Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.

Abstract: A device such as a drum for positioning at least one guide blade row from a plurality of guide blade groups in a turbomachine, the device on the outer circumferential side including at least one flange for attachment to a housing section of the turbomachine and on the inner circumferential side including a plurality of uniformly distributed receptacles for accommodating holding elements of the guide blade groups and a plurality of recesses, the device having a depth-reduced inner circumferential section, relative to the accommodating grooves, between at least two adjoining receptacles, which extends in each case from the one receptacle to the other receptacle; a blade-device combination; a method for assembling such a blade-device combination; and a turbomachine.

Abstract: A method of producing holes in a component, in particular of turbomachines, wherein each hole extends from a first, outer surface to a second, inner surface of the component and wherein the method has, for example, the following steps: (i) producing a 3D model of the actual geometry of the component, at least for the region of the holes; (ii) adapting each hole on the basis of the actual geometry of the component; and (iii) generating a production program for each individual hole. In this way, the process quality and with it the quality of the holes increases, because the offset of holes caused by component tolerances is avoided and the drilling funnels are formed according to specification. Furthermore, drilling defects on account of the offset of holes and/or cores can be avoided. Overlapping holes caused by component tolerances are likewise avoided.

Abstract: A turbomachine including at least one blade-row group that is arranged in the main flow path and at least two rows of blades that are adjacent to each other in the main flow direction, each row having a plurality of blades (38, 40), whereby a narrow cross section and a degree of overlap between the blades of the upstream row of blades and the blades of the downstream row of blades vary starting at the center of the main flow path in the direction of at least one main flow limiter.

Abstract: A method for producing a component of a turbomachine is disclosed. The method includes a) layer-by-layer deposition of a powder component material onto a component platform in a region of a buildup and joining zone, where the deposition takes place in accordance with layer information of the component to be produced; b) local layer-by-layer fusion or sintering of the powder component material by energy supplied in the region of the buildup and joining zone, where the buildup and joining zone is heated to a temperature just below a melting point of the powder component material; c) layer-by-layer lowering of the component platform by a predefined layer thickness; and d) repetition of steps a) to c) until the component is finished. A device for producing a component of a turbomachine is also disclosed.

Abstract: A method for coating a turbine blade of a gas turbine is disclosed. The method includes applying a lacquer coat to a region of the turbine blade, where the lacquer coat includes chromium particles and/or chromium alloy particles, halides, and a binding agent. The method further includes drying the applied lacquer coat at a temperature between 50° C. and 600° C. with disintegration of the binding agent and subsequent reactive connection at a temperature between 900° C. and 1160° C.

Abstract: An inner ring for forming a guide blade ring for a turbomachine is disclosed, composed of at least two one-part ring segments having a plurality of openings, closed on the peripheral side, for accommodating journal bearings on the blade side, the outer diameter of the inner ring being at least 12 times larger than its height. Also disclosed are a guide blade ring having this type of inner ring and a turbomachine having this type of guide blade ring.

Abstract: A blading for a turbomachine, particularly for a gas turbine, wherein thickened areas and depressions are formed and disposed on a lateral wall having a plurality of blades such that at least one depression or thickened area is disposed at a blade pressure side and at least one thickened area or depression is disposed at a blade suction side for each blade of the plurality of blades.

Abstract: A method for producing a rotor, particularly a turbine disk or a turbine ring for a turbine stage of a turbomachine, wherein at least the following steps are carried out: producing a blade ring (14) including a plurality of rotor blades (12), welding adapters (22) together which are disposed in the region of blade footings of the rotor blades (12), wherein at least substantially radial weld seams (30a) having predetermined welding depths are generated, disposing a rotor disk (32) or a rotor ring on the blade ring (14), and welding the rotor disk (32) or the rotor ring to the adapters (22) of the rotor blades (12), wherein at least one further weld seam (30b) is generated. A rotor, particularly a turbine disk or a turbine ring for a turbine stage of a turbomachine is also disclosed.

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: Disclosed is a sealing system for a turbomachine, in particular for a gas turbine, the sealing system being formed in an annular space between a flow-limiting wall of the turbomachine and at least one rotor blade tip of a rotor blade or an outer shroud arranged on the rotor blade tip, and comprising at least one sealing point. The sealing point comprises at least one run-in coating arranged on the rotor blade tip or on the outer shroud in the direction of the flow-limiting wall of the turbomachine. The invention furthermore encompasses a gas turbine comprising the sealing system.

Abstract: Disclosed is a process for producing an alloyed, in particular multiple-alloyed aluminide or chromide layer on a component by alitizing or chromizing. First a green compact layer (9) consisting of a binder (5) and metal particles (7) is deposited on the component (1) to be coated and then alitizing or chromizing is carried out, binder and metal particles being deposited on the component separately from one another, first the binder and then the metal particles. A turbine component produced by this process is also disclosed.

Abstract: A turbomachine stage includes guide vanes and an airfoil platform forming a guide vane cascade, and rotor blades and an airfoil platform forming a rotor blade cascade. Airfoil platforms have cascade regions extending between circumferentially adjacent airfoils, and gap regions which radially and/or axially bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade. A contour of at least one of these gap regions varies in the radial and/or axial direction around the circumference.

Abstract: Disclosed is a process for producing a run-in coating (20, 24, 32, 44) on a component of a turbomachine, in particular of a gas turbine. The run-in coating is applied and produced on the component of the turbomachine by a kinetic cold gas compacting process (K3). The invention also encompasses a run-in coating for a static or rotating component of a turbomachine and a static or rotating component of a turbomachine, in particular of a gas turbine, having at least one run-in coating.

Abstract: Disclosed is an electrode arrangement for the defined rounding or deburring of edges of electrically conductive components, in particular turbine components, by means of electrochemical machining with at least one working electrode (5), which has a tubular electrode carrier, through which an electrolyte inflow line (10) is provided, the electrode carrier having on the front end a closure (13, 18), which is arranged such that the electrolyte inflow line in the axial direction of the electrode carrier is closed, and at least one outlet opening (19) being arranged in the radial direction. Also disclosed is a self-centering electrode arrangement and an installation for the defined rounding or deburring of edges of electrically conductive components by means of electrochemical machining with at least one corresponding electrode arrangement and also a method using the electrode arrangements and the described installation.

Abstract: The present invention relates to a rotating blade (5), in particular for a compressor or turbine stage of a gas turbine, having a radially outer rib arrangement with at least one rib (2), onto which a coating (3) is disposed, whereby, in a meridian section, the coating (3) has an outer contour (3.1), which extends axially outwardly in the radial direction.

Abstract: A method for electrochemically processing a workpiece surface using an electrode, which has at least one effective surface for processing the workpiece surface, and using an electrolyte, wherein the electrolyte is suctioned away from the effective surface. The invention further relates to an electrode, which has at least one electrolyte feed for supplying the electrolyte to the effective area and an electrolyte suctioning system for suctioning the electrolyte away from the effective area.

Abstract: A sealing system and method is disclosed. The sealing system includes a brush sealing element and an accommodating element with an accommodating chamber. The brush sealing element is disposed in the accommodating chamber and a bendable strap is disposed on the accommodating element. The method includes fastening the brush sealing element in the accommodating chamber of the accommodating element. The accommodating element is arranged in a groove of a housing segment and a position of the accommodating element is secured on the housing element by bending the strap of the accommodating element to engage with a side wall of the housing segment.

Abstract: A turbine or compressor stage of a gas turbine, the rotor blade having a radially outer shroud (1) which has a sealing fin array having a first sealing fin (3.1) and a second sealing fin (3.2) which is adjacent to the first sealing fin and connected thereto by a first groove base (10) having a circumferential region (13) of maximum radial height, which is located at a first circumferential position is provided. The sealing fin array has a third sealing fin (3.3) adjacent to the second sealing fin and opposite to the first sealing fin, the third sealing fin being connected to the second sealing fin by a second groove base (20) having a circumferential region (23) of maximum radial height, which is located at a second circumferential position different from the first circumferential position.

Abstract: A blade for a turbomachine is disclosed. At least one cooling channel is configured in the blade neck, whose inlet is disposed near a platform projection on the high-pressure side and whose outlet is disposed in the region of a platform projection on the low-pressure side. An integrally bladed rotor base body having a plurality of these types of blades as well as turbomachine with such a rotor base body is also disclosed.