Abstract:
A rotor blade installation tool for coupling a plurality of rotor blades to a rotor disc wherein each rotor blade extends from the rotor disc to a radially outer blade tip is provided. The tool includes a blade engagement end configured to engage the plurality of rotor blades between the rotor disc and the radially outer blade tip where said blade engagement end comprises an engagement top surface, at least one brace coupled to said blade engagement end at a first end of said at least one brace, and a guide end coupled to a second end of said at least one brace where said guide end comprises a body including a guide end top surface positioned above said engagement top surface.

Description:
BACKGROUND OF THE INVENTION 
   This invention relates generally to gas turbine engines, and more specifically to methods and apparatus for assembling gas turbine engine fan assemblies. 
   At least some known gas turbine engines include a fan for supplying air to a compressor that compresses incoming air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited within a combustion chamber for generating hot combustion gases. The hot combustion gases are channeled downstream to a turbine, which extracts energy from the combustion gases for powering the fan and compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. 
   Known compressors include a rotor assembly that includes 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 the platform, and is used to mount the rotor blade within the rotor assembly to a rotor disk or spool. More specifically, at least some known rotor disks include a plurality of circumferentially-spaced dovetail slots that are each sized to receive a respective one of the plurality of rotor blades therein. Known rotor blade dovetails are generally shaped complementary to the dovetail slot to enable the rotor blade dovetails and the rotor disk slot to mate together and form a dovetail assembly. Adapters may be used to facilitate the mating of the dovetails and the slots. 
   During an installation process, interlocking mid-span dampers extending between adjacent blades, may overlap rather than interlock, if the blades are not inserted substantially simultaneously into the dovetail slots. Know methods of inserting the blade into the dovetails include incremental insertion of each blade in turn until all blades are seated into the dovetail. If, during the installation process, mid-span dampers overlap, the installation process is stopped and the dampers are disengaged before the installation is resumed. If the mid-span dampers become overlapped such that they cannot be disengaged manually, each mid-span damper may need to be non-destructively tested. Because each rotor includes numerous blades and each blade may be handled numerous times during installation, the installation process may be time-consuming and laborious. Additionally, manufacturer requirements may require engines to be removed from an aircraft, or be at least partially disassembled to accommodate the installation process. 
   BRIEF DESCRIPTION OF THE INVENTION 
   In one aspect, a method for assembling a rotor assembly for a gas turbine engine is provided. The method includes providing a plurality of rotor blades that each include a dovetail, providing a rotor disc that includes a plurality of dovetail slots spaced circumferentially about the disc, partially inserting each rotor blade dovetail into a respective rotor dovetail slot, and seating the plurality of rotor blades in the respective rotor dovetail slot substantially simultaneously using an annular blade installation tool. 
   In another aspect, a rotor blade installation tool for installing a plurality of rotor blades onto a rotor disc is provided. The tool includes a blade engagement end, at least one brace coupled to the blade engagement end at a first end of the at least one brace, and a guide end coupled to a second end of the at least one brace. 

   
     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 gas turbine fan disc that may be used with a gas turbine engine, such as the turbine shown in  FIG. 1 ; 
       FIG. 3  is a schematic side view of an exemplary rotor fan blade that may be used with the fan assembly shown in  FIG. 1 ; 
       FIG. 4  is a plan view of an exemplary blade installation tool that may be used to facilitate installing a plurality of rotor blades shown in  FIG. 3 ; 
       FIG. 5  is a side elevation view of the blade installation tool shown in  FIG. 4  taken along line  4 - 4 , also shown in  FIG. 4 ; and 
       FIG. 6  is a perspective view of the blade insertion tool coupled to a gas turbine engine, such as the engine shown in  FIG. 1 . 
   

   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 TFE-731 engine commercially available from Honeywell Aerospace, Phoenix, Ariz. 
   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 directed to drive 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 an exemplary gas turbine fan disc  200  that may be used with a gas turbine engine, such as turbine  10  (shown in  FIG. 1 ). Disc  200  includes a hub  202  that includes a shaft opening  204  extending therethrough. Disc  200  also includes a plurality of circumferentially-spaced dovetail slots  206  that extend from a leading face  208  to a trailing face  210  of disc  200 . 
   In operation, shaft opening  204  is coupled to a shaft (not shown) of engine  10  such that disc  200  is driven through the shaft by compressor  20 . 
     FIG. 3  is an exploded schematic side view of an exemplary rotor fan blade  300  that may be used with fan assembly  12  (shown in  FIG. 1 ). When fully assembled, fan assembly  12  includes a plurality of blades  300  coupled to disc  200 . Blade  300  includes an airfoil  302  that extends between a blade tip  304  and a blade dovetail  306  that is configured to engage one of the plurality of dovetail slots  206  of disc  200 . In the exemplary embodiment, an adapter  308  may be used to facilitate mating of dovetail  306  and slot  206 . Airfoil  302  includes a leading edge  310 , a trailing edge  312 , and a pressure side  314  and a suction side  316  that each extends between leading edge  310  and trailing edge  312 . Suction side  316  includes a first mid-span damper  318  that extends outwardly from suction side  316  and is configured to interlock with a high-pressure side mid-span damper (not shown) coupled to a first adjacent fan rotor blade (not shown). Pressure side  314  includes a second mid-span damper (not shown) that extends outwardly from pressure side  314  and is configured to interlock with a suction-side mid-span damper (not shown) coupled to a second adjacent fan rotor blade (not shown). Each of pressure side  314  and suction side  316  include a platform  320  that extends from leading edge  310  and trailing edge  312  proximate dovetail  306 . 
   During installation, adapter  308  is inserted into slot  206  and dovetail  306  is slid into slot  206  sufficiently to hold adapter  308  in place. An adjacent blade is inserted into a slot adjacent to slot  206  in a similar manner. Each of the plurality of blades is inserted into a predetermined respective slot until all of the plurality of fan rotor blades are inserted into a respective slot just sufficiently to hold respective adapters  308  in place. 
     FIG. 4  is a plan view of an exemplary blade installation tool  400  that may be used to facilitate installing a plurality of rotor blades  300  (shown in  FIG. 3 ).  FIG. 5  is a side elevation view of tool  400  taken along line  4 - 4  (shown in  FIG. 4 ). Tool  400  includes a blade engagement end  402  that includes a central opening  404 . In the exemplary embodiment, end  402  includes a circularly-shaped body having a circularly-shaped opening therethrough. In alternative embodiments, other shaped bodies are contemplated such that engagement end  402  is configured to fulfill the requirements discussed below. Engagement end  402  also includes a pad  406  coupled to an engagement face  408  of engagement end  402 . In the exemplary embodiment, pad  406  is fabricated from a material that is softer than a material from which blade  300  is fabricated from. Pad  406  facilitates protecting blade  300  during an installation process. Additionally, pad  406  transmits an installation force from engagement face  408  to blades  300  during the installation process. Tool  400  includes at least one brace  410  coupled to engagement end  402  to support a guide end  412 . Guide end  412  includes a guide opening  414  therethrough. In the exemplary embodiment, a first end of brace  410  is welded to engagement end  402  such that brace  410  does not interfere with pad  406  and/or any of the plurality of blades  300  during the installation process. A second end of brace  410  is coupled to guide end  412  such that during the installation process engagement end  402  and guide end  412  are substantially co-axially aligned with longitudinal axis  415 . In the exemplary embodiment, four braces  410  are welded to engagement end  402  and guide end  412 . In an alternative embodiment, at least one brace  410  is hingedly coupled to engagement end  402  and guide end  412  such that during non-use engagement end  402  and guide end  412  may not be substantially co-axially aligned. In the exemplary embodiment, engagement end  402  includes a plurality of fastener holes for coupling pad  406  to engagement end  402  using fasteners  416  such as, but not limited to, rivets, nuts and bolts, and pins. In alternative embodiments, pad  406  may be coupled to engagement end  402  using non-fasteners, such as, but not limited to, adhesive, friction fit, and interference fit. In the exemplary embodiment, tool  400  includes at least one handle  418  coupled to brace  410  to facilitate applying manual force to tool  400 . Handle  418  includes a first end  420  coupled to brace  410  and a second opposite end  422  that may be configured for ergonomic manual grasping. Handle  418  may couple to brace  410  perpendicularly. Alternatively, handle  418  may be coupled to brace  410  at an angle that is predetermined to facilitate grasping and applying a force to tool  400 . 
     FIG. 6  is a perspective view of blade insertion tool  400  coupled to a gas turbine engine, such as engine  10  (shown in  FIG. 1 ). During installation in disc  200 , blades  300  are inserted partially into slots  206  as described above. A guide shaft  600  is inserted into a opening in the end of engine shaft  602 . Installation tool is installed onto shaft  600 , threading tool  400  over shaft  600 , engagement end first such that shaft  600  passes through opening  414 . Tool  400  is slid towards blades  300  until pad  406 , if installed, contacts blades  300 . In the exemplary embodiment, engagement end  402  is configured to engage each blade  300  proximate platform  320 . In an alternative embodiment, engagement end  402  is configured to engage each blade  300  between mid-span damper  318  and dovetail  306 . With tool  400  in contact with blades  300 , a manual axial pressure is applied evenly to tool  400  in direction  604  while a manual torque is also applied to tool  400  in direction  606 . Blades  300  slide axially in direction  604  to seat fully in slots  206 . During installation, mid-span dampers  318  interlock with each adjacent mid-span damper. Tool  400  transfers the manual axial pressure from an operator to a substantially simultaneous axial motive force on each blade  300  facilitating preventing interlocking mid-span dampers  318  from stacking-up during the installation process. 
   The above-described blade installation tool is cost-effective and highly reliable for installing fan blades onto a fan rotor such that the blades are seated substantially simultaneously and without mid-span damper overlap. More specifically, the methods and systems described herein facilitate applying a motive force to all blades substantially simultaneously to seat the blades in their respective slots. In addition, the above-described methods and systems facilitate providing a faster and more reliable installation method. As a result, the methods and systems described herein facilitate reducing labor necessary to install fan rotor blades on a fan rotor disc 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.