Abstract:
A method facilitates the assembly of a gas turbine engine. The method of assembly comprises providing a turbine nozzle including an inner band, an outer band, at least one vane extending between the inner and outer bands, and at least one leading edge fillet extending between the at least one vane and at least one of the inner and outer bands, wherein a leading edge of the at least one vane is downstream from the leading edges of the inner and outer bands, and coupling the turbine nozzle within the gas turbine engine such that the leading edge fillet is configured to facilitate minimizing vortex formation along the vane leading edge adjacent at least one of the inner and outer bands.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This invention relates generally to turbine engines and more particularly, to methods and apparatus for assembling gas turbine engines.  
         [0002]     Known gas turbine engines include combustors which ignite fuel-air mixtures which are then channeled through a turbine nozzle assembly towards a turbine. At least some known turbine nozzle assemblies include a plurality of arcuate nozzle segments arranged circumferentially. At least some known turbine nozzles include a plurality of circumferentially-spaced hollow airfoil vanes coupled by integrally-formed inner and outer band platforms. More specifically, the inner band forms a portion of the radially inner flowpath boundary and the outer band forms a portion of the radially outer flowpath boundary.  
         [0003]     Within known engine assemblies, an interface defined between the turbine nozzle and an aft end of the combustor is known as a fish-mouth seal. More specifically, within such engine assemblies, leading edges of the turbine nozzle outer and inner band platforms are generally axially aligned with respect to a leading edge of each airfoil vane extending therebetween. Accordingly, in such engine assemblies, when hot combustion gases discharged from the combustor approach the nozzle vane leading edge, a pressure or bow wave reflects from the vane leading edge stagnation and propagates a distance upstream from the nozzle assembly, causing circumferential pressure variations across the band leading edges and a non-uniform gas pressure distribution. The pressure variations may cause localized nozzle oxidation and/or localized high temperature gas injection, each of which may decrease engine efficiency. Moreover, such pressure variations may also cause the vane leading edge to operate at an increased temperature in comparison to the remainder of the vane.  
       BRIEF SUMMARY OF THE INVENTION  
       [0004]     In one aspect, a method for assembling a gas turbine engine is provided. The method comprises providing a turbine nozzle including an inner band, an outer band, at least one vane extending between the inner and outer bands, and at least one leading edge fillet extending between the at least one vane and at least one of the inner and outer bands, wherein a leading edge of the at least one vane is downstream from the leading edges of the inner and outer bands, and coupling the turbine nozzle within the gas turbine engine such that the leading edge fillet is configured to facilitate minimizing vortex formation along the vane leading edge adjacent at least one of the inner and outer bands.  
         [0005]     In another aspect, a turbine engine nozzle assembly is provided. The turbine engine nozzle assembly includes an outer band, an inner band, at least one vane, and a leading edge fillet. The outer and inner bands each include a leading edge, a trailing edge, and a body extending therebetween. The at least one vane extends between the outer and inner bands. The at least one vane includes a first sidewall and a second sidewall connected together at a leading edge and a trailing edge. The at least one vane leading edge is positioned downstream from the inner and outer band leading edges. The leading edge fillet extends between the at least one vane and at least one of the inner-band and the outer band. The leading edge fillet is configured to facilitate minimizing vortex formation along the vane leading edge adjacent at least one of the inner and outer bands.  
         [0006]     In a further aspect, a gas turbine engine is provided. The engine includes a combustor and a turbine nozzle assembly that is downstream from and in flow communication with the combustor. The nozzle assembly includes an outer band, an inner band, at least one vane extending between the outer and inner bands, and a leading edge fillet. The outer band and inner band each include a leading edge and each is coupled to an aft end of the combustor. The at least one vane includes a first sidewall and a second sidewall connected together at a leading edge and a trailing edge. The at least one vane leading edge is positioned downstream from the inner and outer band leading edges. The leading edge fillet extends between the at least one vane and at least one of the inner band and the outer band. The leading edge fillet is configured to facilitate minimizing vortex formation along the vane leading edge. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0007]      FIG. 1  is a schematic illustration of an exemplary gas turbine engine;  
         [0008]      FIG. 2  is a side view of an exemplary turbine nozzle that may be used with the gas turbine engine shown in  FIG. 1 ;  
         [0009]      FIG. 3  is a perspective view of the turbine nozzle shown in  FIG. 2 ;  
         [0010]      FIG. 4  is an enlarged side view of an exemplary retainer that may be used with the turbine nozzle shown in  FIGS. 2 and 3 ; and  
         [0011]      FIG. 5  is a side view of the turbine nozzle shown in  FIGS. 2 and 3  coupled to a combustor that may be used with the engine shown in  FIG. 1  with the retainer shown in  FIG. 4 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0012]      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 . 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 . In one embodiment, gas turbine engine  10  is an LM2500 engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio. In another embodiment, gas turbine engine  10  is a CFM engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio.  
         [0013]     In operation, air flows through low pressure compressor  12  supplying compressed air from low pressure compressor  12  to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Airflow from combustor  16  is channeled through a turbine nozzle (not shown in  FIG. 1 ) to drive turbines  18  and  20 , prior to exiting gas turbine engine  10  through an exhaust nozzle  24 .  
         [0014]      FIG. 2  is a side view of an exemplary turbine nozzle  50  that may be used with a gas turbine engine, such as turbine engine  10  (shown in  FIG. 1 ).  FIG. 3  is a perspective view of turbine nozzle  50 . In the exemplary embodiment, nozzle  50  is one segment of a plurality of segments that are positioned circumferentially to form a nozzle assembly (not shown) within the gas turbine engine. Nozzle  50  includes at least one airfoil vane  52  that extends between an arcuate radially outer band or platform  54 , and an arcuate radially inner band or platform  56 . More specifically, in the exemplary embodiment, outer band  54  and the inner band  56  are each integrally-formed with airfoil vane  52 .  
         [0015]     Vane  52  includes a pressure-side sidewall  60  and a suction-side sidewall  62  that are connected at a leading edge  64  and at an chordwise-spaced trailing edge  66  such that a cooling cavity  68  is defined between sidewalls  60  and  62 . Vane sidewalls  60  and  62  each extend radially between bands  54  and  56  and in the exemplary embodiment, sidewall  60  is generally concave, and sidewall  62  is generally convex.  
         [0016]     Outer and inner bands  54  and  56  each include a leading edge  70  and  72 , respectively, a trailing edge  74  and  76 , respectively, and a platform body  78  and  80 , respectively, extending therebetween. Airfoil vane(s)  52  are oriented such that outer and inner band leading edges  70  and  72 , respectively, are each a distance d upstream from airfoil vane leading edge  64 . Distance d is variably selected to ensure that leading edges  70  and  72  are upstream from vane leading edge  64 , and to facilitate bands  54  and  56  preventing hot gas injections along vane leading edge  64 , as described in more detail below.  
         [0017]     In the exemplary embodiment, inner band  56  includes an aft flange  90  that extends radially inwardly therefrom. More specifically, flange  90  extends radially inwardly from band  56  with respect to a radially inner surface  92  of band  56 . Inner band  56  also includes a forward flange  94  that extends radially inward therefrom. Forward flange  94  is positioned between inner band leading edge  72  and aft flange  90 , and extends radially inwardly from band  56 . In the exemplary embodiment, an upstream side  100  of forward flange  94  is substantially planar between a radially outermost surface  102  of flange  94  and radially inner surface  92 . Moreover, in the exemplary embodiment, a downstream side  106  of flange  94  includes a shoulder  108 , such that flange downstream side  106  is substantially planar from flange surface  102  to shoulder  108 , and from shoulder  108  to radially inner surface  92 .  
         [0018]     Inner band  56  also includes a plurality of circumferentially-spaced radial tabs  110  that extend radially inwardly therefrom. More specifically, in the exemplary embodiment, the number of radial tabs  110  is the same as the number of vanes  52 . In the exemplary embodiment, each tab  110  includes a substantially parallel upstream and downstream surfaces  120  and  122 , respectively. Radial tabs  110  are spaced a distance d 2  downstream from forward flange  94  such that a retention channel  130  is defined between each radial tab  110  and forward flange  94 .  
         [0019]     In the exemplary embodiment, outer band  54  includes an aft flange  140  that extends generally radially outwardly therefrom. More specifically, flange  140  extends radially outwardly from band  54  with respect to a radially outer surface  142  of band  54 . Outer band  54  also includes a forward flange  144  that extends radially outward therefrom. Forward flange  144  is positioned between outer band leading edge  70  and aft flange  140 , and extends radially inwardly from band  54 . In the exemplary embodiment, an upstream side  146  of forward flange  144  is substantially planar between a radially outermost surface  147  of flange  144  and radially outer surface  142 . Moreover, in the exemplary embodiment, a downstream side  148  of flange  144  includes a shoulder  150 , such that flange downstream side  148  is substantially planar from flange surface  147  to shoulder  150 , and from shoulder  150  to radially outer surface  142 .  
         [0020]     Outer band  54  also includes a plurality of circumferentially-spaced radial tabs  160  that extend radially outwardly therefrom. More specifically, in the exemplary embodiment, the number of radial tabs  160  is the same as the number of vanes  52 . In the exemplary embodiment, each tab  160  includes substantially parallel upstream and downstream surfaces  162  and  164 , respectively. Radial tabs  160  are spaced a distance d 3  downstream from forward flange  144  such that a retention channel  166  is defined between each radial tab  160  and forward flange  144 . In the exemplary embodiment, channels  166  are approximately the same size as channels  130 .  
         [0021]     Turbine nozzle  50  also includes a plurality of leading edge fillets  170 . Fillets  170  are generally larger than fillets used with known turbine nozzles and extend between outer platform  54  and vane  52  in a tip area  180  of each vane leading edge  64 , and between inner platform  56  and vane  52  in a hub area  182  of each vane leading edge  64 . Specifically, within tip area  180 , fillets  170  are blended from vane leading edge  64  across a radially inner surface  184  of outer platform  54  and towards outer band leading edge  70 . Moreover, within hub area  182 , fillets  170  are blended from vane leading edge  64  across a radially outer surface  186  of inner platform  56  and towards inner band leading edge  72 . Accordingly, nozzle vane leading edge  64  is enlarged within both hub area  182  and tip area  180  such that fillets  170  facilitate accelerating the flow passing thereby.  
         [0022]     In the exemplary embodiment, fillets  170  are formed with a plurality of cooling openings  190  that extend through fillets  170  and are configured to discharge cooling air inwardly into the boundary flow flowing over vane  52 . Specifically, each cooling opening  190  is oriented towards a pitch-line of vane  52  and such that openings  190  facilitate energizing the flow momentum in the boundary layer, such that the formation of horseshoe vortices upstream from leading edge  64  is facilitated to be reduced. The reduction in the formation of the horseshoe vortices facilities improving aerodynamic efficiency. Moreover, the plurality of cooling openings  190  also facilitate reducing surface heating and an operating temperature of vane  52 .  
         [0023]     During operation, the location of inner and outer bands  56  and  54 , respectively, with respect to vane leading edge  64  facilitates reducing hot gas injections along vane leading edge  64 . Rather, the combination of enlarged fillets  170  and cooling holes  190  facilitates accelerating the flow and energizing the flow momentum in the boundary layer, such that the formation of horseshoe vortices are facilitated to be reduced. As a result, aerodynamic efficiency is facilitated to be improved and the operating temperature of nozzle airfoil vane  52  is facilitated to be reduced. As such, a useful life of turbine nozzle  50  is facilitated to be extended.  
         [0024]      FIG. 4  is an enlarged side view of an exemplary retainer  200  that may be used with turbine nozzle  50  (shown in  FIGS. 2 and 3 ). In the exemplary embodiment, retainer  200  is known as a spring clip and is configured to facilitate coupling nozzle  50  to an aft end of combustor  16  in a sealing arrangement as described in more detail below. Retainer  200  includes a pair of opposite ends  202  and  204 , and a body  206  extending therebetween. In the exemplary embodiment, body  206  includes an insertion portion  210  and a retention portion  212  that extends integrally from insertion portion  210 .  
         [0025]     Insertion portion  210  is generally U-shaped and extends from end  204  to insertion portion  210 , and retention portion  212  extends from insertion portion  210  to end  204 . Accordingly, insertion portion  210  includes a pair of opposed legs  214  and  216  that are connected by an arcuate portion  218 . In the exemplary embodiment, portion  218  is substantially semi-circular. Arcuate portion  218  has a radius r that is sized to enable legs  214  and  216  to define a width w of retainer  200 , measured with respect to an outer surface  220  and  222  of legs  214  and  216 , respectively, that is narrower than a width, i.e., distance d 2 , of channel  166  or channel  130 . Accordingly, insertion portion  210  is sized for insertion within retention channels  166  and  130 .  
         [0026]     Retention portion  212  includes a first leg  230  that extends obliquely outward from leg  216  to an apex  232  and a second leg  233  that extends obliquely from apex  232  towards leg  214 . As such, a tip  236  of apex  232  is a distance d T  from leg outer surface  222 .  
         [0027]     In the exemplary embodiment, retainer  200  is fabricated from a resilient material that resists deformation. In an alternative embodiment, retainer  200  is fabricated from a shape memory material. In a further alternative embodiment, retainer  200  is fabricated from any material that enables retainer  200  to function as described herein.  
         [0028]      FIG. 5  is a side view of turbine nozzle  50  coupled to combustor  16  using retainer  200 . Combustor  16  includes a combustion zone  240  that is formed by annular, radially inner and radially outer supporting members  244  and  246 , respectively, and combustor liners  250 . Combustor liners  250  shield the outer and inner supporting members from heat generated within combustion zone  240 . More specifically, combustor  16  includes an annular inner liner  256  and an annular outer liner  258 . Liners  256  and  258  define combustion zone  240  such that combustion zone  240  extends from a dome assembly (not shown) downstream to turbine nozzle  50 . Outer and inner liners  258  and  256  each include a plurality of separate panels  260  which include a series of steps  262 , each of which form a distinct portion of combustor liners  250 .  
         [0029]     Each liner  256  and  258  also includes an annular support flange, or aft flange,  270  and  272 , respectively. Specifically, each support flange  270  and  272  couples an aft end  274  and  276  of each respective liner  256  and  258  to supporting members  244  and  246 . More specifically, the coupling of each support flange  270  and  272  to each supporting member  244  and  246  forms an annular gap or fishmouth opening  278 .  
         [0030]     Each support flange  270  and  272  includes a radial portion  280  and a conical datum area  282 . Each radial portion  280  is formed with a plurality of preferential cooling openings or jets  284  that extend therethrough to facilitate discharging cooling air towards nozzle  50 . Air discharged from jets  284  facilitates reducing the formation of horseshoe vortices upstream from vane leading edge  64  and thus facilitates improving aerodynamic efficiency of nozzle  50 . Each conical datum area  282  extends integrally outward and upstream from each radial portion  280  such that conical datum area  282  defines a radially inner portion  286  of each fishmouth opening  278 . A radial outer portion  288  of each fishmouth opening  278  is defined by each supporting member  244  or  246 . Fishmouth opening  278  is used to couple a pair of annular ring interfaces  290  and  291  between combustor  16  and nozzle  50 .  
         [0031]     In the exemplary embodiment, interfaces  290  and  291  are substantially similar and each has a substantially L-shaped cross-sectional profile and includes an upstream edge  292 , a downstream edge  294 , and a body  296  extending therebetween. Body  296  includes a radially inner surface  298  and an opposite radially outer surface  300 . In the exemplary embodiment, interface upstream edge  292  is securely coupled within fishmouth opening  278  and interface downstream edge  294  is inserted within retention channel  166  such that the portion of body inner surface  298  within channel  166  is positioned against the substantially planar portion of nozzle forward flange  144  extending between shoulder  150  and flange surface  147 . Similarly, along inner band  56 , the downstream edge  294  of interface  291  is inserted within retention channel  130  such that the portion of body inner surface  298  within channel  130  is positioned against the substantially planar portion of nozzle forward flange  94  extending between shoulder  108  and flange surface  102 .  
         [0032]     After interfaces  290  and  291  are positioned within channels  166  and  130 , respectively, a retainer  200  is inserted within each retention channel  166  and  130  such that leg outer surface  220  is positioned against a respective radial tab  160  and  110 . More specifically, when fully inserted within channels  166  and  130 , each retainer apex  232  is biased against, and in contact with, interfaces  290  and  291 . Specifically, each retainer  200  is positioned in contact against each interface radially outer surface  300  such that interface radially inner surface  298  is biased in sealing contact within each channel  130  and  166  against each respective nozzle forward flange  94  and  144 . In an alternative embodiment, retainers  200  are not used to couple interfaces  290  and  291  against flanges  94  and  144 , but rather other suitable means for securing interfaces  290  and/or  291  in sealing contact against flanges  94  and  144  may be used, such as, but not limited to, inserting fasteners through radial tabs  110  and/or  166 , or bending radial tabs  110  and  166  against flanges  94  and  144 .  
         [0033]     When the engine is fully assembled, interfaces  290  and  291  provide structural support to combustor  16  and facilitate sealing between combustor  16  and nozzles  50 . As such, a mechanically flexible seal arrangement is provided which provides structural stability and support to the aft end of combustor  16 . Moreover, the assembly of interface rings  290  and  291  between combustor  16  and nozzle  50  is generally less labor intensive and less time-consuming than the assembly of known seal interfaces used with other gas turbine engines.  
         [0034]     In each embodiment, the above-described turbine nozzles include an inner band and an outer band that each extend upstream a distance from the vane leading edge to facilitate reducing hot gas injection along the vane leading edge. Moreover, because each inner and outer band extends upstream from the vane leading edge, each band accommodates enlarged fillets in comparison to known turbine nozzles. The combination of the inner and outer bands, the impingement jets extending through the combustor support flanges, and the cooling openings extending through the fillets facilitates reducing an operating temperature of the nozzle vanes, reducing the formation of horseshoe vortices upstream from each vane leading edge, and improving the aerodynamic efficiency of the nozzle. Moreover, the interface rings extending between the combustor and the turbine nozzle provide structural support to the combustor while being biased in a sealing arrangement with the turbine nozzle. As a result, a useful life of the turbine nozzle is facilitated to be extended in a reliable and cost effective manner.  
         [0035]     Exemplary embodiments of turbine nozzles are described above in detail. The interface rings, fillets, and cooling openings and jets are not limited to use with the specific nozzle embodiments described herein, but rather, the such components can be utilized independently and separately from other turbine nozzle components described herein. Moreover, the invention is not limited to the embodiments of the nozzle assemblies described above in detail. Rather, other variations of nozzles assembly embodiments may be utilized within the spirit and scope of the claims.  
         [0036]     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.