An article has a metallic substrate having a plurality of recesses. A first coating is at least at the recesses and has: a splatted layer; and a columnar layer atop the splatted layer. A second coating is away from the recesses and has: a columnar layer atop the substrate without an intervening splatted layer.

A coating system for coating a part (10), such as a turbine blade or vane, has a mask (14) positioned adjacent to a first portion (16) of the part (10) to be coated and a mechanism (30) for moving the mask (14) relative to the part (10). The mechanism (30) may be a gear mechanism or a magnetic mechanism.

The present invention relates to a relief valve (1) for a turbocharger, in which the crank arm (3) is made of a first material and the shaft (4) is made of a second material different from the first material used for manufacturing the crank arm (3), each of the materials containing a composition that provides the necessary properties according to the application of each component of the relief valve (1). The present invention also relates to a process for manufacturing the relief valve (1), which allows the crank arm (3) and the shaft (4) to be manufactured separately, using different materials for the manufacture of each component.

Disclosed is a method for improving the high-temperature stability of a component, in particular a blade of a turbomachine, formed at least partially from a molybdenum alloy that, besides molybdenum, silicon, boron and titanium, comprises iron and/or yttrium. The method comprises depositing a diffusion barrier layer formed from technically pure molybdenum or tungsten or being an alloy based on molybdenum and/or tungsten at least on an outer surface, which comprises the molybdenum alloy, of the component, and depositing silicon on the diffusion barrier layer to form molybdenum silicides and/or tungsten silicides.

Disclosed is a thermal barrier coating system for components of a turbomachine, especially for high temperature-stressed or hot gas-stressed components of a turbomachine, comprising a ceramic coating of fully or partially stabilized zirconium oxide, and an oxide cover coating which comprises aluminum and at least one element from the group lanthanum, magnesium, silicon, calcium and sodium. The aluminum oxide exists at least partially as free α-Al.sub.2O.sub.3. Also disclosed is a method for producing a corresponding thermal barrier coating system.

A vane device for a gas turbine is provided. The vane device includes a first airfoil having a first suction side and a first pressure side, a second airfoil having a second suction side and a second pressure side, an inner shroud and an outer shroud. The first airfoil and the second airfoil are arranged between the inner shroud and the outer shroud, wherein the first airfoil and the second airfoil are at least partially coated with a MCrAlY coating. At least the first suction side has a first coated surface section which is coated with a thermal barrier coating and which represents at least a part of the total surface of the first suction side. At least the inner shroud or the outer shroud has a further coated surface section which is coated with a further thermal barrier coating. A corresponding method of manufacturing is also provided.

A turbine part, such as a turbine blade or a distributor fin, for example, including a substrate made of superalloy based on monocrystalline nickel, including rhenium and/or ruthenium, and having a γ′-NisAI phase that is predominant by volume and a γ-Ni phase, the part also including a sublayer made of metal superalloy based on nickel covering the substrate, wherein the sublayer has a γ′-NisAI phase that is predominant by volume and wherein the sublayer has an average atomic fraction of aluminium of between 0.15 and 0.25, of chromium of between 0.03 and 0.08, of platinum of between 0.01 and 0.05, of hafnium of less than 0.01 and of silicon of less than 0.01. A process for manufacturing a turbine part including a step of vacuum deposition of a sublayer made of a superalloy based on nickel having predominantly by volume a γ′-NisAI phase, on a substrate made of superalloy based on nickel including rhenium and/or ruthenium.

An article has a metallic substrate having a plurality of recesses. A first coating is at least at the recesses and has: a splatted layer; and a columnar layer atop the splatted layer. A second coating is away from the recesses and has: a columnar layer atop the substrate without an intervening splatted layer.

An embodiment of a line-of-sight coating fixture includes a support structure, a spindle, and a shadow structure. The support structure includes a plurality of compartments disposed below a platter, each compartment having an opening on a periphery of the support structure. Each compartment is adapted to receive and secure a base of a workpiece such that a body of each workpiece to be coated is disposed about a periphery of the support structure and extends above the platter. The spindle is disposed through a center of the platter or support structure for rotating the workpieces thereabout. The shadow structure is disposed about the spindle, inside of the periphery, the shadow structure sized and adapted to shield a portion of each workpiece from line-of-sight coating material.

High temperature stable thermal barrier coatings useful for substrates that form component parts of engines such as a component from a gas turbine engine exposed to high temperatures are provided. The thermal barrier coatings include a multiphase composite and/or a multilayer coating comprised of two or more phases with at least one phase providing a low thermal conductivity and at least one phase providing mechanical and erosion durability. Such low thermal conductivity phase can include a rare earth zirconate and such mechanical durability phase can include a rare earth a rare earth aluminate. The different phases are thermochemically compatible even at high temperatures above about 1200° C.