Patent Publication Number: US-9404379-B2

Title: Gas turbine shroud assemblies

Description:
FIELD OF THE DISCLOSURE 
     Embodiments of the disclosure relate generally to gas turbine engines and more particularly to gas turbine shroud assemblies. 
     BACKGROUND OF THE DISCLOSURE 
     Gas turbines are widely used in industrial and commercial operations. A typical gas turbine includes a compressor at the front, one or more combustors around the middle, and a turbine at the rear. The compressor imparts kinetic energy to the working fluid (e.g., air) to produce a compressed working fluid at a highly energized state. The compressed working fluid exits the compressor and flows to the combustors where it mixes with fuel and ignites to generate combustion gases having a high temperature and pressure. The hot combustion gases flow to the turbine where they expand to produce work. Consequently, the turbine is exposed to very high temperatures due to the hot combustion gases. As a result, the various turbine components, such as the turbine shrouds, typically need to be cooled. Accordingly, there is a need to provide improved shroud cooling systems and methods. 
     BRIEF DESCRIPTION OF THE DISCLOSURE 
     Some or all of the above needs and/or problems may be addressed by certain embodiments of the present disclosure. According to one embodiment, there is disclosed a gas turbine shroud assembly. The assembly may include a shroud structure that defines a first cooling chamber and a second cooling chamber. The assembly may also include a first impingement plate disposed within the first cooling chamber and a second impingement plate disposed within the second cooling chamber. Further, the assembly may include one or more cooling channels formed within the shroud structure. The cooling channels may be configured to connect the first cooling chamber with the second cooling chamber. The assembly may also include a flow of cooling air in communication with the first cooling chamber. In this manner, the flow of cooling air may flow from the first cooling chamber to the second cooling chamber by way of the one or more cooling channels. 
     According to another embodiment, there is disclosed a method. The method may include flowing cooling air into a first cooling chamber defined within a shroud structure. The method may also include flowing the cooling air through a first impingement plate disposed within the first cooling chamber so as to increase the velocity of the flow of cooling air to increase the heat transfer coefficient within the first cooling chamber. Further, the method may include flowing the cooling air through one or more cooling channels formed within the shroud structure to a second cooling chamber defined within the shroud structure. The method may also include flowing the cooling air through a second impingement plate disposed within the second cooling chamber so as to increase the velocity of the flow of cooling air to increase the heat transfer coefficient within the second cooling chamber. 
     Further, according to another embodiment, there is disclosed a gas turbine assembly. The gas turbine assembly may include a rotating blade assembly. The gas turbine assembly may also include a shroud structure positioned about the rotating blade assembly. The shroud structure may define a first cooling chamber and a second cooling chamber. A first impingement plate may be disposed within the first cooling chamber, and a second impingement plate may be disposed within the second cooling chamber. The gas turbine assembly may also include one or more cooling channels formed within the shroud structure. The cooling channels may be configured to connect the first cooling chamber with the second cooling chamber. Further, the gas turbine assembly may include a flow of cooling air in communication with the first cooling chamber. The flow of cooling air may flow from the first cooling chamber to the second cooling chamber by way of the one or more cooling channels. 
     Other embodiments, aspects, and features of the invention will become apparent to those skilled in the art from the following detailed description, the accompanying drawings, and the appended claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Reference will now be made to the accompanying drawings, which are not necessarily drawn to scale, and wherein: 
         FIG. 1  is an example schematic view of a gas turbine engine, according to an embodiment of the disclosure. 
         FIG. 2  is an example schematic cross-sectional view of a gas turbine shroud assembly, according to an embodiment of the disclosure. 
         FIG. 3  is an example schematic view of one or more cooling channels formed within the shroud structure, according to an embodiment of the disclosure. 
         FIG. 4  is an example schematic view of an impingement plate, according to an embodiment of the disclosure. 
     
    
    
     DETAILED DESCRIPTION OF THE DISCLOSURE 
     Illustrative embodiments will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all embodiments are shown. The present disclosure may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Like numbers refer to like elements throughout. 
     Illustrative embodiments are directed to, among other things, gas turbine shroud assemblies. For example,  FIG. 1  depicts an example schematic view of a gas turbine assembly  100  as may be used herein. The gas turbine assembly  100  may include a gas turbine having a compressor  102 . The compressor  102  may compress an incoming flow of air  104 . The compressor  102  may deliver the compressed flow of air  104  to a combustor  106 . The combustor  106  may mix the compressed flow of air  104  with a pressurized flow of fuel  108  and ignite the mixture to create a flow of combustion gases  110 . Although only a single combustor  106  is shown, the gas turbine engine may include any number of combustors  106 . The flow of combustion gases  110  may be delivered to a turbine  112 . The flow of combustion gases  110  may drive the turbine  112  so as to produce mechanical work. The mechanical work produced in the turbine  112  may drive the compressor  102  via a shaft  114  and an external load  116 , such as an electrical generator or the like. 
     The gas turbine engine may use natural gas, various types of syngas, and/or other types of fuels. The gas turbine engine may be any one of a number of different gas turbine engines offered by General Electric Company of Schenectady, N.Y., including, but not limited to, those such as a 7 or a 9 series heavy duty gas turbine engine and the like. The gas turbine engine may have different configurations and may use other types of components. The gas turbine engine may be an aeroderivative gas turbine, an industrial gas turbine, or a reciprocating engine. 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. 
     In certain embodiments, as schematically depicted in  FIG. 2 , the turbine  112  may include a gas turbine shroud assembly  200 . The shroud assembly  200  may form part of the turbine  112 . For example, the shroud assembly  200  may define a hot gas path  202 , in which the flow of combustion gases  110  travels. Moreover, the shroud assembly  200  may be positioned about a rotating blade  204  or the like. In this manner, the flow of combustion gases  110  may drive the rotating blade  204  to produce work. In some instances, as discussed in greater detail below, the shroud assembly  200  may be cooled by a flow of cooling air from the compressor  102  or elsewhere. That is, a flow of cooling air may at least partially flow throughout the shroud assembly  200 . One or more shroud assemblies  200  may be positioned adjacent to one another. For example, the shroud assemblies  200  may be positioned circumferentially adjacent to one another about the rotating blade  204  so as to define a portion of the hot gas path  202 . 
     As the combustion gases  110  travel along the hot gas path  202 , at least a port of the combustion gases  110  may pass between the rotating blade  204  and the shroud assembly  200 . As a result, the shroud assembly  200  may be heated by the combustion gases  110 . In some instances, the leading edge of shroud assembly  200  may become hotter than the trailing edge of the shroud assembly  200 . The systems and methods described herein are configured to cool the shroud assembly  200 . 
     Still referring to  FIG. 2 , the shroud assembly  200  may include a shroud structure  206 . In certain embodiments, the shroud structure  206  may be annular. The shroud structure  206  may include a single unitary structure or a number of structures formed together. Any number of shroud structures  206  may be used. For example, the shroud structure  206  may include an annular shroud support assembly and/or a shroud ring attached thereto. 
     The shroud structure  206  may define a first cooling chamber  208  and a second cooling chamber  210 . That is, the various structural members of the shroud structure  206  may collectively define the first cooling chamber  208  and the second cooling chamber  210 . For example, a first shroud wall  209 , a second shroud wall  211 , an outer shroud portion  213 , and an inner portion  223  may define the first cooling chamber  208 . Likewise, a third shroud wall  219 , the second shroud wall  211 , the outer shroud portion  213 , and the inner portion  223  may define the second cooling chamber  210 . With reference to the flow of hot combustion gases  110 , the first cooling chamber  208  may be positioned upstream of the second cooling chamber  210 . For example, the first cooling chamber  208  may be positioned about a leading edge of the blade  204 , and the second cooling chamber  210  may be positioned about a trailing edge of the blade  204 . The pressure within the first cooling chamber  208  may be greater than the pressure within the second cooling chamber  210 . Any number of cooling chambers may be used herein. 
     The shroud assembly  200  may also include a first impingement plate  212  positioned within the first cooling chamber  208  and a second impingement plate  214  positioned within the second cooling chamber  210 . In some instances, the first impingement plate  212  may be positioned between the first shroud wall  209  and the second shroud wall  211  within the first cooling chamber  208 . In other instances, the second impingement plate  214  may be at least partially supported within the second cooling chamber  210  by a radially extending support member  217  and the third shroud wall  219 . The first impingement plate  212  and the second impingement plate  214  may each include a number of holes  215  therein. In some instances, the holes  215  may include one or more variably sized holes. Moreover, the holes  215  within the first impingement plate  212  and the second impingement plate  214  may be the same size or a different size. That is, the holes  215  within the first impingement plate  212  may be a first size, and the holes  215  within the second impingement plate  214  may be a second size. 
     The shroud assembly  200  may also include one or more cooling channels  216  formed within the shroud structure  206 . For example, the cooling channels  216  may be formed on a surface of the inner shroud portion  223  of the shroud structure  206 . The cooling channels  216  may extend axially between the first cooling chamber  208  to the second cooling chamber  210 . In this manner, the cooling channels  216  may be configured to connect the first cooling chamber  208  with the second cooling chamber  210 . The cooling channels  216  may be configured to cool the inner portion  223 . For example, the cooling channels  216  may extend along the leading edge of the inner portion  223 , which may be hotter than the trailing edge of the inner portion  223 . In this manner, the cooling channels  216  may cool the leading edge of the inner portion  223 . 
     The first cooling chamber  208 , the cooling channels  216 , and the second cooling chamber  210  may collectively define a flow path. For example, as indicated by the dotted lines, the shroud assembly  200  may include a flow of cooling air  218  therethrough. In some instances, the flow of cooling air  218  may be a secondary flow of air supplied by the compressor  102 . However, other sources of cooling air  218  may also be used herein. 
     The flow of cooling air  218  may be in communication with the first cooling chamber  208 . That is, the flow of cooling air  218  may initially enter the first cooling chamber  208 . The flow of cooling air  218  may then pass through the first impingement plate  212  via the holes  215 . The first impingement plate  212  may be configured to create an increase in the velocity of the flow of cooling air  218  within the first cooling chamber  208 . The increase in velocity increases the heat transfer coefficient within the first cooling chamber  208  and facilitates the cooling of the shroud assembly  200 . The flow of cooling air  218  may then flow from the first cooling chamber  208  to the second cooling chamber  210  by way of the cooling channels  216 . The flow of cooling air  218  passing through the cooling channels  216  may facilitate the cooling of the leading edge of the inner shroud portion  223  adjacent to the hot gas path  202 . After entering the second cooling chamber  210 , the flow of cooling air  218  may then pass through the second impingement plate  214  via the holes  215 . The second impingement plate  214  may be configured to create an increase in the velocity of the flow of cooling air  218  within the second cooling chamber  210 . The increase in velocity increases the heat transfer coefficient within the second cooling chamber  210  and facilitates the cooling of the shroud assembly  200 . 
     In some instances, the first cooling chamber  208  may include one or more cooling passages  220  configured to discharge at least a portion of the flow of cooling air  218  into a hot gas path  202  near the leading edge of the blade  204 . In other instances, the second cooling chamber  210  may include one or more exit passages  222  configured to discharge the flow of cooling air  218  from the second cooling chamber into a hot gas path  202  near a trailing edge of the blade  204 . 
       FIG. 3  depicts a schematic view of the inner portion  223  of the shroud assembly  200 . As noted above, the inner portion  223  of the shroud assembly  200  may include a number of cooling channels  216  formed therein. The cooling channels  216  may be any depth and/or any length to enable the passage of cooling air  218  from the first cooling chamber  208  to the second cooling chamber  210 . For example, the cooling channels may extend the entire or partial length of the inner portion  223  of the shroud assembly  200 . Further, the cooling channels  216  may be uniform or otherwise. In some instances, the cooling channels  216  may be positioned about the leading edge of the inner portion  223 . 
       FIG. 4  depicts a schematic view of the first impingement plate  212  of the shroud assembly  200 . As noted above, the first impingement plate  212  may include a number of holes  215  therein. The holes  215  may be uniform or the holes  215  may vary in size. As depicted in  FIG. 4 , the holes  215  about the leading edge of first impingement plate  212  are smaller than the holes about the trailing edge of the first impingement plate  212 . The holes  215  may be any configuration to optimize cooling of the shroud assembly  200 . Similarly, the second impingement plate  214  may include a number of holes  215  therein. The configuration of the holes  215  in the first impingement plate  212  may be the same or different from the configuration of the holes  215  in the second impingement plate  214 . 
     Although embodiments have been described in language specific to structural features and/or methodological acts, it is to be understood that the disclosure is not necessarily limited to the specific features or acts described. Rather, the specific features and acts are disclosed as illustrative forms of implementing the embodiments.