Rolls-royce Deutschland Ltd & Co. Kg

Markus Weinert

Tobias Leymann

Thomas Wunderlich

Thomas Schießl

A blade component of a compressor or turbine stage of a gas turbine, in particular of a gas turbine engine is provided. The blade component has at least two structural elements which can be connected together by means of a connection process, in particular sintering. The at least two structural elements can be coupled and/or connected to at least one means for internal cooling of the blade component, and/or the means for internal cooling of the blade component is arranged in at least one of the least two structural elements. The at least two structural elements, when assembled, can be coupled to a cooling insert inside the blade component, wherein during operation, cooling air flowing into the cooling insert can be guided in targeted fashion via impingement cooling openings onto the inside of the blade component, in particular as impingement flow cooling.

Blade component, method for manufacture of same, and gas turbine

A rotor for an engine is provided. The rotor comprising a rotor base part that has fastening grooves for rotor blades that are arranged in succession around a rotational axis along a circumferential direction, multiple rotor blades that are respectively supported in a form-fit manner inside a corresponding fastening groove by means of a blade root, and at least one securing element for the axial securing&#x2014;with respect to a rotational axis of at least one of the rotor blades at the rotor base part. The at least one securing element has two edges that are arranged at a radial distance to one another and through which the securing element is supported in a form-fit manner at the rotor base part, on the one hand, and, on the other hand, at the at least one rotor blade.

Rotor with overhang at blades for a locking element

A rotor of an aircraft engine includes a disk wheel and at least two turbine blades. A damping device is arranged below a blade platform and inside a cavity between the blades in the radial direction. The cavity is formed by a respective recess at adjoining blades, with respective recesses having an outer guide surface in the radial direction, acting together with the damping device during operation of the rotor. The guide surface extends radially, from the lateral edge area of the blade to the central area of the blade. The respective recess extends circumferentially from a boundary plane defined by the lateral edge area of the blade to a side surface that delimitates the guide surface towards the blade center. The recess of at least one of the adjacent blades completely receives the damping device.

Rotor device of an aircraft engine with a damping device between turbines blades

A cooled turbine runner for a gas turbine, in particular an aircraft engine, with turbine blades that are radially arranged at the circumferential surface of a rotor disk, wherein respectively a turbine blade is inserted with a profiled blade root into a correspondingly profiled disk finger groove at the circumferential surface of the rotor disk. A cooling device with at least one cooling air supply channel is provided, which extends at least substantially axially and at least over a part of the axial length of the blade root, and which has at least one inlet with an inlet opening at an inflow side of the blade root, wherein the inlet is embodied with a projection. A middle axis of the inlet has, in a direction oriented towards the inlet opening and in the area of the inlet opening, an extension component in the rotational direction of the rotor disk.

Cooled turbine runner for an aircraft engine

A rotor of a turbine of a gas turbine, having a rotor hub and a plurality of blade platforms, each carrying at least one turbine blade, and having one each cooling air chamber for passing cooling air formed by adjacent blade platforms. At least one damping element and/or sealing element is arranged radially beneath a platform intermediate gap provided between adjacent blade platforms. The cooling air chamber and/or the platform intermediate gap are designed such that starting from a side of the blade platform facing the adjacent blade platform, at least one outflow pocket open to the platform upper side and connected to the cooling air chamber is formed, and at least one cooling air pocket opened radially inwards to the cooling air chamber is provided in the wall of the cooling air chamber and opens towards the damping element and/or the sealing element or the platform intermediate gap.

Rotor of a turbine of a gas turbine with improved cooling air routing

A stator of a turbine of a gas turbine, having a flow duct in which stator vanes are arranged, each vane having an outer and an inner vane platform, and each having a cooling air chamber formed by adjacent vane platforms for passing through cooling air, characterized in that the resultant cooling air chamber and/or a platform intermediate gap are designed such that starting from a side of the vane platform facing the adjacent vane platform, at least one outflow pocket open to a platform upper side and connected to the cooling air chamber is formed, and that in the wall of the cooling air chamber, at least one cooling air pocket opened radially inwards to the cooling air chamber and limited by an intermediate element or a sealing strip is formed and opens to the platform intermediate gap.

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Rangsdorf, Germany

Markus Weinert