Abstract: An articulating arrangement for a pipe includes a narrow section having a diameter D1 smaller than a diameter D2; a plurality of slits arranged on the narrow section in a spaced apart manner along a central axis of the pipe, each of the plurality of slits including two or more slit sections separated by two or more bridge sections. The plurality of slits includes a set of first slits and a set of second slits angularly arranged, with respect to the set of first slits, along the central axis. The set of first slits is arranged on a first half of the narrow section and the set of second slits is arranged on a second half of the narrow section. A cover member is arranged on the narrow section to seal the plurality of slits. The cover member diametrically aligns with the diameter D2.

Abstract: A pressure-release device has a pressure-release flap arranged in a delimiting wall of a chamber of an aircraft engine which opens in the presence of a predefined pressure in the chamber. The pressure-release flap includes a flap region and a hinge. The flap region is defined by a perforation formed in the delimiting wall which delimits the flap region with respect to a surrounding region of the delimiting wall. The perforation is formed such that, in the presence of the predefined pressure in the chamber, the flap region pivots open relative to the surrounding region, with the perforation rupturing, and the hinge is fixedly connected to the surrounding region and the flap region, such that the hinge holds the flap region on the surrounding region in the event of a rupture of the perforation and a pivoting-open of the flap region.

Abstract: The invention relates to a thrust nozzle for a turbofan engine of a supersonic aircraft, wherein the thrust nozzle includes a thrust nozzle wall, and a flow channel that is delimited radially outwards by the thrust nozzle wall, wherein the flow channel has a nozzle throat surface and a central body that is arranged in a flow channel. According to the invention, the central body forms a bypass channel, which extends within the central body, and which is designed for the gas of the flow channels to flow through.

Abstract: A convergent-divergent nozzle for a turbofan engine of a supersonic aircraft, wherein the nozzle has an inner wall and forms a flow channel with a nozzle throat surface and a nozzle exit surface. A trim ring that is displaceable in the axial direction between a first position and a second position is provided, wherein the inner wall of the nozzle and the trim ring are embodied and positioned with respect to each other in such a manner that the trim ring extends at a radial distance to the inner wall in the first position, wherein a ring-shaped bypass channel is provided, extending between the trim ring and the inner wall of the nozzle, and the trim ring abuts the inner wall in the second position.

Abstract: A guide vane of a gas turbine engine, in particular of an aircraft engine, which has a pressure-side wall, a suction-side wall, a guide vane root, a guide vane tip, a guide vane leading edge area that is impinged by a cooling air flow of a cooling system, a guide vane trailing edge area that is facing away from the guide vane leading edge area, and at least one channel for conducting a fluid to be cooled arranged in an internal space of the guide vane. At that, during operation of the gas turbine engine, a first part of the cooling air flow flows around a pressure-side wall, and a second part of the cooling air flow flows around the suction-side wall, and a third part of the cooling air flow flows through the internal space including the channel. What is further suggested is a gas turbine engine with at least one such static guide vane.

Abstract: An inlet for a gas turbine engine having an inlet lip and an inlet barrel, the inlet lip having a highlight defining a boundary between inner and outer surfaces of the inlet lip. The inlet barrel has a forwardly extending flange that is joined to the inlet lip by overlapping surfaces. The flange interface surface has a conus angle that is divergent from the highlight axis.

Abstract: A jet engine having a flow duct that is arranged in an engine nacelle and is delimited radially on the inside by a central body and radially on the outside by an outer nozzle wall is described. The free cross section of the flow duct is variable by an elastically embodied, radially variable wall of at least one of these components. The radially variable wall at least regionally has an at least approximately wavelike structure and is connected to at least one adjusting device having an adjusting unit. The variable wall of one of these components is variable in a radial direction at two defined control sections, at least one of which defines an axial wall end of the radially variable wall.

Abstract: The invention relates to an engine cowling of an aircraft gas turbine which has a rear area that comprises at least two doors that are arranged at the bottom area of the rear area of the engine cowling and are to be opened by means of overhead hinges, wherein the hinges are arranged so as to be tilted with respect to the engine axis.

Abstract: A device of a turbomachine actuates a setting device to vary a flow cross-section of a flow channel of the turbomachine passable by a fluid flow. The device includes a displacement mechanism having an adjustable actuation appliance and couplable with the setting device, and a drive device for displacing the actuation appliance. The displacement mechanism has a centrifugal force appliance displaceable between a basic position and a maximally displaced working position depending on a number of revolutions of the drive device, wherein the actuation appliance is displaceable depending on the position of the centrifugal force appliance with respect to the drive device. A reset device applies a force to the centrifugal force appliance by which the centrifugal force appliance is pressed in the direction of its basic position.

Abstract: An intake for channeling air flowing past a propeller to an inlet of an aircraft engine that drives the propeller with a drive shaft, the intake including: a static cowling that extends along an axis and that flares outward at an upstream end of the static cowling, and an intake slot that is formed in the static cowling. The intake slot connecting to a passage of the inlet of the aircraft engine, the intake slot including an arched opening that extends less than 360 degrees of a circumference of the static cowling, and the intake slot having a downstream lip with a curved profile that blends into the static cowling.

Abstract: An engine cowling of an aircraft gas turbine with a front area and a rear area that is displaceable in the axial direction, characterized in that the rear area is mounted and supported by means of multiple length-adjustable thrust reversal actuators that are arranged at an angle to one another.

Abstract: A structural element of an aircraft part, in particular an aircraft engine part, with at least partly a double-curvature shape, including a plurality of sets of fibers in textile fabric structure, wherein in at least one region of the structural element the number of fibers in one direction is reduced per area in a flattened out state of the textile fabric structure. It also relates to a method for manufacturing a structural element.

Abstract: An engine cowling of an aircraft gas turbine with a front area and a rear area that is displaceable in the axial direction, characterized in that the rear area is mounted and supported by means of multiple length-adjustable thrust reversal actuators that are arranged at an angle to one another.

Abstract: An engine cowling of an aircraft gas-turbine engine with a thrust-reversing device, with a core engine and a bypass duct surrounding the latter, and with a front cowling enclosing the bypass duct and a rear cowling movable in the axial direction, as well as with cascade elements arranged over the circumference which can be placed by a substantially axial movement into an intermediate space resulting from a movement of the rear cowling relative to the front cowling, characterized in that the cascade elements are connected directly to the rear cowling, that the rear cowling has a rounded outer leading edge and that the cascade element has at its rear area a flow-guiding element, in front of which several thrust-reversing stator vanes are arranged and behind which at least one flow stator vane is arranged whose curvature shape matches the rounding of the outer leading edge of the rear cowling.

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: An engine cowling of an aircraft gas-turbine engine with a core engine and a bypass duct surrounding the latter, with a front cowling enclosing the bypass duct and a rear cowling movable in the axial direction, and with stator vanes arranged in the bypass duct, where recesses for removing fluid from the bypass duct are provided in the area of the stator vanes on the inside of the front cowling, where the fluid discharged through the recesses is routed by means of flow ducts through the front cowling, brought into contact with at least one heat exchanger, and subsequently discharged to the environment.

Abstract: An engine cowling of an aircraft gas turbine having a front cowling and a single rear cowling movable in the axial direction. By a drive device, the rear cowling is movable into different positions including a closed cruise position, a partially opened maximum thrust position and a completely opened thrust-reversing position. The drive device includes at least two actuators arranged in series relative to one another and actuated independently of one another, the actuators having power supply systems separated from one another and separately operating, where one of the actuators is designed for the maximum thrust position and the other actuator for the thrust-reversing position. The rear cowling can be locked in the maximum thrust position or the thrust-reversing position by two independent locking devices.

Abstract: A convergent-divergent nozzle for a turbofan engine of a supersonic aircraft, wherein the nozzle has an inner wall and forms a flow channel with a nozzle throat surface and a nozzle exit surface. A trim ring that is displaceable in the axial direction between a first position and a second position is provided, wherein the inner wall of the nozzle and the trim ring are embodied and positioned with respect to each other in such a manner that the trim ring extends at a radial distance to the inner wall in the first position, wherein a ring-shaped bypass channel is provided, extending between the trim ring and the inner wall of the nozzle, and the trim ring abuts the inner wall in the second position.

Abstract: A convergent-divergent nozzle for a turbofan engine of a supersonic aircraft, wherein the nozzle has an inner wall that delimits a flow channel through the nozzle radially outside, wherein the flow channel has a nozzle throat surface and a nozzle exit surface. The inner wall includes a first group of adjustable segments forming an upstream convergent area of the nozzle, and second group of adjustable segments forming a downstream constant/divergent area of the nozzle. It is provided that the segments of the first group or the segments of the second group are curved towards the flow channel in a convex manner at least in an area that adjoins the other group, forming the nozzle throat surface in the area of the convex curvature and adjoining the segments of the respectively other group at an axial distance to the axial position of the nozzle throat surface.

Abstract: A joint assembly between a first gas turbine engine casing and a second gas turbine casing, the first gas turbine engine casing having a plurality of circumferentially spaced attachment points and the second gas turbine casing having a plurality of circumferentially spaced attachment points for securing the first gas turbine casing and the second gas turbine casing together, wherein each attachment point on the first gas turbine casing is circumferentially offset from a respective attachment point on the second gas turbine casing, and connected by a flexure member that permits limited relative movement of the attachment point on the first gas turbine engine casing and its respective attachment point on the second gas turbine casing.