Abstract: An anti-wear coating, in particular an anti-erosion coating, which is applied to a surface of a component that is stressed under fluid technology and is to be protected, in particular a gas turbine part, is disclosed. The anti-wear coating includes one or more multilayer systems applied in a repeating order to the surface to be coated, and the/each multilayer system includes at least one relatively soft metallic layer and at least one relatively hard ceramic layer. All the layers of the/each multilayer system are based on chromium, and a diffusion barrier layer is applied between the surface to be protected and the multilayer system(s).

Abstract: A thermal barrier layer for metallic components, in particular for gas turbine components which are subject to high temperatures or hot gas, is disclosed. The thermal barrier layer includes an inner contact layer and an outer top layer, where the inner contact layer is applied to a surface of the component via an adhesion-promoting layer that is disposed therebetween, and between the outer top layer and the inner contact layer an intermediate layer is formed.

Abstract: A gas turbine component and a method for producing an anti-erosion coating system are disclosed. The gas turbine component includes a basic material, on which an anti-erosion coating system is provided that is a multilayer system including at least one ductile metal layer and at least one hard, ceramics-containing layer for forming a partial anti-erosion system. At least one anti-corrosion layer that has a lower electrochemical potential than the basic material is provided between the partial anti-erosion system and the basic material, thus providing cathodic corrosion protection.

Abstract: The invention relates to a method for producing a blade tip plating (20) on a blade (10) Tor a lurhomachine, in particular on a high-pressure rotating compressor blade lor a gas turbine, comprising the following steps:—producing a particle composite material (24) having embedded hard material particles (18);—placing the panicle composite material (24) on a solder (30) applied (o the blade tip (16): and—healing the solder (30).

Abstract: A method for manufacturing an integrally bladed rotor (10), in particular for a gas turbine, including the following method steps: preparing a rotor base member (12) having at least one first weld surface (16) and one blade (14) having a second weld surface (18); positioning the rotor base member (12) and the blade (14) in such a way that a join zone (20) is formed between the first and second weld surface (16, 18); and filling the join zone (20) with metal powder (24) and the laser welding or electron beam welding of the metal powder (24). An integrally bladed rotor (10) has a join zone (20) between the rotor base member (12) and blades (14), the join zone (20) being filled with welded metal powder (24).

Abstract: The invention relates to a method for connecting at least one turbine blade (10) to a turbine disk (18) or a turbine ring for a turbine stage of a turbomachine, particularly a thermal gas turbine, wherein first a connecting body (16) is formed on the at least one turbine blade (10) by means of a cold gas spraying method, and the connecting body (16) is subsequently connected to turbine disk (18) or to the turbine ring by means of a fusion-welding method. The invention further relates to a turbine stage for a turbine of a turbomachine as well as a turbomachine having a turbine.

Abstract: The present technology generally relates to a wear-resistant coating, especially for gas turbine components, comprising a horizontally segmented or multilayered structure, i.e. at least one relatively hard, ceramic layer and at least one relatively soft, metallic layer. The ceramic layer and the metallic layer are alternately arranged on top of each other in such a way that an external layer forming an external surface of the wear-resistant coating is embodied as a ceramic layer. According to the invention, at least the external, ceramic layer is segmented in a column-type manner in a vertical direction.

Abstract: The present invention relates to methods for joining a rotationally asymmetrical part, preferably a turbine blade, which is made of a monocrystalline material, to a part, preferably a rotor core, which is made of polycrystalline material, having the method steps of:—joining the rotationally asymmetrical part to an adapter piece made of polycrystalline material by means of rotational friction welding, and—joining the adapter piece to the part made of polycrystalline material. The present invention also relates to a turbine blisk having a rotor core made of polycrystalline material and turbine blades made of monocrystalline material.

Abstract: A blade of a turbomachine such as a gas turbine includes a blade vane with a blade tip and a blade base. A blade tip armor cladding is applied on the blade tip. A coating covers at least the armor cladding and includes at least one multilayer coating system, and preferably plural such coating systems stacked repetitively on one another. Each coating system includes at least two different layers stacked successively on one another, with one layer of a metal material closer to the blade tip and one layer of a ceramic material farther from the blade tip.

Abstract: A wear-resistant coating, in particular an erosion-resistant coating for a component that is exposed to fluidic loads, is disclosed. The wear-resistant coating has one or more multilayer systems applied repeatedly to the surface to be coated, where each of the applied multilayer systems has at least four different layers. A first layer of each multilayer system facing the surface to be coated is made of a metallic material adapted to the composition of the component surface to be coated. A second layer applied to the first layer of each multilayer system is made of a metal alloy material adapted to the composition of the component surface to be coated. A third layer applied to the second layer of each multilayer system is made of a gradated metal-ceramic material and a fourth layer applied to the third layer of each multilayer system is made of a nanostructured ceramic material.

Abstract: The invention relates to an erosion protection coating (11), in particular, for gas turbine components, having a horizontally segmented and/or multi-layered construction. i.e., having at least one relatively hard layer (12) and having at least one relatively soft layer (13), wherein the relatively hard layer or each relatively hard layer as well as the relatively soft layer or each relatively soft layer are disposed on top of one another in an alternating manner, in such a way that an outer-lying layer, which forms an outer surface of the erosion protection coating, is formed as a relatively hard layer (12). According to the invention, the relatively hard layer or each relatively hard layer (12) as well as the relatively soft layer or each relatively soft layer (13) are formed as a ceramic layer in each case.

Abstract: A device and method for the partial coating of a component, particularly for the coating of components of a gas turbine or an aircraft engine, is disclosed. The device includes at least one base receptacle for at least partially receiving the component and a first partial region of the component not to be coated. The device further includes at least one plate-shaped cover that can be positioned in the base receptacle, where the cover includes at least one recess or opening for a second partial region of the component to be coated to pass through. The shape of the recess or opening corresponds to the profile of the component in the region between the partial region not to be coated and the partial region to be coated.

Abstract: A method for producing a blisk (bladed disk) or a bling (bladed ring) of a gas turbine is disclosed. The method includes the following steps: a) producing a turbine blade by joining a blade to an adapter element consisting of a metal material that is suitable for fusion welding, the adapter element being used to form a blade root of the turbine blade, and b) joining the turbine blade to a rotor disk consisting of a metal material that is suitable for fusion welding or to a rotor ring consisting of a metal material that is suitable for fusion welding in such a manner that the turbine blade is arranged on the outer periphery of the rotor disk or of the rotor ring. A component of a gas turbine or of a high-pressure or low-pressure compressor, especially a blisk or bling, and a turbine blade are also disclosed.

Abstract: A decoating device for axially symmetric components, particularly from aircraft engines is provided, the decoating process being carried out through the use of a decoating fluid-which is brought on a localized basis into contact with the component to be decoated, a receptacle for holding the decoating fluid being provided that rotates at least about one axis of rotation, and the decoating fluid surface forming a rotation paraboloid in response to the rotation, the component(s) to be decoated on a localized basis being accommodated in the receptacle and being dipped on the radially outer side into the decoating fluid. A decoating device is thereby devised which makes it possible for radially symmetric components to be decoated in the radially outer region in a simple process.

Abstract: The English-language Abstract from the international application is to be retained and is therefore not duplicated in the specification.

Abstract: A blade of a turbomachine such as a gas turbine includes a blade vane with a blade tip and a blade base. A blade tip armor cladding is applied on the blade tip. A coating covers at least the armor cladding and includes at least one multilayer coating system, and preferably plural such coating systems stacked repetitively on one another. Each coating system includes at least two different layers stacked successively on one another, with one layer of a metal material closer to the blade tip and one layer of a ceramic material farther from the blade tip.

Abstract: The use of titanium-based materials in the manufacturing of gas turbines, especially of engines, and, in particular, of compressors is made possible in that a device is devised that protects the components (guide vanes, guide vane stages, rotor blades, rotor blade stages) that are subject to the action of titanium fire and/or FODs by employing a layer system that includes at least two layers, is situated on the surface of the components to be protected, and is bonded thereto, as the case may be, via an adhesion-promoting layer is disclosed. The outermost layer is ceramic; the layer that is subjacent thereto is of metallic nature. Additional layers can optionally follow, ceramic and metallic layers alternating with one another. Moreover, a method is provided for producing the device.

Abstract: An aircraft engine comprising a compressor and at least one turbine is disclosed. The compressor and the turbine are each provided with vanes, each of which has a blade that forms a suction side and a pressure side, where at least one first of the blades is coated with a protective layer in order to reduce the erosion or wear. The layer is applied such that at least two areas are formed that adjoin each other at a boundary line. The first area has a protective layer that has a substantially constant first thickness and the second area is free of the protective layer or has the protective layer with a substantially constant second thickness, the second thickness differing from the first thickness. The boundary line has at least two points, of which the connecting line differs from the course of the boundary line between the two points.

Abstract: Inlet cone (1) for a jet engine (10) configured as a bypass turbojet engine, which has an engine core area (11) and a casing area (12) that surrounds it, whereby the inlet cone (1) is placed concentrically in the air inlet area of the jet engine (10) and placed on a turnably supported shaft (14), whereby the inlet cone (1) has a conical angle (15) at which solid bodies (16) impinging with the air flow onto the inlet cone (1) ricochet off it in motion paths (17) which lead most of the solid bodies (16) into the casing area (12). Thereby an inlet cone (1) is created in which the entry of solid bodies 16 into the engine core area (11) can be minimized.

Abstract: A method for stripping a component, in particular a gas turbine component, to completely or partially remove a multilayer or sandwich antiwear coating from the surface of the component, the antiwear coating including at least one relatively hard ceramic layer and at least one relatively soft metallic layer is disclosed. According to the present invention, in order to remove the multilayer or sandwich antiwear coating, the component is placed in a bath of an alkaline electrolyte, the component placed in the electrolyte being stripped at a current density of between 1 A/dm2 and 20 A/dm2.