Abstract: An aircraft gas-turbine engine has an engine cowling-nacelle which surrounds a core engine in a tubular way, characterized in that at least one inflow-side part of the engine cowling-nacelle is telescopically movable against the flow direction.

Abstract: A blocker door for a gas turbine engine thrust reverser having a tray with a base and sidewalk extending about the base to define a volume, the volume being dosed by a cover that extends beyond the periphery of the tray. The extension of the cover beyond the periphery provides a sealing feature.

Abstract: A gas turbine engine has a core engine and a nacelle having a cowl translatable along the axis of the engine from a stowed position to a deployed position, A thrust reverser cascade extends circumferentially about the engine axis. To obviate the need for additional blanking cascades the cowl has a forwardly extending tongue that in the deployed position of the cowl is at the same axial position as the thrust reverser cascade.

Abstract: An ejector nozzle tube, having along its length an essentially constant, essentially oval hollow cross-section, with a leading edge nozzle, having a rounded exterior cross-section, being arranged at a flow leading-edge area of the ejector nozzle tube. The leading edge nozzle includes a plurality of grooves issuing to the top and bottom sides of the ejector nozzle tube and connecting to an interior of the ejector nozzle tube formed by the hollow profile. An aircraft engine has an optimized oil heat exchanger with at least one oil cooler disposed in a trailing-edge area of an aerofoil-type structure, with at least one flow entrance area being provided to supply ambient air to the oil cooler disposed in a flow duct. The ejector nozzle tube can be arranged downstream of the oil cooler in the flow duct.

Abstract: The present invention relates to a de-icing device of an aircraft gas-turbine engine with an engine cowling enclosing at least one inflow region, with the engine cowling having a double-walled design and including at least one annular tube element extending in the circumferential direction and being provided with outlet openings for passing hot air to an inflow region, in order to de-ice it, with the tube element in the circumferential direction including multiple individual tube segments which can be attached relative to one another.

Abstract: The present invention relates to an aircraft gas turbine thrust-reversing device with an engine having an engine cowling, the rear area of which can be displaced in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position, resulting in an essentially annular free space towards a forward stationary area of the engine cowling, with the rear area of the engine cowling being functionally coupled to deflecting elements arranged in the forward thrust position within the front area of the engine cowling, with the deflecting elements during displacement of the rear area of the engine cowling being moveable on a partial-circular path facing the central axis of the engine and contactable by their rear end areas with a cowling of the core engine.

Abstract: The present invention relates to a method for manufacturing a body provided with a honeycomb structure, which body is designed preferably axis-symmetrical, with sheet-metal parts being provided with a wave-like structuring by means of a forming process and formed to ring elements, with the sheet-metal parts being shaped and designed in a substantially identical way as regards their structuring, with adjacent sheet-metal parts being arranged offset to one another and then joined together to form the honeycomb structure.

Abstract: A method for manufacturing a sound absorber, the outer wall of which is provided with a plurality of recesses, with funnel-like cone elements each being assigned to the recesses inside the sound absorber, said cone elements having a larger opening facing radially outwards and a smaller opening facing radially inwards, with adjacent cone elements each being provided on a strip-shaped first carrier band, with cup elements being provided radially on the inside relative to the cone elements, and each cup element receiving one cone element, with adjacent cup elements each being provided on a strip-shaped first carrier band, with the first carrier bands being arranged adjacently in a first direction, and the second carrier bands being arranged adjacently in a second direction, with the directions crossing each other and the carrier bands being joined to one another and to the outer wall to form a rigid body.

Abstract: The present invention relates to a gas-turbine exhaust cone having an outer wall, which is provided with a plurality of recesses, a honeycomb-structured layer, which is arranged on the inside of the outer wall and extends along the inside of the outer wall, an inner wall, which extends substantially parallel to the outer wall and is connected to the honeycomb structure, and at least one annular chamber, which adjoins the inner wall and is centered relative to a central axis, with the inner wall being provided with passage recesses 33-connecting the area of the honeycomb structure to the annular chamber.

Abstract: An arrangement (1) for suspending a jet engine (12) to a supporting structure (5) attached to an aircraft fuselage, includes a suspension beam (6) to which the engine casing (13) is fitted by a forward suspension (7), a rearward suspension (8) and at least one thrust rod (9) arranged at an angle relative to the suspension beam (6). In order to provide an arrangement for the suspension of a jet engine (12) which enables weight to be saved and deformation of the jet engine (12) to be reduced, the rearward suspension (8) and the thrust rod (9) are connected to the engine casing (13) by a lightweight framework (11).

Abstract: The present invention relates to an aircraft gas turbine having a fan rotatable about an engine axis in the inflow region of the aircraft gas turbine, with the fan having several fan blades, with each fan blade being moveably mounted on a hub rotatable about the engine axis by means of its blade root on an area which is at the front in the flow direction, and with each fan blade being moveably mounted on an area at the rear in the flow direction, on an adjusting disk axially displaceable relative to the engine axis and non-rotatable with the hub, with the pitch angle of the fan blades being variable by the axial movement of the adjusting disk.

Abstract: A gas-turbine exhaust cone has an outer wall with a plurality of recesses, a honeycomb-structured layer arranged at the inside of the outer wall and extending along said inside of the outer wall, an inner wall connected to the honeycomb structure and extending essentially parallel to the outer wall, and at least one annular chamber centered on a central axis and adjoining the inner wall. The inner wall has passage recesses connecting the area of the honeycomb structure to the annular chamber, with the annular chamber being subdivided in the circumferential direction by at least one partition wall into several chambers. The partition wall is made from a sheetmetal-like material and has a plurality of raised and/or recessed areas in a uniform arrangement formed by shaping of the sheetmetal-like material.

Abstract: An aircraft de-icing device and an engine nacelle of an aircraft gas turbine is provided with a de-icing device, with the engine nacelle 1 in the inflow area being provided with an annular air duct 2, into which heated air is introduced. At least one oil cooler 3 is arranged in the engine nacelle 1, to which engine oil is fed and which is supplied with ambient air, with the heated airflow leaving the oil cooler 3 being passed through the air duct 2.

Abstract: A scoop for a fairing 1 of a core engine 2 of an aircraft gas turbine allows air to be supplied from a bypass flow in a bypass duct 3 to several cooling-air distributors in a core-engine ventilation compartment 4. The scoop includes a first tubular flow duct 5, whose inlet opening 6 is arranged in the bypass duct 3 and which extends through the fairing 1, as well as a second tubular flow duct 7, which at least partly encompasses the first flow duct 5 and whose inlet opening 8 is rearwardly offset relative to the inlet opening 6 of the first flow duct 5 in the direction of flow.

Abstract: A gas-turbine exhaust cone has an outer wall with a plurality of recesses, a honeycomb-structured layer arranged at the inside of the outer wall and extending along said inside of the outer wall, an inner wall connected to the honeycomb structure and extending essentially parallel to the outer wall, and at least one annular chamber centered on a central axis and adjoining the inner wall. The inner wall has passage recesses connecting the area of the honeycomb structure to the annular chamber, with the annular chamber being subdivided in the circumferential direction by at least one partition wall into several chambers. The partition wall is made from a sheetmetal-like material and has a plurality of raised and/or recessed areas in a uniform arrangement formed by shaping of the sheetmetal-like material.

Abstract: The present invention relates to a gas-turbine aircraft engine with a core engine surrounded by a bypass duct, with a substantially annular cooling element being provided in the area of a radially inner wall of the bypass duct downstream of stator vanes of the bypass duct.

Abstract: The present invention relates to an aircraft gas turbine with a core engine 10 which is surrounded by a bypass duct 29 enclosed radially outwards by a bypass wall 30, with a radially inner wall 31 of the bypass duct 29 forming with the radially outer bypass wall 30 a bypass nozzle 32, and with the radially inner wall 31 including an adjusting element 33 extending along the circumference of the inner wall 31 and being deformable radially outwards.

Abstract: The present invention relates to a de-icing device of an aircraft gas-turbine engine with an engine cowling enclosing at least one inflow region, with the engine cowling having a double-walled design and including at least one annular tube element extending in the circumferential direction and being provided with outlet openings for passing hot air to an inflow region, in order to de-ice it, with the tube element in the circumferential direction including multiple individual tube segments which can be attached relative to one another.

Abstract: The present invention relates to an aircraft gas-turbine engine with a core engine surrounded by a bypass duct, with a radially outer engine cowling enclosing the bypass duct and being provided at its rear region with a thrust-reversing device which is moveable relative to the engine cowling, with at least one cooler element extending over at least part of the circumference being arranged in the intermediate area between the engine cowling and the thrust-reversing device.

Abstract: A gas-turbine exhaust cone includes a cone shaped outside cone 1 closed in a flow direction, an inside cone 2 arranged in the outside cone 1 over at least part of the length of the exhaust cone, with the inside cone 2 being connected to the outside cone 1 via at least one helical partition wall 3, thereby forming an interspace 4 between the outside cone 1 and the inside cone 2. The interspace 4 is connected to inlet openings 5 at its inflow side and opens at an exit side 6 towards an interior 7 of the outside cone 1. A dampening element 8 is arranged in the outside cone 1 adjacent to the exit side 6 of the interspace 4.