Abstract:
A method enables a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween to be assembled. The method comprises coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner, and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance

Description:
BACKGROUND OF THE INVENTION  
         [0001]    This application relates generally to gas turbine engines and, more particularly, to combustors for gas turbine engines.  
           [0002]    At least some known gas turbine engines include annular combustors which facilitate reducing nitrogen oxide emissions during gas turbine engine operation. Because of the heat generated within such combustors during operation, at least some known multiple annular combustors include a plurality of multiple dome assemblies that are radially aligned between the combustor dome plate and the combustion chamber. Each dome assembly includes a heat shield to protect the dome plate form excessive heat generated during engine operation.  
           [0003]    At least some known dome assembly heat shields include annular endbodies that extend an axial distance downstream from the heat shield to separate the domes or stages of the combustor to enable primary dilution air to be directed into a pilot stage reaction zone, thus facilitating combustion stability of the pilot stage of combustion at various operating points. However, because the endbodies extend axially towards the combustion chamber, the endbodies are exposed to a high temperature and high acoustic energy environment. Over time, the combination of the high temperatures and high acoustic energy may induce thermal stresses, low cycle fatigue (LCF), and/or high cycle fatigue (HCF) into the heat shield assembly. Continued operation with such stresses may lead to cracking within the heat shield which may shorten the useful life of the combustor.  
           [0004]    To facilitate reducing the effects of exposure to the high temperature and high acoustic energy environment, at least some known heat shield assemblies have employed various design changes to facilitate improving heat shield durability by addressing thermal and LCF failures. Such improvements have included for example, increased impingement cooling flow, surface film cooling, material changes, and/or heat shield contour changes to attempt to stiffen the component. However, such improvements did not completely address HCF failures caused by combustor acoustics. More specifically, due to engine-to-engine operating variation, and manufacturing/assembly tolerances, despite the improvements, at least some known heat shield natural frequencies remain within the combustor acoustic operating range, and over time, may still experience failures due to HCF fatigue.  
         BRIEF SUMMARY OF THE INVENTION  
         [0005]    In one aspect, a method for assembling a gas turbine engine multi-domed combustor including an outer liner and an inner liner that define a combustion chamber therebetween is provided. The method comprises coupling a first dome including a heat shield that includes an annular endbody that extends a first distance axially from the heat shield to the combustor outer liner, and coupling a second dome including a heat shield that includes an annular endbody that extends a second distance axially from the heat shield to the first dome, such that the second dome is radially aligned with respect to the first dome, and wherein the second dome second distance is less than the first dome first distance.  
           [0006]    In another aspect of the invention, an annular combustor for a gas turbine engine is provided. The combustor includes an outer liner, an inner liner, a first dome, and a second dome. The inner liner is spaced radially inwardly from the outer liner to define a combustion chamber therebetween. The first dome includes an outer end coupled to the outer liner and a heat shield including an annular endbody that extends outwardly a first distance axially from the heat shield towards the combustion chamber. The second dome is spaced radially inwardly from, and radially aligned with respect to the first dome. The second dome includes an outer end coupled to an inner end of the first dome, and a heat shield including at least one annular endbody that extends outwardly a second distance from the second dome heat shield. The second distance is less than the first dome first distance.  
           [0007]    In a further aspect, a gas turbine engine including a combustor having a natural combustor acoustic operating range is provided. The combustor includes an outer liner, an inner liner, and a plurality of radially-aligned domes. The outer liner is coupled to the inner liner to define a combustion chamber therebetween. The plurality of domes include at least a first dome and a second dome. The first dome includes a heat shield including an annular endbody that extends a first axial distance from the first dome heat shield. The second dome is radially inward from the first dome and includes a heat shield including an annular endbody extending a second axial distance from the first dome heat shield. The second axial distance is less than the first dome first distance. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0008]    [0008]FIG. 1 is a schematic illustration of a gas turbine engine;  
         [0009]    [0009]FIG. 2 is a cross-sectional view of a combustor that may be used with the gas turbine engine shown in FIG. 1; and  
         [0010]    [0010]FIG. 3 is an enlarged cross-sectional view of a portion of the combustor shown in FIG. 2. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0011]    [0011]FIG. 1 is a schematic illustration of a 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 . Combustor  16  is a lean premix combustor. Compressor  12  and turbine  20  are coupled by a first shaft  21 , and compressor  14  and turbine  18  are coupled by a second shaft  22 . A load (not shown) may also be coupled to gas turbine engine  10  with first shaft  21 . In one embodiment, gas turbine engine  10  is an LM6000 available from General Electric Aircraft Engines, Cincinnati, Ohio.  
         [0012]    In operation, air flows through low pressure compressor  12  and compressed air is supplied from low pressure compressor  12  to high pressure compressor  14 . The highly compressed air is delivered to combustor  30 . Airflow from combustor  16  drives turbines  18  and  20  and exits gas turbine engine  10  through a nozzle  24 .  
         [0013]    [0013]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 cross-sectional view of a portion of combustor  30 . Because a fuel/air mixture supplied to combustor  30  contains more air than is required to fully combust the fuel, and because the air is mixed with the fuel prior to combustion, combustor  30  is a lean premix combustor. Accordingly, a fuel/air mixture equivalence ratio for combustor  30  is less than one. Furthermore, because a gas and a liquid fuel are supplied to combustor  30 , and because combustor  30  does not include water injection, combustor  30  is a dual fuel dry low emissions combustor.  
         [0014]    Combustor  30  includes an annular outer liner  40 , an annular inner liner  42 , and a domed end or dome plate  44  extending between outer and inner liners  40  and  42 , respectively. Outer liner  40  and inner liner  42  are spaced radially inward from a combustor casing  45  and define a combustion chamber  46 . Combustor casing  45  is generally annular and extends downstream from a diffuser  48 . Combustion chamber  46  is generally annular in shape and is disposed radially inward from liners  40  and  42 . Outer liner  40  and combustor casing  45  define an outer passageway  52  and inner liner  42  and combustor casing  45  define an inner passageway  54 . Outer and inner liners  40  and  42  extend to a turbine nozzle  55  disposed downstream from diffuser  48 .  
         [0015]    Combustor domed end  44  includes a plurality of domes  56 . In the exemplary embodiment, domes  56  are arranged in a triple annular configuration. Alternatively, combustor domed end  44  includes a double annular configuration. An outer dome  58  includes an outer end  60  fixedly attached to combustor outer liner  40  and an inner end  62  fixedly attached to a middle dome  64 . Middle dome  64  includes an outer end  66  attached to outer dome inner end  62  and an inner end  68  attached to an inner dome  70 . Accordingly, middle dome  64  is between outer and inner domes  58  and  70 , respectively. Inner dome  70  includes an outer end  72  attached to middle dome inner end  68  and an inner end  74  fixedly attached to combustor inner liner  42 .  
         [0016]    Each dome  56  includes a plurality of premixer cups  80  to permit uniform mixing of fuel and air therein and to channel the fuel/air mixture into combustion chamber  46 . In one embodiment, premixer cups  80  are available from Parker Hannilfin, 6035 Parkland Blvd., Cleveland, Ohio. Combustor domed end  44  also includes an outer dome heat shield  100 , a middle dome heat shield  102 , and an inner dome heat shield  104  to insulate each respective dome  58 ,  64 , and  70  from heat generated within combustion chamber  46 . Heat shields  100 ,  102 , and  104  are radially aligned within engine  10 .  
         [0017]    Outer dome heat shield  100  includes an annular endbody  106  to insulate combustor outer liner  40  from flames burning in an outer primary combustion zone  108 . Endbody  106  extends outwardly an axial distance  110  from a downstream side  112  of heat shield  100  towards combustion chamber  46 . Distance  110  is commonly known as a heat shield wing length. In one embodiment, distance  110  is approximately equal 1.95 inches. In the exemplary embodiment, endbody  106  extends substantially perpendicularly from heat shield  100 .  
         [0018]    Middle dome heat shield  102  includes annular heat shield centerbodies  120  and  122  to segregate middle dome  64  from outer and inner domes  58  and  70 , respectively. Middle dome heat shield centerbodies  120  and  122  are positioned radially outwardly from a middle primary combustion zone  114 , and each extends outwardly an axial distance  126  and  128 , respectively, from a downstream side  130  of heat shield  102  towards combustion chamber  46 . In the exemplary embodiment, endbodies  120  and  122  each extend substantially perpendicularly from heat shield  102 , and as such are substantially parallel outer dome heat shield endbody  106 .  
         [0019]    Middle dome heat shield distance  126  is approximately equal distance  128 . Endbody distances  126  and  128  are shorter than outer dome heat shield endbody length  110 . More specifically, middle dome endbody distances  126  and  128  are at least 0.5 inches shorter than outer dome heat shield endbody length  110 . In the exemplary embodiment, middle dome endbody distances are each equal approximately 1.25 inches.  
         [0020]    Inner dome heat shield  104  includes an annular endbody  140  to insulate combustor inner liner  42  from flames burning in an inner primary combustion zone  142 . Endbody  140  extends outwardly an axial distance  144  from a downstream side  146  of heat shield  100  towards combustion chamber  46 . Endbody distance  144  is approximately equal outer dome heat shield distance  110 . In one embodiment, endbody distance  144  is approximately equal 1.95 inches. In the exemplary embodiment, endbody  106  extends substantially perpendicularly from heat shield  100 .  
         [0021]    During operation of gas turbine engine  10 , as combustor  30  uses radial fuel flow staging to facilitate reducing NOx and CO emissions over the engine operating range, combustor  30  has a natural acoustic operating range. Middle dome heat shield endbodies  120  and  122  facilitate providing additional structural support to middle dome  56 . Specifically, because heat shield endbodies  120  and  122  have a shorted winglength  126  and  128  than outer dome and inner dome endbodies  106  and  140 , respectively, middle dome endbodies  120  and  122  facilitate increasing a stiffness of middle dome heat shield  102  such that the natural frequency of middle dome heat shield  102  is increased above that of the combustor natural acoustic operating range, without adversely impacting engine operability. More specifically, the shortened winglength  126  and  128  does not adversely impact NOx and/or CO emissions, but does facilitate reducing vibrational stresses that may be induced to middle dome  56 . As such, middle dome endbodies  120  and facilitate extending a useful life of combustor  30 .  
         [0022]    The above-described combustor system for a gas turbine engine is cost-effective and reliable. The combustor system includes a combustor including a heat shield that includes at least one endbody that has a shortend winglength in comparison to the other heatshield endbodies. The shortened winglength facilitates reducing vibrational stresses that may be induced to the dome assembly by increasing the natural frequency of the endbody above that of the combustor acoustic operating range, but without adversely affecting engine operability. As a result, the endbody facilitates extending a useful life of the combustor in cost effective and reliable manner.  
         [0023]    Exemplary embodiments of combustor assemblies are described above in detail. The systems are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each combustor assembly component can also be used in combination with other combustor assembly components  
         [0024]    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.