Abstract:
A method enables a combustor for a gas turbine engine to be assembled. The method includes coupling an inner liner to an outer liner such that a combustion chamber is defined therebetween, wherein the outer liner is fabricated from a plurality of panels coupled together, and coupling an outer support radially outward from the outer liner such that an outer passageway is defined between the outer liner and the outer support, wherein the outer support is configured to channel cooling air from the outer passageway towards at least a portion of the outer liner.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to combustors and, more particularly to a method and apparatus for decreasing combustor acoustics.  
         [0002]     At least some known gas turbine engines include a compressor for compressing air which is suitably mixed with a fuel and channeled to a combustor wherein the mixture is ignited for generating hot combustion gases. At least some known combustors include a dome assembly, a cowling, and inner and outer liners to channel the combustion gases to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. The liners are coupled to the dome assembly with the cowling, and extend downstream from the cowling to define the combustion chamber. An outer support is coupled radially outward from the outer liner such that an outer cooling passage is defined radially outward from the outer liner, and an inner support is coupled radially inward from the inner liner such that an inner cooling passage is defined therebetween.  
         [0003]     At least some known liners include a plurality of panels that are serially connected together between the upstream and aft ends of each liner such that the panels define the combustion chamber. Because the panels are exposed to high operating temperatures generated within the combustion chamber, cooling air is channeled through the cooling passages to facilitate reducing the operating temperature of the panels. Specifically, such panels rely on backside convection cooling from the cooling passage flow, as well as hot-side film cooling, to facilitate enhancing extending the useful life of the panels. However, because the aft most panel is radially inward and downstream from the outer support, the aft most panel generally receives significantly less convective cooling air than other liner panels because the outer support directs the cooling downstream through turbine cooling feed windows. Over time, the reduced cooling of the aft panel may cause the aft panel to exceed predetermined operating limits established to facilitate optimizing the useful life of the liner.  
       BRIEF DESCRIPTION OF THE INVENTION  
       [0004]     In one aspect, a method for assembling a combustor for a gas turbine engine is provided. The method comprises coupling an inner liner to an outer liner such that a combustion chamber is defined therebetween, wherein the outer liner is fabricated from a plurality of panels coupled together, and coupling an outer support radially outward from the outer liner such that an outer passageway is defined between the outer liner and the outer support, wherein the outer support is configured to channel cooling air from the outer passageway towards at least a portion of the outer liner.  
         [0005]     In another aspect, a combustor for a gas turbine engine is provided. The combustor includes an inner liner, an outer liner, and an outer support. The outer liner is coupled to the inner liner such that a combustion chamber is defined therebetween. The outer support is coupled radially outward from the outer liner such that an outer cooling passageway is defined radially outward from the outer liner. The outer support is configured to channel cooling air from the outer passageway towards at least a portion of the outer liner.  
         [0006]     In a further aspect, a gas turbine engine is provided. The gas turbine engine includes a combustor including an inner liner, an outer liner, and an outer support. The outer liner coupled to the inner liner such that a combustion chamber is defined therebetween. The outer support is coupled radially outward from the outer liner such that an outer passageway is defined radially outward from the outer liner. The outer support is configured to channel cooling air from the outer passageway towards at least a portion of the outer liner.  
     
    
     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 a combustor that may be used with the gas turbine engine;  
         [0009]      FIG. 3  is an enlarged perspective view of a portion of the combustor shown in  FIG. 2 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0010]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine  10  including a low pressure compressor  12 , a high pressure compressor  14 , and a combustor  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 . In one embodiment, gas turbine engine  10  is an LMS100 engine commercially available from General Electric Company, Cincinnati, Ohio.  
         [0011]     In operation, air flows through low pressure compressor  12  from an upstream side  28  of engine  10 . Compressed air is supplied from low pressure compressor  12  to high pressure compressor  14 . Highly compressed air is then delivered to combustor assembly  16  where it is mixed with fuel and ignited. Combustion gases are channeled from combustor  16  to drive turbines  18  and  20 .  
         [0012]      FIG. 2  is a cross-sectional view of a combustor  30  that may be used with gas turbine engine  10 .  FIG. 3  is an enlarged perspective view of the portion of combustor  30  shown in  FIG. 3 . Combustor  30  includes a dome assembly  32 . A fuel injector  34  extends into dome assembly  32  and injects atomized fuel through dome assembly  32  into a combustion zone or chamber  36  of combustor  30  to form an air-fuel mixture that is ignited downstream of the fuel injector  
         [0013]     Combustion zone  36  is defined by combustor liners  40  that shield components external to combustor  30  from heat generated within combustion zone  36 . Combustion zone  36  extends from dome assembly  32  downstream to a turbine nozzle assembly  41 . Liners  40  include an inner liner  42  and an outer liner  44 . Each liner  42  and  44  is annular and includes a plurality of separate panels  50 . In the exemplary embodiment, each panel  50  includes a series of steps  52 , each of which form a distinct portion of combustor liner  40 .  
         [0014]     Outer liner  44  and inner liner  42  each include a respective aft-most panel  64  and  66 . Panels  64  and  66  are each located at the aft end  68  of combustion zone  36  and are adjacent turbine nozzle assembly  41 . Specifically, each panel  64  and  66  couples an aft end  70  and  72  of each respective liner  44  and  42  to turbine nozzle assembly  41 . Each combustor panel  50  includes a combustor liner surface  80  and an exterior surface  82  that is radially outward from liner surface  80 . Combustor liner surface  80  extends generally from dome assembly  32  to turbine nozzle assembly  41 .  
         [0015]     Each liner  42  and  44  also includes an annular support mount, or aft mount,  90  and  92 , respectively. Specifically, each support mount  90  and  92  couples an aft end  72  and  70  of each respective liner  42  and  44  to turbine nozzle assembly  41  and to a combustor casing  94  extending substantially circumferentially around combustor  30 . More specifically, each support mount  90  and  92  extends radially outward from each respective liner  42  and  44  such that a radially outer cooling passageway  96  and a radially outer cooling passageway  98  are defined between combustor casing  94  and combustor liner  40 . Accordingly, cooling passageway  96  is adjacent liner  42 , and cooling passageway  98  is adjacent liner  44 .  
         [0016]     Each support mount  90  and  92  includes a radial portion  100  and a conical datum area  102 . Each radial portion  100  extends generally axially downstream from each conical datum area  102 . A plurality of turbine cooling feed windows  104  that extend between a radially outer surface  106  and a radially inner surface  108  of each support mount  90  and  92 . In the exemplary embodiment, windows  104  are spaced circumferentially between circumferential sides  110  and  112  of each mount  90  and  92 . During operation, turbine cooling feed windows  104  facilitate channeling cooling air from cooling passageways  96  and  98  towards nozzle assembly  41 .  
         [0017]     Each conical datum area  102  extends between each radial portion  100  and each aft combustor liner panel  64  and  66 . More specifically, each conical datum area  102  extends obliquely downstream, with respect to an axis of rotation of gas turbine engine  10 , from each aft combustor liner panel  64  and  66 . Each conical datum area  102  includes a plurality of cooling openings  120  extending therethrough. In the exemplary embodiment, cooling openings  120  are each substantially circular and are spaced substantially equi-distantly between mount sides  110  and  112 . However, it should be understood that openings  120  may be any desired shape and/or configuration. Accordingly, each opening  120  is aligned obliquely with respect to an outer surface  106  of each mount  90  and  92 .  
         [0018]     During operation, cooling air is channeled into cooling passageways  96  and  98  to facilitate backside cooling of liners  40 . Specifically, cooling air from passageways  96  and  98  facilitates backside convective cooling and film cooling of panels  50 . Because aft panels  64  is downstream from outer mount  92 , if not for openings  120 , mount  92  may inhibit the flow of convective cooling air towards aft panel  64 . However, openings  120  facilitate augmenting backside convection cooling of panels  64 . Specifically, cooling air is channeled from cooling passage  98  through openings  120  wherein it is directed in an oblique direction for impingement against a exterior surface  82  to facilitate convective cooling of panels  64 . More specifically, openings  120  are oriented to ensure cooling air directed therethrough impinges on an axial center of each aft panel  64  to facilitate reducing an operating temperature of each panel  64 . The enhanced cooling of aft panel  64  facilitates increasing a heat transfer coefficient of surface  82 , thus extending the useful life of combustor  30  in comparison to  
         [0019]     The above-described gas turbine engine combustor includes a radial outer support that includes a plurality of cooling openings extending therethrough. The cooling openings enable cooling fluid from the outer cooling passage to be channeled towards the aft most combustor liner panel to facilitate reducing the operating temperature of that panel. Specifically, the cooling openings are oriented to enable cooling air discharged therefrom to impinge against the backside of the aft-most radially-outer liner panel. As such, backside convective and/or impingement cooling of the aft-most panel is facilitated to be increased in a cost-effective and reliable manner.  
         [0020]     Exemplary embodiments of a combustor for a gas turbine engine are described above in detail. The systems and assembly components of the combustor are not limited to the specific embodiments described herein, but rather, components of each system may be utilized independently and separately from other components described herein. Each system and assembly component can also be used in combination with other combustor systems and assemblies or with other gas turbine engine components.  
         [0021]     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.