Abstract:
A method enables a turbine nozzle for a gas turbine engine to be assembled. The method includes providing a turbine nozzle including a plurality of airfoil vanes extending between an inner band and an outer band, wherein the outer band includes at least one mounting system that extends radially outwardly therefrom and includes a rail and at least one hook, and coupling the turbine nozzle into the gas turbine engine using the mounting system such that the turbine nozzle is at least partially supported by at least one hook. The method also includes positioning a seal assembly between at least one hook and the outer band to facilitate reducing radial leakage through the turbine nozzle.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine nozzles and more particularly, to methods and apparatus for assembling gas turbine engine nozzles. 
     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 nozzles arranged circumferentially and configured as doublets. At least some known turbine nozzles include more than two circumferentially-spaced hollow airfoil vanes coupled by integrally-formed inner and outer band platforms. Specifically, the inner band forms a radially inner flowpath boundary and the outer band forms a radially outer flowpath boundary. Additionally, at least some known outer bands include a forward and an aft hook assembly that are used to couple the turbine nozzle within the engine. 
     Forming the turbine nozzle with a plurality of integrally-formed airfoil vanes facilitates improving durability and reducing leakage in comparison to turbine nozzles which include only one airfoil vane. However, when cooling air is channeled to the turbine nozzle, leakage may still occur between circumferentially-adjacent turbine nozzles, which are spaced apart by a gap or interface that facilitates engine assembly, and accommodates thermal expansion between the turbine nozzles. Accordingly, at least some known turbine nozzles include a seal assembly that is positioned radially outwardly from the aft hook assembly to facilitate minimizing leakage through the interface. Over time, thermal cycling may cause degradation of the seals. However, accessing such interface seals may be difficult due to the location of the nozzle. 
     BRIEF SUMMARY OF THE INVENTION 
     In one aspect, a method for assembling a turbine nozzle for a gas turbine engine is provided. The method comprises providing a turbine nozzle including a plurality of airfoil vanes extending between an inner band and an outer band, wherein the outer band includes at least one mounting system that extends radially outwardly therefrom and includes a rail and at least one hook, and coupling the turbine nozzle into the gas turbine engine using the mounting system such that the turbine nozzle is at least partially supported by the at least one hook. The method also comprises positioning a seal assembly between at least one hook and the outer band to facilitate reducing radial leakage through the turbine nozzle. 
     In another aspect of the invention, a turbine nozzle for a gas turbine engine is provided. The nozzle includes an outer band, an inner band, at least one airfoil vane that extends between the inner and outer bands, and a seal assembly. The outer band includes an inside face, an outside face, and an aft hook assembly that extends outwardly from the outside face. The aft hook assembly includes a rail and at least one hook extending outwardly from the rail. The seal assembly is positioned adjacent the outer band hook assembly and is radially inward from at least one hook. 
     In a further aspect, a gas turbine engine is provided. The engine includes at least one turbine nozzle assembly comprising a seal assembly, an outer band, an inner band, and a plurality of airfoil vanes coupled together by the outer and inner bands. The outer band includes a hook assembly that extends radially outwardly from the outer band, and includes a rail and at least one hook extending outwardly from the rail. The seal assembly is positioned radially inwardly from the at least one hook. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic illustration of a gas turbine engine; 
         FIG. 2  is a perspective view of a turbine nozzle that may be used with the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a side perspective view of the turbine nozzle shown in  FIG. 2 ; and 
         FIG. 4  is a partial cross-sectional view of the engine shown in  FIG. 1  including a partial view of the turbine nozzle shown in FIGS.  2  and  3 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a schematic illustration of a gas turbine engine  10  including, in serial flow arrangement, a fan assembly  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 . Engine  10  has an intake side  28  and an exhaust side  30 . In one embodiment, engine  10  is a CF-34 engine commercially available from General Electric Aircraft Engines, Cincinnati, Ohio. 
     In operation, air flows through fan assembly  12  and compressed air is supplied to high-pressure compressor  14 . The highly compressed air is delivered to combustor  16  through a turbine nozzle assembly  32 . Airflow from combustor  16  drives turbines  18  and  20 , and turbine  20  drives fan assembly  12 . Turbine  18  drives high-pressure compressor  14 . 
       FIG. 2  is a perspective view of a turbine nozzle sector  50  that may be used with gas turbine engine  10  (shown in FIG.  1 ).  FIG. 3  is a side perspective view of turbine nozzle sector  50 .  FIG. 4  is a partial cross-sectional view of engine  10  including a partial view of turbine nozzle sector  50 . Nozzle  50  includes a plurality of circumferentially-spaced airfoil vanes  52  coupled together by an arcuate radially outer band or platform  54  and an arcuate radially inner band or platform  56 . More specifically, in the exemplary embodiment, each band  54  and  56  is integrally-formed with airfoil vanes  52 , and nozzle sector  50  includes four airfoil vanes  52 . In one embodiment, each arcuate nozzle sector  50  is known as a four vane segment. 
     Inner band  56  includes an aft flange  60  that extends radially inwardly therefrom. More specifically, flange  60  extends radially inwardly from band  56  with respect to a radially inner surface  62  of band  56 . Inner band  56  also includes a forward flange  64  that extends radially inwardly therefrom. Forward flange  64  is positioned between an upstream edge  66  of inner band  56  and aft flange  60 , and extends radially inwardly from band  56 . 
     Outer band  54  includes a cantilever mounting system  70  that includes a forward retainer  72 , a mid hook assembly  74 , and an aft hook assembly  76 . Cantilever mounting system  70  facilitates supporting turbine nozzle  50  within engine  10  from a surrounding annular engine casing (not shown). Forward retainer  72  extends radially outwardly from an outer surface  80  of outer band  54  and defines a channel  82  that extends continuously in a circumferential direction across a leading edge  84  of outer band  54 . Outer band  54  also includes a trailing edge  86  that is coupled to leading edge  84  by a pair of oppositely-disposed sector ends  87 . 
     Mid hook assembly  74  is positioned aft of forward retainer  72  and in the exemplary embodiment, includes a plurality of circumferentially-spaced and circumferentially-aligned hooks  90  that each extend upstream from a forward rail  92 . Forward rail  92  extends radially outwardly from outer band outer surface  80 . Hook assembly  74  extends in a circumferential direction across outer band outer surface  80  between circumferential ends  87 . 
     Aft hook assembly  76  is positioned aft of mid hook assembly  74 , and as such is between nozzle trailing edge  86  and mid hook assembly  74 . Hook assembly  76  includes an aft rail  94  and a plurality of hooks  96 . Rail  94  extends radially outwardly from outer band outer surface  80  in a circumferential direction across outer band outer surface  80  and between circumferential ends  87 . 
     Hooks  96  do not extend continuously between circumferential ends  87 , but rather hooks  96  are scalloped such that adjacent hooks  96  are spaced a distance  102  apart. Accordingly, a scalloped recessed area  104  is defined between each set of adjacent hooks  96 . Specifically, each recessed area  104  is radially aligned and radially outwardly from a respective airfoil vane  52 . As such, each hook  96  is radially aligned between adjacent vanes  52 . Accordingly, in the exemplary embodiment, nozzle  50  includes four scalloped recessed areas  104 . 
     Airfoil vanes  52  are substantially similar and each includes a first sidewall  110  and a second sidewall  112 . First sidewall  110  is convex and defines a suction side of each airfoil  52 , and second sidewall  112  is concave and defines a pressure side of each airfoil vane  52 . Sidewalls  110  and  112  are joined at a leading edge  114  and at an axially-spaced trailing edge  116  of each airfoil vane  52 , such that a cavity  118  is defined therebetween. Scalloped recessed areas  104  facilitate access to cavities defined within vanes  52 . In one embodiment, inserts (not shown) are inserted within each cavity  118 , and recessed areas  104  facilitate their installation and removal. More specifically, each airfoil trailing edge  116  is spaced chordwise and downstream from each respective airfoil leading edge  114 . First and second sidewalls  110  and  112 , respectively, also extend longitudinally, or radially outwardly, in span from radially inner band  56  to radially outer band  54 . 
     In the exemplary embodiment, each arcuate nozzle portion  50  includes a pair of circumferentially inner airfoil vanes  120  and  122 , and a pair of circumferentially outer airfoil vanes  124  and  126 . Vanes  120 ,  122 ,  124 , and  126  are also oriented substantially parallel to each other. Separation distance  102  and an orientation of vanes  52  are each variably selected to facilitate creating a highly divergent flowpath through nozzle  50 , and to facilitate optimizing aerodynamic accelerating flow through nozzle sector  50 . 
     Engine  10  includes a rotor assembly  140 , such as low pressure turbine  20 , that includes at least one row of rotor blades  142  that is downstream from turbine nozzles  50 . Rotor assembly  140  is surrounded by a rotor shroud  144  that extends circumferentially around rotor assembly  140  and turbine nozzles  50 . Cantilever mounting system  70  couples each turbine nozzle  50  to rotor shroud  144  through a hanger  148  that is supported by and coupled to shroud  144 . More specifically, each hook  96  is slidably coupled within a radially outer channel  150  defined within hanger  148 . 
     Hanger  148  also includes a radially inner channel  152  defined therein. Radially inner channel  152  is radially inward from radially outer channel  150  and each channel  150  and  152  is defined inwardly from a downstream side  154  of hanger  148 . Accordingly, each channel  150  and  152  is adjacent aft hook assembly  76 . Furthermore, when aft hook assembly  76  is coupled to hanger  148 , a cavity  160  is defined between hanger  148 , aft hook assembly  76 , and outer band  54 . 
     A seal assembly  170  is positioned radially inwardly from hooks  96  and extends within hanger radially inner channel  152 . More specifically, seal assembly  170  includes a seal member  172  that extends in sealing contact between hanger  148  and aft hook assembly rail  94 . In one embodiment, seal member  172  extends substantially circumferentially through engine  10  to facilitate minimizing radial leakage past aft hook assembly  76 , as described in more detail below. In an alternative embodiment, seal members  170  are segmented spline seals. In the exemplary embodiment, seal member  172  is a w-seal. 
     During operation, as hot combustion gases flow through nozzle  50 , cooling air is extracted from a high pressure source, such as compressor  14 , and directed at a high pressure into cavity  160 . More specifically, high pressure cooling air circulates through cavity  160  and facilitates cooling outer band  54  and turbine nozzle vanes  52 . Combustion gases flowing through turbine nozzle  50  create an area of low pressure conducive for leakage of high pressure cooling air between hanger  148  and aft hook assembly  76 . However, the relative high pressure of the cooling air causes seal member  172  to expand to facilitate preventing leakage between hanger  148  and aft hook assembly  76 . Moreover, because seal member  172  is radially inward from hooks  96  and is thus closer to the flow path, seal assembly  170  facilitates enhanced sealing in comparison to other known turbine nozzles. The combination of the enhanced sealing and the position of seal member  170  with respect to the flow path, enables aft hook assembly rail  94  to be fabricated with a radial height  190  that is shorter than other known aft rails. In addition, because aft hook assembly  76  is also scalloped, an overall weight of turbine nozzle  50  is reduced in comparison to other known turbine nozzles that do not include recessed areas  104 . As a result, mechanical stresses and thermal stresses induced within nozzle  50  are facilitated to be reduced. 
     The above-described turbine nozzle includes a scalloped aft hook assembly that extends from the aft rail. The hook assembly includes a plurality of recessed areas that are circumferentially spaced across the outer band. The recessed areas not only reduce an overall weight of the turbine nozzle assembly, but also facilitate reducing thermal stresses induced to the turbine nozzle. In addition, the turbine nozzle includes a seal assembly that is positioned radially inwardly from the aft hook assembly. Accordingly, the seal assembly is closer to the flowpath than other known seal assemblies, which facilitates enhanced sealing in comparison to other known seal assemblies, and also permits weight and stress reduction measures, including scalloping and/or removing portions of the radial flange above the sealing surface. As a result, the durability and useful life of the turbine nozzle are facilitated to be increased by the combination of the scalloped hook assembly and the seal assembly. 
     Exemplary embodiments of turbine nozzles are described above in detail. The nozzles are not limited to the specific embodiments described herein, but rather, components of each turbine nozzle may be utilized independently and separately from other components described herein. 
     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.