Abstract: A rotor blade (20) for a rotor blade assembly (10) of a turbomachine (1) is provided, having an inner rotor blade platform (40) which extends axially from the rotor blade (20) with respect to a longitudinal turbomachine axis (2) and has two opposite circumferential end faces (41) and a free axial end (42) whose cross section is radially inwardly and radially outwardly bounded by circular arcs of two concentric circles. The inner rotor blade platform (40) has a cross section of connection (45) with the rotor blade (20) which is bounded radially inwardly by an inner connecting line (46) and radially outwardly by an outer connecting line (47). Each of the connecting lines (46, 47) has a central portion (61) having a convex curvature.

Abstract: The invention relates to a guide vane segment for aircraft gas turbine, which extend along a respective circular arc and together form an annular section, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacently in the peripheral direction, the outer shroud and the two sealing walls form a trough-like profile in longitudinal section, wherein, at the axially front or/and rear sealing wall element(s), at least one relief gap with a main section is provided, which extends substantially radially inward, starting from a radial outer edge of the respective front or/and rear sealing wall element(s) along the sealing wall element. According to the invention, it is proposed that the relief gap has at least one additional section that adjoins the main section radially inward, wherein the additional section is formed by at least one curved subsection.

Abstract: The invention relates to a guide vane segment for aircraft gas turbine, which extend along a respective circular arc and together form an annular section, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacently in the peripheral direction, the outer shroud and the two sealing walls form a trough-like profile in longitudinal section, wherein, at the axially front or/and rear sealing wall element(s), at least one relief gap with a main section is provided, which extends substantially radially inward, starting from a radial outer edge of the respective front or/and rear sealing wall element(s) along the sealing wall element. According to the invention, it is proposed that the relief gap has at least one additional section that adjoins the main section radially inward, wherein the additional section is formed by at least one curved subsection.

Abstract: A guide vane arrangement for a turbomachine, particularly a gas turbine, is disclosed. The guide vane arrangement includes an external shroud band, on which at least two guide vanes protrude radially inward. A pair of axial ribs spaced apart from each other in the peripheral direction and disposed between two adjoining guide vanes in the peripheral direction protrude radially outward from a circumferential surface of the external shroud band and/or axially from a radial flange of the external shroud band.

Abstract: For producing a component, especially a gas turbine component, coated with a wear-protection, corrosion-protection or erosion-protection coating, a method includes the following steps: providing a component (10) to be coated on a component surface (13); at least partially coating the component (11) on its component surface with an at least two-layered protective coating (14), which includes at least one relatively soft layer (15) and at least one relatively hard layer (16); and then surface densifying the at least partially coated component on its coated component surface by ball blasting or shot peening.

Abstract: A turbomachine is disclosed. The turbomachine includes a rotor, a plurality of blades arranged side by side on the rotor in a circumferential direction, and a plurality of sealing apparatuses each disposed on a respective blade shaft of the plurality of blades. Each of the plurality of sealing apparatuses includes first and second bottom plates protruding in a respective axial direction on the respective blade shaft and first and second bulkheads which extend in a radial direction along the respective blade shaft and form a double-walled seal in the axial direction. A respective rounded or parabolic transitional cross-section from the bottom plates to the bulkheads is provided and at least one of the bulkheads is offset with respect to a blade root edge of the respective blade shaft in a direction toward a radial longitudinal axis of the respective blade shaft.

Abstract: A guide vane arrangement for a turbomachine, particularly a gas turbine, is disclosed. The guide vane arrangement includes an external shroud band, on which at least two guide vanes protrude radially inward. A pair of axial ribs spaced apart from each other in the peripheral direction and disposed between two adjoining guide vanes in the peripheral direction protrude radially outward from a circumferential surface of the external shroud band and/or axially from a radial flange of the external shroud band.

Abstract: A turbomachine is disclosed. The turbomachine includes a rotor, a plurality of blades arranged side by side on the rotor in a circumferential direction, and a plurality of sealing apparatuses each disposed on a respective blade shaft of the plurality of blades. Each of the plurality of sealing apparatuses includes first and second bottom plates protruding in a respective axial direction on the respective blade shaft and first and second bulkheads which extend in a radial direction along the respective blade shaft and form a double-walled seal in the axial direction. A respective rounded or parabolic transitional cross-section from the bottom plates to the bulkheads is provided and at least one of the bulkheads is offset with respect to a blade root edge of the respective blade shaft in a direction toward a radial longitudinal axis of the respective blade shaft.

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: 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: 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 component, specifically a component for a gas turbine, having a main body of a base material and a wear protection coating applied to a surface of the main body, specifically a protective coating against corrosion and/or oxidation and/or erosion, is disclosed. A barrier coat is applied between the main body and the wear protection coating to protect the main body during chemical or electrochemical decoating of the component.

Abstract: The invention relates to a method for the production of a component, especially gas turbine component, coated with a wear-protection coating, especially a corrosion-protection coating or erosion-protection coating, with the following steps: a) providing a component (10) to be coated on a component surface (13); b) at least partially coating the component (11) on its component surface with an at least two-layered or at least two-plied wear-protection coating (14), whereby the wear-protection coating (14) encompasses at least one relatively soft layer (15) and at least one relatively hard layer (16); c) surface densifying the at least partially coated component on its coated component surface.

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.