Abstract:
A method for assembling a gas turbine engine is provided. The method includes providing a combustor liner support including a first portion formed with a plurality of circumferentially-spaced recessed areas, and a second portion extending from the first portion. The method also includes coupling the combustor liner support within a combustor such that during engine operations cooling air is channeled through the plurality of recessed areas to facilitate cooling the combustor liner support.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to gas turbine engines, and more particularly, to combustor liner supports used in gas turbine engines.  
         [0002]     At least some known combustors include at least one dome coupled to a combustor liner that defines a combustion zone. More specifically, the combustor liner includes an inner liner and an outer liner that extend from the dome to a turbine nozzle. Each liner is spaced radially inwardly from a combustor casing such that an inner and an outer cooling passageway are defined between each respective liner and the combustor casing. The inner passageway, in particular, enables cooling air to be channeled across an exterior of the inner liner. The cooling air facilitates reducing the temperature of the combustion chamber inner liner to facilitate reducing thermal stresses.  
         [0003]     At least some known combustor inner liners are coupled to a forward inner nozzle support via a bolted joint. Such bolted joints, although structurally sound, may limit the useful life expectancy of the combustor liner. For example bolted joints may result in restricting cooling air flow through the inner passageway. As a result, high loss turns, recirculation zones, and/or cooling air backflow may be generated within the inner passageway, which may cause thermal stresses to be induced to the combustor inner liner. Over time, such stresses may decrease the life cycle fatigue (LCF) of the compressor. Moreover, bolted joints may also increase the overall weight of the combustor, which may adversely affect the performance and the life expectancy of the gas turbine engine as a whole.  
       BRIEF DESCRIPTION OF THE INVENTION  
       [0004]     In one aspect, a method for assembling a gas turbine engine is provided. The method includes providing a combustor liner support including a first portion formed with a plurality of circumferentially-spaced recessed areas, and a second portion extending from the first portion. The method also includes coupling the combustor liner support within a combustor such that during engine operations cooling air is channeled through the plurality of recessed areas to facilitate cooling the combustor liner support.  
         [0005]     In another aspect, a combustor liner support for a gas turbine engine combustor is provided. The combustor liner support includes a first portion including a plurality of circumferentially-spaced recessed areas and a second portion extending from the first portion for coupling the combustor liner support to a portion of the combustor. The plurality of recessed areas are configured to channel cooling air through the combustor liner support.  
         [0006]     In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a nozzle assembly and a combustor assembly coupled upstream from the nozzle assembly. The combustor includes a liner and a liner support, wherein the combustor liner support includes a first portion and a second portion. The first portion includes a plurality of circumferentially-spaced recessed areas configured to channel cooling air through the combustor liner support. The second portion extends from the first portion. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0007]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine;  
         [0008]      FIG. 2  is a cross-sectional view of an exemplary combustor that may be used with the gas turbine engine shown in  FIG. 1 ;  
         [0009]      FIG. 3  is a cross-sectional view of a portion of the combustor shown in  FIG. 2 ;  
         [0010]      FIG. 4  is a perspective view of an exemplary embodiment of a portion of a combustor liner support that may be used with the combustor shown in  FIGS. 2 and 3 ;  
         [0011]      FIG. 5  is a perspective view of a portion of a forward inner nozzle support that may be used with the combustor shown in  FIGS. 2 and 3 ; and  
         [0012]      FIG. 6  is a perspective view of the combustor inner liner support shown in  FIG. 4 , and coupled to the forward inner nozzle support shown in  FIG. 5 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0013]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine  10 . Engine  10  includes a low pressure compressor  12 , a high pressure compressor  14 , and a combustor assembly  16 . Engine  10  also includes a high pressure turbine  18 , and a low pressure turbine  20  arranged in a serial, axial flow relationship. Compressor  12  and turbine  20  are coupled by a first shaft  24 , and compressor  14  and turbine  18  are coupled by a second shaft  26 .  
         [0014]      FIG. 2  is a cross-sectional view of combustor  16  that may be used in gas turbine engine  10 . Combustor  16  includes an annular outer liner  40 , an annular inner liner  42 , and a domed end (not shown) that extends between outer and inner liners  40  and  42 , respectively. Inner liner  42  is coupled to a forward inner nozzle support  43  via a combustor inner liner support  44 . Forward inner nozzle support  43  is spaced radially inward from inner liner  42 . Specifically, in the exemplary embodiment, a plurality of fastener assemblies  45  couple inner liner support  44  substantially flush to forward inner nozzle support  43 . More specifically, combustor inner liner support  44  is coupled to inner liner  42  at a downstream end  46  of combustor  16 . Outer liner  40  and inner liner  42  are spaced inward from a combustor casing  47  such that a combustion chamber  48  is defined between liners  40  and  42 . Outer liner  40  and combustor casing  47  define an outer passageway  52 , and inner liner  42  and forward inner nozzle support  43  define an inner passageway  54 . Combustion chamber  48  is generally annular in shape and is defined between liners  40  and  42 . Outer and inner liners  40  and  42  extend from the domed end, to a turbine nozzle  56  coupled downstream from combustor  16 .  
         [0015]     A plurality of fuel igniters  62  extend through combustor casing  47  and outer passageway  52 , and couple to combustor outer liner  40 . Igniters  62  are bluff bodies that are placed circumferentially around combustor  16  and are downstream from the combustor domed end. Each igniter  62  is positioned to ignite a fuel/air mixture within combustion chamber  48 , and each includes an igniter tube assembly  64  coupled to combustor outer liner  40 . More specifically, each igniter tube assembly  64  is coupled within an opening  66  extending through combustor outer liner  40 , such that each igniter tube assembly  64  is concentrically aligned with respect to each opening  66 . Igniter tube assemblies  64  facilitate maintaining alignment of each respective igniter  62  relative to combustor  16 .  
         [0016]      FIG. 3  is an enlarged cross-sectional view of an embodiment of a portion of combustor  16 . Specifically, within  FIG. 3  inner liner  42  is coupled to forward inner nozzle support  43  via combustor liner support  44 . Combustor liner support  44  includes a first portion  70  and a second portion  72 . Both portions  70  and  72  are annular and extend substantially circumferentially around forward inner nozzle support  43 . Specifically, second portion  72  is coupled to combustor inner liner  42  adjacent combustor downstream end  46 . Portion  70  is coupled substantially flush against forward inner nozzle support  43 , such that combustor inner liner support  44  is positioned within passageway  54 .  
         [0017]     To facilitate an adequate flow of cooling air  73  through passageway  54 , combustor inner liner support  44  is fabricated with a plurality of passageways (described in more detail with respect to  FIG. 4 ) to enable cooling air  73  to pass through passageway  54  substantially uninhibited. Specifically, as cooling air  73  reaches combustor inner liner support  44 , cooling air  73  is divided into an outer flow path  74  and an inner flow path  75 . Outer flow path  74  is channeled between combustor inner liner  42  and combustor inner liner support  44 . Cooling air  73  within flow path  74  is then channeled through the plurality of liner support passageways and between combustor inner liner support second portion  72  and forward inner nozzle support  43 . Inner flow path  75  is channeled through a plurality of passageways defined between adjacent pairs of fastener assemblies  45 . Specifically, within flow path  75 , cooling air  73  is channeled between combustor inner liner support first portion  70  and combustor inner liner support second portion  72 .  
         [0018]      FIG. 4  is a perspective view of an exemplary embodiment of a portion of a combustor inner liner support  44  that may be used with gas turbine engine  10  (shown in  FIG. 1 ).  FIG. 5  is a perspective view of a portion of a forward inner nozzle support  43  that may be used with gas turbine engine  10  (shown in  FIG. 1 ).  FIG. 6  is a perspective view of combustor inner liner support  44  coupled to forward inner nozzle support  43 . Liner support first portion  70  is fabricated with a plurality of recessed areas  76  and a plurality of coupling portions  77 . Both recessed areas  76  and coupling portions  77  extend circumferentially around first portion  70 , and are oriented such that each recessed area  76  is positioned between a pair of coupling portions  77 . Recessed areas  76  provide flow path passageways between circumferentially adjacent pairs of fastener assemblies  45  to facilitate air flow along flow path  75  defined between combustor inner liner support first portion  70  and forward inner nozzle support  43 . In the exemplary embodiment, each coupling portion  77  includes a plurality of fastener openings  78  sized to receive fastener assemblies therein to enable combustor liner support  44  to be coupled substantially flush against forward inner nozzle support  43 .  
         [0019]     Liner support second portion  72  is fabricated with a plurality of apertures  79  that are spaced circumferentially about an outer perimeter of second portion  72 . More specifically, in the exemplary embodiment each aperture  79  is radially outward from a respective first portion coupling portion  77 . Apertures  79  facilitate channeling cooling air  73 , along flow path  74 , and between combustor inner liner support second portion  72  and forward inner nozzle support  43 . Specifically, apertures  79  facilitate directing air flow over fastener assemblies  45 .  
         [0020]     Forward inner nozzle support  43  is frusto-conical, and includes a plurality of attachment portions  80  spaced circumferentially about its perimeter. Each attachment portion  80  may be coupled to a corresponding combustor inner liner support coupling portion  77 . Specifically, each coupling portion  77  and attachment portion  80  is coupled via fastener assemblies  45  to enable combustor inner liner support  44  to couple substantially flush against forward inner nozzle support  43 . Accordingly combustor inner liner support  44  is radially outward from, and extends along an outer circumference of forward inner nozzle support  43 .  
         [0021]     In operation, air flows through low pressure compressor  12  from an upstream side  28  of engine  10 . Compressed air is channeled from low pressure compressor  12  to high pressure compressor  14 . Compressed air is then delivered to combustor assembly  16  wherein it is mixed with fuel and ignited. Combustion gases are channeled from combustor  16  to drive turbines  18  and  20 .  
         [0022]     Airflow exits high pressure compressor  14  at a relatively high velocity and is channeled into combustor  16  wherein the airflow is mixed with fuel and the fuel/air mixture is ignited for combustion using igniters  62 . As airflow enters combustor  16 , cooling air  73  is channeled through combustor inner passageway  54 . Airflow in passageway  54  facilitates cooling inner liners  42 .  
         [0023]     Cooling air  73  is divided at combustor inner liner support  44 . Specifically, a first portion of cooling air  73  is directed along outer flow path  74 , and a second portion of cooling air  73  is directed along inner flow path  75 . Flow path  74  directs cooling air  73  over fastener  45 , and between inner liner  42  and combustor inner liner support second portion  72  to facilitate cooling inner liner  42 . Furthermore, cooling air  73  flows along inner liner  42  and is channeled through apertures  79  to allow cooling air  73  to flow between combustor inner liner support second portion  72  and forward inner nozzle support  43 .  
         [0024]     To facilitate preventing air flow restriction in passageway  54 , flow path  75  directs cooling air  73  between combustor inner liner support first portion  70  and forward inner nozzle support  43 . Specifically, recessed areas  76  define passageways between adjacent pairs of coupling portions  77  and channel cooling air  73  between adjacent coupling portions  77  to facilitate cooling fastener assemblies  45 . After passing through recessed areas  76 , flow path  75  conjoins with flow path  74  between combustor inner liner support second portion  72  and forward inner nozzle support  43 . As such, apertures  79  and recessed areas  76  facilitate providing uninhibited flow of cooling air  73  through inner passageway  54 .  
         [0025]     The above-described methods and apparatus facilitate allowing uninhibited air flow of cooling air through the combustor. Specifically air flow through the recessed areas defined in the combustor inner liner support facilitate preventing high loss turns, recirculation zones, and/or cooling air backflow within the inner passageway of the combustor. Accordingly, the recessed areas facilitate reducing thermal stresses on the combustor inner liner, and also facilitate increasing life cycle fatigue and improving the life span of the gas turbine engine.  
         [0026]     As used herein, an element or step recited in the singular and proceeded with the word “a” or “an” should be understood as not excluding plural said elements or steps, unless such exclusion is explicitly recited. Furthermore, references to “one embodiment” of the present invention are not intended to be interpreted as excluding the existence of additional embodiments that also incorporate the recited features.  
         [0027]     Although the methods and systems described herein are described in the context of fabricating a combustor inner liner support of a gas turbine engine, it is understood that the methods and systems described herein are not limited to combustor inner liner supports or gas turbine engines. Likewise, the combustor inner liner support components illustrated are not limited to the specific embodiments described herein, but rather, components of the combustor inner liner support can be utilized independently and separately from other components described herein.  
         [0028]     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.