US 12,253,006 B2
Composite airfoil assembly having a dovetail portion
Nicholas Joseph Kray, Mason, OH (US); Elzbieta Kryj-Kos, Evendale, OH (US); Tod Winton Davis, Liberty Township, OH (US); and Gary Willard Bryant, Jr., Loveland, OH (US)
Assigned to General Electric Company, Evendale, OH (US)
Filed by GENERAL ELECTRIC COMPANY, Schenectady, NY (US)
Filed on Dec. 27, 2022, as Appl. No. 18/088,957.
Prior Publication US 2024/0209741 A1, Jun. 27, 2024
Int. Cl. F01D 5/28 (2006.01); F01D 5/14 (2006.01); F01D 5/30 (2006.01)
CPC F01D 5/282 (2013.01) [F01D 5/147 (2013.01); F01D 5/3007 (2013.01); F05D 2220/32 (2013.01); F05D 2240/24 (2013.01); F05D 2240/30 (2013.01); F05D 2300/603 (2013.01)] 18 Claims
OG exemplary drawing
 
1. A composite blade assembly for a turbine engine, the composite blade assembly comprising:
a core comprising a composite structure and defining a dovetail portion and a blade portion, with a transition defined between the dovetail portion and the blade portion, the blade portion including an outer wall extending between a root and a tip in a span-wise direction, the core including a first material having a first bulk modulus, the core having a centerline, the dovetail portion having a distal end axially opposite the transition, with respect to the centerline;
a first skin, having at least one composite ply, overlying the core and defining a first radius at the transition, the first skin having a first portion provided along the dovetail portion, and a second portion defining at least a portion of the outer wall, the first skin having a second material having a second bulk modulus, different from the first bulk modulus; and
a second skin, having at least one composite ply, overlying at least a portion of the first portion of the first skin, the second skin defining a second radius at the transition, the second skin extending to the distal end, the second skin including a radial thickness that continuously decreases in size from the distal end to the transition.