Abstract:
A method facilitates assembling a combustor for a gas turbine engine. The method comprises coupling an inner liner to an outer liner such that a combustion chamber is defined therebetween, positioning an outer support a distance radially outward from the outer liner, and positioning an inner support a distance radially inward from the inner liner. The method also comprises forming at least two rows of impingement openings extending through at least one of the inner support and the outer support for channeling impingement cooling air therethrough towards at least one of the inner liner and the outer liner, and forming at least one row of dilution openings extending through at least one of the inner liner and the outer liner for channeling dilution cooling air therethrough into the combustion chamber

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
   The U.S. Government may have certain rights in this invention pursuant to contract number DAAE07-00-C-N086. 

   BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines, more particularly to combustors used with gas turbine engines. 
   Known 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 an inner liner that is coupled to an outer liner such that a combustion chamber is defined therebetween. Additionally, an outer support is coupled radially outward from the outer liner such that an outer cooling passage is defined therebetween, and an inner support is coupled radially inward from the inner liner such that an inner cooling passage is defined therebetween. 
   Within at least some known recuperated gas turbine engines, cooling requirements of turbines may create a pattern factor requirement at the combustor that may be difficult to achieve because of combustor design characteristics associated with recuperated gas turbine engines. More specifically, because of space considerations, such combustors may be shorter than other known gas turbine engine combustors. In addition, in comparison to other known gas turbine combustors, such combustors may include a steeply angled flowpath and large fuel injector spacing. 
   Accordingly, at least some known combustors include a dilution pattern of a single row of dilution jets to facilitate controlling the combustor exit temperatures. The dilution jets are supplied cooling air from an impingement array of openings extending through the inner and outer supports. However, because of cooling considerations downstream from the combustor and because of the limited number and relative orientation of such impingement and dilution openings, such combustors may only receive only limited dilution air from such openings. 
   BRIEF DESCRIPTION OF THE INVENTION 
   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, positioning an outer support a distance radially outward from the outer liner, and positioning an inner support a distance radially inward from the inner liner. The method also comprises forming at least two rows of impingement openings extending through at least one of the inner support and the outer support for channeling impingement cooling air therethrough towards at least one of the inner liner and the outer liner, and forming at least one row of dilution openings extending through at least one of the inner liner and the outer liner for channeling dilution air therethrough into the combustion chamber. 
   In another aspect, a combustor for a gas turbine engine is provided. The combustor includes an inner liner, an outer liner, an outer support, and an inner support. The outer liner is coupled to the inner liner to define a combustion chamber therebetween. The outer support is radially outward from the outer liner such that an outer passageway is defined between the outer support and the outer liner. The inner support is radially inward from the inner liner such that an inner passageway is defined between the inner support and the inner liner. At least one of the inner support and the outer support includes at least two rows of impingement openings arranged in an array and extending therethrough for channeling impingement cooling air towards at least one of the inner liner and the outer liner. At least one of the inner liner and the outer liner includes at least one row of dilution openings extending therethrough for channeling dilution air into the combustion chamber. 
   In a further aspect, a gas turbine engine including a combustor is provided. The combustor includes at least one injector, an inner liner, an outer liner, an outer support, and an inner support. The inner liner is coupled to the outer liner to define a combustion chamber therebetween. The inner and outer liners further define an injector opening, and the injector extends substantially concentrically through the injector opening. The outer support is spaced radially outward from the outer liner. The inner support is spaced radially inward from the inner liner. At least one of the inner support and the outer support includes at least two rows of impingement openings arranged in an array and extending therethrough for channeling impingement cooling air towards at least one of the inner liner and the outer liner. At least one of the inner liner and the outer liner includes at least one row of dilution openings extending therethrough for channeling dilution air into the combustion chamber. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic of a gas turbine engine. 
       FIG. 2  is a cross-sectional illustration of a portion of an annular combustor used with the gas turbine engine shown in  FIG. 1 ; 
       FIG. 3  is a roll-out schematic view of a portion of the combustor shown in  FIG. 2  and taken along area  3 ; 
       FIG. 4  is a roll-out schematic view of a portion of the combustor shown in  FIG. 2  and taken along area  4 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a schematic illustration of a gas turbine engine  10  including a compressor  14 , and a combustor  16 . Engine  10  also includes a high pressure turbine  18  and a low pressure turbine  20 . Compressor  14  and turbine  18  are coupled by a first shaft  24 , and turbine  20  drives a second output shaft  26 . Shaft  26  provides a rotary motive force to drive a driven machine, such as, but, not limited to a gearbox, a transmission, a generator, a fan, or a pump. Engine  10  also includes a recuperator  28  that has a first fluid path  29  coupled serially between compressor  14  and combustor  16 , and a second fluid path  31  that is serially coupled between turbine  20  and ambient  35 . In one embodiment, the gas turbine engine is an LV100 engine available from General Electric Company, Cincinnati, Ohio. In the exemplary embodiment, compressor  14  is coupled by a first shaft  24  to turbine  18 , and powertrain and turbine  20  are coupled by a second shaft  26 . 
   In operation, air flows through high pressure compressor  14 . The highly compressed air is delivered to recouperator  28  where hot exhaust gases from turbine  20  transfer heat to the compressed air. The heated compressed air is delivered to combustor  16 . Airflow from combustor  16  drives turbines  18  and  20  and passes through recouperator  28  before exiting gas turbine engine  10 . In the exemplary embodiment, during operation, air flows through compressor  14 , and the highly compressed recuperated air is delivered to combustor  16 . 
     FIG. 2  is a cross-sectional illustration of a portion of an annular combustor  16 .  FIG. 3  is a roll-out schematic view of a portion of combustor  16  and taken along area  3  (shown in  FIG. 2 ).  FIG. 4  is a roll-out schematic view of a portion of combustor  16  and taken along area  4  (shown in  FIG. 2 ). Combustor  16  includes an annular outer liner  40 , an outer support  42 , an annular inner liner  44 , an inner support  46 , and a dome  48  that extends between outer and inner liners  40  and  44 , respectively. 
   Outer liner  40  and inner liner  44  extend downstream from dome  48  and define a combustion chamber  54  therebetween. Combustion chamber  54  is annular and is spaced radially inward between liners  40  and  44 . Outer support  42  is coupled to outer liner  40  and extends downstream from dome  48 . Moreover, outer support  42  is spaced radially outward from outer liner  40  such that an outer cooling passageway  58  is defined therebetween. Inner support  46  also is coupled to, and extends downstream from, dome  48 . Inner support  46  is spaced radially inward from inner liner  44  such that an inner cooling passageway  60  is defined therebetween. 
   Outer support  42  and inner support  46  are spaced radially within a combustor casing  62 . Combustor casing  62  is generally annular and extends around combustor  16 . More specifically, outer support  42  and combustor casing  62  define an outer passageway  66  and inner support  46  and combustor casing  62  define an inner passageway  68 . Outer and inner liners  40  and  44  extend to a turbine nozzle  69  that is downstream from liners  40  and  44 . 
   Combustor  16  also includes a dome assembly  70  which includes an air swirler  90 . Specifically, air swirler  90  extends radially outwardly and upstream from a dome plate  72  to facilitate atomizing and distributing fuel from a fuel nozzle  82 . When fuel nozzle  82  is coupled to combustor  16 , nozzle  82  circumferentially contacts air swirler  90  to facilitate minimizing leakage to combustion chamber  54  between nozzle  82  and air swirler  90 . 
   Combustor dome plate  72  is mounted upstream from outer and inner liners  40  and  44 , respectively. Dome plate  72  contains a plurality of circumferentially spaced air swirlers  90  that extend through dome plate  72  into combustion chamber  54  and each include a center longitudinal axis of symmetry  76  that extends therethrough. Fuel is supplied to combustor  16  through a fuel injection assembly  80  that includes a plurality of circumferentially-spaced fuel nozzles  82  that extend through air swirlers  90  into combustion chamber  54 . More specifically, fuel injection assembly  80  is coupled to combustor  16  such that each fuel nozzle  82  is substantially concentrically aligned with respect to air swirlers  90 , and such that nozzle  82  extends downstream into air swirler  90 . Accordingly, a centerline  84  extending through each fuel nozzle  82  is substantially co-linear with respect to air swirler axis of symmetry  76 . 
   Because of the steeply angled flowpath  100  defined within combustor  16 , circumferential spacing between adjacent fuel nozzles  82  and air swirlers  90 , and downstream component cooling requirements, combustion gases generated within combustor  16  are cooled prior to being discharged from combustor  16  to enable combustor  16  to maintain a pre-determined pattern factor. Combustor pattern factor is generally defined as:
 
 PF =( T 4     peak   −T 4     avg )/( T 4     avg   −T   35 )
 
where T 4  refers to the combustor exit temperature, T35 refers to the combustor inlet temperature, and T 4   peak  refers to the maximum temperature measured, and T 4   avg . refers to the average of the temperatures measured. Pattern factor is a measure of the distortion in combustor exit temperature and generally, a lower value is more desirable.
 
   Accordingly, combustor outer and inner liners  40  and  44 , each include a plurality of dilution jets  110  to facilitate locally cooling combustion gases generated within combustion chamber  54 , and to provide radial and circumferential exit temperature distribution. In the exemplary embodiment, dilution jets  110  are substantially circular and extend through liners  40  and  44 . More specifically, outer liner  40  includes a plurality of primary larger diameter dilution openings  120 , a plurality of smaller diameter dilution openings  122 , and a plurality of secondary dilution openings  124 . Openings  120 ,  122 , and  124  extend circumferentially around combustor  16 . 
   Smaller diameter outer primary dilution openings  122  are positioned substantially axially downstream with respect to air swirler centerline  76  at pre-determined distances D 1  downstream from dome  72 . More specifically, in the exemplary embodiment, smaller outer primary dilution openings  122  are positioned downstream from dome plate  72  at a distance D 1  that is approximately equal 0.65 combustor passage heights h 1 . Combustor passage heights h 1  is defined as the measured distance between outer and inner liners  40  and  44  at combustor chamber upstream end  74 . 
   Larger diameter outer primary dilution openings  120  have a larger diameter d 2  than a diameter d 3  of smaller diameter outer primary dilution openings  122 , and are positioned between adjacent air swirlers  90  at the same axial locations as openings  122 . In one embodiment, larger diameter openings  120  have a diameter d 2  that is approximately equal 0.307 inches, and smaller diameter openings  122  have a diameter d 3  that is approximately equal 0.243 inches. Accordingly, each opening  120  is between a pair of circumferentially adjacent openings  122 . 
   Outer secondary dilution openings  124  each have a diameter d 4  that is smaller than that of openings  120  and  122 , and are each located at a predetermined axial distance D 5  aft of openings  120  and  122 . In one embodiment, openings  124  have a diameter d 4 that is approximately equal  0.168 inches. More specifically, in the exemplary embodiment, openings  124  are approximately 0.25 passage heights h 1  downstream from openings  120  and  122 . In addition, each secondary dilution opening  124  is positioned downstream from, and between, a pair of circumferentially adjacent primary dilution openings  120  and  122 . 
   Inner liner  44  also includes a plurality of dilution jets  110  extending therethrough. More specifically, inner liner  44  includes a plurality of inner primary dilution openings  130  which each have a diameter d 6  that is smaller than a diameter d 2  and d 3  of respective outer primary dilution openings  120  and  122 . In one embodiment, openings  130  have a diameter d 6  that is approximately equal 0.228 inches. Each inner primary dilution opening  130  is circumferentially aligned with each outer secondary dilution opening  124  and between adjacent outer primary dilution openings  120  and  122 . More specifically, in the exemplary embodiment, inner primary dilution openings  130  are positioned downstream from dome plate  72  at a distance D 8  that is approximately equal 0.70 combustor passage heights h 1 . Accordingly, because primary dilution jets  120  and  122 , and  130  are not opposed, enhanced mixing and enhanced circumferential coverage is obtained between dilution jets  110  and mainstream combustor flow. Accordingly, the enhanced mixing facilitates reducing combustor exit temperature distortion and, thus reduces pattern factor. 
   A number of dilution jets  110  is variably selected to facilitate achieving a desired radial and circumferential exit temperature distribution from combustor  16 . More specifically, combustor  16  includes an equal number of outer primary dilution openings  120  and  122 , outer secondary dilution openings  124 , and inner primary dilution openings  130 . In the exemplary embodiment, combustor  16  includes eighteen larger diameter outer primary dilution openings  120 , eighteen smaller diameter outer primary dilution openings  122 , and thirty-six inner primary dilution openings  130 . More specifically, the number of outer primary dilution openings  120  and  122 , outer secondary dilution openings  124  is selected to be twice the number of fuel injectors  82  fueling combustor  16 . 
   Outer primary dilution openings  120  and  122 , and outer secondary dilution openings  124  receive air discharged through impingement openings or jets  140  formed within outer support  42 . Specifically, openings  140  are arranged in an array  144  that facilitates maximizing the cooling airflow available for impingement cooling of outer liner  40 . Within array  144 , openings  140  extend circumferentially around outer support  42 , but do not extend into pre-designated interruption areas  146  defined across outer support  42 . More specifically, each interruption area  146  is formed radially outward from outer primary dilution openings  120  and  122 , and outer secondary dilution openings  124  to facilitate avoiding variable interaction between impingement and dilution jets  140  and  110 , respectively, either by entrainment or by ejector effect. 
   Similarly, inner primary dilution openings  130  receive air discharged through impingement jets or openings  140  formed within inner support  46 . Specifically, opening array  144  facilitates maximizing the cooling airflow available for impingement cooling of inner liner  44 . Within array  144 , openings  140  extend circumferentially across inner support  46 , but do not extend into pre-designated interruption areas  150  defined across support  46 . More specifically, each interruption area  150  is formed radially outward from inner primary dilution openings  130  to facilitate avoiding variable interaction between impingement and dilution jets  140  and  110 , respectively, either by entrainment or by ejector effect. 
   Impingement jets  140  also supply airflow to multi-hole film cooling openings  160  formed within outer and inner liners  40  and  44 , respectively. More specifically, openings  160  are oriented to discharge cooling air therethrough for film cooling liners  40  and  44 . Accordingly, the number of impingement jets  140  is selected to facilitate maximizing the amount of cooling airflow supplied to liners  40  and  44 . In the exemplary embodiment, the number of impingement jets  140  is a multiple of the number of dilution jets  110 . More specifically, the number of impingement jets  140  and dilution jets  110  are selected to ensure that the pressure differential across impingement holes  140  in outer and inner supports  42  and  46 , respectively, approximately matches the pressure differential across the film cooling openings  160  and across dilution openings  120 ,  122 ,  124 , and  130 . 
   During operation, impingement cooling air is directed through impingement jets  140  towards outer and inner liners  40  and  44 , respectively, for impingement cooling of liners  40  and  44 . The cooling air is also channeled through dilution jets  110  and through film cooling openings  160  into combustion chamber  54 . More specifically, airflow discharged from openings  160  facilitates film cooling of liners  40  and  44  such that an operating temperature of each is reduced. Airflow entering chamber  54  through jets  110  facilitates radially and circumferentially cooling a temperature of the combustor flow path such that a desired exit temperature distribution is obtained. As such, the reduced combustor operating temperatures facilitate extending a useful life of combustor  16  and the desired exit temperature distribution facilitates extending a useful life to turbine hardware downstream of combustor  16 . 
   The above-described dilution and impingement jets provide a cost-effective and reliable means for operating a combustor. More specifically, each support includes a plurality of impingement jets that channel cooling air radially inward for impingement cooling of the combustor outer and inner liners. The outer and inner liners each include a plurality of dilution jets and film cooling openings which channel air inward into the combustion chamber. As a result, at least some of the impingement cooling air film cools the liners, and the remaining impingement cooling air is directed inward to facilitate radially and circumferentially cooling the combustor flow path such that a desired exit temperature distribution is obtained. 
   An exemplary embodiment of a combustion system is described above in detail. The combustion system components illustrated are not limited to the specific embodiments described herein, but rather, components of each combustion system may be utilized independently and separately from other components described herein. For example, the impingement jets and/or dilution jets may also be used in combination with other engine combustion systems. 
   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.