Abstract: A turboprop aircraft engine includes at least one outer air intake 1 and at least one downstream inner duct system 2 for the delivery of air to a compressor, and the inner duct system 2 spirals around the center axis 3 of the engine.

Abstract: Turbojet powerplant with at least one compressor (1), at least one combustion chamber (2), a high-pressure turbine (3) and a low-pressure turbine (4), characterised in that a heat exchanger (5) is arranged between the compressor (1) and the combustion chamber (2), in that at least one hot-gas line (6) branches off from an area downstream of the high-pressure turbine (3) and is connected to the heat exchanger (5), and in that at least one cold-gas line (7) connects the heat exchanger (5) with an area upstream of the low-pressure turbine (4).

Abstract: The exchange, removal or addition, as applicable, of media and/or power is provided between the individual shafts of an engine, between individual engines and between the engines and the aircraft. Thus, additional degrees of freedom are provided, enabling engine parameters to be addressed in terms of a reduction or avoidance of negative resonances or beats. It also provides an ability to alter thrust from engines of a multi-engined aircraft to reduce rudder trim.

Abstract: The present invention relates to a combustion chamber head for a gas turbine with at least one combustion chamber wall 2,3, a metering panel 1, at least one heat shield 4, at least one sleeve 5 and a cowling 6, characterized in that the metering panel 1 forms one part with the combustion chamber outer wall 2 and the combustion chamber inner wall 3, in that the heat shield 4 is mounted onto the metering panel 1 from the downstream side of the metering panel 1, in that the sleeve 5 is mounted to the metering panel 1 from the upstream side, and in that the cowling 6 is separably attached to the metering panel 1.

Abstract: This invention relates to a method and an apparatus for the cooling of the casing (1) of the turbines (2, 3) of jet engines, in which cooling air is diverted from a bypass flow and supplied to the outer side of the casing via an inlet duct (8) provided with a shut-off element (5). According to the present invention, the cooling air is supplied to a first chamber (6) in which it is divided by volume. One portion of the cooling air is issued to the casing (1) via orifice holes (10), while another portion is ducted via several tubes (9) to a second chamber (7) which annularly encloses the casing (1) in the area of a low-pressure turbine (3).

Abstract: A pre-swirl nozzle carrier for an aircraft gas turbine with a body (1) is provided with a fitted seat (2) for the location of the pre-swirl nozzle carrier on a main carrier (3) and which features several pre-swirl nozzles (4), characterized in that the body (1) is extended radially inward, that it has several secondary pre-swirl nozzles (6) on its radial inward area (5) and that it is provided with an inner seal area (7) on its inner radial side.

Abstract: This invention relates to a bleed valve of a compressor, in particular a compressor of a bypass aero-engine, with a guiding device provided downstream of a valve body in the bleed duct, said guiding device directing the bleed airflow from a compressor duct to a bypass duct in such a manner that the direction of the flow of the bleed airflow is imparted with a component which is unidirectional with the airflow carried in the bypass duct. The present invention provides a dissipation screen in the bleed duct between the valve body and the guiding device to reduce the energy contained in the bleed airflow, thereby decreasing the loading of the bypass duct. Moreover, the present invention provides for an attenuation chamber downstream of the dissipation screen as viewed in the direction of the bleed airflow.

Abstract: The invention relates to lobe-type mixer tube for a bypass jet engine in which a bypass flow delivered by a fan is mixed with a hot gas stream issuing from the core engine, said lobe-type mixer tube having essentially circumferentially equally spaced gutter-shaped, radially moderately alternately inward and outward directed ducts to guide both the bypass flow and the hot gas stream, said ducts each having an end formed by an essentially plane extreme section and the respective extreme sections of some ducts when viewed in a longitudinal section extending in the direction of flow being inclined vis-a-vis the extreme sections of the other ducts. In this arrangement the radially outward directed ducts, when viewed circumferentially, alternately issue at points offset one from the other.

Abstract: This invention relates to a hydraulic seal arrangement between two, more particularly co-rotating shafts of especially a gas turbine engine, where the radially outer shaft when viewed with reference to the shafts' axis of rotation has an annulus extending radially outward along its circumference into which the radially inner shaft when viewed with reference to the shafts' axis of rotation projects with a web that extends radially outward across its circumference, and which annulus can through an inlet area be filled siphon-fashion under centrifugal effect with a hydraulic medium at least in the area of the free end of the web when the shaft(s) is (are) rotating. In accordance with the present invention, means are provided for exhausting part of the hydraulic medium, said means branching off from an annulus area that is farthest away from the inlet area.

Abstract: This invention relates to a cooling air ducting system in the high-pressure turbine section of a gas turbine engine, in which a portion of the air flow which exits from the compressor section of the engine and bypasses the combustion chamber of the gas turbine is ducted via a first pre-swirl system provided in a diaphragm into a pre-swirl chamber arranged upstream of the first stage turbine rotor disk and is fed from this pre-swirl chamber, in particular, to the air-cooled blades of this rotor disk for cooling purposes, and in which another portion of this air flow is applied to the face of the first stage turbine rotor disk via a second air transfer system which, when viewed in the radial direction, is arranged further inward than the first pre-swirl system. According to this invention, the second air transfer system is again provided as a pre-swirl system connecting to the pre-swirl chamber.

Abstract: In a process for the manufacture of a blade/vane of a turbomachine, the blade/vane is made of a metallic enveloping structure and a supporting structure made of composite material provided inside of the enveloping structure and suitably joined to the enveloping structure. A metal-felt weave is welded or soldered/brazed to the sheet making up the enveloping structure, whereupon this sheet-weave-assembly is shaped by hydroforming or an internal high-pressure forming process to provide the enveloping structure of the blade/vane. Finally, the composite material making up the supporting structure is introduced into the cavity of the enveloping structure.

Abstract: An intermediate-stage seal carrier for a high-pressure jet-engine turbine includes a supporting area for vane retainers, wherein the supporting area includes only one annular wall upstream of and abutting the vane retainers.

Abstract: This invention relates to a gas-turbine engine which is provided in the turbine interior with a bearing chamber for a turbine shaft, the bearing chamber being sealed by sealing elements, and with a scavenge line for the lubricating oil which is supplied to the bearing chamber, in particular is sprayed onto the bearing(s) arranged therein. In accordance with the present invention, the single scavenge line, which is connected to a suitable lubricating-oil collecting compartment, features such an ample cross-section and the sealing elements have such a high sealing effect that the pressure in the bearing chamber is essentially equal to the pressure in the lubricating-oil-collecting compartment and maximally half the pressure in the turbine interior.

Abstract: This invention relates to a method for the detection of a shaft failure in a turbomachine with the object of initiating thereupon an appropriate speed-limiting action, more particularly a rapid fuel shut-off on an aero gas-turbine system, in which a torque-exerting turbine rotor and a torque-recipient unit are connected via the shaft (3) to be monitored for failure, said shaft being supported at its ends in at least two roller bearings (6, 7). In this method, the rotational frequencies (fn1, fn2) of the two shaft ends of the shaft compared with each other continually and essentially in real time, with a failure of the shaft (3) inferred if the rotational frequency (fn2) of the roller bearing (7) on the side of the turbine rotor exceeds the rotational frequency (fn1) of the roller bearing (6) on the side of the torque-recipient unit.

Abstract: This invention relate to a retaining arrangement for rotor blades of axial-flow turbomachinery which features blade roots slots on a rotor disk for the accommodation and radial fixation of roter blades and which further features a continues, circumferential retaining ring for the axial fixation of the rotor blades in the blade root slots and a locking arrangement for the axial retention of the retaining ring. The retaining arrangement is characterized in that the retaining ring has recesses which correspond with sections of the blade root slots and projecting sections on its outer circumference. The present invention provides a cost-effective retaining arrangements for blade of axial-flow turbomachinery which provides for better control of the leakage flows in the blade root area, generally improved flow conditions in the inter-stage area, lower centrifugal forces and ease of assembly.

Abstract: This invention relates to an arrangement for the connection of two rotor disks of an axial-flow turbomachine arranged one behind the other, more particularly two turbine rotor disks of an aero-engine, in which connecting arms protruding from the rotor disks in the axial direction and facing each other are separably joined together by way of several bolts distributed along their circumference. According to the present invention, these bolts are arranged in the radial direction. In a preferential embodiment, each connecting arm features an alternation of prongs extending in the axial direction and of adjacent recesses, with the prongs of the connecting arm of one rotor disk engaging the recesses of the connecting arm of the other rotor disk. A retaining ring subdivided into individual ring segments lodges against the inner side of the connecting arms and is connected to at least one of the two connecting arms by means of bolts.

Abstract: A depth and other characteristics of a ferromagnetic impurity in a workpiece of nonmagnetic material can be determined after demagnetizing of the workpiece by magnetizing the workpiece in a uniform field, preferably to saturation, and thereafter taking measurements of the field strength from the impurity at two different distances, forming a quotient or ratio of the measured values and determining the depth from a curve in which the signal ratio or quotient is plotted against the depth.

Abstract: In a staged gas turbine combustion engine, each combustion chamber of an annular array of chambers includes at least a pilot and a main combustion stage each having an fuel injection nozzle; the method includes feeding fuel continuously in operation to the pilot stage nozzle and controlling the fuel flow to the main stage nozzle through a pulsing and dosing valve to control the fuel flow delivered to the main stage fuel nozzle.

Abstract: A rotor blade airfoil of an axial-flow turbomachine, for a fan or a compressor high-pressure stage of a gas turbine system, configured in the tip area to improve supersonic performance. The blade afflux edge has a forward-back sweep. The tip region of the blade efflux edge can be designed essentially similar to the blade afflux edge, with the tip region extending across the outer 15% to 25% of the blade span.

Abstract: A turbofan aircraft engine has a core engine, which is arranged in a gondola, and a fan which delivers a cold air current through a bypass flow duct to a mixer within which a low-pressure turbine cone is provided and on which the hot-gas current emerging from the low-pressure turbine of the core engine is mixed with the cold-air current. A portion of the cold-air current is branched off by an ejector effect and, as a cooling air current for the low pressure turbine housing, arrives in a ring gap between the housing and a covering surrounding this housing. A portion of this cooling air current is guided by hollow outlet guide blades of a follower guide wheel of the low-pressure turbine for sealing off the rearmost ring gap between the rotor and the stator of the low-pressure turbine, while the largest portion of the cooling air current reaches the ejector.