Abstract: A diffuser for a radial compressor of a turbomachine is provided. The diffuser has a plurality of diffuser channels, wherein the diffuser channels extend across a radial area of the diffuser across a bent area into an axial area of the diffuser, wherein, in the radial area of the diffuser, the diffuser channels have diffuser walls in particular at v-shaped blades that are bent in the movement direction of the radial compressor or are straight.

Abstract: A device for adjusting a gap between the housing of an impeller and the impeller in a radial compressor of a turbomachine is provided. The device comprises at least one mechanical setting device by means of which the housing of the impeller can be axially adjusted in a targeted manner for the purpose of adjusting the gap, wherein the at least one setting device can be actuated via an adjusting ring, and a transmission of a force and/or a torsional moment from the adjusting ring can be effected via at least one adjustment lever device, in particular a thread lever, or a tilting lever.

Abstract: A fan blade for an aircraft engine, including a leading edge, a trailing edge, a suction side, a pressure side and a blade tip, is provided. The fan blade has a large-area elastomer layer which takes up at least 20% of the surface of the suction side of the fan blade.

Abstract: A turbofan engine for a civil supersonic aircraft, including: an engine intake; a multi-stage fan that is arranged behind the engine intake; a core engine that comprises a compressor, a combustion chamber and a turbine; a primary flow channel that leads through the core engine; a secondary flow channel that leads past the core engine; an adjustable converging-diverging thrust nozzle that forms a nozzle throat area and a nozzle outlet area, wherein at least the nozzle throat area is adjustable; and a thrust reverser that is integrated in the adjustable converging-diverging thrust nozzle.

Abstract: A turbofan engine for a civil supersonic aircraft, including: a multi-stage fan, wherein each stage of the multi-stage fan has a fan rotor with rotor blades and a fan stator with stator blades; an inlet flow channel that leads through the fan rotor of the first stage of the multi-stage fan; a primary flow channel that leads through a core engine of the turbofan engine; and a secondary flow channel that leads past the core engine. The inlet flow channel divides into the primary flow channel and the secondary flow channel. It is provided that the fan rotors of all stages of the fan are arranged in the inlet flow channel before it is divided into the primary flow channel and the secondary flow channel.

Abstract: An aircraft gas turbine has a core engine and a bypass duct surrounding the latter, where the core engine includes in its inflow zone a booster, in the area of which at least one booster bleed duct is provided for supplying an airflow from the booster into the bypass duct, and a stator vane row arranged in the bypass duct downstream of a fan. A vane root of the stator vane is designed as a structural element mounting the core engine, and the booster bleed duct is provided inside the vane root and issues downstream of the stator vane into the bypass duct.

Abstract: A device for adjusting a gap between the housing of an impeller and the impeller in a radial compressor of a turbomachine is provided. The device comprises at least one mechanical setting device by means of which the housing of the impeller can be axially adjusted in a targeted manner for the purpose of adjusting the gap, wherein the at least one setting device can be actuated via an adjusting ring, and a transmission of a force and/or a torsional moment from the adjusting ring can be effected via at least one adjustment lever device, in particular a thread lever, or a tilting lever.

Abstract: A diffuser for a radial compressor of a turbomachine is provided. The diffuser has a plurality of diffuser channels, wherein the diffuser channels extend across a radial area of the diffuser across a bent area into an axial area of the diffuser, wherein, in the radial area of the diffuser, the diffuser channels have diffuser walls in particular at v-shaped blades that are bent in the movement direction of the radial compressor or are straight.

Abstract: A fan blade for an aircraft engine, including a leading edge, a trailing edge, a suction side, a pressure side and a blade tip, is provided. The fan blade has a large-area elastomer layer which takes up at least 20% of the surface of the suction side of the fan blade.

Abstract: An aircraft gas-turbine engine having a fan mounted on a fan shaft and including several fan blades, where the fan shaft is mounted on an engine structure by means of an upstream fan bearing, when seen in the flow direction, and a downstream bearing, where at least one predetermined breaking point is provided in the area of the fan bearing, characterized in that at least one shock-absorbing element designed as metal fabric damper is arranged at a radial gap to have a parallel effect relative to the predetermined breaking point.

Abstract: A method for balancing and assembling a turbine rotor 2 including at least one turbine disk 3 and a compressor rotor 15 of a gas turbine. A front cover disk 6 is fastened by connections 10 to the turbine disk 3, and the turbine disk 3 is balanced together with the cover disk 6. The turbine disk 3 and the cover disk 6 are braced by a clamping device 19, 20. The bolted connections 10 are then removed and the turbine disk 3 and the cover disk 6 are connected to the compressor rotor 15 by bolted connections 10. The clamping device 19, 20 is then removed.

Abstract: An inductive sensor and to a method for manufacturing an inductive sensor is provided. The inductive sensor comprises a winding body, an induction coil made of a high-temperature resistant metal alloy and wound onto the winding body, and a metal casing at least partially surrounding the induction coil. Here, the induction coil is compressed with metal oxide powder inside the metal casing.

Abstract: A method for balancing and assembling a turbine rotor 2 including at least one turbine disk 3 and a compressor rotor 15 of a gas turbine. A front cover disk 6 is fastened by connections 10 to the turbine disk 3, and the turbine disk 3 is balanced together with the cover disk 6. The turbine disk 3 and the cover disk 6 are braced by a clamping device 19, 20. The bolted connections 10 are then removed and the turbine disk 3 and the cover disk 6 are connected to the compressor rotor 15 by bolted connections 10. The clamping device 19, 20 is then removed.

Abstract: The present invention relates to an aircraft gas turbine having a core engine and a bypass duct surrounding the latter, where the core engine includes in its inflow zone a booster, in the area of which at least on booster bleed duct is provided for supplying an airflow from the booster into the bypass duct, and a stator vane row arranged n the bypass duct downstream of a fan, characterized in that a vane root of the stator vane is designed as a structural element mounting the core engine, and that the booster bleed duct is provided inside the vane root and issues downstream of the stator vane into the bypass duct.