Abstract: The invention relates to a pre-swirl nozzle system in a gas turbine, in particular in an aircraft engine, having at least one pre-swirl nozzle, through which cooling air can flow, characterized in that the at least one pre-swirl nozzle has a region with a constant first cross section and a downstream region with a second cross section, which expands in the flow direction, in particular expands in a strictly monotonic manner.

Abstract: The invention relates to a mixing device for a primary fluid flow in a first pipe with at least one secondary fluid flow, characterized in that a mixing point of the fluid flows is arranged in each case at the connection of the first pipe to at least one second pipe for the at least one secondary fluid flow, and an orifice plate device for a targeted reduction in the flow cross section is arranged upstream of the mixing point in the at least one second pipe, or the first pipe has a wall section with an orifice plate device for the targeted reduction in the flow cross section, through which orifice plate device the at least one secondary fluid flow flows into the first pipe during operation. The invention furthermore relates to a mixing method and to an aircraft engine.

Abstract: The invention relates to a measuring device for an aircraft engine, characterized by at least one probe device for measuring a physical and/or chemical state in at least one measuring space within the aircraft engine, wherein the at least one measuring space is fluidically connected to a cavity, and at least one air-conducting device, which is fluidically coupled to the cavity in such a manner that a fluid flow, in particular a gas flow, can be removed from the at least one cavity to a pressure sink. The invention also relates to an aircraft engine and to a measuring method.

Abstract: An air supply system for a gas turbine engine including a first duct which is connected or capable of being connected to a first compressor bleed air supply of a gas turbine engine; a second duct which downstream of the first compressor bleed air supply is connected or capable of being connected to a second compressor bleed air supply of the gas turbine engine; a nozzle by way of which air from the second duct is capable of being blown into the first duct; and an exhaust air duct having an opening which downstream of the nozzle is disposed in the first duct in such a manner that air blown by way of the nozzle into the first duct can flow out of the first duct through the opening into the exhaust air duct. A gas turbine engine and an aircraft are furthermore provided.

Abstract: A description is given of a gas turbine engine of an aircraft, having an engine core which comprises at least one compressor and at least one turbine, through which a core air flow is passed and which are rotatably mounted in the region of bearings. Part of the core air flow flows out of the engine core as a partial air flow into a region situated radially inside the engine core. A device which at least partially deflects the flow of the partial air flow in such a way that a static pressure in the region downstream of the device is lower than upstream of the device is provided in the flow path of the partial air flow, in the transitional region between the engine core and the radially inner region, and thus an axial bearing force starting from a surface on which the lower pressure acts is reduced.

Abstract: The invention relates to a measuring device for an aircraft engine, characterized by at least one probe device for measuring a physical and/or chemical state in at least one measuring space within the aircraft engine, wherein the at least one measuring space is fluidically connected to a cavity, and at least one air-conducting device, which is fluidically coupled to the cavity in such a manner that a fluid flow, in particular a gas flow, can be removed from the at least one cavity to a pressure sink. The invention also relates to an aircraft engine and to a measuring method.

Abstract: An air supply system for a gas turbine engine including a first duct which is connected or capable of being connected to a first compressor bleed air supply of a gas turbine engine; a second duct which downstream of the first compressor bleed air supply is connected or capable of being connected to a second compressor bleed air supply of the gas turbine engine; a nozzle by way of which air from the second duct is capable of being blown into the first duct; and an exhaust air duct having an opening which downstream of the nozzle is disposed in the first duct in such a manner that air blown by way of the nozzle into the first duct can flow out of the first duct through the opening into the exhaust air duct. A gas turbine engine and an aircraft are furthermore provided.

Abstract: A valve system for a fluid line system in an aircraft engine, which fluid line system has at least one fluid line wherein the at least one fluid line has at least one check valve, wherein the valve position of an actuator in the at least one check valve is changeable, in particular automatically changeable, in dependence on the pressure ratios in each case acting on the at least one check valve. The valve system has a monitoring means for recording the respective valve position, in particular the open position and/or the closed position of the actuator, in dependence on at least one measurement value, wherein a signal is output in dependence on the recorded valve position. The valve system furthermore has a means for setting a minimum required sealing-air stream. The invention also relates to a valve control method.

Abstract: A sensor arrangement with a sensor element for measuring at least one physical and/or chemical fluid characteristic in a turbomachine is provided. The sensor element detects the at least one fluid characteristic inside a non-contact seal, in particular a labyrinth seal, between a rotor stage and a stator stage, wherein during operation the sensor element is in contact with the fluid flow along the flow path inside the labyrinth seal.