Abstract:
A method for assembling a turbine nozzle for a gas turbine engine facilitates improving cooling efficiency of the turbine nozzle. The method includes providing a hollow doublet including a leading airfoil and a trailing airfoil coupled by at least one platform, wherein each airfoil includes a first sidewall and a second sidewall that extend between a respective leading and trailing edge. The method also includes inserting an insert into at least one of the airfoils, wherein the insert includes a first sidewall including a first plurality of cooling openings that extending therethrough, and a second sidewall including a second plurality of cooling openings extending therethrough, and wherein the first plurality of cooling openings facilitate more cooling of the airfoil than the second plurality of cooling openings.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine nozzles and more particularly, to methods and apparatus for cooling 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 lease some known turbine nozzle assemblies include a plurality of nozzles arranged circumferentially and configured as doublets. A turbine nozzle doublet includes a pair of circumferentially-spaced hollow airfoil vanes coupled by integrally-formed inner and outer band platforms. 
     The doublet type turbine nozzles facilitate improving durability and reducing leakage in comparison to non-doublet turbine nozzles. Furthermore, turbine nozzle doublets also facilitate reducing manufacturing and assembly costs. In addition, because such turbine nozzles are subjected to high temperatures and may be subjected to high mechanical loads, at least some known doublets include an identical insert installed within each airfoil vane cavity to distribute cooling air supplied internally to each airfoil vane. The inserts include a plurality of openings extending through each side of the insert. 
     In a turbine nozzle, the temperature of the external gas is higher on the pressure-side than on the suction-side of each airfoil vane. Because the openings are arranged symmetrically between the opposite insert sides, the openings facilitate distributing the cooling air throughout the airfoil vane cavity to facilitate achieving approximately the same operating temperature on opposite sides of each airfoil. However, because of the construction of the doublet, mechanical loads and thermal stresses may still be induced unequally across the turbine nozzle. In particular, because of the orientation of the turbine nozzle with respect to the flowpath, typically the mechanical and thermal stresses induced to the trailing doublet airfoil vane are higher than those induced to the leading doublet airfoil vane. Over time, continued operation with an unequal distribution of stresses within the nozzle may shorten a useful life of the nozzle. 
     BRIEF SUMMARY OF THE INVENTION 
     In one aspect of the invention, a method for assembling a turbine nozzle for a gas turbine engine is provided. The method includes providing a hollow doublet including a leading airfoil vane and a trailing airfoil vane coupled by at least one platform, wherein each airfoil vane includes a first sidewall and a second sidewall that extend between a respective leading and trailing edge. The method also includes inserting an insert into at least one of the airfoil vanes, wherein the insert includes a first sidewall including a first plurality of cooling openings that extend therethrough, and a second sidewall including a second plurality of cooling openings extending therethrough. 
     In another aspect, a method of operating a gas turbine engine is provided. The method includes directing fluid flow through the engine using at least one turbine airfoil nozzle that includes a leading airfoil and a trailing airfoil coupled by at least one platform that is formed integrally with the leading and trailing airfoils, and wherein each respective airfoil includes a first sidewall and a second sidewall that extend between respective leading and trailing edges to define a cavity therein. The method also includes directing cooling air into the turbine airfoil nozzle such that the nozzle trailing airfoil is cooled more than the leading airfoil. 
     In a further aspect of the invention, a turbine nozzle for a gas turbine engine is provided. The nozzle includes a pair of identical airfoil vanes coupled by at least one platform formed integrally with the airfoil vanes. Each airfoil vane includes a first sidewall and a second sidewall that are connected at a leading edge and a trailing edge, such that a cavity is defined therebetween. The nozzle also includes at least one insert that is configured to be inserted within the airfoil vane cavity and includes a first sidewall and a second sidewall. The insert first sidewall includes a first plurality of openings extending therethrough for directing cooling air towards at least one of the airfoil vane first and second sidewalls. The insert second sidewall includes a second plurality of openings that extend therethrough for directing cooling air towards at least one of the airfoil vane first and second sidewalls. The first plurality of openings are configured to facilitate lower metal temperatures therefrom than the second plurality of openings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic illustration of a gas turbine engine; 
         FIG. 2  is an exploded perspective forward-looking-aft view of turbine nozzle that may be used with the gas turbine engine shown in  FIG. 1 ; and 
         FIG. 3  is an exploded perspective aft-looking-forward view of the turbine nozzle shown in FIG.  2 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  is a schematic illustration of a gas turbine engine  10  including 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, or upstream, side  28  and an exhaust, or downstream, side  30 . In one embodiment, engine  10  is a CF6-80 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 . Airflow from combustor  16  is discharged through a turbine nozzle assembly (not shown in  FIG. 1 ) that includes a plurality of nozzles (not shown in  FIG. 1 ) and used to drive turbines  18  and  20 . Turbine  20 , in turn, drives fan assembly  12 , and turbine  18  drives high-pressure compressor  14 . 
       FIG. 2  is an exploded perspective forward-looking-aft view of turbine nozzle  50  that may be used with gas turbine engine  10  (shown in FIG.  1 ).  FIG. 3  is an exploded perspective aft-looking-forward view of turbine nozzle  50 . Nozzle  50  is known as a doublet and includes a pair of circumferentially-spaced airfoil vanes  52  coupled together by an arcuate radially outer band or platform  56  and an arcuate radially inner band or platform  54 . More specifically, in the exemplary embodiment, each band  54  and  56  is formed integrally with airfoil vanes  52 . 
     Inner band  54  includes a retention flange  60  that extends radially inwardly therefrom. More specifically, flange  60  extends substantially perpendicularly from band  54  with respect to a radially outer surface  62  of flange  60 . Outer band  56  also includes a retention flange  64  that extends radially outwardly therefrom, and a leading edge flange  66  that also extends radially outwardly therefrom. More specifically, outer band retention flange  64  and leading edge flange  66  extend substantially perpendicularly from band  56  with respect to a radially inner surface  68  of band  56 . Surfaces  62  and  68  define a radially outer and radially inner boundary for a flowpath through nozzle  50 . 
     Airfoil vanes  52  are identical and include a leading airfoil vane  76  and a trailing airfoil vane  78 . Each airfoil vane  52  includes a first sidewall  80  and a second sidewall  82 . First sidewall  80  is convex and defines a suction side of each airfoil vane  76  and  78 , and second sidewall  82  is concave and defines a pressure side of each airfoil vane  76  and  78 . Sidewalls  80  and  82  are joined at a leading edge  84  and at an axially-spaced trailing edge  86  of each airfoil vane  76  and  78 . More specifically, each airfoil trailing edge  86  is spaced chordwise and downstream from each respective airfoil leading edge  84 . 
     First and second sidewalls  80  and  82 , respectively, extend longitudinally, or radially outwardly, in span from radially inner band  54  to radially outer band  56 . Additionally, first and second sidewalls  80  and  82 , respectively, define a cooling chamber  90  within each airfoil vane  52 . More specifically, chamber  90  is bounded by an inner surface  92  and  94  of each respective sidewall  80  and  82 , and extends through each band  54  and  56 . 
     Each cooling chamber  90  is sized to receive an insert  100  therein. More specifically, lead airfoil chamber  90  is sized to receive a lead insert  102 , and trailing airfoil chamber  90  is sized to receive a trailing insert  104  therein. Inserts  102  and  104  are substantially similar and each includes a respective key feature  110  and  112 , and an identical attachment flange  114 . Flange  114  extends from a radially outer end  116  of each insert  102  and  104 , and enables each insert  102  and  104  to be secured within each respective cooling chamber  90 . In one embodiment, flange  114  is brazed to radially outer band  56 . In another embodiment, flange  114  is welded to radially outer band  56 . 
     Key features  110  and  112  extend through flange  114  at each insert radially outer end  116 . Specifically, key features  110  and  112  are unique to each respective insert  102  and  104 , and are sized to be received in a mating slot (not shown) that extends through nozzle radially outer band  56 . More specifically, key features  110  and  112  prevent lead insert  102  from being inadvertently inserted within trailing airfoil vane  78 , and prevent trailing insert  104  from being inadvertently inserted within leading airfoil vane  76 . 
     Each insert  102  and  104  has a cross sectional profile that is substantially similar to that of a respective airfoil vane  76  and  78 . More specifically, each insert  102  and  104  includes a first sidewall  120  and  122 , respectively, and a second sidewall  124  and  126 . Accordingly, each insert first sidewall  120  and  122  is adjacent each respective airfoil vane first sidewall  80  when each insert  102  and  104  is installed within each respective cooling chamber  90 . Each insert first sidewall  120  and  122  is convex and defines a suction side of each respective insert  102  and  104 , and each insert second sidewall is concave and defines a pressure side of each respective insert  102  and  104 . Respective pairs of insert sidewalls  120  and  124 , and  122  and  126 , are joined at respective leading edges  128  and  130 , and at respective trailing edges  132  and  134 . 
     Lead insert first sidewall  120  defines a suction side of lead insert  102  and includes a first plurality of openings  140  that extend therethrough to a cavity  142  defined therein. Lead insert second sidewall  124  includes a second plurality of openings  144  that extend therethrough to cavity  142 . First and second sidewall openings  140  and  144  of insert  102  are biased to facilitate cooling a suction side  80  of lead airfoil vane  76 , more than a pressure side  82  of lead airfoil vane  76 . In the exemplary embodiment, the plurality of first sidewall openings  140  are greater than that required to achieve substantially equal surface temperatures when compared to the plurality of second sidewall openings  144 . The ratio of ninety first sidewall openings  140  to ninety-seven second sidewall openings  144  results in biased cooling and is in contrast to known inserts which have a ratio of seventy-six first sidewall openings to one hundred thirty-seven second sidewall openings which results in cooling all four airfoil sidewalls substantially equally. In an alterative embodiment, the larger volume of air is facilitated because insert first sidewall  120  includes openings  140  which are larger in diameter than corresponding openings  144  extending through insert second sidewall  124 . It should be noted that the arrangement of openings  140  and  144  with respect to each respective sidewall  120  and  124  is variable. Furthermore, the number and size of openings  140  and  144  is also variable. 
     Trailing insert first sidewall  122  defines a suction side of trailing insert  104  and includes a first plurality of openings  150  that extend therethrough to a cavity  152  defined therein. Trailing insert second sidewall  126  includes a second plurality of openings  154  that extend therethrough to cavity  152 . First sidewall openings  150  permit a larger volume of cooling air to pass therethrough than second sidewall openings  154 . More specifically, insert  104  is biased to facilitate cooling a suction side  80  of trailing airfoil vane  78 , more than a pressure side  82  of trailing airfoil vane  78 . In the exemplary embodiment, the larger volume of air is facilitated because the plurality of first sidewall openings  150  outnumber the plurality of second sidewall openings  154 . More specifically, in the exemplary embodiment, first sidewall  122  includes one hundred forty-two openings  150 , and second sidewall  126  includes ninety-seven openings  154 . In an alterative embodiment, the larger volume of air is facilitated because insert first sidewall  122  includes openings  150  which are larger in diameter than corresponding openings  154  extending through insert second sidewall  126 . It should be noted that the arrangement of openings  150  and  154  with respect to each respective sidewall  122  and  126  is variable. Furthermore, the number and size of openings  150  and  154  is also variable. 
     Each nozzle  50  is in flow communication with a cooling system (not shown) that directs cooling air into each airfoil vane cooling chamber  90  for internal cooling of nozzle airfoil vanes  52 . Specifically, the cooling system directs cooling air into each airfoil vane insert  100 , which in-turn, channels the cooling air for cooling airfoil vanes  52 . In addition to being biased to facilitate cooling a suction side of each respective airfoil vane  76  and  78 , nozzle inserts  100  are biased to facilitate cooling trailing airfoil vane  78  more than lead airfoil vane  76 . More specifically, trailing insert openings  150  and  154  are biased such that a larger volume cooling air is directed towards trailing airfoil vane  78  through trailing insert  104  than is directed through lead insert  102  towards lead airfoil vane  76 . In the exemplary embodiment, the larger volume of air is facilitated because the plurality of trailing airfoil vane first sidewall openings  150  outnumber the plurality of, lead airfoil vane first sidewall openings  140 . In an alternative embodiment, the larger volume of air is facilitated by varying the size of trailing airfoil vane openings  150  in comparison to lead airfoil vane openings  140 . 
     During operation, cooling air is routed through the cooling system into nozzle  50 , which may not be thermally loaded or mechanically stressed equally between adjacent airfoil vanes  76  and  78 . More specifically, due to gas loading, thermal variations, and mechanical loading, more mechanical and thermal stresses are induced and transmitted through trailing airfoil vane  78  than through lead airfoil vane  76 . Because nozzle inserts  102  and  104  provide nozzle  50  with a cooling scheme that may be customized to particular applications, cooling air supplied to nozzle  50  is allocated more to a suction side  80  of the airfoil vanes  52  than to a pressure side  82  of the airfoil vanes  52 . Accordingly, as cooling air is channeled into nozzle  50 , inserts  102  and  104  direct cooling air towards a respective nozzle airfoil vane  76  and  78 . The cooling air exits outwardly from each nozzle airfoil vane  52  through a plurality of airfoil trailing edge openings (not shown), and thermal stresses induced within each individual airfoil vane  76  and  78  are facilitated to be reduced. Furthermore, by biasing the cooling airflow to cool trailing airfoil vane  78  more than lead airfoil vane  76 , thermal stresses across nozzle  50  are facilitated to be controlled. As a result, although a maximum temperature on each airfoil vane concave surface is increased, the thermal stresses induced in nozzle  50  are facilitated to be controlled to counteract the mechanical stresses, thus facilitating increasing a useful life of nozzle  50 . 
     The above-described turbine nozzle includes a pair of inserts that enable a cooling scheme for the nozzle to be customized to particular applications. Specifically, the inserts bias the distribution of cooling air supplied to the nozzle more to the suction side of each of the airfoil vanes, and more to the trailing airfoil vane in the doublet. As a result, the inserts facilitate controlling thermal stresses induced within the nozzle, and thus, facilitate increasing the useful life of the nozzle in a cost-effective and reliable manner. 
     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.