The present disclosure provides a fir tree coupling for gas turbine engine parts comprising a load beam having a longitudinal axis, a base, a first side, and a second side, a rail extending from the base of the load beam between the first side and the second side, a tooth running parallel to the longitudinal axis disposed on the first side of the load beam. The rail may comprise at least one of, a convex sidewall having a convex curvature, a concave sidewall having a concave curvature, or a vertical sidewall extending perpendicular to the base. The rail may comprise a sidewall comprising a sidewall step wherein the sidewall has a step cut into a portion of the rail. The rail may comprise a tapered sidewall wherein the tapered sidewall extends at an angle to the base.

The present disclosure provides a fir tree coupling for gas turbine engine parts comprising a load beam having a longitudinal axis, a rounded base, a first side, and a second side, wherein the rounded base has a radius of curvature from the first side to the second side, a tooth running parallel to the longitudinal axis and disposed on the first side of the load beam. The fir tree coupling may comprise a channel through the rounded base across a portion of the radius of curvature from the first side to the second side. The channel may comprise a sidewall having a sidewall step cut into a portion of the channel sidewall.

A load absorbing system that may include a rotor blade retention system is provided. The load absorbing system may include a block, a first retainer plate, and a deformable core. The block may be selectively positioned alongside a dovetail groove. The block may have a first face directed away from the blade root and an axially-spaced second face directed toward the blade root. The first retainer plate may be attached to the second face of the block and axially positioned between the block and the axially-directed surface of the blade root. The deformable core may be positioned between the block and the first retainer plate.

A metal leading edge includes a nose positioned along the leading edge of a fan blade airfoil body. The metal leading edge also includes a first edge extending axially aftward from the nose along a pressure side of the fan blade airfoil body. The metal leading edge further includes a second edge extending axially aftward from the nose along a suction side of the fan blade airfoil body. The first edge and the second edge forming a notch at the conjunction of the first edge, the second edge, and the nose. The metal leading edge also includes a nose length extending from a nose tip to the notch. The nose length at a first radial location is different from the nose length at a second radial location.

The invention relates to a turbine engine blade, particularly made of composite material, including: on the one hand, an airfoil (10) which exhibits: a leading edge (12), a trailing edge (14) opposite to the leading edge (12), intrados (16) and extrados (18) lateral walls which connect the leading edge (12) to the trailing edge (14), et on the other hand, a structural reinforcement (20) including a base (24) extending into two fins (26, 28) and designed to be applied to the leading edge (12) and the lateral walls (16, 18) of the airfoil (10), characterized in that the fins (26, 28) of the structural reinforcement (20) and/or the lateral walls (16, 18) of the airfoil (10) are shaped to maintain an assembly space (30) with nonzero thickness between at least one of the fins (26, 28) and the airfoil (10) when the structural reinforcement (20) is in place on the airfoil (10), said assembly space (30) extending from a free end (27, 29) of said fin (26, 28) toward the base (24) of the structural reinforcement (20).

A method of additively manufacturing is provided. The method may include successively depositing and fusing together layers of a superalloy powder mixture comprised of a base material powder and a eutectic powder, to build up an additive portion, which eutectic powder has a solidus temperature lower than the solidus temperature of the base material powder. The method may also include heat treating the additive portion at a temperature greater than 1200° C. to heal cracks and/or fill pores and to homogenize the alloy of which the additive portion is comprised. The additive portion alloy has a chemistry defined by the superalloy powder mixture. The base material powder may be formed of a nickel-base superalloy with an aluminum content by weight of at least 1.5%. The eutectic powder may be a nickel-base alloy including by weight about 6% to about 11% chromium, about 5% to about 9% titanium, and about 9% to about 13% zirconium, with balance nickel as its primary components.

A fan blade includes a metallic body, a first composite cover, and a second composite cover. The metallic body may have a first side, a second side, a plurality of first retention slots, and a plurality of second retention slots, in accordance with various embodiments. The first and second retention slots may extend from the first side to the second side of the metallic body. The first composite cover may be coupled to the first side of the metallic body and may include a plurality of first fingers that extend through the first retention slots and are coupled to the second side of the metallic body. The second composite cover may be coupled to the second side of the metallic body and may include a plurality of second fingers that extend through the second retention slots and are coupled to the first side of the metallic body.

A gas turbine engine includes: a turbine section including a casing extending circumferentially about a plurality of turbine blades and having at least one seal member coated with an abradable coating. At least one turbine blade has sides and a tip and at least one seal member is located adjacent to the tip of the at least one turbine blade. The tip of the at least one turbine blade has a wear resistant layer and an abrasive coating disposed on the wear resistant layer. The wear resistant layer has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A gas turbine engine includes an engine static structure extending circumferentially about an engine centerline axis; a compressor section, a combustor section, and a turbine section within the engine static structure. At least one of the compressor section and the turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil. A tip of the at least one airfoil is metal having a wear resistant coating and the at least one seal member is coated with an abradable coating. The wear resistant coating is formed as a layer in a base metal surface of the airfoil, has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A fixture assembly including a first fixture portion; a second fixture portion that interfaces with the first fixture portion; and a bladder assembly mounted to the second fixture portion to face the first fixture portion. A method of manufacturing a fan blade includes inserting a blade body and a cover into a fixture; and deploying a bladder assembly within the fixture to press the cover into the blade body.