The present invention concerns a turbine part comprising a substrate made of nickel-based monocrystalline superalloy, comprising chromium and at least one element chosen among rhenium and ruthenium, the substrate having a - phase, an average mass fraction of rhenium and of ruthenium greater than or equal to 4% and an average mass fraction of chromium less than or equal to 5% and preferably less than or equal to 3%, a sub-layer covering at least a part of a surface of the substrate, characterised in that the sublayer has a - phase and an average atomic fraction of chromium greater than 5%, of aluminium between 10% and 20% and of platinum between 15% and 25%.

A coating system for a surface of a superalloy component is provided. The coating system includes a MCrAlY coating on the surface of the superalloy component, where M is Ni, Fe, Co, or a combination thereof. The MCrAlY coating generally has a higher chromium content than the superalloy component. The MCrAlY coating also includes a platinum-group metal aluminide diffusion layer. The MCrAlY coating includes Re, Ta, or a mixture thereof. Methods are also provided for forming a coating system on a surface of a superalloy component.

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

An assembly is provided for a turbine engine with a flowpath. This assembly includes a fuel source and an engine component. The engine component forms a peripheral boundary of the flowpath. The engine component includes a component internal passage. The engine component is configured to receive fuel from the fuel source. The engine component is configured to crack at least some of the fuel within the component internal passage thereby cooling the engine component and providing at least partially cracked fuel. The assembly is configured to direct the at least partially cracked fuel into the flowpath for combustion.

Methods for repairing a thermal barrier coating deposited on a component with localized spallation of the thermal barrier coating includes depositing a primer slurry on a thermally grown oxide of the component exposed by the localized spallation, depositing a ceramic slurry on the primer slurry, and heating the primer slurry and the ceramic slurry. The primer slurry includes a primer that includes at least one of a metal and a metal oxide. The ceramic slurry includes a ceramic material, a ceramic slurry binder material, and a ceramic slurry fluid carrier. Heating the primer slurry and the ceramic slurry forms a first chemical bond between the primer and the thermally grown oxide and a second chemical bond between the primer and the ceramic material.

A method for coating a substrate having a cavity structure, in particular a cooling structure, inside the substrate, wherein the cavity structure includes openings in the surface of the substrate. At least one bonding layer, in particular a diffusion layer, or at least one metallic layer is applied onto the substrate, in particular onto the surface of the substrate, and subsequently at least one thermal protection layer is applied onto the at least one diffusion layer by using a plasma spray physical vapour deposition (PS-PVD) method, a hollow cathode sputtering method or a suspension plasma spray (SPS) method.

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

A turbine temperature estimation system controls a valve in a cooling passage to control the flow rate of cooling air supplied to a turbine component on the basis of its temperature. The system includes a coating layer formed on a surface of a component of the gas turbine; a measuring unit to supply an electric current to the coating layer and to measure a change in a resistance value of the coating layer; and a controller to estimate a temperature of the coating layer on the basis of the resistance value. The coating layer includes a heat shielding material and a resistive material whose resistance value changes with temperature. A cooling passage supplies cooling air to cool the turbine component, and the controller controls an opening of the cooling passage according to a voltage value of the coating layer.

An aircraft engine incudes: a catalyst, e.g. platinum, applied to turbine blades (580c, 590c) and/or to a catalytic grid downstream of the engine's combustion chamber. An exhaust fluid additive injection system is incorporated upstream of the catalyst via a line 560, in a stator blade within the engine. The catalyst is used to reduce NOx emissions from the engine by a Selective Catalytic Reduction reaction.

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.