Abstract:
A method facilitates assembling a rotor assembly for gas turbine engine. The method comprises providing a first rotor blade that includes an airfoil, a platform, a shank and a dovetail, coupling the first rotor blade to a rotor shaft using the dovetail, and coupling a second rotor blade to the rotor shaft such that a shank cavity is defined between the first and second blades. The method also comprises inserting a seal pin into the horizontal platform seal pin slot such that a gap defined between the first and second rotor blade platforms are substantially sealed wherein the seal pin includes a first end, a second end and a substantially cylindrical body extending therebetween and sized to frictionally engage the slot, wherein at least one of the first and second ends has a cross-sectional area that is smaller than a cross-sectional area of the body.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for assembling gas turbine engine rotor assemblies.  
         [0002]     At least some known rotor assemblies include at least one row of circumferentially-spaced rotor blades, which are known as buckets in some applications. 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, and is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. At least some known blades are hollow and include an internal cooling cavity that is defined at least partially by the airfoil, platform, shank, and dovetail.  
         [0003]     During operation, a clearance between circumferentially-adjacent blades with a row of blades, may cause a platform seal pin positioned between each blade to bind during initial engine operations and/or during transient operations. Such binding may cause the platform seal pin to deform, may induce cracking within the platform, and/or may cause the seal between the shank area of the blade and the hot gas path to become ineffective. An increase in the sealing effectiveness may increase the life of the blade by facilitating minimizing thermal stresses. Accordingly, within at least some known gas turbine engines, cylindrical pins, machined to mate with a corresponding notch formed in the end cover plates of the blade have been used to facilitate reducing binding of the pins. However, such pins have also been shown to bind in operation.  
       BRIEF SUMMARY OF THE INVENTION  
       [0004]     In one embodiment, a method for assembling a rotor assembly for gas turbine engine is provided. The method comprises providing a first rotor blade that includes an airfoil, a platform, a shank that extends radially inward from the platform and includes a horizontal platform seal pin slot and a dovetail that extends radially inward from the shank, coupling the first rotor blade to a rotor shaft using the dovetail, and coupling a second rotor blade to the rotor shaft such that a shank cavity is defined between the first and second blades. The method also comprises inserting a seal pin into the horizontal platform seal pin slot such that a gap defined between the first and second rotor blade platforms are substantially sealed wherein the seal pin includes a first end, a second end and a substantially cylindrical body extending therebetween and sized to frictionally engage the slot, wherein at least one of the first and second ends has a cross-sectional area that is smaller than a cross-sectional area of the body.  
         [0005]     In another embodiment, a gas turbine engine rotor assembly is provided. The rotor assembly includes a rotor shaft, a first blade, a second blade, and a seal pin. The first blade is coupled to the rotor shaft, and includes a first platform and a first shank extending radially inward from the platform. The first shank includes at least one sidewall including a seal pin slot. The second blade includes a second platform and a second shank extending radially inward from the second platform. The second blade is coupled to the rotor shaft adjacent the first blade such that a gap is defined between the first and second platforms, and such that a shank cavity is defined between the first and second shanks. The seal pin is inserted within the seal pin slot, and includes a first end, a second end, and a substantially cylindrical body extending therebetween. At least one of the first end and the second end has a cross-sectional area that is smaller than the body first cross-sectional area.  
         [0006]     In a further embodiment, a rotor blade seal pin for a gas turbine engine rotor assembly including a rotor shaft and a plurality of circumferentially-spaced rotor blades coupled to the rotor shaft is provided. Each rotor blade includes a platform and a shank, wherein the shank extends radially inward from the platform. The rotor blade seal pin comprises a first end and a second end, and a substantially cylindrical body having a first cross-sectional area sized for frictional engagement with a rotor blade seal pin slot formed adjacent to the platform. At least one of the first end and the second end has a second cross-sectional area that is smaller than the body first cross-sectional area. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0007]      FIG. 1  is schematic illustration of a gas turbine engine;  
         [0008]      FIG. 2  is a schematic view of a downstream side of an exemplary rotor disk that may be used with the gas turbine engine shown in  FIG. 1 ;  
         [0009]      FIG. 3  is an enlarged perspective view of a rotor blade shown in  FIG. 1  and viewed from a first side of the rotor blade;  
         [0010]      FIG. 4  is an enlarged side schematic view of an exemplary horizontal platform seal pin that may be used with the rotor blade shown in  FIG. 3 ; and  
         [0011]      FIG. 5  is an enlarged view of an end of the seal pin shown in  FIG. 4 . 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0012]      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  coupled via a single rotor or shaft  18 . In an alternative embodiment, shaft  18  is segmented into a plurality of shaft segments (not shown), 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 7FA+e gas turbine engine commercially available from General Electric Company, Greenville, S.C.  
         [0013]     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 .  
         [0014]      FIG. 2  is a schematic view of a downstream side of an exemplary rotor disk  36  that may be used with gas turbine engine  10  (shown in  FIG. 1 ). Rotor disk  36  includes a plurality of blade slots  38  defined therein and sized to receive a blade  40 , as illustrated in two of the plurality of blade slots  38  shown in  FIG. 2 . In the exemplary embodiment, adjacent blades  40  are substantially identical and each extends radially outward from rotor disk  36  and includes an airfoil  42 , a platform  44 , a shank  46 , and a dovetail  48 . In the exemplary embodiment, airfoil  42 , platform  44 , shank  46 , and dovetail  48  are collectively known as a bucket.  
         [0015]     Airfoil  42  extends radially inward from platform  44 , and shank  46  extends radially inward from platform  44 . Shank  46  includes a trailing edge radial seal pin slot  50  that extends generally radially through shank  46  between platform  44  and dovetail  48 . More specifically, in the exemplary embodiment, trailing edge radial seal pin slot  50  is defined within a downstream sidewall  52  of shank  46  and is adjacent a convex sidewall  54  of shank  46 .  
         [0016]     Shank seal pin slot  50  is sized to receive a radial seal pin  56  to facilitate sealing between adjacent rotor blade shanks  46  when adjacent rotor blades  40  are coupled within rotor disk  36 . A horizontal platform seal pin  58  is positioned within a horizontal platform seal pin slot (not shown in  FIG. 2 ) to facilitate sealing shank  46  from hot combustion gases  28 .  
         [0017]      FIG. 3  is an enlarged perspective view of rotor blade  40  viewed from a first side  44  of rotor blade  40 . In one embodiment, blade  40  is a newly cast blade  40 . In an alternative embodiment, blade  40  is a blade  40  that has been retrofitted to include the features described herein.  
         [0018]     When coupled within rotor assembly  10 , each rotor blade  40  is coupled to rotor disk  36  and as such, 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).  
         [0019]     Each airfoil  42  includes a first sidewall  70  and a second sidewall  72 . First sidewall  70  is convex and defines a suction side of airfoil  42 , and second sidewall  72  is concave and defines a pressure side of airfoil  42 . Sidewalls  70  and  72  are joined together at a leading edge  74  and at an axially-spaced trailing edge  76  of airfoil  42 . More specifically, airfoil trailing edge  76  is spaced chord-wise and downstream from airfoil leading edge  74 .  
         [0020]     First and second sidewalls  70  and  72 , respectively, extend longitudinally or radially outward in span from a blade root  78  positioned adjacent platform  44 , to an airfoil tip (not shown). The airfoil tip defines a radially outer boundary of an internal cooling chamber (not shown) that is defined within blades  40 . More specifically, the internal cooling chamber is bounded within airfoil  42  between sidewalls  70  and  72 , and extends through platform  44  and through shank  46  and at least partially into dovetail  48 .  
         [0021]     Platform  44  extends between airfoil  42  and shank  46  such that each airfoil  42  extends radially outward from each respective platform  44 . Shank  46  extends radially inwardly from platform  44  to dovetail  48 , and dovetail  48  extends radially inwardly from shank  46  to facilitate securing rotor blades  40  to rotor disk  36 . Platform  44  also includes an upstream side or skirt  90  and a downstream side or skirt  92  which are connected together with a pressure-side edge (not shown) and an opposite suction-side edge  96 . When rotor blades  40  are coupled within the rotor assembly, a gap  97  is defined between adjacent rotor blade platforms  44 , and accordingly is known as a platform gap.  
         [0022]     Shank  46  includes a substantially concave sidewall (not shown) and a substantially convex sidewall  54  connected together at an upstream sidewall  124  and a downstream sidewall  126  of shank  46 . Accordingly, the shank concave sidewall 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  98  is defined between adjacent rotor blade shanks  46 .  
         [0023]     In the exemplary embodiment, a forward angel wing  130  and an aft angel wing  132  each extend outwardly from respective shank sides  124  and  126  to facilitate sealing forward and aft angel wing buffer cavities (not shown) defined within the rotor assembly. In addition, a forward lower angel wing  134  also extends outwardly from shank side  124  to facilitate sealing between buckets  40  and the rotor disk. More specifically, forward lower angel wing  134  extends outwardly from shank  46  between dovetail  48  and forward angel wing  130 .  
         [0024]     In the exemplary embodiment, a portion  184  of platform  44  is chamfered or tapered along platform suction-side edge  96 . In an alternative embodiment, platform  44  does not include chamfered portion  184 . More specifically, chamfered portion  184  extends across a platform radially outer surface  186  adjacent to platform downstream skirt  92 .  
         [0025]     In the exemplary embodiment, shank  46  includes a leading edge radial seal pin slot  200  and a trailing edge radial seal pin slot  50 . In an alternative embodiment, shank  46  may include only one, or neither, of slots  200  and  50 . Specifically, each seal pin slot  200  and  50  extends generally radially through shank  46  between platform  44  and dovetail  48 . More specifically, leading edge radial seal pin slot  200  is defined within shank upstream sidewall  124  adjacent shank convex sidewall  54 , and trailing edge radial seal pin slot  50  is defined within shank downstream sidewall  126  adjacent shank convex sidewall  54 .  
         [0026]     Each shank seal pin slot  200  and  50  is sized to receive a radial seal pin  56  therein to facilitate sealing between adjacent rotor blade shanks  46  when rotor blades  40  are coupled within rotor assembly  10 . Although leading edge radial seal pin slot  200  is sized to receive a radial seal pin  56  therein, in the exemplary embodiment, when rotor blades  40  are coupled within the rotor assembly, a seal pin  56  is only positioned within trailing edge seal pin slot  50 , and slot  200  remains empty.  
         [0027]     Shank  46  also includes a horizontal platform seal pin slot  202  that extends generally axially through shank  46  between shank sides  124  and  126 . More specifically, horizontal platform seal pin slot  202  is defined between shank convex sidewall  54  and platform  44  and is substantially parallel to axis  30 . Horizontal platform seal pin slot  202  is sized to receive a horizontal platform seal pin  58  therein to facilitate sealing a low pressure side of shank  46  from combustion gases  28 . Horizontal platform seal pin slot  202  is defined by a pair of opposed radially-spaced sidewalls  210  and  212 , and extends generally axially between shank sides  124  and  126 . In the exemplary embodiment, sidewalls  210  and  212  are substantially parallel.  
         [0028]      FIG. 4  is an enlarged side schematic view of an exemplary horizontal platform seal pin  58  that may be used with gas turbine engine  10  (shown in  FIG. 1 ).  FIG. 5  is an enlarged view of a first end  400  of pin  58 . Horizontal platform seal pin  58  includes end  400 , a second end  402 , and a substantially cylindrical body  404  extending therebetween. Body  404  has an outer peripheral surface  405  and is generally symmetric about a longitudinal axis  406 .  
         [0029]     First end  400  includes a first end face  408  and second end  402  includes a second end face  410 . In the exemplary embodiment, each end face  408  and  410  is substantially planar and extends obliquely with respect to longitudinal axis  406 . In alternative embodiments, at least one of end face  408  and/or  410  is formed substantially perpendicularly to longitudinal axis  406 . In another alternative embodiment, at least one of end face  408  and/or  410  is formed non-planarly. In the exemplary embodiment, a first flat  412  extends from first end face  408  generally axially toward second end  402  a first distance  414 , such that a substantially planar face is formed by face  408 . In an alternative embodiment, a second flat  418 , having a substantially planar face, is formed such that the faces of flats  418  and  412  are substantially parallel. Second flat  418  extends from first end face  408  axially toward second end  402  a second distance  420 .  
         [0030]     In the exemplary embodiment, a third flat  422  extends from second end face  410  axially toward first end  400  a third distance  424  forming a substantially planar face. In an alternative embodiment, a fourth flat  426 , having a substantially planar face, is formed such that the faces of flats  422  and of flat  426  are substantially parallel. Fourth flat  426  extends from second end face  410  axially toward first end  400  a fourth distance  428 .  
         [0031]     In the exemplary embodiment, a portion of body  404  milled to form flats  412 ,  418 ,  422 , and  426  is approximately 20 mils. In alternative embodiments, other dimensions may be selected. Flats  412 ,  418 ,  422 , and  426  are formed and function similarly, and as such, only flat  412  is described below. Referring to  FIG. 5 , in the exemplary embodiment, each flat  412  includes a radius portion  430  and an adjacent chamfer portion  432 . Radius portion  430  is formed by a diameter of the mill tool used to form flat  412 , and a chamfer portion  432  is formed to substantially eliminate sharp edges that may result from the milling and/or other machining processes. Radius portion  430  and chamfer portion  432 , together form a generally tapered surface extending between flat  412  and an outer peripheral surface  405  of body  404 .  
         [0032]     During assembly of turbine  14 , a horizontal platform seal pin  58  is inserted generally axially into horizontal platform seal pin slot  202  to facilitate sealing a path for combustion gas flow between platforms  92  of each pair of adjacent blades  40  and the shank cavity. During transient operation and engine startup procedures, operating conditions in the path of combustion gases  28  may change relatively rapidly, for example, a temperature of combustion gases may increase or decrease. Such temperature changes cause a temperature gradient across components of blades  40  and rotor disk  36 , which causes the components to expand or contract, generally at differing rates than adjacent mating components due to material differences. Expansion or contraction of the components may cause a relative motion between adjacent components, such as for example, blade platforms  92 . Horizontal platform seal pin  58  may also move relative to horizontal platform seal pin slot  202  during these temperature transients. During such movement outer peripheral surface  405  slides in frictional engagement with sidewalls  210  and  212 . If during the sliding process, horizontal platform seal pin  58  binds in horizontal platform seal pin slot  202 , for example, by an edge of horizontal platform seal pin  58  engaging sidewalls  210  and  212  such that the edge digs in or gouges sidewalls  210  and  212 , which prevents horizontal platform seal pin  58  from sliding within horizontal platform seal pin slot  202 . In such case, horizontal platform seal pin  58  may deform, additional stress may be applied to horizontal platform seal pin slot  202  such that cracks are initiated in the vicinity of horizontal platform seal pin slots  202 . In accordance with one embodiment of the present invention, the ability of horizontal platform seal pin  58  to engage sidewalls  210  and  212  in a non-slidable manner is facilitated being reduced by removing portions of body  404  to form flats  412 ,  418 ,  422 , and  426  and forming an incline surface between outer peripheral surface  405  and flats  412 ,  418 ,  422 , and  426 .  
         [0033]     The above-described platform seal pin provides a cost-effective and highly reliable method for sealing a gap between adjacent blade platforms and the shank cavity. More specifically, thermal and mechanical stresses induced within the platform, and the operating temperature of the platform is facilitated to be reduced. Accordingly, platform cracking is also facilitated to be reduced. As a result, the rotor blade horizontal seal pin 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.  
         [0034]     Exemplary embodiments of rotor blade seal pins and rotor assemblies are described above in detail. The rotor blade seal pins are not limited to the specific embodiments described herein, but rather, features of each rotor blade seal pin may be utilized independently and separately from other components described herein. For example, each rotor blade seal pin feature 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 and rotor configurations.  
         [0035]     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.