Abstract: A method for manufacturing a component having a defined geometry includes: a) defining a pre-component geometry including interim shape elements and additional, sacrificial elements for supporting interim elements; b) on a base plate, depositing multiple layers of a powder including a material from which the pre-component will be manufactured; c) sintering the powder to form the pre-component to the defined geometry; d) removing at least some of the sacrificial elements from the pre-component; e) subjecting the remaining pre-component from step d) to a HIP step; and f) removing remaining sacrificial elements from the pre-component product of step e) to provide a component to the defined component geometry. In the definition of the pre-component geometry, the interim elements differ from the corresponding final shape elements in the defined component geometry such that during the HIP step, the interim shape elements adjust to form final shape elements in the defined component geometry.

Abstract: A cooled turbine runner, in particular a high-pressure turbine runner for an aircraft engine, with turbine blades that are radially arranged at a circumferential surface of a rotor disk, wherein respectively one turbine blade with a profiled blade root is inserted into a correspondingly profiled disk finger groove at the circumferential surface of the rotor disk, and wherein a cooling device is provided with at least one cooling air supply channel that extends at least substantially axially and at least over a part of the axial length of the blade root, and with at least one cooling channel that branches off from the same and extends in the interior of the turbine blade up to an outlet opening at its surface. At an inflow side of the blade root, a plug with a cooling air passage is arranged in the cooling air supply channel, wherein the cooling air passage has a geometry that forms a micro-compressor.

Abstract: A turbine, in particular a high-pressure turbine, for an aircraft engine, with a housing at which turbine guide vanes are circumferentially arranged, wherein the turbine guide vanes have at least one interior space through which cooling air flows during operation of the turbine. At least one turbine guide vane has a cooling air passage in the area of an inner wall with respect to the radial direction of the turbine, via which the interior space of the turbine guide vane can be supplied with cooling air.

Abstract: A method for manufacturing a component having a defined geometry includes: a) defining a pre-component geometry including interim shape elements and additional, sacrificial elements for supporting interim elements; b) on a base plate, depositing multiple layers of a powder including a material from which the pre-component will be manufactured; c) sintering the powder to form the pre-component to the defined geometry; d) removing at least some of the sacrificial elements from the pre-component; e) subjecting the remaining pre-component from step d) to a HIP step; and f) removing remaining sacrificial elements from the pre-component product of step e) to provide a component to the defined component geometry. In the definition of the pre-component geometry, the interim elements differ from the corresponding final shape elements in the defined component geometry such that during the HIP step, the interim shape elements adjust to form final shape elements in the defined component geometry.

Abstract: A cooled turbine runner, in particular a high-pressure turbine runner for an aircraft engine, with turbine blades that are radially arranged at a circumferential surface of a rotor disk, wherein respectively one turbine blade with a profiled blade root is inserted into a correspondingly profiled disk finger groove at the circumferential surface of the rotor disk, and wherein a cooling device is provided with at least one cooling air supply channel that extends at least substantially axially and at least over a part of the axial length of the blade root, and with at least one cooling channel that branches off from the same and extends in the interior of the turbine blade up to an outlet opening at its surface. At an inflow side of the blade root, a plug with a cooling air passage is arranged in the cooling air supply channel, wherein the cooling air passage has a geometry that forms a micro-compressor.

Abstract: A turbine, in particular a high-pressure turbine, for an aircraft engine, with a housing at which turbine guide vanes are circumferentially arranged, wherein the turbine guide vanes have at least one interior space through which cooling air flows during operation of the turbine. At least one turbine guide vane has a cooling air passage in the area of an inner wall with respect to the radial direction of the turbine, via which the interior space of the turbine guide vane can be supplied with cooling air.

Abstract: The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, on which means for duct-type guidance of cooling air extending from a front suction-side area of the blade tip to a rear area of the blade tip are provided, and to a method for cooling a blade tip of a turbine rotor blade of a gas turbine, where air from a hot gas flow is guided from a front suction-side area of a blade tip to a rear area of the blade tip through a duct-type guidance.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.

Abstract: In impingement air cooling of gas turbine components, cooling air velocity packs of a certain amplitude and a given frequency are applied to impingement air openings, with intervallic annular swirl structures being formed which penetrate a cross-flow and hit a component to be cooled with high intensity, thus providing for efficient cooling. In order to obtain annular swirl structures with optimum cooling effect, the Strouhal number, which is determined by a ratio of amplitude, frequency of the velocity packs and size of impingement air cooling openings, ranges between 0.2 and 2.0, and preferably between 0.8 and 1.2.

Abstract: The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct (14) extending radially, relative to a rotary axis of the gas-turbine engine, inside the airfoil and air-supply ducts (12) issuing into said cooling duct, characterized in that the cooling duct (14) extends into the blade root (6) in order to generate close to the wall a cooling airflow moved at high circumferential velocity and radially in helical form and that in the area of the blade root (6) at least one nozzle-shaped air-supply duct (12) issues into the cooling duct (14) tangentially or with a tangential velocity component.

Abstract: In impingement air cooling of gas turbine components, cooling air velocity packs of a certain amplitude and a given frequency are applied to impingement air openings, with intervallic annular swirl structures being formed which penetrate a cross-flow and hit a component to be cooled with high intensity, thus providing for efficient cooling. In order to obtain annular swirl structures with optimum cooling effect, the Strouhal number, which is determined by a ratio of amplitude, frequency of the velocity packs and size of impingement air cooling openings, ranges between 0.2 and 2.0, and preferably between 0.8 and 1.2.