Blade leading edge tip rib

Abstract

A rotor blade for a gas turbine engine includes a leading edge tip rib projecting outwardly from an airfoil of the blade at a tip region thereof. The tip rib continuously surrounds a leading edge of the airfoil and extends rearwardly from the leading edge along respective pressure and suction side surfaces to thereby alter the blade tip leakage vortex structure and strength, resulting in a stage efficiency benefit.

1 . A rotor blade for a gas turbine engine, comprising:
 an attachment; and   an airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip defining a chord line, and the pressure side surface adjacent to the tip being substantially planar; and   wherein the tip region includes a tip rib projecting outwardly from the airfoil, the tip rib continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface, to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.   
     
     
         2 . The rotor blade as defined in  claim 1  wherein the tip rib has a lateral extension in a range between 5% and 30% as thick as the maximum airfoil thickness along the mean camber line. 
     
     
         3 . The rotor blade as defined in  claim 1  wherein the tip rib has a lateral extension which is 25% as thick as the maximum airfoil thickness along the mean camber line. 
     
     
         4 . The rotor blade as defined in  claim 1  wherein each of the first and second axial extensions is 35% of the length of the chord line immediately downstream of the leading edge. 
     
     
         5 . The rotor blade as defined in  claim 1  wherein tip rib is flush with the tip of the airfoil. 
     
     
         6 . The rotor blade as defined in  claim 1  wherein the chord line is substantially parallel to the pressure side surface. 
     
     
         7 . The rotor blade as defined in  claim 1  wherein a chord line increases as the airfoil extends from the attachment to the tip. 
     
     
         8 . The rotor blade as defined in  claim 1  wherein the first and second axial extensions of the tip rib are equal. 
     
     
         9 . The rotor blade as defined in  claim 1  wherein the first and second axial extensions of the tip rib are different. 
     
     
         10 . The rotor blade as defined in  claim 3  wherein the lateral extension on the respective leading edge, pressure side surface and suction side surface are equal. 
     
     
         11 . The rotor blade as defined in  claim 3  wherein the lateral extension varies on the leading edge, pressure side surface and suction side surface. 
     
     
         12 . A gas turbine engine, comprising:
 a compressor section;   a combustor section; and   a turbine section;   wherein the turbine section includes a plurality of rotors having a plurality of radially disposed rotor blades, each of the rotor blades including an attachment and an airfoil, the airfoil defining a pressure side surface and a suction side surface, extending between the attachment and a tip, a base region disposed adjacent to the attachment, a tip region, and a transition region located between the base region and the tip region, the tip region defining a chord line and the pressure side surface in the tip region being substantially planar; and   wherein the tip region includes a tip rib projecting outwardly from the airfoil, continuously surrounding a leading edge of the airfoil, extending from the leading edge rearwardly along the respective pressure side surface and suction side surface to define a first axial extension on the pressure side surface and a second axial extension on the suction side surface, each of the first and second axial extensions being in a range between 5% and 35% of the length of the chord line immediately downstream of the leading edge.   
     
     
         13 . The gas turbine engine as defined in  claim 12  wherein the tip rib has a lateral extension in a range between 5% and 30% as thick as the maximum airfoil thickness along the mean camber line.