Abstract: A shroud arrangement for a fluid-flow machine includes a blade root connected to at least one blade/vane end of a row of rotor blades or stator vanes, or is connected to at least one such blade/vane end. A shroud is connected to the blade root on the side of the blade root facing away from the blades/vanes. The blade root on the side facing away from the blades/vanes has two blade root fingers extending substantially in the radial direction and axially spaced from one another, the blade root fingers forming between them a recess being continuous in the circumferential direction. An element or area of the shroud for fastening to the blade root is arranged in the recess. The shroud has two webs extending into the recess and enclosed by the blade root fingers, with the two webs forming substantially a V-shaped arrangement.

Abstract: A structural assembly for a fluid-flow machine includes a main flow path boundary, a row of relatively rotating blades with a gap existing between the blade ends and the main flow path boundary. A secondary flow duct is connected to the main flow path via the two openings spaced apart in the flow direction. A structural assembly has at least one support component and at least one replaceable plug connected directly or indirectly to the support component. The replaceable plug includes a part-section of a secondary flow duct, where the part-section complements at least one further part-section of the secondary flow duct extending outside the plug in the structural assembly to form a secondary flow duct which is continuous between its openings.

Abstract: A method for determining a machining result during surface machining of components, has the following method steps of: providing a component, applying at least one device, which changes under pressure, to the component, machining the surface of the component provided with the at least one device, evaluating the machining operation on the basis of the change in the at least one device as a result of the surface machining of the component. At least one device is in the form of a film which changes at least one property during the surface machining of the component.

Abstract: A structural assembly for a fluid-flow machine includes a main flow path boundary a row of relatively rotating blades with a gap existing between the blade ends and the main flow path boundary. A secondary flow duct is connected to the main flow path via two openings. A structural assembly has at least one support component and at least one insertion component. A structure extending in the circumferential direction and receiving or holding at least one insertion component along the circumference is provided in the support component. Each insertion component forms with at least some of its faces at least part of the main flow path boundary. Each secondary flow duct is jointly limited along at least part of its course by faces of at least two components of the structural assembly.

Abstract: A structural assembly for a fluid-flow machine includes: a main flow path boundary and at least one row of relatively rotating blades with a gap existing between blade ends of the at least one row of blades and the main flow path boundary. At least one secondary flow duct has in the main flow path boundary one opening each at ends spaced apart in the flow direction, such that the secondary flow duct is connected to the main flow path via the two openings. The structural assembly has at least two components connected to one another, i.e. at least one support component and at least one connecting component, where the support component at least partially forms the main flow path boundary and where the connecting component forms or surrounds at least one part-section of the secondary flow duct.

Abstract: A gas turbine oil supply system and a method for the operation of a bearing oil supply system for a gas turbine includes an oil tank 1 and a supply line 2 branching off from the oil tank 1. An electric oil pump 3 is arranged in the supply line 2. A first controllable valve 4, by which the oil flow can be recirculated into the oil tank 1 via a return line 5, is arranged downstream of the oil pump 3. If a temperature of the oil is below a predetermined value prior to turbine start-up, the oil is diverted from supplying the bearings and recirculated through the tank to heat the oil to a desired level.

Abstract: The present invention relates to a combustion chamber for a gas turbine with a combustion chamber head and a heat shield which are designed in one piece, with the heat shield being provided with at least one recess, in which is arranged a burner or an annular sleeve enclosing the burner, with a burner seal being provided between the heat shield and the burner, characterized in that the heat shield is provided in the area of the recess with an annular groove, in which is arranged at least one elastic sealing element forming the burner seal.

Abstract: A retaining segment for axially securing at least one rotor blade on a disk wheel of a rotor device of a jet engine is operatively connectable to at least one rotor blade and to a disk wheel. In a first peripheral area of the retaining segment that is laterally disposed in the circumferential direction in the installed state, a projecting region is provided, and in a second lateral peripheral area facing away from the first peripheral area, a receiving region is provided. The projecting region at least partially overlaps a receiving region of a retaining segment that is identical in design and arranged adjacently in the circumferential direction, and for performing frictional work with the receiving region. A rotor device has rotor blades arranged via blade roots substantially in the axial direction inside recesses of a disk wheel and secured by several such retaining segments.

Abstract: An intermediate casing for a gas-turbine engine has an outer ring (3), an inner ring (1) and an intermediate ring (2) mutually supported by mutually aligned outer and inner supporting struts (4, 5). A radial drive shaft (6), positioned through inner and outer supporting struts (4.1, 5.1) aligned at a 6-o? clock position, connects to an auxiliary gear drive (7). Four additional outer supporting struts (5.2, 5.3, 5.4, 5.5) are spaced apart 72°. A mounting bracket connects to the outer ring at a side facing an aircraft fuselage in areas of two of the additional outer supporting struts. The arrangement of supporting struts and type of mounting bracket attachment, saves weight, improves flow in the bypass duct and increase maintenance and installation access at the bottom side of the engine to nearly 144°.

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

Abstract: With a method for finish-machining the blade tips of rotor wheels made in BLISK design, the rotor wheel or multi-stage compressor rotor (1), is machined—blade by blade and stage by stage—at the blade tips (2) by a conical cutter (5) whose cutting edge (6) is oriented tangentially to the outside surface of the rotor wheel/compressor rotor and routed on the blade tip essentially along the centerline thereof, with the cutting force being introduced perpendicular to the tool axis and in the stiffest direction of the blade (3), i.e. the centerline (11) of the blade cross-section and the perpendicular axis extending from the centerline. The cutting action and the force introduction of the conical cutter at the blade tip allow production of blade tips with a geometrically precisely defined edge without setting the respective blades into vibration.

Abstract: A fluid flow machine has a main flow path (“MFP”) 2 with a blade row 5 therein, a blade end of a blade row being connected to the MFP confinement and a peripheral chamber 7 arranged near this blade end outside the MFP confinement. An outlet 6 is arranged near the fixed blade end near a blade suction side can issue fluid from peripheral chamber 7 onto the surface of the MFP confinement into the MFP. The fluid jet is oriented essentially tangentially to the contour of the MFP confinement when viewed in the meridional plane (x-r plane) and essentially parallel to the local tangent to the skeleton line of the nearest profile, when viewed in the plane established by circumferential direction u and meridional direction m.

Abstract: A running-gap control system of an aircraft gas turbine with a core engine including a turbine whose blade rows have a running gap 13 to the turbine casing 12, with the turbine casing 12 being surrounded by a core-engine ventilation compartment 15 enclosed by a fairing 2 of the core engine, with the fairing 2 forming an inner wall of a bypass duct 1. Air from the bypass duct 1 can be introduced via at least one inlet nozzle 5 into a cooling-air distributor 7 by a control system 6, 10, 11 and the air is subsequently returned to the bypass flow 3.

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: With filmless radiographic inspection of components by means of digital X-ray technology, an uneven surface geometry of the component is smoothened by defining a digital virtual smoothening layer for better, preferably automated, recognition of defects, where the digital radiation signals generated by an X-ray detector are overlaid with digitized surface measurement signals, so that a change in absorption and intensity of radiation due to the surface topography of the component, i.e. due to an uneven surface, is compensated for and only a density caused by internal material defects is represented in the X-ray image.

Abstract: In an arrangement for discharging oil-venting air (oil air) separated by a lubricating oil de-aeration system, kinetic energy of the oil-venting air (O) is increased to a static pressure which exceeds the exhaust-gas flow (A) pressure by a diffuser and the oil-venting air is then mixed with the exhaust-gas flow and discharged with the latter. The venting line (10) coming from the lubricating oil de-aeration system issues into a diffuser (13) which is integrated into the exhaust cone (9) enveloped by the exhaust-gas flow of the engine and whose oil-air outlet opening (14) is connected to the exhaust-gas flow either directly or via an attenuation chamber (18) provided in the exhaust cone. Provision is such made for discharging oil air from the engine to the atmosphere without contaminating visible engine parts and for obtaining a weight reduction due to the reduced length of the venting line.

Abstract: A compressor blade 1 of a gas-turbine engine has a flow-exposed leading edge 2. A component 3, made of wear-resistant material and featuring a higher erosive wear resistance than the base material, is attached to at least part of a length of the leading edge 2 The component 3 forming the flow-exposed leading edge 2 is strip-shaped, has an essentially rectangular cross-section and, with reference to a sectional plane perpendicular to the radial direction, is arranged only on an inflow side of the compressor blade profile in a partial area of a sectional surface of the inflow area.

Abstract: The present invention describes an aircraft gas turbine with a first rotatable shaft and a second shaft arranged coaxially thereto and which at least in an area close to the shaft end is non-rotatably connected to the first shaft. Recesses are provided in shaft areas at a distance from the connecting area between the shafts, said recesses overlapping one another at least in some areas when a defined twist of the first shaft relative to the second shaft is exceeded. The shafts are assigned at least one lever element designed rotatable and engaging in the recesses of the shafts in the radial direction in the area of an end of the shafts and swivelling about a pivot point. A second lever arm of the lever elements is, when the first lever arm is engaged in the recesses, operatively connectable to an electric sensor device, in the area of which a sensor signal equivalent to the defined twist of the first shaft can be generated.

Abstract: The present invention relates to a gas-turbine combustion chamber having a combustion chamber wall including a tile carrier, on which wall tiles are mounted at a distance to form an impingement cooling gap, where the tile carrier has impingement cooling holes and the tile is provided with effusion holes, where the tile has on its side facing the tile carrier a surface structure which raises from the surface of the tile and extends in the direction of the tile carrier.

Abstract: In a method for beam welding on components with a laser beam or electron beam generated by a beam source, the heat treatment of the welded component required to remove stresses is integrated into the welding process. In a work cycle combined with the welding process, a regulated heat supply to a selected region takes place, according to its residual heat resulting from the welding process and the predicted stresses in that region, using the residual heat remaining from the welding process following the welding process from the same or other beam source(s). The welding region is cooled in a controlled manner, so that welded components which are likely to be subject to high stress in use, provided for example for a aircraft engine, can be made available without inherent stresses and with the desired microstructure in a single—combined—work step.