The present disclosure relates to a layer stack and methods of preparing the same for use as an oxidation and chemical barrier with superalloy substrates, including Ni, Ni—Co, Co, and Ni-aluminide based substrates. The layer system can be applied to a substrate in a single physical vapor deposition process with no interruption of vacuum conditions.

Airfoils for gas turbine engines are described. The airfoils include an airfoil body having a pressure and suction side walls and an internal rib and defining a cavity between, at least, the pressure side wall, the suction side wall, and the internal rib. A baffle insert is arranged within the cavity and includes a sealing portion arranged adjacent to the internal rib of the airfoil body with a gap therebetween. A seal element is arranged within the cavity and located in the gap. The seal element is free to move relative to each of the baffle insert and the airfoil body within the gap. When a pressure differential exists across the seal element, the seal element will sealingly engage between sealing portion of the baffle insert and the internal rib of the airfoil body to block a flow through the gap.

A casting method includes: forming a seed, the seed having a first end and a second end, the forming including bending a seed precursor; placing the seed second end in contact or spaced facing relation with a chill plate; contacting the first end with molten material; and cooling and solidifying the molten material so that a crystalline structure of the seed propagates into the solidifying material. The forming further included reducing a thickness of the seed proximate the first end relative to a thickness of the seed proximate the second end.

Embodiments of the disclosure provide a turbomachine component, including: a base portion configured for mounting on a rotor; an airfoil portion having a first end coupled to the base portion, and a second end opposite the first end. A creep resistance of the airfoil portion is greater than the base portion, and a fracture toughness of the airfoil portion is less than the base portion. A tip portion may be coupled to the second end of the airfoil portion. A creep resistance of the tip portion is less than the airfoil portion and greater than the base portion. A fracture toughness of the tip portion is less than the base portion and greater than the airfoil portion.

A turbomachine airfoil element comprises an airfoil having: an inboard end; an outboard end; a leading edge; a trailing edge; a pressure side; and a suction side. A span between the inboard end and the outboard end is 1.75-2.20 inches. A chord length at 50% span is 1.05-1.35 inches. At least two of: a first mode resonance frequency is 2400±10% Hz; a second mode resonance frequency is 4950±10% Hz; a third mode resonance frequency is 7800±10% Hz; a fourth mode resonance frequency is 8700±10% Hz; and a fifth mode resonance frequency is 12500±10% Hz.

There is provided a cobalt-based alloy product comprising: in mass %, 0.08-0.25% C; more than 0.04% and 0.2% or less N, the total amount of C and N being more than 0.12% and 0.28% or less; 0.1% or less B; 10-30% Cr; 5% or less Fe and 30% or less Ni, the total amount of Fe and Ni being 30% or less; W and/or Mo, the total amount of W and Mo being 5-12%; 0.5% or less Si; 0.5% or less Mn; 0.5 to 2 mass % of an M component being a transition metal other than W and Mo and having an atomic radius of more than 130 pm; and the balance being Co and impurities. The product comprises matrix phase crystal grains, in which particles of MC carbides, M(C,N) carbonitrides and/or MN nitrides including the M component are precipitated at an average interparticle distance of 0.13-2 μm.

Embodiments of the present disclosure generally relate to oxide layer compositions for turbine engine components and methods for depositing the oxide layer compositions. In one or more embodiments, a turbine engine component includes a superalloy substrate and a bond coat disposed over the superalloy substrate. The turbine engine component includes an oxide layer disposed over the bond coat, where the oxide layer includes aluminum oxide and a metal dopant. The turbine engine component includes a thermal barrier coating disposed over the oxide layer.

A sealing assembly positioned between a first component having a first interface and a second component having a second interface is provided. The sealing assembly includes a flat bracket fixedly attached to the first interface to define an extension and a floating seal including a body portion, a first U-shaped channel arranged to engage the extension, and a second U-shaped channel inhibiting movement of the floating seal in an axial direction while allowing movement in a radial direction, and the second U-shaped channel allowing movement in the axial direction and inhibiting movement in the radial direction.

The present disclosure relates to turbine blades adapted for use in gas turbine engines. In particular, this disclosure is directed to turbine blades that include components made from ceramic matrix composite materials and that incorporate abradable materials.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. Radial cooling channels in the trailing edge portion of the airfoil permit radial flow of a cooling fluid through the trailing edge portion. Each radial cooling channel has a first end at a lower surface at a root edge of the trailing edge portion or at an upper surface at a tip edge of the trailing edge portion and a second end opposite the first end at the lower surface or the upper surface. A method of making a turbine component and a method of cooling a turbine component are also disclosed.