Patent Publication Number: US-6984112-B2

Title: Methods and apparatus for cooling gas turbine rotor blades

Description:
BACKGROUND OF THE INVENTION 
   This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies. 
   At least some known rotor assemblies include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform. Each rotor blade also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. Known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail. 
   During operation, because the airfoil portions of the blades are exposed to higher temperatures than the dovetail portions, temperature mismatches may develop at the interface between the airfoil and the platform, and/or between the shank and the platform. Over time, such temperature differences and thermal strain may induce large compressive thermal stresses to the blade platform. In addition, if the blade platform generally is fabricated with a greater stiffness than the airfoil, such thermal strains may also induce thermal deformations to the airfoil, as the airfoil is displaced in response to the stresses induced to the shank and platform. Moreover, over time, the increased operating temperature of the platform may cause platform oxidation, platform cracking, and/or platform creep deflection, which may shorten the useful life of the rotor blade. 
   To facilitate reducing the effects of the high temperatures, within at least some known rotor blades, at least one of the pressure side and/or suction sides of the platform is formed with a recessed slot which facilitates channeling airflow from a shank cavity defined between adjacent rotor blades for use in cooling the platform trailing edge of an adjacent circumferentially-spaced rotor blade. Although such slots do facilitate reducing an operating temperature of an adjacent rotor blade platform trailing edge, such slots may induce stresses into the rotor blade in which they are formed. 
   BRIEF DESCRIPTION OF THE INVENTION 
   In one aspect, a method for fabricating a rotor blade for a gas turbine engine is provided. The method comprises providing a rotor blade that includes an airfoil, a platform, a shank, and a dovetail, wherein the shank extends between the platform and the dovetail, and wherein the platform extends between the airfoil and the shank, wherein the platform includes a leading edge side and a trailing edge side connected together by a pair of opposing sidewalls. The method also comprises forming an undercut in a portion of the platform to facilitate cooling the trailing edge side of the platform during operation, and forming a purge slot in a portion of the platform to facilitate channeling downstream towards the platform trailing edge side. 
   In another aspect, a rotor blade for a gas turbine is provided. The rotor blade includes a platform, an airfoil, a shank, and a dovetail. The platform includes a radially outer surface and a radially inner surface. The platform radially inner surface includes an undercut and a purge slot formed therein. The purge slot is for channeling cooling air downstream therefrom. The undercut facilitates cooling a portion of the platform during engine operation. The airfoil extends radially from the platform radially outer surface. The shank extends radially from the platform radially inner surface, and the dovetail extends from the shank for coupling the rotor blade within the gas turbine engine. 
   In a further aspect, a rotor assembly for a gas turbine engine is provided. The rotor assembly includes a rotor shaft and a plurality of circumferentially-spaced rotor blades that are coupled to the rotor shaft. Each of the rotor blades includes an airfoil, a platform, a shank, and a dovetail. The airfoil extends radially outward from the platform, and the platform includes a radially outer surface and a radially inner surface. The shank extends radially inward from the platform, and the dovetail extends from the shank for coupling each rotor blade to the rotor shaft. At least a first of the rotor blades includes an undercut and a purge slot defined within a portion of the first rotor blade platform. The undercut facilitates cooling the platform, and the purge slot facilitates channeling air downstream past the shank. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic illustration of an exemplary gas turbine engine; 
       FIG. 2  is a perspective view of an exemplary rotor blade that may be used with the gas turbine engine shown in  FIG. 1 ; 
       FIG. 3  is a perspective view of the rotor blade shown in FIG.  2  and viewed from an opposite end of the rotor blade; 
       FIG. 4  is a side view of a portion of the rotor blade shown in  FIG. 3 ; and 
       FIG. 5  is a cross-sectional view of a portion of the rotor blade shown in  FIG. 4  taken along line  5 — 5 . 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a schematic illustration of an exemplary gas turbine engine  10  coupled to an electric generator  16 . In the exemplary embodiment, gas turbine system  10  includes a compressor  12 , a turbine  14 , and generator  16  arranged in a single monolithic rotor or shaft  18 . In an alternative embodiment, shaft  18  is segmented into a plurality of shaft segments, wherein each shaft segment is coupled to an adjacent shaft segment to form shaft  18 . Compressor  12  supplies compressed air to a combustor  20  wherein the air is mixed with fuel supplied via a stream  22 . In one embodiment, engine  10  is a 6FA+e gas turbine engine commercially available from General Electric Company, Greenville, S.C. 
   In operation, air flows through compressor  12  and compressed air is supplied to combustor  20 . Combustion gases  28  from combustor  20  propels turbines  14 . Turbine  14  rotates shaft  18 , compressor  12 , and electric generator  16  about a longitudinal axis  30 . 
     FIGS. 2 and 3  are each perspective views of an exemplary rotor blade  40  that may be used with gas turbine engine  10  (shown in FIG.  1 ). And viewed from an opposite sides of blade  40 .  FIG. 4  is a side view of a portion of rotor blade  40 , and  FIG. 5  is a cross-sectional view of a portion of rotor blade  40  taken along line  5 — 5 . When blades  40  are coupled within a rotor assembly, such as turbine  14  (shown in FIG.  1 ), each rotor blade  40  is coupled to a rotor disk (not shown) that is rotatably coupled to a rotor shaft, such as shaft  18  (shown in FIG.  1 ). In an alternative embodiment, blades  40  are mounted within a rotor spool (not shown). In the exemplary embodiment, blades  40  are identical and each extends radially outward from the rotor disk and includes an airfoil  60 , a platform  62 , a shank  64 , and a dovetail  66 . In the exemplary embodiment, airfoil  60 , platform  62 , shank  64 , and dovetail  66  are collectively known as a bucket. 
   Each airfoil  60  includes first sidewall  70  and a second sidewall  72 . First sidewall  70  is convex and defines a suction side of airfoil  60 , and second sidewall  72  is concave and defines a pressure side of airfoil  60 . Sidewalls  70  and  72  are joined together at a leading edge  74  and at an axially-spaced trailing edge  76  of airfoil  60 . More specifically, airfoil trailing edge  76  is spaced chord-wise and downstream from airfoil leading edge  74 . 
   First and second sidewalls  70  and  72 , respectively, extend longitudinally or radially outward in span from a blade root  78  positioned adjacent platform  62 , to an airfoil tip  80 . Airfoil tip  80  defines a radially outer boundary of an internal cooling chamber (not shown) defined within blade  40 . More specifically, the internal cooling chamber is bounded within airfoil  60  between sidewalls  70  and  72 , and extends through platform  62  and through shank  64  and into dovetail  66 . 
   Platform  62  extends between airfoil  60  and shank  64  such that each airfoil  60  extends radially outward from each respective platform  62 . Shank  64  extends radially inwardly from platform  62  to dovetail  66 , and dovetail  66  extends radially inwardly from shank  64  to facilitate securing rotor blades  40  and  44  to the rotor disk. Platform  62  also includes an upstream side or skirt  90  and a downstream side or skirt  92  which are connected together with a pressure-side edge  94  and an opposite suction-side edge  96 . 
   Shank  64  includes a substantially concave sidewall  120  and a substantially convex sidewall  122  connected together at an upstream sidewall  124  and a downstream sidewall  126  of shank  64 . Accordingly, shank sidewall  120  is recessed with respect to upstream and downstream sidewalls  124  and  126 , respectively, such that when buckets  40  are coupled within the rotor assembly, a shank cavity  128  is defined between adjacent rotor blade shanks  64 . 
   In the exemplary embodiment, a forward angel wing  130  and an aft angel wing  132  each extend outwardly from respective shank sides  90  and  92  to facilitate sealing forward and aft angel wing buffer cavities (not shown) defined within the rotor assembly. In addition, a forward coverplate  134  also extends outwardly from respective shank sides  124  and  126  to facilitate sealing between buckets  40  and the rotor disk. More specifically, coverplate  134  extends outwardly from shank  64  between dovetail  66  and forward angel wing  130 . 
   In the exemplary embodiment, a platform undercut or trailing edge recessed portion  140  is defined within platform  62 . Specifically, platform undercut  140  is defined within platform  62  between a platform radially inner surface  142  and a platform radially outer surface  144 . More specifically, platform undercut  140  is defined within platform downstream skirt  92  at an interface  150  defined between platform pressure-side edge  94  and platform downstream skirt  92 . Accordingly, when adjacent rotor blades  40  are coupled within the rotor assembly, undercut  140  facilitates improving trailing edge cooling of platform  62  such that the low cycle fatigue life of blade  40  is improved. 
   Platform  62  also includes a recessed portion or purge slot  160 . More specifically, slot  160  is only defined within platform radially inner surface  142  along platform suction-side edge  96  between shank upstream and downstream sidewalls  124  and  126 . Moreover, a channel  166  is formed adjacent slot  160  for receiving a damper pin  168  therein when each rotor blade  40  is coupled within the rotor assembly. 
   Purge slot  160 , as described in more detail below, facilitates channeling cooling air from shank cavity  128  to facilitate increasing an amount of cooling air supplied to an undercut  140  formed on a circumferentially-adjacent rotor blade  40 . 
   An overall size, shape, and location of slot  160  with respect to blade  40  varies depending on flow requirements necessary to ensure adequate cooling flow to platform undercut  140 . A relative location of purge slot  160  is empirically determined relative to a datum W and to an aft surface  170  of downstream skirt  92 . More specifically, in the exemplary embodiment, purge slot  160  is a distance D 1  aft of a datum W and a distance D 2  upstream from skirt surface  170 . In the exemplary embodiment, distance D 1  is approximately 0.765 inches and distance D 2  is approximately 0.48 inches. 
   A relative size and shape of purge slot  160  is also empirically determined to facilitate optimizing cooling air flow to trailing edge undercut  140 . In the exemplary embodiment, purge slot  160  has a substantially elliptically-shaped cross-sectional area and is formed with a pre-determined radius of curvature R 1  such that purge slot  160  has a width W 1 . In an alternative embodiment, purge slot  160  has a non-elliptically shaped cross-sectional area. More specifically, in the exemplary embodiment, purge slot  52  radius of curvature R 1  is approximately equal to 0.145 inches, and purge slot width W 1  is approximately equal 0.265 inches. 
   Furthermore, purge slot  160  is formed with a depth D 3  measured with respect to platform side  94  that facilitates ensuring an adequate amount of cooling air is channeled past damper pin  168  when blade  40  is coupled within the rotor assembly. In the exemplary embodiment, depth D 3  is approximately equal to 0.169 inches. As is known in the art, damper pins  168  are inserted within channel  166  to facilitate coupling adjacent rotor blades  40  together. More specifically, when damper pin  168  is inserted within groove  166 , purge slot  160  is such that a flow gap  180  is defined between slot  160  and damper pin  168 . In one embodiment, gap  180  has a width W 5  that is at least approximately equal 0.051 inches wide to enable cooling air to enter purge slot  160  and be channeled around damper pin  168 . 
   During operation, wheel space cooling flow enters a first rotor blade shank cavity  128  and is channeled around damper pin  166  and discharged from purge slot  160  to facilitate increasing cooling flow to undercut  140  facilitates reducing an operating temperature of platform  62  and also reducing thermal stresses induced to blade  40 . In addition, the enhanced cooling also facilitates increasing the fatigue capability of blade  40 . 
   In addition, the combination of purge slot  160  and undercut  140  facilitates preventing crack initiation within platform  62  or between platform  62  and airfoil  60 . Accordingly, when adjacent rotor blades  40  are coupled within the rotor assembly, the combination of undercut  140  and purge slot  160  facilitates improving trailing edge cooling of platform  62  such that the low cycle fatigue life of blade  40  is improved. Moreover, because undercut  140  extends through the load path of blade  40 , mechanical stresses induced to platform downstream skirt  92  are also facilitated to be reduced, thus extending the useful life of rotor blade  40 . 
   The above-described rotor blades provide a cost-effective and highly reliable method for supplying cooling air to facilitate reducing an operating temperature of the rotor blade platform. More specifically, the purge slot facilitates ensuring an adequate flow of cooling air is channeled to the trailing edge platform undercut, such that the operating temperature of the platform is facilitated to be reduced. Accordingly, platform oxidation, platform cracking, and platform creep deflection is also facilitated to be reduced. As a result, the platform purge slot facilitates extending a useful life of the rotor assembly and improving the operating efficiency of the gas turbine engine in a cost-effective and reliable manner. 
   Exemplary embodiments of rotor blades and rotor assemblies are described above in detail. The rotor blades are not limited to the specific embodiments described herein, but rather, components of each rotor blade may be utilized independently and separately from other components described herein. For example, each rotor blade component can also be used in combination with other rotor blades, and is not limited to practice with only rotor blade  40  as described herein. Rather, the present invention can be implemented and utilized in connection with many other blade cooling configurations. 
   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.