Abstract:
A gas turbine engine compressor including a stator assembly and a method of assembling the same are provided. The method includes providing a compressor casing including at least two stator vane casing rails extending from the casing, coupling a rail liner within each respective casing rail, and coupling a stator vane assembly including two dovetails, and at least two stator vanes coupled together within the casing rails within the liner such that a first dovetail is received within a first casing rail and a first rail liner, and a second dovetail is received within a second casing rail and a second rail liner.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines, and more particularly, to methods and apparatus for assembling gas turbine engine compressors. 
   At least some known gas turbine engines include, in serial flow arrangement, a compressor, a combustor, a high pressure turbine, and a low pressure turbine. The compressor, combustor and high pressure turbine are sometimes collectively referred to as the core engine. Compressed air is channeled from the compressor to the combustor where it is mixed with fuel and ignited. The combustion gasses are channeled to the turbines which extract energy from the combustion gasses to power the compressors and to produce useful work to propel an aircraft in flight or to power a load, such as an electrical generator. 
   Known compressors include a rotor assembly and a stator assembly. Known rotor assemblies include a plurality of rows of circumferentially-spaced rotor blades that extend radially outward from a shaft or disk. Known stator assemblies may include a plurality of stator vanes which extend circumferentially between adjacent rows of rotor blades to form a nozzle for directing air passing therethrough towards downstream rotor blades. More specifically, known stator vanes extend radially inward from a compressor casing between adjacent rows of rotor blades. 
   In at least some compressors, each stator vane is unitarily formed with an airfoil and platform that are mounted through an integrally-formed dovetail to the compressor casing. To facilitate assembly of the stator vanes to the casing, a small amount of clearance is permitted between a casing dovetail or vane rail and the vane platform. However, the clearance enables a small degree of relative motion between the vane platform and the casing vane rail. Over time, continued movement between the stator vanes and the casing rail may cause vane platform and/or casing wear. Such relative movement of the stator vanes may be enhanced by vibrations generated during engine operation. 
   To facilitate reducing wear between the casing and vane platform, at least some stator assemblies are coated with wear coatings or lubricants. Other known compressors use casing rail liners, and/or vane springs to facilitate reducing such wear. However, known wear coatings may not be useful in some single vane applications, and known vane springs may not be suitable for use with vanes that include air bleed holes. Moreover, known rail liners are only useful in a limited number of engine designs. 
   BRIEF DESCRIPTION OF THE INVENTION 
   In one aspect, a method for assembling a gas turbine engine compressor is provided. The method includes providing a compressor casing including at least one stator vane casing rail extending from the casing, coupling a rail liner to the casing rail, and coupling a stator vane assembly including at least two stator vanes coupled together to the casing rail within the liner. 
   In another aspect, a stator vane assembly for a gas turbine engine is provided that includes a plurality of circumferentially-spaced stator vane doublets. Each doublet includes a pair of stator vanes coupled together at a respective outer stator vane platform of each vane. Each stator vane platform is configured to slidably couple each doublet to a vane rail extending from a compressor casing that extends at least partially circumferentially around the stator vane assembly. 
   In another aspect, a compressor for a gas turbine engine is provided. The compressor includes a casing including a plurality of stator vane rails. The casing defines an axial flow path for the compressor. A rotor is positioned within the flow path. The rotor includes a plurality of rows of circumferentially-spaced rotor blades. A stator vane assembly extends between adjacent rows of the plurality of rows of rotor blades. Each stator vane assembly includes a plurality of circumferentially-spaced stator vane doublets received within the vane rail. Each stator vane doublet includes a pair of stator vanes coupled together at a respective outer stator vane platform of each vane. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic illustration of a gas turbine engine; 
       FIG. 2  is a cross sectional view of a compressor suitable for use with the engine shown in  FIG. 1 ; 
       FIG. 3  is a perspective view of an exemplary stator vane doublet suitable for use in the compressor shown in  FIG. 2 ; and 
       FIG. 4  is a cross sectional view of the stator vane doublet shown in  FIG. 3  mounted in a compressor casing. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1  is a schematic illustration of a gas turbine engine  10  including a low pressure compressor  12 , a high pressure compressor  14 , and a combustor  16  that defines a combustion chamber (not shown). 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 rotor shaft  24 , and compressor  14  and turbine  18  are coupled by a second rotor shaft  26 . In one embodiment, engine  10  is a CF6 engine available from General Electric Aircraft Engines, Cincinnati, Ohio. 
   In operation, air flows through low pressure compressor  12  and compressed air is supplied from low pressure compressor  12  to high pressure compressor  14 . The highly compressed air is delivered to combustor  16 . Airflow from combustor  16  drives rotating turbines  18  and  20 . 
     FIG. 2  is a cross-sectional illustration of a portion of a compressor  30  that may be used with gas turbine engine  10 .  FIG. 3  illustrates an exemplary stator vane doublet  80 . In an exemplary embodiment, compressor  30  is a high pressure compressor. Compressor  30  includes a rotor assembly  32  and a stator assembly  34  that are positioned within a casing  36  that defines a flowpath  38 . The rotor assembly  32  defines an inner flowpath boundary  40  of the flowpath  38 . Stator assembly  34  defines an outer flowpath boundary  42  of flowpath  38 . Compressor  30  includes a plurality of stages with each stage including a row of circumferentially-spaced rotor blades  50  and a row of stator vane assemblies  52 . In an exemplary embodiment, rotor blades  50  are coupled to a rotor disk  54 . Specifically, each rotor blade  50  extends radially outwardly from rotor disk  54  and includes an airfoil  56  that extends radially from an inner blade platform  58  to a blade tip  60 . 
   Stator assembly  34  includes a plurality of rows of stator vane assemblies  52  with each row of vane assemblies  52  positioned between adjacent rows of rotor blades  50 . The compressor stages are configured for cooperating with a motive or working fluid, such as air, such that the motive fluid is compressed in succeeding stages. Each row of vane assemblies  52  includes a plurality of circumferentially-spaced stator vanes  66  that each extends radially inward from casing  36  and includes an airfoil  68  that extends from an outer vane platform  70  to a vane tip  72 . Airfoil  68  includes a leading edge  73  and a trailing edge  74 . In an exemplary embodiment, stator vanes  66  have no inner platform. Compressor  30  includes one stator vane row per stage, some of which are bleed stages  76 . 
   At bleed stages  76 , vane assembly  52  includes a plurality of circumferentially-spaced stator vane doublets  80 . As shown in  FIG. 3 , stator vane doublet  80  includes a pair of stator vanes  66  joined at abutting edges  82  of their respective outer stator vane platforms  70  to form a vane segment. The joined platforms  70  are configured to be received in a vane rail  88  formed in compressor casing  36  as will be described. The stator vane doublet  80  includes two airfoils  68  joined together through a brazing process and has a circumferential width W. In an exemplary embodiment, stator vanes  66  are joined by a gold-nickel braze material. Each stator vane platform  70  includes an inwardly facing surface  84  that defines a portion of outer flowpath boundary  42  in compressor  30 . At bleed stage  76 , stator vane doublet  80  includes a bleed hole  86  formed in the joined vane platforms  70  between airfoils  68 . Bleed holes  86  bleed off a portion of the motive fluid for use in cooling one or more stages of HP turbine  18 . 
     FIG. 4  illustrates a cross sectional view of stator vane doublet  80  mounted within casing  36 . Casing  36  includes casing vane rails  88  that each includes a vane platform engagement surface  90 . Stator vane platform  70  includes dovetails  92  that are received in casing vane rails  88 . In an exemplary embodiment, a vane rail liner  94  is mounted within casing vane rails  88  and stator vane doublets  80  are received within vane rail liner  94 . Vane rail liner  94  provides a sacrificial wear surface between casing vane rails  88  and stator vane platform dovetails  92 . 
   In operation, stator vane doublet  80  provides a vane segment that has a circumferential width W that is sufficiently large to substantially reduce a range of relative movement between stator vane platforms  70  of stator vanes  66  and casing vane rails  88 . The reduced allowable movement reduces an amount of wear experienced between casing vane rails  88  and stator vane platforms  70 . In an exemplary embodiment, vane rail liner  94  and stator vane doublet  80  cooperate to further reduce the range of relative movement between stator vane doublet  80  and casing vane rail  88 . Vibration from the coupled stator vane airfoils  68  partially cancel each other so that with stator vane doublet  80 , vibration transmitted to joined platforms  70  is reduced. 
   Stator vanes  66  are joined to form vane doublets  80 . In forming vane doublets  80 , at least a portion of abutting edges  82  of stator vane platforms  70  of stator vanes  66  is first nickel-plated. The stator vanes  66  are then mounted in a precision tack welding fixture (not shown) that has a curvature substantially corresponding to a curvature of casing vane rail  88  and tack welded. The tack welded stator vanes  66  are then placed in a carbon member (not shown) to hold the desired shape during the braze furnace cycle. The tack welded stator vanes  66  are then brazed along outer vane platforms  70  using a gold-nickel braze alloy to form stator vane doublet  80 . The gold-nickel braze provides ductility and temperature stability in the braze joint necessary for durability of the joint during engine operation. After brazing, the stator vane doublet  80  is re-aged in the carbon member to restore metallurgical properties. 
   Assembly of vane doublet  80  into compressor casing  36  is accomplished by mounting a casing vane rail liner  94  on casing vane rail  88  and mounting vane doublet  80  within vane rail liner  94 . The extended platform length of vane doublet  80  together with casing vane rail liner  88  take up excess clearance in casing vane rail  88  which facilitates reducing a vibration response of vane doublet  80  with respect to individual vanes  66 . 
   The above described compressor assembly provides a cost effective and reliable means for reducing stator vane platform to casing vane rail wear. More specifically, the compressor assembly employs stator vane doublets at the compressor bleed stages. The stator vane doublets provide vane segment that have a circumferential width that is sufficiently large to substantially reduce the amount of allowable movement between stator vane platforms and the casing vane rails. The reduced allowable movement reduces the amount of wear experienced between the casing vane rails and the stator vane platforms. A vane rail liner further reduces movement between the stator vane doublet and casing vane rail and provides a sacrificial surface which can be easily replaced. Vibration from the coupled stator vane airfoils also partially cancels each other so that with the stator vane doublet, vibration transmitted to the joined platforms is reduced. 
   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.