Patent application number | Description | Published |
20080286107 | Course of leading edges for turbomachine components - The course of the leading edges of turbomachine components, such as rotor blades and stator vanes is defined mathematically exactly and repeatedly as well as aerodynamically advantageously by the respective axial coordinate in the direction of the machine axis in relation to the blade height in percent, extending from the blade tip as per equation (1): | 11-20-2008 |
20090044543 | Jet engine with compressor air circulation and method for operating the jet engine - In a jet engine with compressor air circulation provided for stabilizing the flow conditions, the compressed hot compressor air tapped from the flow path of the compressor ( | 02-19-2009 |
20090208324 | Casing structure for stabilizing flow in a fluid-flow machine | 08-20-2009 |
20090238677 | Centrifugal compressor with air extraction and return at the casing - A centrifugal compressor includes a rotor | 09-24-2009 |
20090238682 | Compressor stator with partial shroud - A gas-turbine axial compressor has a casing | 09-24-2009 |
20090301102 | Gas-turbine compressor with bleed-air tapping - A gas-turbine compressor has a casing | 12-10-2009 |
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 |
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 |
20100111702 | HUB CONE FOR AN AIRCRAFT ENGINE - A hub cone ( | 05-06-2010 |
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 |
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 |
20110011058 | TURBOFAN ENGINE - With a turbofan engine the inner sidewall ( | 01-20-2011 |
20110014037 | Axial-flow compressor with a flow pulse generator - Axial-flow compressor including, within a compressor casing ( | 01-20-2011 |
20110014058 | PROPELLER - On a propeller, more particularly one for aircraft applications, an efflux slot ( | 01-20-2011 |
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 |
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 |
20110211947 | Bypass duct of a turbo engine - A guide blade ring ( | 09-01-2011 |
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 |
20110255964 | BYPASS DUCT OF A TURBOFAN ENGINE - On a turbofan engine, at least one of the downstream guide vanes ( | 10-20-2011 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
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 |
20150027127 | COMBUSTION CHAMBER TILE OF A GAS TURBINE - The present invention relates to a combustion chamber tile of a gas turbine with a plate-like basic element which is provided with at least one effusion cooling hole extending through the basic element from a surface of one side to the other side, with the effusion cooling hole being designed, from the one side of the basic element and beginning from an inlet opening, substantially at right angles to the surface over part of its length, wherein the effusion cooling hole inside the basic element has a straight section and is then provided with an offset section. | 01-29-2015 |
20150030787 | LASER DEPOSITIONING DEVICE AND METHOD FOR PRODUCING A COMPONENT BY DIRECT LASER DEPOSITIONING - The present invention relates to a laser depositioning device with a machine bed, onto the surface of which can be deposited a powder material, using at least one first laser for layer by layer melting of the powder material, characterized in that at least one second laser is provided to melt on the powder material layer by layer, with the second laser having a higher power than the first laser, as well as to a method for producing a component provided with a supporting structure. | 01-29-2015 |