Abstract: A rotor blade for a turbomachine, in particular in an aircraft engine, with a cooling arrangement for cooling a surface inside the rotor blade by means of a cooling medium, in particular cooling air. The rotor blade having an impingement cooling device with a plurality of impingement cooling openings for deflecting the cooling medium flowing in the interior of the impingement cooling device onto the surface that is to be cooled by means of impingement cooling inside the rotor blade and that is located outside of the impingement cooling device, so that the surface can be cooled by means of an impingement cooling through a cooling medium than exits from the impingement cooling openings, wherein the impingement cooling device is movably mounted with respect to the rotor blade.

Abstract: An adaptive aircraft engine including a first, inner bypass duct inside a core engine of the aircraft engine and a second, outer bypass duct at least partially surrounding the first bypass duct, and adaption means, in particular adjustable nozzles for altering a flowed-through cross-section of the core engine with the first bypass duct and for altering the flowed-through cross-section of the second bypass duct depending on the flying speed. An aircraft having at least one adaptive aircraft engine is also provided.

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 rotor blade for a turbomachine, in particular in an aircraft engine, with a cooling arrangement for cooling a surface inside the rotor blade by means of a cooling medium, in particular cooling air. The rotor blade having an impingement cooling device with a plurality of impingement cooling openings for deflecting the cooling medium flowing in the interior of the impingement cooling device onto the surface that is to be cooled by means of impingement cooling inside the rotor blade and that is located outside of the impingement cooling device, so that the surface can be cooled by means of an impingement cooling through a cooling medium than exits from the impingement cooling openings, wherein the impingement cooling device is movably mounted with respect to the rotor blade.

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: An adaptive aircraft engine including a first, inner bypass duct inside a core engine of the aircraft engine and a second, outer bypass duct at least partially surrounding the first bypass duct, and adaption means, in particular adjustable nozzles for altering a flowed-through cross-section of the core engine with the first bypass duct and for altering the flowed-through cross-section of the second bypass duct depending on the flying speed. An aircraft having at least one adaptive aircraft engine is also provided.

Abstract: A gas turbine engine is disclosed comprising a shaft driven by a turbine, the turbine directly driving a lower speed compressor via a shaft and driving a higher speed compressor via the shaft and a multiplier power gearbox. The gearbox comprises a ring gear, a plurality of planet gears and planet carrier and a sun gear. The input from the turbine to the gearbox is to the carrier and the output from the gearbox to drive the higher speed compressor is from the sun gear or the ring gear.

Abstract: A gas-turbine combustion chamber arrangement includes a flame tube, a diffuser element arranged upstream of the flame tube in the flow direction, the diffuser element including an annular duct, and an axial compressor arranged upstream of the diffuser element. The diffuser element features an annular guide vane area in which guide vanes are arranged, which for redirecting an incoming flow are provided at an angle (?) in a range between 28° and 32° relative to a central axis of the gas turbine. Downstream of the guide vane area, a diffuser area is arranged, the diffuser area not being provided with flow-guiding elements affecting the flow, where burners arranged in the annular combustion chamber are provided with their burner axes at an angle (?) between 40° and 50° relative to the central axis.

Abstract: A gas-turbine combustion chamber arrangement includes at least one centrifugal compressor as well as one centripetal annular combustion chamber, with a stator vane arrangement being provided between the centrifugal compressor and the annular combustion chamber. The stator vane arrangement for diverting the air flowing out of the centrifugal compressor is designed at an angle ? of 20° to 30°, preferably 25°, relative to the engine axis, so the airflow is passed at essentially this angle ? to the combustion chamber. The inflow area into the combustion chamber for supplying air is designed at an angle of 20° to 30°, preferably 25°, relative to the meridional plane, and the center axes of the burners or of the injection nozzles of the combustion chamber are arranged inclined at an angle ? of 30° to 40°, preferably 35°, relative to a meridional plane passing through the engine axis.

Abstract: A running-gap control system of an aircraft gas turbine with a core engine including a turbine whose blade rows have a running gap 13 to the turbine casing 12, with the turbine casing 12 being surrounded by a core-engine ventilation compartment 15 enclosed by a fairing 2 of the core engine, with the fairing 2 forming an inner wall of a bypass duct 1. Air from the bypass duct 1 can be introduced via at least one inlet nozzle 5 into a cooling-air distributor 7 by a control system 6, 10, 11 and the air is subsequently returned to the bypass flow 3.

Abstract: A gas-turbine combustion chamber arrangement includes at least one centrifugal compressor as well as one centripetal annular combustion chamber, with a stator vane arrangement being provided between the centrifugal compressor and the annular combustion chamber. The stator vane arrangement for diverting the air flowing out of the centrifugal compressor is designed at an angle ? of 20° to 30°, preferably 25°, relative to the engine axis, so the airflow is passed at essentially this angle ? to the combustion chamber. The inflow area into the combustion chamber for supplying air is designed at an angle of 20° to 30°, preferably 25°, relative to the meridional plane, and the center axes of the burners or of the injection nozzles of the combustion chamber are arranged inclined at an angle ? of 30° to 40°, preferably 35°, relative to a meridional plane passing through the engine axis.

Abstract: An aircraft gas turbine engine includes a core engine 1 having a high pressure turbine 6 and a downstream low pressure turbine 7. A bypass duct 18 surrounds the core engine 1. A mixer 19 is arranged in an inlet portion of the low pressure turbine 7, into which a bypass flow 20 from the bypass duct 18 and a core flow 22 from the core engine 1 are supplied.

Abstract: A turboprop propulsion unit includes at least one pusher propeller 5, 6 driven by an aircraft gas-turbine engine, with the aircraft gas-turbine engine being arranged in front of the pusher propeller 5, 6 in a direction of flight. A turbine outlet area 9 is arranged at the front in the direction of flight and a compressor area 14 faces towards the pusher propeller 5, 6.

Abstract: A gas-turbine combustion chamber arrangement includes a flame tube, a diffuser element arranged upstream of the flame tube in the flow direction, the diffuser element including an annular duct, and an axial compressor arranged upstream of the diffuser element. The diffuser element features an annular guide vane area in which guide vanes are arranged, which for redirecting an incoming flow are provided at an angle (?) in a range between 28° and 32° relative to a central axis of the gas turbine. Downstream of the guide vane area, a diffuser area is arranged, the diffuser area not being provided with flow-guiding elements affecting the flow, where burners arranged in the annular combustion chamber are provided with their burner axes at an angle (?) between 40° and 50° relative to the central axis.

Abstract: An aircraft propeller-engine layout includes at least one engine (10) arranged in a tail (2) of a fuselage (1) of an aircraft and has at least one propeller arrangement, with a shaft of the propeller (6, 7) being connected to a shaft (21) of the engine (10) via at least one transfer shaft (12).

Abstract: The air exit hole of the vent line (6) connected to a venting apparatus for the lubricating oil system of a jet engine is arranged behind the nozzle throat (2) on the periphery of the exiting engine jet (4). The exit hole is tangentially arranged on the periphery of the engine jet, or slightly enters the rim area of the engine jet. The vent line extends under the protection of an aerodynamically shaped fairing (5, 11). This arrangement of the air exit, while being simply designed, cost-effective and weight-saving, provides for clean, invisible discharge of air from the oil venting apparatus.

Abstract: The invention refers to an aircraft gas turbine engine including a core engine 1 comprising at least a high pressure turbine 6 and a downstream low pressure turbine 7, a bypass duct 18 surrounding the core engine 1, as well as a mixer 19 arranged in an inlet portion of the low pressure turbine 7, into which a bypass flow 20 from the bypass duct 18 and a core flow 22 from the core engine 1 are supplied.

Abstract: A turboprop propulsion unit includes at least one pusher propeller 5, 6 driven by an aircraft gas-turbine engine, with the aircraft gas-turbine engine being arranged in front of the pusher propeller 5, 6 in a direction of flight. A turbine outlet area 9 is arranged at the front in the direction of flight and a compressor area 14 faces towards the pusher propeller 5, 6.