An airfoil assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil including an airfoil section extending from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extending between a tip portion and the root section in a radial direction. The airfoil section defines a pressure side and a suction side separated in a circumferential direction. A root mount includes first and second mount members secured to circumferentially opposed sides of the root section. Arcuate circumferential faces of the first and second mount members are dimensioned to pivotably mount the root section to a hub in an installed position. A method of assembly is also disclosed.

A gas turbine engine, having: a disk; a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with a platform of an adjacent blade that is secured to the disk; a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades; and a damper restraint located on an interior surface of each platform, wherein the damper restraint extends into the cavity and is a raised feature configured to contact a peripheral edge portion of a damper seal when it is adjacent to the gap defined by the platforms of the blades.

A gas turbine engine includes a turbine section having static turbine vane rings and turbine wheel assemblies. The vane rings are arranged to direct combustion products toward blades included in the turbine wheel assemblies to cause the turbine wheel assemblies to rotate.

A gas turbine engine, having: a disk; a plurality of blades secured to the disk, each of the plurality of blades having a platform located between a root portion and an airfoil portion of the blade, wherein the platform of one of the plurality of blades is configured to define a cavity with a platform of an adjacent blade that is secured to the disk; a damper seal located in the cavity and positioned adjacent to a gap defined by edges of the platforms of the blades; and a damper restraint located on an interior surface of each platform, wherein the damper restraint extends into the cavity and is a raised feature configured to contact a peripheral edge portion of a damper seal when it is adjacent to the gap defined by the platforms of the blades.

The present disclosure provides devices related to aircraft engine fan assemblies and blade lock retainers. In various embodiments, a blade lock retainer is formed from sheet metal and comprises an annular ring portion, an outer retainer tab portion, and an inner tab portion. In various embodiments, the outer retainer tab portion is disposed on an outer circumference of the annular ring portion, oriented substantially perpendicular to the annular ring portion, and extends in an aft direction from the annular ring portion. In various embodiments, the inner retainer tab portion is disposed on an inner circumference of the annular ring portion and radially aligned with the outer retainer tab portion, oriented substantially perpendicular to the annular ring portion, and extends in a forward direction from the annular ring portion.

A gas turbine engine includes a turbine section having static turbine vane rings and turbine wheel assemblies. The vane rings are arranged to direct combustion products toward blades included in the turbine wheel assemblies to cause the turbine wheel assemblies to rotate.