Patent Publication Number: US-2007122280-A1

Title: Method and apparatus for reducing axial compressor blade tip flow

Description:
TECHNICAL FIELD AND BACKGROUND OF THE INVENTION  
      This invention relates to a method and apparatus for reducing axial compressor tip flow in airfoils, such as blades and vanes. Blade tip flow in the compressor area of a turbine engine results in loss of compressor efficiency and stall margin. In addition, flow recirculation in seal cavities along the inner flow path between the vanes and blades also degrades compressor performance. One prior art solution for reducing tip flow is to reduce the blade tip clearance. This is done by a variety of means, including control of the casing and vane interface using mechanical and/or thermal methods. These methods can cause tip rubbing, excess wear and loss of engine efficiency.  
     BRIEF DESCRIPTION OF THE INVENTION  
      According to one aspect of the invention, a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing is provided, and comprises the step of providing in the airfoil a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip. Air is extracted and pressurized from a region proximate the base of the airfoil, introduced into the channel and conveyed through the channel to the airfoil tip. The air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.  
      Another aspect of the invention provides a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing, comprising the steps of providing a first radially-extending channel having an inlet opening proximate a base of the airfoil on a leading edge side thereof, and an exit opening on the airfoil tip, and providing a second radially-extending channel having an inlet opening proximate the base of the airfoil on a trailing edge side thereof, and an exit opening on the airfoil tip. The air is extracted from a region proximate the base of the airfoil into the channel and pumped through the channel to the airfoil tip. The air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.  
      In another aspect of the invention a turbine machine compressor airfoil is provided, comprising a airfoil base, a airfoil tip, and an air flow channel extending radially from an air inlet opening in the airfoil proximate the airfoil base to an exit opening in the airfoil tip for providing a air blockage against an axial flow of air from the pressure side of the airfoil to the suction side of the airfoil to thereby reduce compressor airfoil tip flow. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
      The invention is described below in conjunction with the following drawings, in which:  
       FIG. 1  is a fragmentary cross-section of an axial flow compressor section of a turbine engine illustrating one embodiment of the invention;  
       FIG. 2  is an enlarged fragmentary cross-section of a portion of the compressor shown in  FIG. 1 ;  
       FIG. 3  is a fragmentary cross-section of the compressor section of a turbine engine illustrating another embodiment of the invention;  
       FIG. 4  is an enlarged fragmentary cross-section of a portion of the compressor shown in  FIG. 3 ;  
       FIG. 5  is a fragmentary cross-section of the compressor section of a turbine engine illustrating yet another embodiment of the invention; and  
       FIG. 6  is an enlarged fragmentary cross-section of a portion of the compressor shown in  FIG. 5 . 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT AND BEST MODE  
      Referring now specifically to the drawings, a partial section of the axial compressor section of a turbine engine T 1  illustrating a method and apparatus for controlling axial compressor blade tip flow according to the present invention is illustrated in  FIG. 1 . The turbine engine “T 1 ” includes compressor blades  10 - 14  and intermediately-positioned stator vanes  15 - 19  in a casing C 1 . The compressor blades  10 - 14  include respective leading edges  10 A- 14 A.  
      As is shown in  FIG. 2 , blade  10  is shown in enlarged detail for clarity, and is also exemplary of blades  11 - 14 . Air is extracted and pressurized from the area of the leading edge side  10 A of the blade  10  through holes  10 B in a disk  20 . The holes  10 B communicate with a channel  10 C that extends radially outwardly through the blade  10  to the tip where it exits through holes  10 D. Note that the channel  10 C may branch out before exiting the tip of the blade  10 . The size of the channel  10 C and the location and number of the branches is determined empirically based on blade size, shape and volume, and engine performance, rating, tip clearance and similar factors. Note in the drawings that the tip clearance is sufficiently small in relation to the scale of the drawings that actual representation of the tip clearance cannot be shown.  
      Referring now to  FIG. 3 , a turbine engine “T 2 ” includes compressor blades  30 - 34  and intermediately-positioned stator vanes  35 - 39  in a casing C 2 . The compressor blades  30 - 34  include respective trailing edges  30 A- 34 A.  
      In  FIG. 4 , blade  31  is shown in enlarged detail for clarity, and is exemplary of blades  30  and  32 - 34 . Air is extracted from the area of the trailing edge side  31 A of the blade  31  through holes  31 B in the disk rim  40 . The holes  31 B communicate with a channel  31 C that extends radially outwardly through the blade  31  to the tip where it preferably branches before exiting through holes  31 D.  
      Referring now to  FIG. 5 , a turbine engine “T 3 ” includes compressor blades  50 - 54  and intermediately-positioned stator vanes  55 - 59  in a casing C 3 . The compressor blades  50 - 54  include respective leading edges  50 A- 54 A and respective trailing edges  50 B- 54 B.  
       FIG. 6  illustrates a blade  52  that is shown in enlarged detail for clarity, and is exemplary of blades  51  and  52 - 54 . Air is extracted from both the areas of the leading edge side  52 A and trailing edge side  52 B of blade  52  through holes  52 C and  52 D in the disk rim  60 . The holes  52 C and  52 D communicate with channels  52 E and  52 F, respectively, that extend radially outwardly through the blade  52  to the tip, where they preferably branch before exiting through holes  52 G.  
      In each of the embodiments described above, the air discharged at the blade tip reduces or prevents blade tip flow by aerodynamically blocking air flow in the region of the tip clearance between the blade tip and the casing. Air from the inner flow path is brought to the tip clearance, as described above, to form this air block. The pressure of the extracted air increases due to the compressor rotor pumping and, when exiting the blade at the tip, resists air flow across the blade tip from the pressure side to the suction side.  
      The methods described above can be applied to both low pressure compressors (boosters) and high pressure compressors. There is no chargeable flow loss when these methods are utilized. Furthermore, by reducing air flow by aerodynamic air blockage rather than by a tight running clearance between the blade tips and the casing, a larger assembly clearance between the blade tips and the casing can be established and maintained. Blade tip rubs are thus reduced, as is recirculation in the inner flow path between the vane and the blade. The extracted air is continuously pumped from the inner flow path to the blade tip, thus providing a continuous air blockage to the blade tip at all times during engine operation.  
      The methods described in this application also have application in blisk (blade integrated disk), skewed or circumferential dovetailed blades.  
      A method and apparatus for controlling axial compressor blade tip flow is described above. Various details of the invention may be changed without departing from its scope. Furthermore, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation—the invention being defined by the claims.