Abstract: The invention regards a gas turbine engine control system and a method for limiting turbine overspeed in case of a shaft failure. The control system includes: an overspeed protection system that activates an activation member in case a shaft failure is detected; a fuel limiting mechanism coupled with the activation member, wherein the fuel limiting mechanism is configured to limit the fuel supply to the gas turbine engine combustor if the activation member is activated; a variable stator vane mechanism which is configured to adjust variable stator vanes of a compressor of the gas turbine engine in their rotational position, the variable stator vanes having a closed position which blocks air flow through the compressor.

Abstract: An epicyclic gear system for a gas turbine engine includes a planet carrier with at least one structural member, on which a planet gear is pivot-mounted by a bearing that is radially arranged between the planet gear and the structural member. Furthermore, a roller bearing device is arranged radially between the planet gear and the bearing. A relative movement between the planet gear and the bearing is prevented by the roller bearing device, if a friction torque in the region of the bearing is less than or equal to a threshold value.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: The production of engine parts with a complex geometrical structure. More particularly, a method for producing a complex part, comprising making available a first component, having a thermal expansion coefficient of the first component; a first joining surface; and a first bearing surface; making available a second component, having a thermal expansion coefficient of the second component; a second joining surface; a second bearing surface; and making available a jacket element, having a thermal expansion coefficient of the jacket element; and a jacket-element bearing surface; and heating the first component, the second component and the jacket element from a first temperature to a second temperature in order to carry out a joining process on the first component and the second component. Furthermore, a part, in particular for a gas turbine engine for an aircraft, and to a gas turbine engine of this kind.

Abstract: The proposed solution concerns in particular an engine, with at least two valve devices for controlling bleed air in dependence on an operating state of the engine, the at least two valve devices in each case being adjustable between an open position and a closed position. A flow space is provided for at least one fluid flow occurring in the flow space in dependence on a position of the at least two valve devices. Arranged in the flow space is at least one fluid sensor, by means of which at least one position of at least one of the at least two valve devices can be determined on the basis of the fluid in the flow space.

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: A blade wheel of a turbomachine, which blade wheel has a multiplicity of blades which are suitable and provided for extending radially in a flow path of the turbomachine, wherein the blades form a blade entry angle and a blade exit angle. Provision is made whereby the blade wheel forms N blocks of blades, where N?2, wherein the blades of a block have in each case the same blade entry angle and the same blade exit angle, and the blades of at least two mutually adjacent blocks have a different blade entry angle and/or a different blade exit angle. According to a further aspect of the invention, partial gaps that the blades form in relation to an adjacent flow path boundary are varied in mutually adjacent blocks.

Abstract: A rotor for an electric machine is provided. Permanent magnets are disposed on the rotor such that the permanent magnets cannot carry out any tangential movement. The rotor may have four shoulders that extend away from a shell surface of the rotor in a radial direction. Primary permanent magnets that are disposed so as to be mutually contiguous and contiguous to the respective shoulders and accordingly cannot be displaced are in each case located between the shoulders. Secondary permanent magnets are disposed on the shoulders such that an annulus of magnets that is free of gaps and voids and is not movable in the tangential direction is formed.

Abstract: The present invention provides a rotor, an axial flux electrical machine and a hybrid-electrical or electrical air craft. The rotor for the axial flux electrical machine comprises first sections of a first material and second sections of a second material. The first sections and second sections are arranged in alternating order and ring-shape. The arrangement is characterized by third sections which form interface areas between the first sections and second sections, whereas the third sections are comprising the first material and the second material in such a manner that the first section and the second section are connected force-fitted.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, which includes a primary combustion zone with a primary combustion zone depth and a secondary combustion zone. A ratio of the primary combustion zone depth to the lean burn fuel injector head tip diameter is less than 2.4.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, which includes a primary combustion zone with a primary combustion zone depth and a secondary combustion zone. A ratio of the primary combustion zone depth to the lean burn fuel injector head tip diameter is less than 2.4.

Abstract: The proposed solution concerns in particular an engine assembly with a casing part extending along a central axis, at least one cover part which at least partially covers an outside of the casing part, and for this, in relation to the central axis, extends along a circumferential direction which points along a circular path around the central axis, and at least one connecting element via which the at least one cover part is fixed relative to the casing part and which compensates for different thermal expansions of the casing part and the at least one cover part.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: A fuel injector including: a plurality of air swirler passages; at least one fuel supply passage arranged to supply fuel into at least one of the air swirler passages; and at least one cavity separating an exterior of the fuel supply passage from a body of the fuel injector; wherein the cavity is at least partially filled with a thermally insulating material.

Abstract: The disclosure relates to an electric machine and in particular to the monitoring of the electric machine for the presence of a fault, (e.g., in the stator windings). A monitoring unit is provided, wherein the monitoring unit measures the multiphase electrical time signals transmitted from or to the machine and with the aid of a Hilbert filter determines substantially in real time the envelopes and the phase positions of the individual phases of the time signal. The envelopes corresponding to the different phases or the corresponding phase positions are compared with one another by way of forming differences and, in the event that one or more of the differences deviate(s) from a specified expectation value, the presence of a fault is inferred. The approach allows significantly increased operational reliability of the electric machine to be achieved in particular.

Abstract: Aspects of the disclosure regard an exhaust nozzle of a gas turbine engine which includes an outer nozzle wall, a flow channel, a centerbody arranged in the flow channel, at least two struts connecting the centerbody to the nozzle wall, a thrust reverser unit, and a plurality of actuators, wherein each actuator is associated with a strut for displacing the strut in the axial direction. The struts are connected to a structure of the outer nozzle wall that forms part of the thrust reverser unit. For allowing axial relative movement between the struts and the outer nozzle wall, each strut includes a sliding element extending radially from the radial outer end of the strut, wherein the sliding element is arranged in a receiving slot that extends in the axial direction in the nozzle wall. It is provided that the sliding element comprises an interaction structure that interacts with one of the actuators for effecting relative axial movement between the strut and the nozzle wall.

Abstract: A gearbox for a gas turbine engine with a rotatably mounted structural component that is formed with an at least approximately rotationally symmetrical duct into which, starting from its radially inner area, oil from an oil supply which is fixedly attached at the housing can be introduced. Starting from the oil supply, an introduction direction of the oil into the duct encloses an angle of 90° or an angle of between 60° and 90° with the axial extension direction of the duct. Additionally, in the circumferential direction of the duct, the introduction direction of the oil encloses an angle with the radial extension direction that is larger than or equal to 0° and smaller than 90°.

Abstract: An exhaust nozzle of a gas turbine engine includes: a nozzle wall, a centerbody arranged in a flow channel, and two struts connecting the centerbody to the wall. One of the struts is connected to the wall by a coupling arrangement that includes two first brackets and a third bracket, the brackets being spaced in an axial direction and being connected either directly to the wall or to a sliding element that is arranged in a displaceable manner in the wall. The brackets each have a first, highest stiffness in a first direction and smaller stiffnesses in a second and third direction. The brackets are oriented such that with the two first brackets, the first direction is aligned with a circumferential direction of the nozzle and that with the third bracket, the first direction is aligned with the axial direction of the nozzle.