shown). For the transition duct  200 , the inlet ring  202  is generally cylindrical, the aft frame is an arc-shaped rectangular opening, and the duct body  204  transitions between these two openings. 
     Positioned radially outward of the inlet ring  202  and encompassing the inlet ring  202  is a bellmouth  210 . The bellmouth  210  provides an inlet  212  through which cooling air is provided to the combustion system, as depicted in  FIG. 9 . That is, air for cooling and combustion passes through the bellmouth  210 . The inlet  212  further encourages compressed air to enter the bellmouth  210  with a flared inlet  214 . The flared inlet, which is flared outward and away from the bellmouth  210 , helps to direct compressed air from the region around the duct body  204  and into bellmouth  210  by providing a wider opening to receive compressed air. 
     Extending radially between and attached to the inlet ring  202  and the bellmouth  210  is a plurality of struts  216 . The assembly of the inlet ring  202 , bellmouth  210 , and plurality of struts  216  is secured to the duct body  204  and can be an integral assembly, such as a weldment, brazed joints, or an integral one-piece casting. Each of struts  216  further comprises a leading edge  218 , an opposing trailing edge  220 , and a body  222  having a thickness therebetween. The leading edge  218  of strut  216  is located towards the flared inlet  214 . Since the struts  216  are positioned in a region of relatively cool air, and therefore do not need to be cooled, the struts  216  are solid. However, in an alternate configuration of the disclosure, the struts  216  can be hollow in order to reduce weight or should it be desired to inject a fluid through the struts. 
     The configuration of the struts  216  can vary depending on specific engine and combustor operating conditions. For example, in an embodiment of the disclosure, the plurality of struts  216  have a rounded leading edge  218  and a rounded trailing edge  220  with a constant thickness to the strut  216  therebetween. This configuration is depicted in  FIG. 4 . In an alternate embodiment of the disclosure, the leading edge  218  can be rounded, with the thickness of the strut tapering so that the trailing edge  220  is thinner than the leading edge  218 . In yet another embodiment of the present disclosure, the thickness of the struts  216  taper to a reduced thickness proximate the leading edge  218  and the trailing edge  220 . 
     For the embodiment of the disclosure depicted in  FIGS. 2-10 , the plurality of the struts  216  are oriented generally parallel with respect to an axis A-A. That is, the air flow enters the inlet  212 , passes through the struts  216  and then flows in a direction generally parallel to the orientation of the struts  216 . This air exits the bellmouth  210 , cools a combustion system, and is injected into the combustion liner where it is used in a combustion process. In an alternate embodiment, the plurality of struts  216  can be oriented at an angle relative to the axis A-A extending through the inlet ring  202 , thus imparting a swirl to the airflow passing between the struts  216 . In a further embodiment of the disclosure, the struts  216  can be curved where each of the struts  216  have an airfoil-like cross sectional shape, which can also be used to impart a swirl to the airflow. 
     In addition to the directional orientation of the struts  216 , the quantity and spacing of the struts  216  between the inlet ring  202  and bellmouth  210  can also vary. In the embodiment of the disclosure depicted in  FIGS. 2-10 , the plurality of struts  216  are equally spaced about the perimeter of the inlet ring  202 . However, in alternate configurations, the spacing between the struts  216  can be non-uniform. For example, depending on the flow of compressed air into the inlet  212  and desired distribution of cooling flow, one configuration may include large gaps between struts  216  or certain regions having struts  216  removed. Such larger gaps between struts  216  can permit more air to flow through these regions, thus increasing cooling flow to certain areas around the combustor. 
     The bellmouth  210  is described herein as an integral part of the transition duct  200 . However, it is to be understood that the bellmouth  210  could also be a separate component attached to the transition duct  200 . Where a separate bellmouth is used, the bellmouth can be attached to the inlet of a transition duct by a slip fit including a spring between the inner diameter of bellmouth and an outer diameter of the transition duct. 
     The present disclosure also provides a method of increasing airflow to a gas turbine combustor. Accordingly, a transition duct  200  having an inlet ring  202 , a duct body  204  connected to the inlet ring  202 , and an aft frame  208  connected to the duct body  204  is provided. The duct body  204  also comprises a bellmouth  210  positioned radially outward of and encompassing the inlet ring  202  and a plurality of struts  216  positioned between the bellmouth  210  and the inlet ring  202 . Referring now to  FIG. 9 , a flow sleeve  230  is provided and coupled to the transition duct  200 , such that the bellmouth  210  engages a flow sleeve aft end  232 . A combustion liner  240  engages the inlet ring  202  of the transition duct  200 , thereby forming a passage  242  between the combustion liner  240  and the flow sleeve  230 . 
     In operation, a flow of air from the engine compressor is provided to a compressor discharge plenum (not shown). This air can serve to cool the transition duct  200  and is then directed into the bellmouth  210  at inlet  212 , where it passes between struts  216 , which serve to properly orient and distribute the flow of compressed air in the passage  242 . This air flow then continues through the passage  242 , along an outer surface of the combustion liner  240 , and to an inlet of the combustor. 
     As a result of the bellmouth  210  and the plurality of struts  216  coupled to the inlet ring  202 , air for cooling the combustion liner  240  is more evenly distributed along an outer surface of the combustion liner  240 , thereby eliminating the need for the openings  102  and injector ports  104  in the flow sleeve of the prior art of  FIG. 1 . Eliminating these openings and injector ports in the flow sleeve allows for a further reduction of pressure drop across the combustion system and avoids cross-flow of different cooling air flows as seen in the prior art and other combustor designs. The airflow is also more evenly distributed to the inlet of the combustor, which will improve combustion efficiency and reduce combustion dynamics. 
     Although a preferred embodiment of this disclosure has been provided, one of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure. Since many possible embodiments may be made of the disclosure without departing from the scope thereof, it is to be understood that all matter herein set forth or shown in the accompanying drawings is to be interpreted as illustrative and not in a limiting sense. 
     From the foregoing, it will be seen that this disclosure is one well adapted to attain all the ends and objects hereinabove set forth together with other advantages which are obvious, and which are inherent to the structure. 
     It will be understood that certain features and subcombinations are of utility and may be employed without reference to other features and subcombinations. This is contemplated by and is within the scope of the claims.