Abstract: The disclosure relates to an electrical drive device having: an inverter including an inverter unit for each phase; a control unit configured to control the inverter units by application of vector control; and a rotating electrical machine having a stator that includes a plurality of phase windings connected to the inverter units. Each of the phase windings includes a first part-winding and an electrically isolated second part-winding. The inverter units include a first phase module and a second phase module. The phase modules deliver the electrical phase assigned to the respective inverter unit in a separate and a mutually electrically isolated manner. The first part-winding is electrically connected to the first phase module and the second part-winding is electrically connected to the second phase module.

Abstract: A gas turbine engine for an aircraft including a unit supplied with oil from a first oil circuit and a second oil circuit. The first oil circuit and the second oil circuit each are fluidly coupled with at least one inlet and with at least one outlet of the unit and with at least one inlet and with at least one outlet of an oil tank. The first oil circuit and the second oil circuit are configured to receive oil from the oil tank and to direct the received oil to the unit. The oil tank is incorporating offset outlets to each of the oil circuits. The offset outlet of the second oil circuit is positioned higher in the oil tank than the offset outlet of the first oil circuit.

Abstract: An exhaust nozzle of a gas turbine engine which includes a nozzle wall, the nozzle wall including an upstream, fixed structure and a downstream, translatable structure that is translatable relative to the fixed structure, a flow channel which is limited radially outwards by the nozzle wall, a centerbody arranged in the flow channel, at least one strut connecting the centerbody to the nozzle wall, a thrust reverser unit that comprises blocking doors, a first actuation system for deployment of the blocking doors into a deployed position for thrust reversal, and a second actuation system for translating the translatable structure between a stowed, upstream position and a fully deployed, downstream position. It is provided that the at least one strut is connected to the fixed structure, that the blocking doors are connected to the translatable structure, and that the first actuation system and the second actuation system are independent actuation systems.

Abstract: A bearing assembly for a gas turbine engine comprises a bearing; a bearing bracket, which holds the bearing and is secured by a predetermined breaking device on a connecting element, which can be connected or is connected to a support structure of the gas turbine engine; and a clutch for transmitting a torque from a first clutch element connected in a fixed manner to the rotor of the bearing to a second clutch element supported on the bearing bracket, wherein the clutch elements are spaced apart when the predetermined breaking device is intact and can be brought into contact with one another by destruction of the predetermined breaking device. A gas turbine engine and a method are furthermore provided.

Abstract: A combustion chamber assembly for an engine with at least one tile component, on the cold side of which facing away from a combustion space, a positioning aid with at least two positioning elements, wherein a defined position relative to the combustion chamber component is predefined for the tile component via the contact of the at least two positioning elements on at least one contact face of the combustion chamber component in two spatial axes.

Abstract: An exhaust nozzle of a gas turbine engine which includes an outer nozzle wall, a flow channel which is limited radially outwards by the nozzle wall, a centerbody arranged in the flow channel, and exactly one strut connecting the centerbody to the nozzle wall. The strut is connected to the nozzle wall by means of a connecting structure that is displaceable in the axial direction of the outer nozzle wall. At least one actuator is provided interacting with the connecting structure or the outer nozzle wall for displacing the strut in the axial direction.

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 gearbox assembly for a gas turbine engine comprises a planetary gearbox having at least one ring gear and at least one planet gear, which exerts a force on the ring gear in the direction of a force vector as it rolls on said ring gear; and a holding device for fastening the at least one ring gear on another structure, having a first section, which extends in the axial direction on one side of the force vector and/or of a straight-line extension thereof, and having a second section, which extends in the axial direction on the other side of the force vector and/or of the straight-line extension thereof. A gas turbine engine and a method for producing a gearbox assembly are furthermore made available.

Abstract: The proposed solution relates to a stator vane assembly for an engine, having at least one stator vane of a stator vane row and a casing for the at least one stator vane row, wherein the at least one stator vane is mounted adjustably on the casing by means of a bearing journal, which is rotatably mounted in a bearing opening in the casing and passes through this bearing opening along a longitudinal axis. A section of a sealing element, on which at least one sealing ridge extending radially in relation to the longitudinal axis is formed and/or which has a nonlinear slot passing longitudinally through the section, is provided within the bearing opening.

Abstract: A gas turbine engine for an aircraft includes an engine core, a fan and a gearbox which receives an input from a turbine shaft and which outputs drive for the fan for driving the fan at a lower rotational speed than the turbine shaft. The gearbox has a sun gear which is driven by the turbine shaft. The turbine shaft and the sun gear form a spline connection which is lubricated with oil. The turbine shaft has, on its radially inner side, at least one indentation which is suitable for receiving oil, and the turbine shaft has at least one bore which extends from the indentation to the radially outer side of the turbine shaft and which is suitable for transporting oil from the at least one indentation to the radially outer side of the turbine shaft as the turbine shaft rotates.

Abstract: A ventilation device for controlling an air flow through at least one air inlet into an interior space of a gas turbine engine, comprises: at least one actuable closure device for the at least one air inlet, which has an open position for exposing the air inlet and can be transferred by actuation into a closed position for closing the air inlet; and a fire extinguishing device having an extinguishing agent arranged in a reservoir, wherein the fire extinguishing device is connected or can be connected to the interior space and the closure device in such a way that activation of the fire extinguishing device both leads to admission of extinguishing agent from the reservoir into the interior space and brings about actuation of the closure device to transfer the closure device into the closed position.

Abstract: A planetary gear with a sun wheel, a hollow wheel and a planetary carrier on which a planetary wheel is rotatably mounted. In the axial direction on a first side of the planetary carrier, the sun wheel and hollow wheel include connection areas for coupling the sun wheel and hollow wheel to rotatable or torque-proof areas of an engine. The planetary carrier has a connection area for attaching to rotatable or torque-proof areas of the engine on its opposite second side. The structural component stiffnesses of the sun wheel, planetary carrier, the hollow wheel and the planetary wheel are adjusted to each other such that, during operation they have twistings in the axial direction of the planetary gear that respectively qualitatively correspond to each other between the connection areas and side areas facing away from the connection areas due to the respectively applied torques.

Abstract: An engine nacelle for a turbofan engine, having a nacelle wall and an engine inlet, wherein the nacelle wall has a stationary downstream section and an upstream section that is displaceable in the axial direction, and the displaceable upstream section is displaceable between a first upstream position and a second downstream position. A ring-shaped seal is provided, formed between the stationary section and the displaceable section of the nacelle wall, wherein, in the first position of the displaceable section, the engine nacelle forms a ring-shaped additional flow channel that extends from the outer side of the nacelle wall to the engine inlet and via which ambient air can flow into the engine inlet, and the ring-shaped additional flow channel is closed in the second position of the displaceable section, wherein the stationary section, the ring-shaped seal and the displaceable section adjoin each other.

Abstract: A combustion chamber assembly for an engine includes a wall element fixed to a combustion chamber structure and a chamber between the wall element and the structure, the chamber being supplied with air through impingement-cooling openings in the structure and connected to the combustion space by film-cooling openings in the wall element. Two cooling-air holes formed in the structure generate a cooling-air flow toward the combustion space and past the wall element. The wall element has a flow guide device for generating at least one scavenging-air flow directed between the two cooling-air holes. A sum of flow cross sections of film-cooling openings in the wall element and of the flow guide device yields a larger area than a sum of flow cross sections of all wall element impingement-cooling openings via which air is guided through the structure into the chamber and to the rear side of the wall element.

Abstract: A gas turbine engine, includes: an engine core including a turbine, compressor, and shaft system connecting the turbine to the compressor, and forming a torque path therebetween. The shaft system is axially located by a thrust bearing located forward of the turbine, and the engine is configured, in the event of a shaft break which divides the shaft system into a front portion located by the thrust bearing and a rear portion unlocated by the thrust bearing, the rear portion is free to move axially rearwardly under a gas load. The engine further includes a shaft break detector having a forward speed sensor configured to measure a rotational speed of the front portion of the shaft system, and a rear microwave sensor configured to measure a rotational speed of the rear portion of the shaft system, wherein a shaft break can be detected based on differences in the measured speeds.

Abstract: A method for in-flight monitoring of the health of an aero gas turbine engine including in flow series a compressor system, a core gas generator including a combustor, and a turbine, and further including a shaft system connecting the turbine to the compressor system and forming a torque path there between. The method includes: measuring a first rotational speed of a forward portion of the shaft system; measuring a second rotational speed of a rear portion of the shaft system; measuring other operational parameters of the engine; calculating from the measured rotational speeds the power delivered by the torque path from the turbine to the compressor system; calculating from the other measured operational parameters the power delivered to the turbine by the core gas generator; correlating the power delivered by the torque path at a given time with the power delivered to the turbine at that time.

Abstract: A gas turbine engine, in particular an aircraft engine, including: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.

Abstract: Disclosed is a seal for an engine firewall. The seal including a proximal part having a recess; a distal part having a notch; and a flexible intermediate part integrally extending between the proximal part and the distal part.

Abstract: A structural subassembly for a compressor of a turbomachine, which has a rotor having a plurality of blades, which extend radially in a flow path of the turbomachine, and a compressor casing, which forms a flow path boundary, which delimits the flow path through the turbomachine radially on the outside. In this subassembly, the compressor casing has casing structuring adjoining the rotor. It is envisaged that the compressor casing has a plurality of circumferential segments, which extend in the circumferential direction, wherein only one or only some of the circumferential segments forms or form casing structuring.

Abstract: A planetary gearbox, a splined sleeve, a gas turbine engine with a planetary gearbox and a method for producing a torque-proof connection between a planetary carrier and a support element. The splined sleeve is formed with an outer substantially cylindrical shell surface area and an inner conical shell surface area, and is inserted between the planetary carrier and the support element for creating the torque-proof connection. The diameter of the inner conical shell surface area decreases in the axial direction, starting from an end of the splined sleeve in the direction of a second end of the splined sleeve. The outer cylindrical shell surface area is formed in a stepped manner. The outer cylindrical shell surface area comprises a first cylindrical section and a second cylindrical section which has a diameter larger than the diameter of the first cylindrical section. Between the sections, a conical area is provided.