Gas turbine engine fan blade with axial lean

Abstract

A fan blade for a gas turbine engine is provided with forward axial lean. The fan blade may have a substantially straight leading edge. The geometry of the fan blade results in a lower susceptibility to flutter, thereby allowing a gas turbine engine comprising such a fan blade to operate over a wider range of operating conditions.

We claim: 
     
         1 . A fan blade for a gas turbine engine, the gas turbine engine defining axial, radial and circumferential directions, the fan blade comprising:
 an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span wherein, when viewed along a circumferential direction:   the angle (α) formed between the radial direction and a straight line (AC) drawn between the leading edge at the root and at the tip is in the range of from −6° and −0.2°; and   the angle (α (P 1 , P 2 )) formed between the radial direction and a line drawn between any two points (P 1 , P 2 ) on the leading edge that have a difference in radius of at least 5% of the blade span is in the range of from −6° and 0°,   where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade.   
     
     
         2 . A fan blade for a gas turbine engine according to  claim 1 , wherein:
 when viewed along a circumferential direction, the maximum perpendicular distance (e) between any point on the leading edge and a straight line drawn between the leading edge at the root and at the tip is 2% of the blade span.   
     
     
         3 . A fan blade for a gas turbine engine according to  claim 1 , wherein the fan blade comprises:
 a platform; and   a root portion, wherein   the root portion extends between the platform and the root of the aerofoil portion.   
     
     
         4 . A fan blade for a gas turbine engine according to  claim 3 , wherein the radial extent of the root portion is no more than 7% of the span of the aerofoil portion. 
     
     
         5 . A fan blade for a gas turbine engine according to  claim 1 , wherein the fan blade comprises a tip portion that extends at least radially away from the tip of the aerofoil portion. 
     
     
         6 . A fan blade for a gas turbine engine according to  claim 5 , wherein the radial extent of the tip portion is no more than 7% of the span of the aerofoil portion. 
     
     
         7 . A fan blade for a gas turbine engine according to  claim 1 , wherein:
 for two points on the leading edge that are radially closer to the tip than to the root and have a difference in radius of at least 5% of the blade span, the axial position of the radially outer point is forward of the axial position of the radially inner point.   
     
     
         8 . A fan blade for a gas turbine engine according to  claim 7 , wherein:
 for two points on the leading edge that are radially closer to the tip than to the root and have a difference in radius of at least 2% of the blade span, the axial position of the radially outer point is forward of the axial position of the radially inner point.   
     
     
         9 . A fan blade for a gas turbine engine according to  claim 1 , wherein the aerofoil portion has a trailing edge extending from the root the tip, wherein, when viewed along a circumferential direction the angle (β) formed between the radial direction and a straight line (BD) drawn between the trailing edge at the root and at the tip is in the range of from −20° and −5°, where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade. 
     
     
         10 . A fan stage for a gas turbine engine comprising:
 a hub; and   a plurality of fan blades according to  claim 1 , wherein:   the fan blades extend radially form the hub.   
     
     
         11 . A fan stage for a gas turbine engine according to  claim 10 , wherein:
 the ratio of the radius of the position where the leading edge of one of the fan blades meets the hub to the outermost radial extent of the leading edge of the fan blade is less than 0.33.   
     
     
         12 . A gas turbine engine comprising a fan blade according to  claim 1 . 
     
     
         13 . A gas turbine engine comprising a fan stage according to  claim 10 . 
     
     
         14 . A gas turbine engine comprising:
 a fan having a plurality of fan blades according to  claim 1 ;   a turbine; and   a gearbox, wherein:   the fan is driven from the turbine via the gearbox, in order to reduce the rotational speed of the fan stage compared with the driving turbine stage.   
     
     
         15 . A gas turbine engine according to  claim 12  with a specific thrust of less than 100 N/Kg/s. 
     
     
         16 . A gas turbine engine according to  claim 13  with a specific thrust of less than 100 N/Kg/s. 
     
     
         17 . A gas turbine engine according to  claim 14  with a specific thrust of less than 100 N/Kg/s. 
     
     
         18 . A method of manufacturing a fan stage for a gas turbine engine comprising:
 providing a fan hub; and   attaching a plurality of fan blades the fan hub using linear friction welding, the fan blades comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span wherein, when viewed along a circumferential direction,the angle (α) formed between the radial direction and a straight line (AC) drawn between the leading edge at the root and at the tip is in the range of from −6° and −0.2°; and   the angle (α (P 1 , P 2 )) formed between the radial direction and a line drawn between any two points (P 1 , P 2 ) on the leading edge that have a difference in radius of at least 5% of the blade span is in the range of from −6° and 0°,   where a negative angle indicates that the respective line has an axial component that is in the same direction as the axial component of the direction from a trailing edge to the leading edge of the blade.