Gas turbine engine airfoil tip recesses

Abstract

A gas turbine engine airfoil extending from an airfoil base to an airfoil tip at a free end of the airfoil and from a leading edge to a trailing edge of the airfoil has at least one recess extending into and circumferentially completely through the airfoil tip and located inwardly of the leading and trailing edges. The recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs and may be a circular scallop having a scallop radius and a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. The airfoil may be on a radial blade, a stator vane, or an impeller blade. An annular tip seal may surround a plurality of such airfoil tips with an airfoil tip clearance therebetween.

What is claimed: 
     
         1 . A gas turbine engine component comprising:
 an airfoil extending from an airfoil base to an airfoil tip at a free end of the airfoil,   the airfoil extending downstream from a leading edge to a trailing edge of the airfoil,   at least one recess extending into and circumferentially completely through the airfoil tip, and   the recess located inwardly of the leading and trailing edges of the airfoil.   
     
     
         2 . The component as claimed in  claim 1  wherein the recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs. 
     
     
         3 . The component as claimed in  claim 2  wherein the recess is a circular scallop having a scallop radius. 
     
     
         4 . The component as claimed in  claim 2  wherein the recess has a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. 
     
     
         5 . A gas turbine engine compressor blade comprising:
 an airfoil extending radially outwardly from an airfoil base located on a gas turbine engine rotor to an airfoil tip at a free end of the airfoil,   the airfoil extending downstream from a leading edge to a trailing edge of the airfoil,   at least one recess extending into and circumferentially completely through the airfoil tip, and   the recess located inwardly of the leading and trailing edges of the airfoil.   
     
     
         6 . The gas turbine engine compressor blade as claimed in  claim 5  wherein the recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs. 
     
     
         7 . The gas turbine engine compressor blade as claimed in  claim 6  wherein the recess is a circular scallop having a scallop radius. 
     
     
         8 . The gas turbine engine compressor blade as claimed in  claim 6  wherein the recess has a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. 
     
     
         9 . The gas turbine engine compressor blade as claimed in  claim 6  wherein the gas turbine engine compressor blade is a radial compressor blade or an impeller compressor blade. 
     
     
         10 . A gas turbine engine stator vane comprising:
 a vane airfoil extending radially inwardly from a base of the vane airfoil at an outward end of the stator vane to a vane airfoil tip of the vane airfoil at a radial inward end of the stator vane,   the vane airfoil extending downstream from a leading edge to a trailing edge of the vane airfoil,   at least one recess extending into and circumferentially completely through the vane airfoil tip, and   the recess located inwardly of the leading and trailing edges of the vane airfoil.   
     
     
         11 . The gas turbine engine stator vane as claimed in  claim 10  wherein the recess is located in an area of the vane airfoil tip subject to high tip vibratory stress and or rubs. 
     
     
         12 . The gas turbine engine stator vane as claimed in  claim 11  wherein the recess is a circular scallop having a scallop radius. 
     
     
         13 . The gas turbine engine stator vane as claimed in  claim 11  wherein the recess has a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. 
     
     
         14 . A gas turbine engine assembly comprising:
 a plurality of airfoils,   each of the airfoils extending from an airfoil base to an airfoil tip at a free end of the airfoil,   each of the airfoils extending downstream from a leading edge to a trailing edge of the airfoil,   an airfoil tip clearance between airfoil tips and an annular tip seal surrounding the airfoil tips,   at least one recess extending into and circumferentially completely through each of the airfoil tips, and   the recess located inwardly of the leading and trailing edges of the airfoil.   
     
     
         15 . The assembly as claimed in  claim 14  wherein the recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs. 
     
     
         16 . The assembly as claimed in  claim 15  wherein the recess is a circular scallop having a scallop radius. 
     
     
         17 . The assembly as claimed in  claim 15  wherein the recess has a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. 
     
     
         18 . The assembly as claimed in  claim 14  further comprising:
 a plurality of blades including the airfoils, 
 the airfoils extending radially outwardly from the airfoil base to the airfoil tip, 
 the airfoil base mounted on a gas turbine engine rotor, and 
 wherein the annular tip seal is a blade rub land. 
 
     
     
         19 . The assembly as claimed in  claim 18  wherein the recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs. 
     
     
         20 . The assembly as claimed in  claim 18  wherein the recess is a circular scallop having a scallop radius. 
     
     
         21 . The assembly as claimed in  claim 14  further comprising:
 the airfoils being vane airfoils on stator vanes cantilevered from and fixed to an engine casing at radial outward ends of the stator vane, 
 the airfoils unsupported at free radial inward ends of the airfoils, and 
 wherein the annular tip seal is a rotor seal land on a rotor of the assembly. 
 
     
     
         22 . The assembly as claimed in  claim 21  wherein the recesses are located in an area of the airfoil tips subject to high tip vibratory stress and or rubs. 
     
     
         23 . The assembly as claimed in  claim 21  wherein the recesses are circular scallops having scallop radii. 
     
     
         24 . The assembly as claimed in  claim 14  further comprising:
 the airfoils being impeller airfoils extending outwardly from impeller airfoil bases on a rotor disc portion of an impeller to impeller airfoil tips at free ends of the impeller airfoils, and 
 wherein the annular tip seal is an annular centrifugal blade tip shroud. 
 
     
     
         25 . The assembly as claimed in  claim 24  wherein the recesses are located in an area of the airfoil tips subject to high tip vibratory stress and or rubs. 
     
     
         26 . The assembly as claimed in  claim 24  wherein the recesses are circular scallops having scallop radii. 
     
     
         27 . A method of reducing vibratory stress at an airfoil tip at a free end of an airfoil of a gas turbine engine component, the method comprising determining an area or areas of high tip vibratory stress and machining, cutting, or otherwise forming at least one recess or scallop extending into and circumferentially completely through the airfoil tip in the area or areas of high tip vibratory stress. 
     
     
         28 . The as claimed in  claim 27  wherein the recess or scallop is formed to a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess or scallop. 
     
     
         29 . The as claimed in  claim 27  wherein:
 the airfoil tip and airfoil are from pluralities of airfoil tips and airfoils from a single stage or circumferential row of blades or vanes surrounded by annular tip seals, 
 the determining an area or areas of high tip vibratory stress includes running or rotating a rotor containing the blades or an annular tip seal surrounding the vanes respectively and observing nicks or scratches in the annular seal or airfoil tips, and 
 forming the recesses or scallops in area or areas of high tip vibratory stresses adjacent the nicks or scratches.