Abstract:
A liner cooling assembly for a gas turbine system includes a liner having an outer surface and an inner surface, the inner surface defining an interior region. Also included is a sleeve disposed radially outwardly of the outer surface of the liner, the sleeve and the outer surface of the liner defining a channel configured to receive a cooling flow. Further included is a cooling jacket extending circumferentially around at least a portion of the outer surface of the liner and disposed between the liner and the sleeve, wherein the cooling jacket comprises at least one support operably coupled to the cooling jacket and the outer surface of the liner.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The subject matter disclosed herein relates to gas turbine systems, and more particularly to a liner cooling assembly for such gas turbine systems. 
         [0002]    A combustor section of a gas turbine system typically includes a combustor chamber disposed relatively adjacent a transition piece, where a hot gas passes from the combustor chamber through the transition piece to a turbine section. At least a portion of the combustor chamber is often surrounded by a flow sleeve, while at least a portion of the transition piece is surrounded by an impingement sleeve. The flow sleeve typically includes a plurality of apertures for providing impingement cooling for portions of a liner of the combustor. An additional airflow passes from a region defined by the impingement sleeve and the transition piece to a region defined by the flow sleeve and the combustor liner. The impingement cooling of the liner of the combustor is achieved by cooling jets that are pushed onto the liner in a direction relatively perpendicular to the additional airflow flowing from the region proximate the impingement sleeve to the region proximate the flow sleeve. The additional airflow, perpendicular to the impingement jet is one disruption, among other inefficiencies, resulting in a reduced cooling efficiency of the combustor liner. Conventionally, it is called cross flow. In general, the cooling air is directly used for the combustion air. Otherwise, a much more complex system has to be developed to deliver and discharge the spent cooling air. An associated expense of using combustion air for cooling is the loss of kinetic energy, which is commonly referred to as pressure loss. Reduction of cooling air immediately results in the benefits of lower pressure loss, which will produce higher overall gas turbine efficiency and cost savings. Operation capability of a large scale gas turbine is defined by emissions and thermo-acoustic response of a combustor. Control of cooling air will reduce air temperature variation that limits operability. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0003]    According to one aspect of the invention, a liner cooling assembly for a gas turbine system includes a liner having an outer surface and an inner surface, the inner surface defining an interior region. Also included is a sleeve disposed radially outwardly of the outer surface of the liner, the sleeve and the outer surface of the liner defining a channel configured to receive a cooling flow. Further included is a cooling jacket extending circumferentially around at least a portion of the outer surface of the liner and disposed between the liner and the sleeve, wherein the cooling jacket comprises at least one support operably coupled to the cooling jacket and the outer surface of the liner. 
         [0004]    According to another aspect of the invention, a combustor liner cooling assembly for a gas turbine system includes a combustor liner defining a combustor chamber. Also included is a flow sleeve surrounding at least a portion of the combustor liner, the flow sleeve and the combustor liner defining a channel configured to receive a cooling flow. Further included is a cooling jacket surrounding at least a portion of the combustor liner and disposed between the combustor liner and the flow sleeve, wherein the cooling jacket comprises at least one support extending radially inwardly from the cooling jacket toward the combustor liner. 
         [0005]    According to yet another aspect of the invention, a combustor liner cooling assembly for a gas turbine system includes a combustor liner defining a combustor chamber. Also included is a flow sleeve surrounding at least a portion of the combustor liner, the flow sleeve and the combustor liner defining a channel configured to receive a cooling flow. Further included is a cooling jacket integrally formed with the combustor liner, wherein the cooling jacket includes a body portion disposed radially outwardly of the combustor liner. The cooling jacket also includes a plurality of support members fixedly connected to an inner surface of the body portion and an outer surface of the combustor liner. The cooling jacket further includes a plurality of apertures extending through the body portion of the cooling jacket. 
         [0006]    These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
         [0007]    The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
           [0008]      FIG. 1  is a partial, schematic illustration of a combustor section of a gas turbine system; 
           [0009]      FIG. 2  is an enlarged view of section II of  FIG. 1 , illustrating a liner cooling assembly; 
           [0010]      FIG. 3  is an enlarged, schematic illustration of section III of  FIG. 2 , illustrating a support of the cooling assembly; 
           [0011]      FIG. 4  is a schematic cross-sectional view taken along lines IV-IV of  FIG. 2  of a portion of the cooling assembly having a plurality of apertures disposed therein; and 
           [0012]      FIG. 5  is a cross-sectional view taken along lines V-V of  FIG. 4 . 
       
    
    
       [0013]    The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings. 
       DETAILED DESCRIPTION OF THE INVENTION 
       [0014]    The terms “axial” and “axially” as used in this application refer to directions and orientations extending substantially parallel to a center longitudinal axis of a turbine system. The terms “radial” and “radially” as used in this application refer to directions and orientations extending substantially orthogonally to the center longitudinal axis of the turbine system. The terms “upstream” and “downstream” as used in this application refer to directions and orientations relative to an axial flow direction with respect to the center longitudinal axis of the turbine system. 
         [0015]    With reference to  FIG. 1 , a partial schematic illustrates a combustor section of a gas turbine system and is referred to generally with numeral  10 . The combustor section  10  includes a transition piece  12  defining a transition region  14  that is at least partially surrounded by an impingement sleeve  16  disposed radially outwardly of the transition piece  12 . Upstream thereof, proximate a forward end  18  of the impingement sleeve  16  is a combustor liner  20  defining a combustor chamber  22 . The combustor liner  20  is at least partially surrounded by a flow sleeve  24  disposed radially outwardly of the combustor liner  20 . A forward sleeve  26  is located at the junction between the forward end  18  of the impingement sleeve  16  and an aft end  28  of the flow sleeve  24 . 
         [0016]    Although the combustor liner  20  and the transition piece  12  are described above and illustrated as being distinct, separate components, it is to be appreciated that a single, integrated liner may define the combustor chamber  22  and the transition region  14 . In such an embodiment, a single sleeve may be employed to surround the liner, rather than two separate sleeves, such as the flow sleeve  24  and the impingement sleeve  16  described above. 
         [0017]    The combustor section  10  uses a combustible liquid and/or gas fuel, such as a natural gas or a hydrogen rich synthetic gas, to run the gas turbine system. The combustor chamber  22  is configured to receive and/or provide an air-fuel mixture, thereby causing a combustion that creates a hot pressurized gas through the transition piece  12  into the turbine section (not illustrated), causing rotation of the turbine section. The presence of the hot pressurized gas increases the temperature of the combustor liner  20  surrounding the combustor chamber  22 , particularly proximate a downstream end  30  of the combustor liner  20 . To overcome issues associated with excessive thermal exposure to the combustor liner  20 , a plurality of apertures  32  within the flow sleeve  24  are arranged to provide impinged air in the form of a plurality of cooling jets onto the combustor liner  20 . A cross-flow  36  flows relatively perpendicularly to the plurality of cooling jets. Specifically, the cross-flow  36  flows from a region defined by the impingement sleeve  16  and the transition region  14  to a region defined by the flow sleeve  24  and the combustor liner  20 . 
         [0018]    Referring now to  FIG. 2 , an enlarged view of the region defined by the flow sleeve  24  and the combustor liner  20  is shown in greater detail. Although the following description is made with reference to the region defined by the flow sleeve  24  and the combustor liner  20 , as noted above, it is contemplated that exemplary embodiments relate to the region defined by the impingement sleeve  16  and the transition piece  12 . Yet other embodiments include a region defined by a single sleeve and a single, integrated liner defining the transition region  14  and the combustor chamber  22 . 
         [0019]    Disposed within an annular channel  38  defined by the flow sleeve  24  and the combustor liner  20  is a cooling jacket  40  that includes a body portion  42  extending circumferentially around at least a portion of an outer surface  44  of the combustor liner  20 , which also includes an inner surface  45 . The body portion  42  includes a body portion inner surface  46  and a body portion outer surface  48 . At least one, but typically a plurality of support members  50  are disposed between the body portion  42  of the cooling jacket  40  and the outer surface  44  of the combustor liner  20 . Each of the plurality of support members  50  are operably connected to the cooling jacket  40  and typically are integrally formed with the cooling jacket  40 . The plurality of support members  50  are also typically operably connected to the combustor liner  20 , with the operable connection comprising any suitable fastening structure, such as a mechanical fastener or a weld, for example. Additionally, in one embodiment, the cooling jacket  40  is integrally formed with the combustor liner  20  by a fixed connection between the plurality of support members  50  and the combustor liner  20 . 
         [0020]    The plurality of support members  50  may be formed in various geometric configurations, with an exemplary geometric configuration comprising an airfoil-shaped member that is configured to interact with a first cooling flow  52  that is split from a second cooling flow  54 . The first cooling flow  52  is directed between the cooling jacket  40  and the combustor liner  20 , while the second cooling flow  54  is directed between the cooling jacket  40  and the flow sleeve  24 . It is also contemplated that the plurality of support members  50  may be of various alternative geometries, such as a cylindrical member, for example. Irrespective of the precise geometric configuration, the plurality of support members  50  may be disposed in numerous arrangements. Typically, the plurality of support members  50  are disposed at a plurality of axial locations and circumferentially spaced from one another. The plurality of support members  50  can be used to reduce the cross flow effects from the first cooling flow  52 . The plurality of support members  50  can be more sophisticated, as will be discussed below with reference to  FIG. 3  or simply a wall-like structure as shown in  FIG. 5 . 
         [0021]    The cooling jacket  40  includes at least one, but typically a plurality of apertures  56  extending through the body portion  42  of the cooling jacket  40 . The plurality of apertures  56  provide additional impinged air in the form of convective cooling streams  58  that are in close proximity to the outer surface  44  of the combustor liner  20 , thereby enhancing the convective cooling of the combustor liner  20 . 
         [0022]    Referring now to  FIGS. 3 and 4 , enlarged views of the annular channel  38 , as well as the combustor liner  20  and the cooling jacket  40 , are shown in greater detail. As illustrated, one or more of the plurality of support members  50  may include a hollow portion  60  configured to receive a portion of the second cooling flow  54 . Injection of the second cooling flow  54  into the hollow portion  60  of the plurality of support members  50  provides a cooling effect on the plurality of support members  50 , which conductively cools the combustor liner  20  to which the plurality of support members  50  are operably connected to. The portion of the second cooling flow  54  that is circulated within the plurality of support members  50  may be expelled through a hole  62  extending from the hollow portion  60  to the annular channel  38 . Furthermore, the hole  62  may be aligned to expel the second cooling flow  54  toward the outer surface  44  of the combustor liner  20 , which enhances the convective cooling effect that is already provided by the plurality of apertures  56  disposed within the body portion  42  of the cooling jacket  40 . 
         [0023]    The plurality of support members  50  may be arranged in a staggered arrangement to form a torturous path for the first cooling flow  52  to flow through. Such an arrangement includes positioning portions of the plurality of support members  50  in relative circumferential alignment with at least one of the plurality of apertures  56  disposed in the body portion  42  of the cooling jacket  40 , thereby diverting the first cooling flow  52  to reduce a disturbance of the convective cooling streams  58  generated by the plurality of apertures  56 . The convective cooling streams  58  more efficiently cool targeted locations of the combustor liner  20 . Additionally, the diversion of the first cooling flow  52  increases the average velocity of the first cooling flow  52 , which increases the convective heat transfer associated with the flowing of the first cooling flow  52  over the combustor liner  20 . 
         [0024]    While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.