Rolls-Royce Deutschland Ltd & Co KG Patent applications |
Patent application number | Title | Published |
20150369355 | ACCESSORY GEARBOX DEVICE - The present invention describes an accessory gearbox with a shaft incorporated into a casing and connectable at least in one lateral end area to an accessory unit, with the shaft at least in the lateral end area being designed as a hollow shaft having an inner toothing. To lubricate the inner toothing in the area of the casing at least one baffle device is provided which is designed for guiding hydraulic fluid present in the area of the casing, under the effect of gravity, in the direction of at least one opening provided in the shaft. | 12-24-2015 |
20150367459 | ROLLING TOOL DEVICE - The present invention proposes a rolling tool device for compression rolling of, in particular, blade elements of a rotor area of a jet engine provided with a tool carrier to which two pliers-type bodies are rotatably connected about a joint pivot bearing relative to the tool carrier. The pliers-type bodies are each provided with a rolling area, and a distance between the rolling areas is variable in dependence of a rotary movement of the pliers-type bodies. The pliers-type bodies in the zone of the rolling areas are each provided with a part, where main axes of the parts have an extension component in the direction of a rotary axis of the pivot bearing. | 12-24-2015 |
20150367458 | ROLLING TOOL DEVICE - The present invention proposes a rolling tool device for compression rolling of, in particular, blade elements of a rotor area of a jet engine with a tool carrier to which two pliers-type bodies are rotatably connected, with the pliers-type bodies each being provided with a rolling area. A distance between the rolling areas is variable in dependence of a rotary movement of the pliers-type bodies. A separate spacer device is provided by means of which a minimum distance between the rolling areas can be defined. | 12-24-2015 |
20150086344 | ASSEMBLY FOR A FLUID FLOW MACHINE - 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 and the main flow path boundary. At least one secondary flow duct having one opening each is provided in the main flow path boundary 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 one support component and at least one insertion component. The support component includes a recess extending in the circumferential direction that receives the at least one insertion component such that the support component surrounds the at least one insertion component largely on its sides not facing the main flow path, and where the insertion component completely surrounds or forms at least one secondary flow duct. | 03-26-2015 |
20140366555 | ACCESSORY MOUNTING FOR A GAS TURBINE ENGINE - A gas turbine engine assembly is connected to a pylon for mounting the gas turbine engine to an aircraft. The assembly has a frame supporting at least one accessory independently of the gas turbine engine. The frame is attached to the pylon at forward and rearward engine mounting locations. The frame comprises at least one hollow tube and at least one duct is arranged to supply coolant into the at least one hollow tube and the at least one hollow tube is arranged to supply coolant to the at least one accessory. | 12-18-2014 |
20140363283 | SHROUD ARRANGEMENT FOR A FLUID FLOW MACHINE - 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. | 12-11-2014 |
20140363277 | ASSEMBLY FOR A FLUID FLOW MACHINE - 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. | 12-11-2014 |
20140356144 | ASSEMBLY FOR A FLUID FLOW MACHINE - 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. | 12-04-2014 |
20140356143 | ASSEMBLY FOR A FLUID FLOW MACHINE - 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. | 12-04-2014 |
20140352450 | METHOD FOR DETERMINING A MACHINING RESULT DURING SURFACE MACHINING OF COMPONENTS - 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. | 12-04-2014 |
20140318148 | BURNER SEAL FOR GAS-TURBINE COMBUSTION CHAMBER HEAD AND HEAT SHIELD - 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. | 10-30-2014 |
20140314578 | SECURING SEGMENT FOR THE VIBRATION DAMPING OF TURBINE BLADES AND ROTOR DEVICE - 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. | 10-23-2014 |
20140241860 | FLIGHT GAS TURBINE WITH A FIRST ROTATABLE SHAFT - 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. | 08-28-2014 |
20140238030 | IMPINGEMENT-EFFUSION COOLED TILE OF A GAS-TURBINE COMBUSTION CHAMBER WITH ELONGATED EFFUSION HOLES - 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. | 08-28-2014 |
20140234096 | TURBOMACHINE COMPONENT WITH AN EROSION AND CORROSION RESISTANT COATING SYSTEM AND METHOD FOR MANUFACTURING SUCH A COMPONENT - A turbomachine component of a stationary turbomachine includes a substrate made of high alloyed 10% to 18% chromium steels or titanium alloys or nickel base alloys or cobalt base alloys with a substrate surface and an erosion and corrosion resistant coating system. The coating system includes a first layer, which is deposited on the substrate surface of the turbomachine component and acts as a corrosion resistant layer and a second layer. The second layer is deposited on the first layer and acts as an erosion resistant layer, wherein the first layer is a Zr single layer and the second layer is a W/WC multilayer. | 08-21-2014 |
20140234080 | DEVICE AND METHOD FOR BLEEDING COMPRESSOR AIR IN A TURBOFAN ENGINE - A device for bleeding compressor air in a turbofan engine includes a low-pressure compressor of a primary duct, a fan stator arranged in a secondary duct downstream of a fan, and an arrangement for discharging compressor air from the low-pressure compressor into the secondary duct. The arrangement includes air guiding devices to guide compressor air into the secondary duct upstream of the fan stator relative to the flow direction in the secondary duct, past the stator vanes of the fan stator in a separate duct in the axial direction and is only mixed with the air of the secondary duct downstream of the fan stator. The separate duct is formed by a plurality of hollow structures extending in the longitudinal direction each between two adjacent stator vanes of the fan stator, with the hollow structures having a closed circumference in cross-section. | 08-21-2014 |
20140161591 | LUBRICANT SYSTEM - An aircraft engine having a fan or propeller stage mounted to a planetary gearing system through a planetary gearing output drive shaft is provided with a lubricant system having a seal with lubricant delivery chamber for receiving a lubricant and a sealing surface sealing against a rotatable shaft, the rotatable shaft has a hollow interior and an opening in registration with the delivery chamber. The lubricant system has a lubricant pump in fluidical communication with and downstream of the opening. | 06-12-2014 |
20140150442 | GAS TURBINE CENTRIPETAL ANNULAR COMBUSTION CHAMBER AND METHOD FOR FLOW GUIDANCE - A gas-turbine combustion chamber arrangement includes at least one centrifugal compressor as well as one centripetal annular combustion chamber, with a stator vane arrangement being provided between the centrifugal compressor and the annular combustion chamber. The stator vane arrangement for diverting the air flowing out of the centrifugal compressor is designed at an angle α of 20° to 30°, preferably 25°, relative to the engine axis, so the airflow is passed at essentially this angle α to the combustion chamber. The inflow area into the combustion chamber for supplying air is designed at an angle of 20° to 30°, preferably 25°, relative to the meridional plane, and the center axes of the burners or of the injection nozzles of the combustion chamber are arranged inclined at an angle β of 30° to 40°, preferably 35°, relative to a meridional plane passing through the engine axis. | 06-05-2014 |
20140144146 | TILE FASTENING ARRANGEMENT OF A GAS-TURBINE COMBUSTION CHAMBER - The present invention relates to a tile fastening arrangement of a gas-turbine combustion chamber having a combustion chamber wall with tiles fastened to said combustion chamber wall at a distance from the latter, characterized in that the tile is provided with an annular flange arranged on its side assigned to the combustion chamber wall, said flange being dimensioned to match a recess of the combustion chamber wall and fastened to the combustion chamber wall by means of a fastening element which engages in the combustion chamber wall. | 05-29-2014 |
20140144145 | GAS TURBINE COMBUSTION CHAMBER - The present invention relates to a gas-turbine combustion chamber with an outer combustion chamber wall concentric to a gas-turbine center axis and with an inner combustion chamber wall, with several burners arranged spread over the circumference of the combustion chamber, and with air inlet recesses which in at least one radial plane are provided spread over the circumference on an outer combustion chamber wall and on an inner combustion chamber wall, where the burner is designed to form a flow provided with a swirl and where air inlet recesses assigned to a burner are dimensioned in differing size in order to generate airflows of differing size, characterized in that at least one of the respective air inlet recesses is designed for supplying air in the flow direction of the swirl of the combustion chamber flow and is provided with flow guidance walls. | 05-29-2014 |
20140141894 | Shaft of a gas-turbine engine, in particular a radial shaft or a shaft arranged at an angle to the machine axis - A radial shaft of a gas-turbine engine which is made up of +/−45° layers, zero layers and +/−30° layers of carbon fiber composite and connected to load-input end pieces via sinusoidal connecting areas. | 05-22-2014 |
20140140820 | AEROENGINE SEALING ARRANGEMENT - A bifurcation aerofoil having a leading edge portion and flanks where the leading edge portion is mounted to an inner wall of the bypass duct using a floating seal which permits circumferential movement of the leading edge. A cowl has a wall that provides the flanks of the aerofoil. On closing of the cowl into its flight position any contact between the flanks and the leading edge portion enables realignment of the leading edge portion to the inner wall of the bypass duct. | 05-22-2014 |
20140130964 | Method for manufacturing a shaft of a gas-turbine engine, in particular a radial shaft or a shaft arranged at an angle to the machine axis - A method for manufacturing a radial shaft of a gas-turbine engine which is made up of +/−45° layers, zero layers and +/−30° layers of carbon fiber composite and connected to load-input end pieces via sinusoidal connecting areas. | 05-15-2014 |
20140130501 | COMBUSTION CHAMBER TILE OF A GAS TURBINE - A combustion chamber tile of a gas turbine includes a bolt for mounting the combustion chamber tile on a combustion chamber wall, with the combustion chamber tile being designed substantially plate-like and having on one side at least one mounting element, on which the bolt, which is provided as a separate component, is positively anchored. | 05-15-2014 |
20140127010 | NOZZLE WITH GUIDING DEVICES - The present invention relates to a nozzle with a nozzle surface area and a nozzle rim, on which first and second guiding devices are alternatingly provided in the circumferential direction, where the first guiding devices are of the nozzle-type design and the second guiding devices are of the diffuser-type design. The first guiding devices each have a first azimuthal guide wall and two wall elements, with the first azimuthal guide wall forming a first trailing edge and two first edges to the wall elements. The second guiding devices each have a second azimuthal guide wall and two wall elements, with the second azimuthal guide wail forming a second trailing edge and two second edges to the wall elements. A wall element connects a first guiding device and a second guiding device. The wall elements have a curved course in the axial direction. | 05-08-2014 |
20140127005 | METHOD FOR PRODUCING A COMPONENT, COMPONENT AND TURBOMACHINE HAVING A COMPONENT - A method for manufacturing a metallic component especially configured and designed for a turbomachine includes a) for a precision-casting process, a wax model with a wax structure is produced, subsequently b) the tip of the wax structure is thermally and/or mechanically treated such that a region with an undercut is formed on the wax structure, subsequently c) the metallic component is manufactured from the wax model in the precision-casting process, with a component structural element with an undercut forming on a surface of the component, and d) the component structural element is provided at least partially with a ceramic coating, a plastic-containing coating, in particular a fibre composite layer and/or a plastic component. | 05-08-2014 |
20140127003 | Stator vane adjusting device of a gas turbine - A stator vane adjusting device of a gas turbine has a plurality of stator vanes each rotatable about a radial axis 44 and arranged in at least two radial planes, as well as at least one stator vane adjusting ring connected to the respective stator vanes and rotatable in the circumferential direction by at least one actuating device. The actuating device includes a crankshaft element rotatable about a stationary pivot axis by an actuator. A first lever is articulated by a joint to the stator vane adjusting ring, with its free end being connected by a joint to a center area of a second lever, the second lever being mounted at its one end on a stationary pivot point and at its other end being linked by a joint to a third lever, which is mounted by a joint at its free end on the crankshaft element. | 05-08-2014 |
20140127001 | GAS-TURBINE ENGINE WITH A TELESCOPIC AIR INLET IN THE ENGINE LINING - 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. | 05-08-2014 |
20140124289 | NOZZLE WITH GUIDING DEVICES - A nozzle has a nozzle surface area and a nozzle rim, on which first and second guiding devices are alternatingly provided in the circumferential direction, where the first guiding devices are of the nozzle-type design and the second guiding devices are of the diffuser-type design. The first guiding devices each have a first azimuthal guide wall and two wall elements. The second guiding devices each have a second azimuthal guide wall and two wall elements. A wall element connects a first guiding device and a second guiding device. At least some of the first azimuthal guide walls of the first guiding device and at least some of the second azimuthal guide walls of the second guiding device have differing axial lengths, so that first and second trailing edges thereof have differing axial positions. | 05-08-2014 |
20140119942 | TURBINE ROTOR BLADE OF A GAS TURBINE - The present invention relates to a turbine rotor blade of a gas turbine with a blade tip, said blade tip having at least on its suction side, extending from a stagnation point on the blade leading edge to an intersection point of the suction-side profile line of the blade with a trailing-edge circle, an overhang which is substantially zero at the stagnation point and at the intersection point and which has a maximum value at around 40% of the running length of the suction-side overhang. | 05-01-2014 |
20140119883 | BLEED FLOW PASSAGE - A liner wall insert is provided for a compressor rotor stage of a gas turbine engine. Several liner wall inserts are provided radially outboard of the tips of the rotor blades. The liner wall inserts have bleed flow channels formed therein. The bleed flow channels are arranged to remove flow from a trailing edge region of the stage and re-inject the bleed flow at an upstream region. The re-injected bleed flow alters the flow field around the tips of the rotor blades, for example the tip leakage flow. Thus, the bleed flow is used to improve the efficiency of the compressor rotor stage, and thus of the gas turbine engine. | 05-01-2014 |
20140116026 | AEROENGINE THRUST REVERSER ARRANGEMENT - 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. | 05-01-2014 |
20140116025 | AEROENGINE THRUST REVERSER ARRANGEMENT - 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. | 05-01-2014 |
20140096533 | Bearing chamber venting system for an aircraft engine and method for providing a required pressure ratio at bearing chamber seals of an air-sealed bearing chamber - A bearing chamber venting system for an aircraft engine includes at least one bearing chamber air-sealed by sealing air via bearing chamber seals, a vent line connected to the bearing chamber, via which an oil/air mixture present in the bearing chamber is vented out, and an oil separator connected to the vent line. An oil return line is connected to the oil separator via which oil separated from the mixture is discharged. An air outlet line is connected to the oil separator, via which cleaned air is passed to the environment. An air ejector is arranged in the air outlet line to eject gas supplied to the air ejector into the air outlet line at a speed which is greater than the speed of the air flowing in the air outlet line and passed to the environment out of the oil separator. | 04-10-2014 |
20140079539 | Turbine blade of a gas turbine with swirl-generating element and method for its manufacture - The present invention relates to a turbine blade of a gas turbine with an airfoil arranged on a blade root and having at least one cooling air duct running in the longitudinal direction of the turbine blade, arranged inside the turbine blade and extending through the blade root, characterized in that at least one swirl-generating element is arranged in the transitional area between blade root and airfoil in the cooling air duct, with the swirl-generating element including an outer ring and several swirl-generating stator vanes arranged thereon, which are connected to a centric area, as well as to a method for its manufacture. | 03-20-2014 |
20140064952 | ASSEMBLY OF AN AXIAL TURBOMACHINE AND METHOD FOR MANUFACTURING AN ASSEMBLY OF THIS TYPE - The present invention relates to an assembly of an axial turbomachine, comprising at least one outlet guide vane of a compressor and a diffuser arranged downstream of the outlet guide vane in the flow direction. It is provided that the outlet guide vane is connected to the compressor and that the diffuser is connected to the combustion chamber, without there being a direct mechanical connection between the diffuser and the outlet guide vane. The invention furthermore relates to a method for manufacturing an assembly of this type. | 03-06-2014 |
20140064928 | Engine casing of an aircraft gas turbine having sound-absorbing elements in the fan inflow region - The present invention relates to an engine casing of an aircraft gas turbine having a radially inner honeycomb-structured layer arranged on the engine casing in the flow direction upstream of a fan in the area of an inflow-side air inlet, characterized in that sound-absorbing elements are arranged radially outside at least one axial section of the honeycomb-structured layer, said elements extending in the axial direction and being arranged annularly next to one another in the circumferential direction. | 03-06-2014 |
20140050561 | OIL SUPPLY SYSTEM AND METHOD FOR SUPPLYING OIL FOR A TURBOPROP ENGINE - An oil supply system for a propeller turbine engine includes a first oil circuit for supplying the turbomachine and a second oil circuit for supplying the propeller gearbox and the high-pressure pump associated with a propeller adjusting device. The two oil circuits are integrated into a common oil tank and connected to one another by means of an oil supply line in the flow direction downstream of the first and the second oil conveying pumps in such a way that a limited oil volumetric flow can flow from the first oil circuit into the second oil circuit, and only in this direction, if the pressure difference between the first and the second oil circuits falls below a certain level in the event of a pressure reduction in the second oil circuit. In this way, a sufficiently high initial pressure of the oil is produced at the high-pressure pump, even at low propeller speeds and during transient propeller adjustment processes, so that the function of the propeller adjusting device is not impaired and the high-pressure pump is not damaged. Limiting the oil volumetric flow in the oil supply line connecting the two oil circuits prevents that the turbomachine is inadequately supplied with oil, while oil is being removed from the first oil circuit. | 02-20-2014 |
20140033731 | Method for fuel temperature control of a gas turbine - The present invention relates to a method for controlling the fuel temperature of a gas turbine, where parameters are determined as input values, where the parameters are compared with emission-optimized nominal values and an optimum fuel temperature is determined, and where the fuel to be supplied to a combustion chamber is heated or cooled. | 02-06-2014 |
20140033723 | UNKNOWN - The present invention relates to a gas-turbine combustion chamber with mixing air orifices, with a combustion chamber wall, with tiles arranged on the inside of the combustion chamber wall at a certain distance from said combustion chamber wall as well as with mixing air orifices passing through the combustion chamber wall and the tiles, characterized in that the mixing air orifices are formed by chutes which are designed tube-like and pass through the tiles, and that several chutes are formed on a mixing air wall element extending at least around part of the circumference of the combustion chamber. | 02-06-2014 |
20140030109 | low-Modulus Gas-Turbine Compressor Blade - The present invention relates to an aircraft gas-turbine compressor blade having an airfoil with a leading edge, with at least the area of the leading edge being made from a low-modulus titanium alloy. | 01-30-2014 |
20140030107 | DECOUPLED COMPRESSOR BLADE OF A GAS TURBINE - The present invention relates to a compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic, which is fastened by means of a blade root to a disk, as well as a metallic leading-edge element, which is arranged on the leading-edge side of the airfoil and partially encompasses the latter, with the leading-edge element itself being fastened to the disk. | 01-30-2014 |
20140030106 | Compressor blade of a gas turbine as well as method for manufacturing said blade - The present invention relates to a compressor blade of a gas turbine having an airfoil made of a fiber-reinforced plastic and a leading-edge element connected to said airfoil, characterized in that the leading-edge element includes two partial elements, which are fabricated as separate elements and connected to one another, with the leading-edge element being connected to the airfoil essentially using the clamping effect of the two partial elements, as well as to a manufacturing method. | 01-30-2014 |
20140026700 | ACCESSORY GEARBOX DEVICE FOR A JET ENGINE - The present invention proposes an accessory gearbox for an engine having at least one accessory gearbox shaft connectable to an auxiliary unit and driveable by a drive shaft, said drive shaft being operatively connectable to an engine shaft of the engine. At least one auxiliary unit can be arranged on the drive shaft and/or in the area of the operative connection between the drive shaft and the engine shaft and/or on an accessory gearbox shaft connectable to a further auxiliary unit. | 01-30-2014 |
20140026592 | ASSEMBLY FOR A JET ENGINE OF AN AIRCRAFT - The present invention proposes a structural unit for an aircraft engine having at least one fuel pump of a fuel circuit and at least one hydraulic fluid pump of a hydraulic fluid circuit, where the structural unit can be coupled to an accessory gearbox shaft of an accessory gearbox of the engine. | 01-30-2014 |
20140026534 | OIL SUPPLY SYSTEM FOR AN AIRCRAFT ENGINE - In an oil supply system for an aircraft engine, including at least one oil circuit, into which an oil tank having a closing valve associated with the outlet opening of the oil tank into a suction line, an oil pump and at least one consumer to be supplied with oil by means of a pressure line are integrated, a negative pressure relief line having a shut-off valve incorporated is connected to the suction line, said negative pressure relief line for ensuring pressure relief in the suction line being connectable after shutdown of the engine to the pressure line or to the ambient air. Via a volume flow limiter designed as a diaphragm and alternatively incorporated into said negative pressure relief line, the suction line can also be connected permanently to the pressure line. A negative pressure occurring in the suction line after shutdown of the engine when the closing valve is closed and the plastic deformation of the suction line as well as a reduction of the pump suction capacity are thus avoided. | 01-30-2014 |
20130326876 | METHOD FOR REPAIRING COMPRESSOR OR TURBINE DRUMS - During the repair of welded two-stage or multi-stage compressor or turbine drums, which after welding are machined in the weld area down to a height A, a first material coating with a height B extending on both sides from the root face is applied to the machined weld area, and a second and lesser material coating is applied by laser welding to the area of plasma pockets formed on the intact rotor disk, with the sum of the heights A and B corresponding to an original height C in the weld area prior to its final machining. Then the defective rotor disk is detached in the root face and through the middle of the first material coating, and after that a new rotor disk already validated in the new construction with the root face validated in the new construction is welded on using appropriate welding parameters, and the weld area is machined, in line with the parameters already used for the new construction, down to the original height A. It is thus possible to provide repaired compressor or turbine drums matching the original component at low expense. | 12-12-2013 |
20130318945 | AVIATION GAS TURBINE THRUST REVERSING DEVICE - 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. | 12-05-2013 |
20130309091 | OIL SUPPLY SYSTEM FOR A PROPELLER TURBINE ENGINE - An oil supply system for a propeller turbine engine including a propeller, a propeller gearbox connected to the propeller and a propeller adjusting device for altering the pitch angle of the propeller blades. The system includes an oil circuit and a propeller main pump for supplying the propeller gearbox and a propeller high-pressure pump supplying the propeller adjusting device with oil. The system includes an oil accumulator for providing, in the event of an undersupply of oil by the propeller main pump, oil contained in the oil accumulator to the inlet of the propeller high-pressure pump such that the pump inlet pressure at the inlet of the propeller high-pressure pump does not fall short of a certain minimum value. The oil accumulator is incorporated into the inflow to the propeller high-pressure pump to be continuously supplied and charged with oil of the propeller main pump. | 11-21-2013 |
20130306403 | METHOD FOR THE PRODUCTION OF A SOUND ABSORBER, ESPECIALLY FOR A GAS TURBINE EXHAUST CONE - 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. | 11-21-2013 |
20130306402 | METHOD FOR THE PRODUCTION OF A SOUND ABSORBER, ESPECIALLY FOR A GAS TURBINE EXHAUST CONE - 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. | 11-21-2013 |
20130306401 | SOUND ABSORBER FOR A GAS TURBINE EXHAUST CONE, AND METHOD FOR THE PRODUCTION THEREOF - 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. | 11-21-2013 |
20130302143 | COOLING DEVICE FOR A JET ENGINE - The present invention relates to a cooling device for a jet engine having an axial compressor with several compressor stages including a rotor with rotor blades, a stator with stator vanes and an annulus. In order to reduce the temperature of the components at the outlet of the high-pressure compressor by simple measures and hence to increase the efficiency of a jet engine, a slot-like branch opening surrounding the rotor for a cooling airflow diverted from the main airflow into a first cavity upstream of the rotor is provided upstream of the last compressor stage of the axial compressor, with passage openings being arranged in the rotor for passing on the diverted cooling airflow from the first cavity into a second cavity downstream of the rotor. | 11-14-2013 |
20130266470 | METHOD FOR THE MANUFACTURING HIGH-TEMPERATURE RESISTANT ENGINE COMPONENTS - For near net shape manufacturing of high-temperature resistant engine components of geometrically complex design consisting of an intermetallic phase, a low melting-point metallic phase in the molten state or in a temperature range near the molten state is mixed with a high melting-point metallic phase provided as a metal powder, and the mixture is mechanically treated under the effect of kneading and shear forces, thereby heating it up and reducing its viscosity. In a subsequent injection moulding process the engine component substantially matching the final contour is formed and mechanically finish-machined, if required, and afterwards subjected to a heat treatment for creating an intermetallic phase. | 10-10-2013 |
20130266469 | METHOD FOR NEAR NET SHAPE MANUFACTURING OF HIGH-TEMPERATURE RESISTANT ENGINE COMPONENTS - For near net shape manufacturing of a high-temperature resistant component of complex design a high melting-point part of an intermetallic phase provided as a metal powder is mixed with a binder, and from the feedstock such formed a green compact substantially matching the final contour is produced by metal injection moulding, into the pores of said compact that remain after removal of the binder the low melting-point part of the intermetallic phase is infiltrated. The brown compact thereby created is mechanically processed, if required, and subjected to a specific heat treatment depending on the metallic phases used in order to create the intermetallic phase. This permits engine components consisting of intermetallic phases and having a geometrically complex structure to be manufactured cost-efficiently. | 10-10-2013 |
20130266424 | STATOR VANE ADJUSTING DEVICE OF A GAS TURBINE - The present invention relates to a stator vane adjusting device of a gas turbine having a plurality of stator vanes each swivellable about a radial axis and arranged in at least two radial planes, as well as at least two stator vane adjusting rings connected to the respective stator vanes and rotatable in the circumferential direction by at least one actuating device, characterized in that the actuating device is connected to the stator vane adjusting rings by means of a first transmission device and that a second transmission device, which is not coupled to the actuating device, is arranged essentially opposite to the first transmission device, with the second transmission device being connected to the stator vane adjusting rings. | 10-10-2013 |
20130259732 | METHOD FOR PRODUCING ENGINE COMPONENTS WITH A GEOMETRICALLY COMPLEX STRUCTURE - The present invention relates to a method for manufacturing thermally stressed engine components having a geometrically complex structure by metal injection moulding of metal powder mixed with a binder, by which method individual parts of the engine component are produced as separately moulded green compact sections and then as debindered brown compact sections which are joined together to form a two-part or multi-part brown compact and sintered in the assembled state, with the brown compact sections having differing shrinkage properties in the sintering process, depending on the type and size of the metal powder used, with at least one more heavily shrinking first brown compact section being automatically pressed against at least one second brown compact section during sintering of the assembled brown compact. It is provided that connecting elements in the form of positively engaging projections and recesses are provided at the joining surfaces of the brown compact sections to be joined and having differing shrinkage in such a way that during sintering of the assembled brown compact, the brown compact section with projections undergoes a greater shrinkage than the brown compact section with recesses. | 10-03-2013 |
20130230383 | Aircraft gas turbine with adjustable fan - 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. | 09-05-2013 |
20130219982 | ROLLING TOOL DEVICE - The present invention describes a rolling tool device for compression rolling of, in particular, blade elements of a rotor area of a jet engine provided with a tool carrier. The tool carrier is connectable to a carrier spindle. Furthermore, two pliers-type bodies are rotatably connected to the tool carrier. The pliers-type bodies are each provided with a rolling area, with a distance between the rolling areas being variable in dependence of the rotary movement of the pliers-type bodies. In accordance with the present invention, an axis of the carrier spindle in the state connected to the tool carrier passes between the rolling areas through a contact point present at a distance between the rolling areas equal to zero. | 08-29-2013 |
20130216391 | METHOD FOR THE PRODUCTION OF A ONE-PIECE ROTOR AREA AND ONE-PIECE ROTOR AREA - The present invention describes a method for the production of a one-piece rotor area, preferably of a jet engine. The rotor area includes an annular base body and several, circumferentially distributed blade elements extending essentially radially from the base body. Residual stresses are imparted to the blade elements in surface-near areas by way of roller compression using a rolling tool introduced between the blade elements. During roller compression, one each area of a blade element is arranged between areas of the rolling tool, with longitudinal sides of the blade element being simultaneously roller-compressed. According to the present invention, the rolling tool is radially introduced between the blade elements and the surfaces of the blade elements are roller-compressed, thus at least the blade elements having a roller-compressed surface. | 08-22-2013 |
20130206676 | MEASURING DEVICE HAVING A PRESSURE SENSOR - The present invention relates to a measuring device having a pressure sensor, with the pressure sensor having a measuring cell designed to detect a pressure. It is provided that the measuring cell of the pressure sensor is arranged in or adjacent to a protective cell filled with a measuring fluid that can be coupled to a fluid to be measured via at least one separating membrane. It is furthermore provided that a damping device is arranged in the protective cell between the separating membrane and the measuring cell, said damping device including a restrictor and a downstream volume expansion means that provides an increased volume between the restrictor and the measuring cell in the event of a pressure increase in the measuring fluid. | 08-15-2013 |
20130205788 | UNKNOWN - The present invention relates to a premix burner of a combustion chamber of a gas turbine with at least one annular duct for supplying air and fuel, including a radially outer and a radially inner combustion chamber wall relative to a burner central axis and with at least one swirler arranged in the duct, said swirler including several flow-guiding elements distributed around the circumference of the duct cross-section, characterized in that at least one radially inner duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type. | 08-15-2013 |
20130205753 | Aircraft gas turbine thrust-reversing device - An aircraft gas turbine thrust-reversing device has an engine cowling, the rear area of which is displaceable 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 space to a front and stationary area of the engine cowling. The rear area of the engine cowling is coupled to deflecting elements and blocker doors, which in the forward thrust position are arranged completely inside the front area of the engine cowling. A drive element is provided between the front area and the rear area, which effects the axial displacement of the rear area. The drive element is a two-stage drive element, effecting in a first stage an axial displacement by an axial partial displacement path, and in a second stage an axial displacement by the full axial displacement path. | 08-15-2013 |
20130202088 | Method for Radiographically Inspection a Component by Means of X-ray Beams Using a Smoothing Agent and Smoothing Agent for Carrying Out the Method - The present invention relates to a method for radiographically inspecting a component by means of X-rays, where at least one component surface to be radiographed is provided with a surface structure, with at least the surface provided with the surface structure being smoothed by means of a smoothing material to level out the surface structure, with at least one organic compound and at least one metal powder being used as smoothing material, with the X-ray absorption behaviour of the smoothing material essentially equaling the X-ray absorption behaviour of the material of the component, as well as to a smoothing material for carrying out the method in accordance with claim. | 08-08-2013 |
20130199187 | Gas-turbine combustion chamber having non-symmetrical fuel nozzles - The present invention relates to an annular gas-turbine combustion chamber having a radially outer and a radially inner combustion chamber wall relative to a machine axis, a combustion chamber head and a combustion chamber outlet nozzle, where the combustion chamber head includes several fuel nozzles spread over its circumference for supplying air and fuel, the latter exiting in an outlet surface of the fuel nozzles, where the respective fuel nozzle has a burner axis which is vertical to the outlet surface and where the intersections of the burner axes with the outlet surfaces define a circular burner centerline around the engine axis, characterized in that a cross-sectional area of the fuel nozzle radially outside the burner centerline is identical to a cross-sectional area radially inside the burner centerline. | 08-08-2013 |
20130199040 | Device and method for treatment of high-pressure turbine blades of a gas turbine - The present invention relates to a device and a method for in-situ treatment of turbine blades/vanes of a gas turbine, with the treatment device being insertable into a combustion chamber of the gas turbine via at least one access aperture and placeable on cooling-air recesses of the turbine blades, with the treatment device at its end including a tool head with at least one tool and at least one jet nozzle, with the jet nozzle being connected to a medium hose for externally supplying a treatment fluid, with the treatment device being provided with an actuating device for exactly guiding and/or placing and/or moving the tool head. | 08-08-2013 |
20130195610 | NOISE-REDUCED TURBOMACHINE - Turbomachine with an annular main flow duct ( | 08-01-2013 |
20130192232 | Annular combustion chamber of a gas turbine - The present invention relates to an annular combustion chamber of a gas turbine with—relative to the engine axis—a radially outer combustion chamber wall and a radially inner combustion chamber wall, with the combustion chamber walls forming an annular combustion space, with a combustion chamber head having a plurality of fuel nozzles and air inlet openings, with the respective central axes of the fuel nozzles forming an envelope rotationally symmetrical to the engine axis, the envelope dividing the combustion chamber into an annular and radially outer area and an annular and radially inner area, with the radially outer area and the radially inner area having the same volumes. | 08-01-2013 |
20130186707 | UNKNOWN - An acoustic absorber includes a wall provided with a plurality of apertures as well as a substantially non-perforated second wall, with the first wall and the second wall being spacedly arranged to one another, and at least one honeycomb structure being provided between the walls, with the honeycomb structure having a substantially cylindrical recess in at least one area, with a funnel element opening to the first and second walls being provided in the recess, and having a height greater than the distance of the walls, with the second wall being designed pot-like in the area of the funnel element. | 07-25-2013 |
20130180227 | Fastening element and de-icing device of an aircraft gas-turbine engine - The present invention relates to a fastening element, in particular to its use in a de-icing device of an aircraft gas-turbine engine, for connecting two components, with the fastening element ensuring a connection of the components with a predetermined relative movability to each other, with the fastening element including two struts arranged at an angle to each other, where two first end areas, spacedly arranged to each other, can be fastened to one of the components, and the two other second end areas can be connected to each other and fastened to the other component. | 07-18-2013 |
20130166056 | METHOD FOR SELECTING A GEOMETRY OF A BLADE - A method selects from a plurality of predetermined blade geometries, a blade geometry for a blade wheel for a turbomachine, with for the predetermined blade geometries at least one characteristic value identifying an aerodynamic property of the blade geometry and at least one characteristic value identifying a structural mechanism of the blade geometry being filed in a memory. An evaluation unit ascertains for each blade geometry a total value, which is calculated from the assigned characteristic values. The evaluation unit selects at least that blade geometry whose total value has an extreme value of all computed total values of the blade geometries. At least one characteristic value identifying producibility of the respective blade geometry is filed in the memory additionally to the blade geometries. The evaluation unit also incorporates while ascertaining the total value of the respective blade geometry the characteristic value identifying producibility. | 06-27-2013 |
20130163848 | METHOD AND APPARATUS FOR INSPECTION OF COMPONENTS - The quality of a dry film lubricant coating or molybdenum based undercoat coating provided on the root of a gas turbine fan blade is determined by a method which includes providing an image of the coating taken under controlled conditions which include diffusing light through a translucent screen. The difference between an RGB value of a group of pixels from the image compared with a known RGB value is determined and used to calculate the quality of the coating. | 06-27-2013 |
20130160459 | ACCESSORY MOUNTING FOR A GAS TURBINE - Accessories are mounted to a gas turbine engine from a frame attached to a pylon which is used to mount the engine to an airframe. The frame independently mounts the accessories so that engine vibrations are not transmitted to the accessories. | 06-27-2013 |
20130157000 | COMPONENT AND TURBOMACHINE HAVING A COMPONENT - A component, especially contrived and designed for being used in a turbomachine, includes a high-temperature coating being arranged above a base of the component. The base has at least one structural element for connecting it to the high-temperature coating, with the cross-section of the at least one structural element having at least three different widths, i.e. a base width at the lower end of the at least one structural element, a center width above it, and a tip width above that, where on average the center width is greater than or equal to the base width, but less than four times the base width, in particular less than or equal to three times the base width. | 06-20-2013 |
20130156626 | METHOD FOR MANUFACTURING A PART BY METAL INJECTION MOULDING - A method for manufacturing a part includes metal injection molding of metal powder mixed with a binder to produce individual components of the part as separately molded green compact sections which are then debindered to form brown compact sections. At least one of the brown compact sections is subjected to a pre-sintering process to undergo a first shrinkage. The pre-sintered brown compact section and a further brown compact section are joined together to form a multi-part brown compact which is subsequently subjected to a main sintering process, where the pre-sintered brown compact section undergoes less shrinkage than the further brown compact section to draw together and firmly connect the pre-sintered brown compact section and the further brown compact section. | 06-20-2013 |
20130142467 | METHOD AND DEVICE FOR ADJUSTING THE BEARING PLAY IN A CERAMIC HYBRID BEARING - A method and a device for adapting a bearing clearance in a ceramic-hybrid bearing having rolling elements in the form of balls or rollers made from a ceramic material and an outer and an inner bearing shell made from ferromagnetic material. In order to be in a position to adapt the bearing clearance also during operation, a magnetic flux is introduced into an outer side of the stationary, non-rotating bearing shell and exits again at the opposite outer side of the bearing shell, with a magnetic field and an associated magnetic flux being generated, and that a deformation of the bearing shell in the sense of a decrease or increase in the circumference of the bearing shell is brought about due to the resulting magnetostrictive effect from a change in the magnetic flux. | 06-06-2013 |
20130139578 | PRESSURE-MEASURING DEVICE AND PRESSURE-MEASURING METHOD FOR A TURBOMACHINE - A pressure-measuring device and a pressure-measuring method for measuring a pressure in a combustion chamber of a turbomachine are provided. The pressure-measuring device includes at least one sensor device for measuring an ion current at at least one point inside the combustion chamber. | 06-06-2013 |
20130129505 | BEARING DEVICE AND TURBOMACHINE HAVING A BEARING DEVICE - A bearing device, for example in a turbomachine is provided. Between bearing races of the anti-friction bearings and a casing component positioned adjacent the bearing races is arranged a device for converting an axial force between the anti-friction bearings into a torque at the bearing races in the manner of a ball screw or planetary screw assembly. The bearing races are connected to one another by a feedback mechanism such that a torque at one bearing race can be converted into an opposite torque at the other bearing race. | 05-23-2013 |
20130129504 | BEARING DEVICE AND TURBOMACHINE HAVING A BEARING DEVICE - A bearing device with an arrangement of at least two bearings is provided. A filling made of an amorphous/solid active medium as a mechanism of compensating for axial loads is arranged between a bearing race of at least one of the bearings and a surrounding component of the bearing arrangement. | 05-23-2013 |
20130124114 | METHOD FOR DETERMINING A FLOW BEHAVIOUR OF A MEDIUM - A method for determining a flow behavior of a medium uses at least one analysis device placed inside the medium. The at least one analysis device is freely moveable in the medium and supplies data characterizing at least one property of the medium flow behavior to a data evaluation device. The at least one property of the medium flow behavior is determined by the data evaluation device on the basis of the data in selected areas or in all areas through which the at least one analysis device flows with the medium. | 05-16-2013 |
20130098179 | ACCESSORY GEARBOX DEVICE FOR A JET ENGINE - The present invention proposes an accessory gearbox ( | 04-25-2013 |
20130091704 | METHOD FOR REPAIRING ROTOR BLADES - In a method for repairing worn or damaged rotor blades of an aircraft engine, laser cladding for the reconstruction of the blade tips is not performed in a continuous welding process along the blade edge but during a cycle-by-cycle rotation of the rotor blades attached to a rotor disk, in only one limited section of the respective blade edge for each cycle. The method reduces time expenditure for the repair of the blades and improves the tribological properties of the blade material in the area of the repaired blade tip. | 04-18-2013 |
20130086908 | GAS TURBINE COMBUSTION CHAMBER ARRANGEMENT OF AXIAL TYPE OF CONSTRUCTION - A gas-turbine combustion chamber arrangement includes a flame tube, a diffuser element arranged upstream of the flame tube in the flow direction, the diffuser element including an annular duct, and an axial compressor arranged upstream of the diffuser element. The diffuser element features an annular guide vane area in which guide vanes are arranged, which for redirecting an incoming flow are provided at an angle (α) in a range between 28° and 32° relative to a central axis of the gas turbine. Downstream of the guide vane area, a diffuser area is arranged, the diffuser area not being provided with flow-guiding elements affecting the flow, where burners arranged in the annular combustion chamber are provided with their burner axes at an angle (β) between 40° and 50° relative to the central axis. | 04-11-2013 |
20130057093 | GENERATOR AND ACCESSORY GEARBOX DEVICE WITH A GENERATOR - The present invention proposes a generator for arrangement on a shaft of an accessory gearbox of an engine with a stator and with a rotor which can be coupled to a shaft of the accessory gearbox of the engine and which is rotatably mounted relative to the stator, where a stator area receiving the stator can be separated from a rotor area receiving the rotor. The rotor can be supplied with cooling medium in the rotor area. It furthermore proposes an accessory gearbox of an engine with a drive shaft operatively connectable to a main shaft of the engine and with at least one generator arranged on a shaft of the accessory gearbox. | 03-07-2013 |
20130055716 | Gas-turbine combustion chamber with a holding means of a seal for an attachment - Gas-turbine combustion chamber with a combustion chamber head made from a metallic material and mounting at least one burner, and with a combustion chamber wall made from a ceramic material, where at least one igniter plug or other combustion chamber attachments such as acoustic dampers, sensors or valves are arranged in a recess of the combustion chamber wall, and where in the area of the recess a seal is arranged that is mounted by means of a metallic holding means from another component than the CMC combustion chamber wall. | 03-07-2013 |
20130045093 | METHOD FOR THE MANUFACTURE OF A COMPONENT FOR HIGH THERMAL LOADS, A COMPONENT PRODUCIBLE BY THIS METHOD AND AN AIRCRAFT ENGINE PROVIDED WITH THE COMPONENT - A method for manufacturing a thermally deformable component for high thermal loads, includes: providing a first area of the component with a first metallic material by a generative laser process, or making the first area of the first metallic material; providing a second area of the component with a second metallic material by a generative laser process, or making the second area of the second metallic material; where at least one of the metallic materials is deposited by the generative laser process, and a ratio of a linear expansion coefficient α | 02-21-2013 |
20130042627 | COMBUSTION CHAMBER HEAD OF A GAS TURBINE WITH COOLING AND DAMPING FUNCTIONS - A combustion chamber head of a gas turbine has a substantially annular combustion chamber outer wall | 02-21-2013 |
20130025962 | Gas-turbine exhaust cone with three-dimensionally profiled partition wall and plate-type wall element - 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. | 01-31-2013 |
20130011246 | Gas-Turbine Aircraft Engine With Structural Surface Cooler - 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. | 01-10-2013 |
20130008147 | Aircraft gas turbine with variable bypass nozzle - The present invention relates to an aircraft gas turbine with a core engine | 01-10-2013 |
20120325354 | Apparatus and method for the creation of an impingement jet generating annular swirls as well as turbomachine with an apparatus of this type - The present invention relates to an apparatus and a method for the creation of an impingement jet (P) generating annular swirls ( | 12-27-2012 |
20120321477 | Rotor device for a jet engine with a disk wheel and several rotor blades - The present invention proposes a rotor device for a jet engine with a disk wheel and several rotor blades connected to said disk wheel, with the rotor blades being arranged in each case via a blade root substantially in the axial direction inside recesses of the disk wheel. Several locking segments are provided in the recesses of the disk wheel for axially locking the rotor blades, said locking segments interacting on the one hand with grooves of the rotor blades and on the other hand with at least one groove of the disk wheel. In the area of the groove of at least one of the rotor blades between the groove of the rotor blade and a locking segment there is a positive fit preventing a relative movement between the rotor blades and the locking segments. | 12-20-2012 |
20120304658 | Segment component in high-temperature casting material for an annular combustion chamber, annular combustion chamber for an aircraft engine, aircraft engine and method for the manufacture of an annular combustion chamber - The present invention relates to a segment component in high-temperature casting material for an annular combustion chamber of an aircraft engine, characterized by a combustion-chamber wall which in operation shields a fuel flame extending along a burner axis from the environment, with the combustion-chamber wall having a bulge which points in a direction facing away from the burner axis. The invention furthermore relates to an annular combustion chamber, an aircraft engine with an annular combustion chamber as well as a method for the manufacture of an annular combustion chamber. | 12-06-2012 |
20120298802 | De-icing device of an aircraft gas-turbine engine - 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. | 11-29-2012 |
20120288359 | Gas-turbine engine with bleed-air tapping device - The present invention relates to a gas-turbine engine with at least one compressor and at least one bleed-air tapping device, which includes an annular duct in a radially outer wall of a flow duct, and with an annular closing element, which is arranged in the region of the annular duct and can be moved in a substantially axial direction from a closed position to an open position, with the closing element having an annular flow divider projection which in the open position projects in the flow duct. | 11-15-2012 |
20120285138 | AIRCRAFT GAS-TURBINE ENGINE WITH OIL COOLER IN THE ENGINE COWLING - 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. | 11-15-2012 |
20120282082 | Gas-turbine balancing device - The present invention relates to a gas-turbine balancing device with at least one annular groove provided on the outer circumference of at least one intermediate or high-pressure rotor of an intermediate or high-pressure compressor, with at least one balancing element being arranged in the annular groove, and with the balancing element being moveable in the circumferential direction along the annular groove and provided with a fixing device. | 11-08-2012 |
20120269619 | FLUID-FLOW MACHINE - A fluid-flow machine includes at least one rotor having a rotary element with a plurality of rotor blades arranged on the rotary element, and a circumferential casing having a central axis and surrounding the rotor. The circumferential casing or a part connected thereto has an annular space surface on the inside, which delimits a flow duct of the fluid-flow machine radially outwards. The annular space surface has a structuring at least in one area adjoining a rotor on the circumferential side. At least one structuring of the annular space surface has, relative to the central axis of the circumferential casing, a circumferentially asymmetrical design. | 10-25-2012 |
20120266426 | Method and apparatus for surface strengthening and/or smoothing of an integrally bladed rotor area of a jet engine - The present invention describes a method and an apparatus ( | 10-25-2012 |
20120251324 | Rotor of an axial compressor stage of a turbomachine - The present invention relates to a rotor of an axial compressor stage of a turbomachine featuring a rotor assembly with a rotary axis, forming on its circumference a blade ring with a radially outer ring surface, and several rotor blades arranged on the blade ring. It is provided that the ring surface between two adjacent rotor blades has at least in a partial area a changing radius relative to the rotary axis of the rotor assembly both in the axial direction and in the circumferential direction. | 10-04-2012 |
20120251312 | Stator of an axial compressor stage of a turbomachine - The present invention relates to a stator of an axial compressor stage of a turbomachine featuring a radially outer blade ring forming an outer ring surface, a radially inner blade ring forming an inner ring surface, and several stator blades connected to the blade rings. It is provided that the outer ring surface and/or the inner ring surface has at least in a partial area a changing radius relative to a central axis of the stator both in the axial direction and in the circumferential direction. | 10-04-2012 |
20120247110 | DEVICE FOR MIXING FUEL AND AIR OF A JET ENGINE - A device for mixing fuel and air of a jet engine includes at least one air-carrying duct and at least one further fuel-carrying duct. Guide elements extending over the duct height and distributed over the circumference of the duct are provided at least in the air-carrying duct, in the area of which guide elements a twist can be imparted to the air flowing in the air-carrying duct in order to improve mixing between the air and the fuel downstream of the ducts. The guide elements are provided at a trailing edge with at least one downstream extending projection and/or upstream extending recess designed at least approximately or at least in some areas with a sharp edge. | 10-04-2012 |
20120240595 | COMBUSTION CHAMBER HEAD WITH HOLDING MEANS FOR SEALS ON BURNERS IN GAS TURBINES - A combustion chamber head of a gas turbine has a base plate | 09-27-2012 |
20120240583 | SEGMENTED COMBUSTION CHAMBER HEAD - A combustion chamber head for a gas turbine, with the gas turbine having a substantially annular outer combustion chamber wall, at least one substantially annular inner combustion chamber wall, and several burners | 09-27-2012 |
20120222396 | JET ENGINE DEVICE WITH A BYPASS DUCT - A jet engine has a bypass duct limited by an inner wall and an outer wall and inside which a fluid flows. Between the inner and outer walls of the bypass duct a support unit is provided that includes strut-like support elements connected at opposite ends to the inner and outer walls, respectively. Central longitudinal planes of the support elements describe in the areas of the support elements facing the inner wall a positive acute angle with an engine axis, and in the areas of the support elements facing the outer wall a negative acute angle with the engine axis. Flow cross-sections are each enlarged in the area between the side surfaces of the support elements each describing an acute angle with the walls, starting from the areas facing the fluid flow in the direction of the areas of the support elements facing away from the fluid flow. | 09-06-2012 |
20120204723 | Centrifugal oil separator for an aircraft engine - On a centrifugal oil separator for an aircraft engine for separating the oil particles contained in the vent air, a metallic foam body ( | 08-16-2012 |
20120174588 | GAS-TURBINE LEAN COMBUSTOR WITH FUEL NOZZLE WITH CONTROLLED FUEL INHOMOGENEITY - A gas-turbine lean combustor includes a combustion chamber ( | 07-12-2012 |
20120148396 | FLUID-FLOW MACHINE - BLADE WITH HYBRID PROFILE CONFIGURATION - A blade for use in a fluid-flow machine, where at least one of the meridional flow line profile sections MSLi and MSLo is provided at a distance of max. 35% of the main flow path width from the respective inner and outer main flow path boundaries. The blade has in at least parts of a central area a multi-profile configuration such that at least two partial profiles ( | 06-14-2012 |
20120125523 | REPAIR METHOD FOR A SANDWICH-LIKE COMPONENT - A gas-turbine casing includes a sandwiched structure of a wall | 05-24-2012 |
20120117973 | GAS-TURBINE COMBUSTION CHAMBER WITH A COOLING-AIR SUPPLY DEVICE - A gas-turbine combustion chamber has a cooling-air supply device delivering cooling air to a combustion chamber wall | 05-17-2012 |
20120094777 | DRIVE SHAFT, IN PARTICULAR RADIAL SHAFT FOR A GAS-TURBINE ENGINE - A drive shaft, in particular a radial shaft for a gas-turbine engine, includes a metallic and hollow shaft shank | 04-19-2012 |
20120076665 | COOLED TURBINE BLADES FOR A GAS-TURBINE ENGINE - The present invention relates to a cooled turbine blade for a gas-turbine engine having at least one cooling duct ( | 03-29-2012 |
20120067101 | FORGING TOOL - A forging tool for precision forging of components of intermetallic or high-temperature stable phases with high yield stresses and shapeable at temperatures up to 1400° C. is made of graphite with a low-melting metal or a low-melting metal alloy infiltrated into its open-pored cavities, where metal carbides are created by heat treatment and form with the graphite a two-phase material hardened by subsequent quenching. The tool features high strength thanks to the yield stress increasing as the temperature increases at forging temperatures up to 1400° C., and is oxidation-resistant. It is electrically conductive, and has a low heat capacity, so that rapid inductive heating of the tool involving only low energy expenditure, short forging cycles and an inexpensive isothermic shaping process are possible. It has good lubrication properties, low wear and low manufacturing costs. | 03-22-2012 |
20120067025 | BLOOM MIXER FOR A TURBOFAN ENGINE - An adjustable bloom mixer for a turbofan engine for setting a mixing ratio, adapted to the respective flight condition, between cold airflow in the bypass duct and hot airflow in the core flow duct includes an air-guiding element ( | 03-22-2012 |
20120045032 | METHOD FOR FILMLESS RADIOGRAPHOC INSPECTION OF COMPONENTS - 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. | 02-23-2012 |
20120034333 | Injection moulding tool for the manufacture of a hybrid component - Injection moulding tool for the manufacture of a plastic element sheathed with a sheet-metal shell ( | 02-09-2012 |
20120033787 | METHOD FOR RADIOGRAPHIC INSPECTION OF COMPONENTS - During the radiographic inspection of components by X-rays for better detection of defects, preferably automated, an uneven surface topography ( | 02-09-2012 |
20120020802 | COMPRESSOR BLADE OF A GAS-TURBINE ENGINE WITH A SELF-SHARPENING LEADING-EDGE STRUCTURE - A compressor blade | 01-26-2012 |
20120017663 | METHOD FOR THE MANUFACTURE OF A WORKPIECE WITH DEFINED SURFACE - A method for manufacturing a workpiece | 01-26-2012 |
20120014780 | FAN DOWNSTREAM GUIDE VANES OF A TURBOFAN ENGINE - Fan downstream guide vane profiles have an optimized form of skeleton line angle distribution in an area situated between an upper and a lower limitation as well as a specific thickness distribution superimposed on the respective skeleton line angle distribution. Such guide vanes are characterized by lower pressure losses and a larger working range than the known downstream guide vanes, thereby reducing fuel consumption of the engine and increasing the operating stability thereof. | 01-19-2012 |
20120011827 | BLEED AIR OUTLET IN THE BYPASS DUCT OF A TURBOFAN ENGINE - A bleed air outlet provided in the bypass duct of a turbofan engine includes a bleed air tube protruding into the bypass duct and a cover having a plurality of air outlet openings provided in the top of the cover, with the cover being conceived as an elongate, essentially oval, shell-like aerodynamic fairing element ( | 01-19-2012 |
20120006614 | GAS-TURBINE EXHAUST CONE - A gas-turbine exhaust cone includes a cone shaped outside cone | 01-12-2012 |
20110308966 | METHOD FOR MANUFACTURING BLISKS - A method for manufacturing a blisk for a gas turbine, in particular an aircraft gas turbine, includes generating blade profiles from an outer contour of a forged disk by milling and/or electrochemical machining. A robot-controlled, mechanical rework is performed of blade areas, in areas of the leading and trailing edges, the annulus, the fillet, the platform and the blade tip. A specified contour according to the engineering drawing of the blisk blade areas is referenced and an actual contour is determined by visualization. A difference between the specified contour and the actual contour is then calculated. The blade areas are then finish-machined and polished in a special processing machine according to a program prepared on the basis of the calculated difference. | 12-22-2011 |
20110288740 | Engine synchronization method - This invention relates to an engine synchronization method for aircraft equipped with at least two gas-turbine engines, where the respective speeds N | 11-24-2011 |
20110271837 | Centrifugal oil separator for an aircraft engine - A centrifugal oil separator ( | 11-10-2011 |
20110271765 | Method and apparatus for the detection of defects in the raceways of bearing shells and in the rolling elements of ceramic hybrid bearings - This invention relates to a method and an apparatus for the detection of defects in the raceways of bearing shells and the rolling elements of ceramic hybrid bearings. In order to enable the detection of such defects without falsifying the measuring values of the ceramic hybrid bearings installed, the present invention provides that the outer or inner ferromagnetic bearing shell (outer ring | 11-10-2011 |
20110271681 | LEAN PREMIX BURNER OF A GAS-TURBINE ENGINE PROVIDED WITH A FLOW-GUIDING ELEMENT - This invention relates to a lean premix burner of a gas-turbine engine with an annular central body | 11-10-2011 |
20110271680 | LEAN PREMIX BURNER OF A GAS-TURBINE ENGINE PROVIDED WITH A CONCENTRIC ANNULAR CENTRAL BODY - This invention relates to a lean premix burner of a gas-turbine engine with an annular central body | 11-10-2011 |
20110255964 | BYPASS DUCT OF A TURBOFAN ENGINE - On a turbofan engine, at least one of the downstream guide vanes ( | 10-20-2011 |
20110219782 | AERODYNAMICALLY SHAPED SUPPORTING AND/OR FAIRING ELEMENT IN THE BYPASS DUCT OF A GAS-TURBINE ENGINE - An aerodynamically shaped supporting and fairing element ( | 09-15-2011 |
20110215172 | Aircraft engine with optimized oil heat exchanger - This invention relates town ejector nozzle tube | 09-08-2011 |
20110211947 | Bypass duct of a turbo engine - A guide blade ring ( | 09-01-2011 |
20110209458 | Aircraft gas turbine engine - The invention refers to an aircraft gas turbine engine including a core engine | 09-01-2011 |
20110182499 | METHOD FOR DETERMINING THE SURFACE COVERAGE OBTAINED BY SHOT PEENING - In a method for determining the surface coverage obtained by shot peening to ensure uniform and complete strengthening of the surface of components, in particular blisk blades, a shot-peened surface topography is digitalized by an optical digital recording unit. A three-dimensional height profile is then prepared by measuring and evaluation software which includes both indentations and excrescences due to shot peening and also roughnesses due to manufacturing, which are smaller than the excrescences and indentations. The roughnesses are subsequently filtered out from the height image by a software filter using mathematical methods. A height diagram with the indentations situated below a zero line is established, with the size of these indentations being calculated in relation to the total area in the height diagram and the extent of coverage of the entire shot-peened surface being determined therefrom. | 07-28-2011 |
20110179844 | METHOD AND APPARATUS FOR SURFACE STRENGTHENING OF BLISK BLADES - For surface strengthening of blisk blades, a blade area of the blisk is completely inserted into a water bath ( | 07-28-2011 |
20110173990 | INTERMEDIATE CASING FOR A GAS-TURBINE ENGINE - An intermediate casing for a gas-turbine engine has an outer ring ( | 07-21-2011 |
20110146224 | ARRANGEMENT FOR THE DISCHARGE OF OIL-VENTING AIR ON A GAS-TURBINE ENGINE - 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 ( | 06-23-2011 |
20110146223 | ARRANGEMENT FOR THE DISCHARGE OF EXHAUST AIR SEPARATED FROM THE LUBRICATING OIL DE-AERATION SYSTEM OF A GAS-TURBINE ENGINE - With an arrangement for the discharge of oil-contaminated exhaust air separated from the lubricating oil de-aeration system of a gas-turbine engine and led to the atmosphere via a venting line, the venting line ( | 06-23-2011 |
20110139925 | ARRANGEMENT FOR THE SUSPENSION OF A JET ENGINE TO A SUPPORTING STRUCTURE - An arrangement ( | 06-16-2011 |
20110127728 | SEALING RINGS FOR A LABYRINTH SEAL - In sealing rings for a labyrinth seal that are arranged on a rotationally symmetrical component and in frictional contact with a stationary run-in layer, a part of the sealing ring ( | 06-02-2011 |
20110121136 | AIRCRAFT DE-ICING DEVICE AND ENGINE NACELLE OF AN AIRCRAFT GAS TURBINE WITH DE-ICING DEVICE - 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 | 05-26-2011 |
20110116908 | GAS TURBINE ENGINE WITH AN ARRANGEMENT FOR MEASURING THE SHAFT ROTATION SPEED - A gas turbine engine includes an arrangement for measuring shaft rotation speed, including a laser source ( | 05-19-2011 |
20110070074 | GAS TURBINE WITH A SHROUD AND LABYRINTH-TYPE SEALING ARRANGEMENT - A gas turbine with a shroud and labyrinth-type sealing arrangement has a casing | 03-24-2011 |
20110038666 | ENGINE SHAFT OF HYBRID DESIGN - On an engine shaft of hybrid design with an externally toothed power transmission element ( | 02-17-2011 |
20110027091 | Axial-flow compressor, more particularly one for an aircraft gas-turbine engine - On an axial-flow compressor, more particularly one for an aircraft gas-turbine engine, including at least one rotor disposed in a casing and having compressor blades extending from a rotor hub as well as one stator, a slot-type recess ( | 02-03-2011 |
20110016883 | Cross-sectional profile for the struts or the fairing of struts and service lines of a turbofan engine - The aerodynamically shaped, symmetrical cross-sectional profile for the struts or the fairing ( | 01-27-2011 |
20110014058 | PROPELLER - On a propeller, more particularly one for aircraft applications, an efflux slot ( | 01-20-2011 |
20110014057 | ENGINE BLADE WITH EXCESSIVE LEADING EDGE LOADING - A free end of a blade of a fluid flow machine has a skeleton line camber distribution having an excessive value to a relative skeleton line camber of at least α*=0.35 for a related running length of s*=0.1 in a blade profile flow line section between the free end and a blade section at 30% of a main flow path width from the free end. S* is a local running length relative to a total running length of the profile skeleton line and α* is an angular change of the skeleton line relative to a total camber of the skeleton line from a leading edge to a related running length s*. The skeleton line camber distribution runs between leading edge point V (s*=0, α*=0) and trailing edge point H (s*=1, α*=1). | 01-20-2011 |
20110014040 | FLUID FLOW MACHINE WITH BLADE ROW GROUP - A fluid flow machine has a main flow path in at least one stage, with a rotor arrangement and a stator arrangement, resulting in an increased rotor-stator constriction ratio QRS, which satisfies the following equation: | 01-20-2011 |
20110014037 | Axial-flow compressor with a flow pulse generator - Axial-flow compressor including, within a compressor casing ( | 01-20-2011 |
20110011093 | GAS-TURBINE COMBUSTION CHAMBER WITH STARTER FILM FOR COOLING THE COMBUSTION CHAMBER WALL - A gas turbine combustion chamber has a starter film for cooling the combustion chamber wall, and a combustion chamber head, into which cooling air can be introduced and which is confined to the combustion chamber by a heat shield ( | 01-20-2011 |
20110011058 | TURBOFAN ENGINE - With a turbofan engine the inner sidewall ( | 01-20-2011 |
20110005858 | NOISE-REDUCED AIRCRAFT ENGINE AND METHOD FOR REDUCING NOISE EMISSIONS OF AN AIRCRAFT ENGINE - A method for reducing noise emissions of an aircraft engine ( | 01-13-2011 |
20110005233 | COMBUSTION CHAMBER HEAD OF A GAS TURBINE - A combustion chamber head of a gas turbine has a confinement enclosing a dampening volume ( | 01-13-2011 |
20100303629 | FLUID FLOW MACHINE WITH A BLADE ROW GROUP FEATURING A MERIDIONAL EDGE DISTANCE - A main flow path of a fluid flow machine includes N adjacent member blade rows firmly arranged relative to each other in both a meridional direction m and a circumferential direction u. A number of the member blade rows, N, is greater than/equal to 2 and (i) designates a running index with values between 1 and N. A trailing edge HK (i) of a blade of the member blade row (i) is spaced from a leading edge VK(i+1) of a blade of the adjacent, downstream member blade row (i+1) by a meridional edge distance D in a meridional plane established by axial direction x and radial direction r. A value of D along a height of the main flow path increases towards the main flow path confinement at least along a part of the area between the main flow path center and the main flow path confinement. | 12-02-2010 |
20100288648 | METHOD AND APPARATUS FOR ETCHING THE SURFACES OF INTEGRALLY BLADED ROTORS - For electrolytically etching the surfaces of integrally bladed rotors (blisks) ( | 11-18-2010 |
20100287772 | METHOD FOR SURFACE STRENGTHENING AND SMOOTHENING OF METALLIC COMPONENTS - With a method for surface strengthening and smoothening of metallic components, in particular of rotors or rotor drums ( | 11-18-2010 |
20100259013 | ABRADABLE LABYRINTH SEAL FOR A FLUID-FLOW MACHINE - An abradable labyrinth seal for a fluid-flow machine seals a sealing gap between a stationary carrier ( | 10-14-2010 |
20100258199 | INTAKE CONE IN A FIBER COMPOUND MATERIAL FOR A GAS-TURBINE ENGINE - The intake cone for a gas-turbine engine is wound in one piece from fiber compound material with fiber layers crossing one another. It is provided with a fiber compound belt, wound in the circumferential direction of the intake cone, in the connecting area with a mounting flange, which is of segmented design and thus features reduced circumferential stiffness, of a metallic retaining ring attached to a fan rotor disk. The intake cone is easily manufacturable with almost constant wall thickness and in high quality, can reliably be subjected to non-destructive testing and, in combination with the mounting flange, which is flexible in the radial direction, ensures safe connection of the two components despite the different thermal and elastic behavior of the respective materials. | 10-14-2010 |
20100232954 | BYPASS DUCT OF A TURBOFAN ENGINE - In the area of the support struts and/or the aerodynamic fairings downstream of the stator vanes, the cross-section of the bypass duct of a turbofan engine is enlarged such that the pressure variations caused by the stagnation effect of the installations and reacting on the fan are reduced, enabling the fan to be operated with more efficiency and stability and finally the losses of the overall system and the fuel consumption to be reduced. The cross-sectional enlargement is accomplished by modifying the course of the wall in a limited area, actually by gradually enlarging the flow cross-section in the bypass duct in the axial and in the circumferential direction, with this enlargement being confined to the area around the leading edge of the support struts and/or the aerodynamic fairings. | 09-16-2010 |
20100224348 | METHOD FOR THE MANUFACTURE OF AN EJECTOR NOZZLE TUBE - A method for manufacturing an ejector nozzle tube, includes forming an essentially rectangular plate-type blank | 09-09-2010 |
20100223905 | SCOOP OF A RUNNING-GAP CONTROL SYSTEM OF AN AIRCRAFT GAS TURBINE - A scoop for a fairing | 09-09-2010 |
20100215481 | RUNNING-GAP CONTROL SYSTEM OF AN AIRCRAFT GAS TURBINE - A running-gap control system of an aircraft gas turbine with a core engine including a turbine whose blade rows have a running gap | 08-26-2010 |
20100212285 | TURBOPROP PROPULSION UNIT WITH PUSHER PROPELLER - A turboprop propulsion unit includes at least one pusher propeller | 08-26-2010 |
20100192588 | METHOD FOR THE PROVISION OF A COOLING-AIR OPENING IN A WALL OF A GAS-TURBINE COMBUSTION CHAMBER AS WELL AS A COMBUSTION-CHAMBER WALL PRODUCED IN ACCORDANCE WITH THIS METHOD - A cooling-air opening in produced in a wall ( | 08-05-2010 |
20100176097 | ARRANGEMENT FOR THE REPAIR OF THE BLADES OF BLISK DRUMS BY MEANS OF LASER DEPOSITION WELDING - An arrangement for the repair of BLISK blades damaged at their leading and trailing edges by use of laser deposition welding, includes a laser source and laser optics for generating a focused laser beam ( | 07-15-2010 |
20100175256 | METHOD FOR THE MANUFACTURE OF THE BLADE TIPS OF ROTOR WHEELS MADE IN BLISK DESIGN - With a method for finish-machining the blade tips of rotor wheels made in BLISK design, the rotor wheel or multi-stage compressor rotor ( | 07-15-2010 |
20100161107 | MACHINE CONTROL FOR A NUMERICALLY CONTROLLED MACHINE TOOL - On a numerically controlled machine tool a sub-program is integrated into the given program for controlling the machine tool by which a periodic change of at least one cutting tool parameter is effected in one processing cycle, or also in several processing cycles, or in certain length and diameter zones of the workpiece characterized by tool vibrations and leading to inferior workpiece quality. By way of the periodic cutting parameter change, self-excited tool vibrations are avoided in these sensitive machining areas, thereby enabling tool wear to be reduced and workpiece quality improved. Furthermore, this intervention into the machining program is documentable and identically reproducible for each workpiece—independently of the expertise and personal judgement of the operator. | 06-24-2010 |
20100155526 | AIRCRAFT WITH TAIL PROPELLER-ENGINE LAYOUT - An aircraft propeller-engine layout includes at least one engine ( | 06-24-2010 |
20100150696 | FAN CASING FOR A JET ENGINE - A fan casing ( | 06-17-2010 |
20100148449 | SEGMENTED SEALING LIPS FOR LABYRINTH SEALING RINGS - A labyrinth sealing ring for a labyrinth seal of a turbine disk | 06-17-2010 |
20100148027 | SENSOR BRACKET FOR AT LEAST ONE SENSOR ON A GAS TURBINE - A sensor bracket ( | 06-17-2010 |
20100143140 | FLUID FLOW MACHINE WITH SIDEWALL BOUNDARY LAYER BARRIER - A fluid flow machine has a main flow path in which at least one row of blades ( | 06-10-2010 |
20100140230 | METHOD FOR THE MANUFACTURE OF A WELDED ROTOR FOR A GAS-TURBINE ENGINE - With a method for manufacturing welded rotors for a turbine, especially a gas-turbine engine, in which two or more rotor disks are joined to each other by conventional welding processes using welds extending radially to the rotor axis and the weld zone is subsequently thermally treated at a certain temperature to relieve residual tensile stresses by relaxation, the weld is set to a significantly lower non-relaxatory temperature level than the heat-affected zone adjoining the weld so that, as a result of the high temperature gradient, a residual compressive stress or at least a substantially reduced residual tensile stress is impressed on the weld. Compared to conventionally heat treated and welded rotors, improved strength properties in the weld zone and an increased service life are obtained as a result of the reduced tensile stresses. | 06-10-2010 |
20100139278 | METHOD AND APPARATUS FOR THE OPERATION OF A TURBOPROP AIRCRAFT ENGINE PROVIDED WITH PUSHER PROPELLERS - A method and an apparatus is disclosed for the operation of a turboprop aircraft engine provided with pusher propellers. In order to reduce thermal loading of the pusher propellers impaired by the hot exhaust-gas flow of the engine and increase the service life of the pusher propellers, cold air from the environment outside of the aircraft engine is fed into, and mixed with, the hot exhaust-gas flow passing the pusher propellers and their connecting structure before the hot exhaust-gas flow reaches the pusher propellers. | 06-10-2010 |
20100139241 | FLOW DIVIDER FOR A FAN ENGINE - A flow divider ( | 06-10-2010 |
20100129651 | HYBRID COMPONENT FOR A GAS-TURBINE ENGINE - A hybrid component for a gas-turbine engine includes a supporting structure ( | 05-27-2010 |
20100126662 | METHOD FOR THE MANUFACTURE OF HYBRID COMPONENTS FOR AIRCRAFT GAS TURBINES - When manufacturing hybrid components, especially the fan blades or stator vanes of an aircraft gas turbine including a metallic enveloping structure and a supporting structure in fiber-composite material, the mating faces of the metallic structure are cleaned and roughened with plasma prior to fitting the fiber-composite material and the metallic molecules activated so that high attractive forces take effect and an intimate, extremely strong adhesive connection is produced between the metal and the fiber-composite material. Fan blades or stator vanes of the fan structure of a gas-turbine engine which are manufactured according to this method are capable of transmitting high loads and feature a long service life. | 05-27-2010 |
20100126589 | DEVICE FOR THE PREVENTION OF HIGH OIL TANK PRESSURES UNDER NEGATIVE G CONDITIONS - An aircraft gas-turbine oil tank has a breather outlet tube | 05-27-2010 |
20100124487 | MULTI-VANE VARIABLE STATOR UNIT OF A FLUID FLOW MACHINE - A fluid flow machine has a main flow path which is confined by a hub ( | 05-20-2010 |
20100122518 | OIL SYSTEM HEATING FOR AIRCRAFT GAS TURBINES - An oil pre-heating apparatus for an aircraft gas turbine has a suction line | 05-20-2010 |
20100113171 | ENGINE SHAFT IN THE FORM OF A FIBER-COMPOSITE PLASTIC TUBE WITH METALLIC DRIVING AND DRIVEN PROTRUSIONS - On a low-pressure turbine shaft made of fiber-composite material, the connection to the metallic driven protrusion or the driving protrusion ( | 05-06-2010 |
20100113170 | ENGINE SHAFT FOR A GAS-TURBINE ENGINE - A low-pressure turbine shaft ( | 05-06-2010 |
20100111702 | HUB CONE FOR AN AIRCRAFT ENGINE - A hub cone ( | 05-06-2010 |
20100109184 | METHOD FOR THE MANUFACTURE OF AN ENGINE SHAFT - For the manufacture of a tubular low-pressure turbine shaft made of a fiber-composite material with metallic driven/driving protrusions ( | 05-06-2010 |
20100098536 | FLUID FLOW MACHINE WITH RUNNING GAP RETRACTION - A fluid flow machine includes a main flow path which is confined by a hub ( | 04-22-2010 |
20100098530 | COMPRESSOR FOR A GAS TURBINE - A compressor for a gas turbine, in particular an aircraft gas turbine, has a rotor hub carrying rotor blades, a stator equipped with stator vanes, a shroud associated to the stator vanes, and an arrangement providing sealing between the shroud and rotor hub to prevent leakage. To achieve almost complete suppression of leakage air and, concurrently, simplification of design and manufacture, the sealing arrangement ( | 04-22-2010 |
20100098527 | FLUID FLOW MACHINE WITH PERIPHERAL ENERGIZATION NEAR THE SUCTION SIDE - A fluid flow machine has a main flow path (“MFP”) | 04-22-2010 |
20100051112 | INTAKE CONE FOR A GAS-TURBINE ENGINE - An intake cone for a gas-turbine engine has an essentially conical body | 03-04-2010 |
20100034637 | FLUID FLOW MACHINE - A fluid flow machine has a main flow path | 02-11-2010 |
20100024214 | METHOD FOR THE MANUFACTURE OF METALLIC COMPONENTS - When manufacturing slender metallic components ( | 02-04-2010 |
20100021310 | METHOD FOR IMPROVING THE FLOW CONDITIONS ON THE PROPELLER OR FAN OF AN AIRCRAFT ENGINE AND ACCORDINGLY DESIGNED HUB CONE - At a propeller ( | 01-28-2010 |
20100014960 | GAS-TURBINE ENGINE WITH VARIABLE STATOR VANES - A gas-turbine engine includes stator vanes ( | 01-21-2010 |
20100014956 | FLUID FLOW MACHINE FEATURING A GROOVE ON A RUNNING GAP OF A BLADE END - A fluid-flow machine has at least one row of blades | 01-21-2010 |
20100005776 | FUEL SUPPLY SYSTEM FOR A GAS-TURBINE ENGINE - On a fuel supply system for a gas-turbine engine with a high-pressure supply line and a fuel metering unit, the fuel distribution system connected to staged or non-staged burners includes only one single fuel line ( | 01-14-2010 |
20100000226 | TURBOFAN ENGINE WITH AT LEAST ONE APPARATUS FOR DRIVING AT LEAST ONE GENERATOR - A turbofan engine ( | 01-07-2010 |
20100000198 | GAS TURBINE WITH AT LEAST ONE MULTI-STAGE COMPRESSOR UNIT INCLUDING SEVERAL COMPRESSOR MODULES - A gas turbine includes at least one multi-stage compressor unit | 01-07-2010 |
20090317232 | BLADE SHROUD WITH APERTURE - A fluid flow machine includes a main flow path in which at least one row of blades ( | 12-24-2009 |
20090314004 | TURBOPROP ENGINE WITH AN APPARATUS FOR THE GENERATION OF A COOLING AIRFLOW - A turboprop engine ( | 12-24-2009 |