Patent Publication Number: US-6905309-B2

Title: Methods and apparatus for reducing vibrations induced to compressor airfoils

Description:
BACKGROUND OF THE INVENTION 
     This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for reducing vibrations induced to rotor blades. 
     Gas turbine engine rotor blades typically include airfoils having leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between the airfoil root and the tip. An inner flowpath is defined at least partially by the airfoil root, and an outer flowpath is defined at least partially by a stationary casing. For example, at least some known compressors include a plurality of rows of rotor blades that extend radially outwardly from a disk or spool. 
     Known compressor rotor blades are cantilevered adjacent to the inner flowpath such that a root area of each blade is thicker than a tip area of the blades. More specifically, because the tip areas are thinner than the root areas, and because the tip areas are generally mechanically unrestrained, during operation wake pressure distributions may induce chordwise bending or other vibration modes into the blade through the tip areas. In addition, vibrational energy may also be induced into the blades by a resonance frequency present during engine operation. Continued operation with chordwise bending or other vibration modes may limit the useful life of the blades. 
     To facilitate reducing tip vibration modes, and/or to reduce the effects of a resonance frequency present during engine operations, at least some known vanes are fabricated with thicker tip areas. However, increasing the blade thickness may adversely affect aerodynamic performance and/or induce additional radial loading into the rotor assembly. Accordingly, other known blades are fabricated with a shorter chordwise length in comparison to other known blades. However, reducing the chord length of the blade may also adversely affect aerodynamic performance of the blades. 
     BRIEF SUMMARY OF THE INVENTION 
     In one aspect a method for fabricating a rotor blade for a gas turbine engine is provided. The method comprises forming an airfoil including a first side wall and a second side wall that each extend in radial span between an airfoil root and an airfoil tip, and wherein the first and second side walls are connected at a leading edge and at a trailing edge, and forming a winglet that extends outwardly from at least one of the airfoil first side wall and the airfoil second side wall, such that a radius extends between the winglet and at least one of the airfoil first side wall and the second side wall. 
     In another aspect, an airfoil for a gas turbine engine is provided. The airfoil includes a leading edge, a trailing edge, a tip, a first side wall that extends in radial span between an airfoil root and the tip, wherein the first side wall defines a first side of said airfoil, and a second side wall connected to the first side wall at the leading edge and the trailing edge, wherein the second side wall extends in radial span between the airfoil root and the tip, such that the second side wall defines a second side of the airfoil. The airfoil also includes a winglet extending outwardly from at least one of said first side wall and said second side wall such that a radius extends between said winglet and at least least one of said first and second side walls. 
     In a further aspect, a gas turbine engine including a plurality of rotor blades is provided. Each rotor blade includes an airfoil having a leading edge, a trailing edge, a first side wall, a second side wall, and at least one winglet that extends outwardly from at least one of the first side wall and the second side wall such that a radius is formed between the winglet and at one of said first and second side walls. The airfoil first and second side walls are connected axially at the leading and trailing edges, and the first and second side walls also extend radially from a blade root to an airfoil tip. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is schematic illustration of a gas turbine engine; 
         FIG. 2  is a perspective view of a rotor blade that may be used with the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a partial perspective view of the rotor blade shown in  FIG. 2 , and viewed from an opposite side of the rotor blade; 
         FIG. 4  is a cross-sectional view of the rotor blade shown in  FIG. 3  and taken along line  4 — 4 ; 
         FIG. 5  is a cross-sectional view of the rotor blade shown in  FIG. 3  and taken along line  5 — 5 ; 
         FIG. 6  is a cross-sectional view of an alternative embodiment of a rotor blade that may be used with the gas turbine engine shown in  FIG. 1 . 
     
    
    
     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 , a low pressure turbine  20 , and a booster  22 . Fan assembly  12  includes an array of fan blades  24  extending radially outward from a rotor disc  26 . Engine  10  has an intake side  28  and an exhaust side  30 . In one embodiment, the gas turbine engine is a GE90 available from General Electric Company, 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 (not shown in  FIG. 1 ) from combustor  16  drives turbines  18  and  20 , and turbine  20  drives fan assembly  12 . 
       FIG. 2  is a partial perspective view of a rotor blade  40  that may be used with a gas turbine engine, such as gas turbine engine  10  (shown in  FIG. 1 ).  FIG. 3  is a partial perspective view of rotor blade  40  viewed from an opposite side of rotor blade  40 .  FIG. 4  is a cross-sectional view of rotor blade  40  taken along line  4 — 4 .  FIG. 5  is a cross-sectional view of rotor blade  40  taken along line  5 — 5 . In one embodiment, a plurality of rotor blades  40  form a high pressure compressor stage (not shown) of gas turbine engine  10 . Each rotor blade  40  includes an airfoil  42  and an integral dovetail  43  used for mounting airfoil  42  to a rotor disk (not shown) in a known manner. Alternatively, blades  40  may extend radially outwardly from a disk (not shown), such that a plurality of blades  40  form a blisk (not shown). 
     Each airfoil  42  includes a first contoured side wall  44  and a second contoured side wall  46 . First side wall  44  is convex and defines a suction side of airfoil  42 , and second side wall  46  is concave and defines a pressure side of airfoil  42 . Side walls  44  and  46  are joined at a leading edge  48  and at an axially-spaced trailing edge  50  of airfoil  42 . More specifically, airfoil trailing edge  50  is spaced chordwise and downstream from airfoil leading edge  48 . First and second side walls  44  and  46 , respectively, extend longitudinally or radially outward in span from a blade root  52  positioned adjacent dovetail  43 , to an airfoil tip  54 . 
     A winglet  70  extends outwardly from second side wall  46 . In an alternative embodiment winglet  70  extends outwardly from first side wall  44 . In a further alternative embodiment, a first winglet extends outwardly from second side wall  46  and a second winglet extends outwardly from first side wall  44 . Accordingly, winglet  70  is contoured to conform to side wall  46  and as such follows airflow streamlines extending across side wall  46 . In the exemplary embodiment, winglet  70  extends in a chordwise direction substantially across side wall  46 , such that winglet  70  is substantially flush with side wall  46  adjacent leading edge  48  and adjacent trailing edge  50 . Alternatively, the winglet is aligned in a non-chordwise direction with respect to side wall  46 . More specifically, in the exemplary embodiment, winglet  70  extends chordwise substantially between airfoil leading and trailing edges  48  and  50 , respectively. Alternatively, the winglet extends to only one of airfoil leading or trailing edges  48  and  50 , respectively. In a further alternative embodiment, winglet  70  extends only partially along side wall  46  between airfoil leading and trailing edges  48  and  50 , respectively, and does not extend to either leading or trailing edges  48  and  50 , respectively. 
     Winglet  70  has a non-rectangular cross-sectional profile and is aerodynamically-shaped with respect to side wall  46  such that a first radius R 1  and a second radius R 2  extend between winglet  70  and side wall  46 . In the exemplary embodiment, winglet  70  also includes an arcuate outer surface  90  that extends between first radius R 1  and a second radius R 2 . More specifically, first radius R 1  extends along winglet  70  to provide a smooth transition between winglet  70  and airfoil tip  54 , and second radius R 2  extends along winglet  70  to provide a smooth transition between winglet  70  and root  52 . In the exemplary embodiment, first radius R 1  is larger than second radius R 2 . A geometric configuration of winglet  70 , including a relative position, size, and length of winglet  70  with respect to blade  40 , can vary and is selected based on operating and performance characteristics of blade  40 . 
     Winglet  70  facilitates stiffening airfoil  42  such that a natural frequency of vibration of airfoil  42  is increased to a frequency that is not present within gas turbine engine  10  during normal engine operations. Accordingly, modes of vibration that may be induced into similar airfoils that do not include a winglet  70 , are facilitated to be substantially eliminated by winglet  70 . More specifically, winglet  70  enables a provides a technique for tuning chordwise mode frequencies out of the normal engine operating speed, such that a desired frequency margin may be achieved. In addition, winglet  70  also facilitates strengthening blade  40  without providing frequency margin. 
     Moreover, during assembly of airfoil  42 , the cross-sectional shape of winglet  70  enables winglet  70  to be formed integrally with airfoil  42  with reduced manufacturing costs compared to other geometric shapes. Specifically, the combination of winglet first radius R 1 , second radius R 2 , and arcuate outer surface  90 , enable winglet  70  to be formed using an eletro-chemical machining (ECM) process with a radial electrolyte flow. More specifically, the smooth transition formed by each radius R 1  and R 2  between winglet  70  and airfoil  42  facilitates the ECM electrode flowing smoothly and continuously over winglet  70  without cavitation or flow disruption. The ECM process facilitates blade  40  being manufactured with reduced costs and time in comparison to other known blade manufacturing methods. 
     Energy induced to airfoil  42  is calculated as the dot product of the force of the exciting energy and the displacement of airfoil  42 . More specifically, during operation, aerodynamic driving forces, i.e., wake pressure distributions, are generally the highest adjacent airfoil tip  54  because tip  54  is generally not mechanically constrained. However, winglet  70  stiffens and increases a local thickness of airfoil  42 , such that the displacement of airfoil  42  is reduced in comparison to similar airfoils that do not include winglet  70 . Accordingly, because winglet  70  increases a frequency of airfoil  42  and reduces an amount of energy that is induced to airfoil  42 , airfoil  42  receives less aerodynamic excitation and less harmonic input from wake pressure distributions. In addition, because winglet  70  is positioned radial distance  102  from tip  54 , rib  70  will not contact the stationary shroud. Furthermore, because first radius R 1  is larger than second radius R 2 , first radius R 1  facilitates reducing stress concentrations between winglet  70  and airfoil  42 , thus improving the strength and useful life of blade  40 . 
       FIG. 6  is a cross-sectional view of an alternative embodiment of a rotor blade  200  that may be used with gas turbine engine  10  (shown in  FIG. 1 ). Rotor blade  200  is substantially similar to rotor blade  40  (shown in  FIGS. 2–5 ) and components in rotor blade  200  that are identical to components of rotor blade  40  are identified in  FIG. 6  using the same reference numerals used in  FIGS. 2–5 . Specifically, in one embodiment, rotor blade  200  is identical to rotor blade  40  with the exception that rotor blade  200  includes a second winglet  202  in addition to winglet  70 . More specifically, in the exemplary embodiment, winglet  202  is identical to rib  70  but extends across side wall  44  rather than side wall  46 . 
     Winglet  202  extends outwardly from first side wall  44  and is contoured to conform to side wall  44 , and as such, follows airflow streamlines extending across side wall  44 . In the exemplary embodiment, winglet  202  extends in a chordwise direction substantially across side wall  44 , such that winglet  202  is substantially flush with side wall  44  adjacent leading edge  48  and adjacent trailing edge  50 . Alternatively, winglet  202  is aligned in a non-chordwise direction with respect to side wall  46 . More specifically, in the exemplary embodiment, winglet  202  extends chordwise substantially between airfoil leading and trailing edges  48  and  50 , respectively. Alternatively, winglet  202  extends to only one of airfoil leading or trailing edges  48  and  50 , respectively. In a further alternative embodiment, winglet  202  extends only partially along side wall  46  between airfoil leading and trailing edges  48  and  50 , respectively, and does not extend to either leading or trailing edges  48  and  50 , respectively. 
     A geometric configuration of winglet  202 , including a relative position, size, and length of winglet  202  with respect to blade  40 , is variably selected based on operating and performance characteristics of blade  40 . In one embodiment, winglet  202  is positioned radial distance  102  from airfoil tip  54 , and as such is substantially radially aligned with winglet  70 . In another embodiment, winglet  202  is not radially aligned with respect to winglet  70 . 
     The above-described rotor blade is cost-effective and highly reliable. The rotor blade includes a winglet that extends outwardly from at least one of the airfoil surfaces. The winglet facilitates tuning chordwise mode frequencies of the blade out of the normal engine operating speed range. Furthermore, the stiffness of the winglet facilitates decreasing an amount of energy induced to each respective airfoil. Moreover, the winglet facilitates improving performance of the airfoil relative to an airfoil having substantially less tip chord. As a result, a winglet is provided that facilitates maintaining aerodynamic performance of a blade, while providing aeromechanical stability to the blade, in a cost effective and reliable manner. 
     Exemplary embodiments of blade assemblies are described above in detail. The blade assemblies are not limited to the specific embodiments described herein, but rather, components of each assembly may be utilized independently and separately from other components described herein. Each rotor blade component can also be used in combination with other rotor blade components. 
     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.