US 11,834,964 B2
Low radius ratio fan blade for a gas turbine engine
Li Zheng, Niskayuna, NY (US); Scott Roger Finn, Evendale, OH (US); Nicholas J. Kray, West Chester, OH (US); Trevor H. Wood, Clifton Park, NY (US); and Kishore Ramakrishnan, Rexford, NY (US)
Assigned to GENERAL ELECTRIC COMPANY, Schenectady, NY (US)
Filed by General Electric Company, Schenectady, NY (US)
Filed on Nov. 24, 2021, as Appl. No. 17/535,291.
Prior Publication US 2023/0160312 A1, May 25, 2023
Int. Cl. F01D 5/30 (2006.01); F01D 5/28 (2006.01)
CPC F01D 5/3007 (2013.01) [F01D 5/282 (2013.01); F05D 2220/32 (2013.01); F05D 2230/60 (2013.01); F05D 2240/30 (2013.01)] 21 Claims
OG exemplary drawing
 
1. A fan blade for a gas turbine engine, comprising:
an airfoil including a leading edge and a trailing edge extending between a root and a tip of the airfoil, a distance between the leading edge and the trailing edge defining a flow path length, a radially innermost point of the leading edge defining a leading edge hub point, an axially forwardmost point of the leading edge defining a forward leading edge point, the forward leading edge point positioned forward of the leading edge hub point, the leading edge hub point positioned proximal to a central axis of the gas turbine engine, the forward leading edge point positioned distal to the central axis of the gas turbine engine; and
a root section including an axial dovetail having a pair of opposed pressure faces, an axial length of the axial dovetail less than the flow path length, the axial dovetail including a fore end and an aft end, a portion of the root of the airfoil extending axially between the fore end and the aft end, a top edge of the axial dovetail radially inward of the portion of the root of the airfoil, the top edge of the axial dovetail radially outward of the leading edge hub point.