Patent application number | Description | Published |
20120156047 | AIRCRAFT ENGINE BLADING - A blading for a turbine, in particular a gas turbine, is disclosed. The blades of the blading in a section near the tip have a distribution ratio (t/l) of at least 0.70, in particular at least 0.9, and/or at most 0.97, in particular at most 0.95. A downstream flow angle (α) is at most 167°, in particular at most 165°, and at least 155°, in particular at least 160°. In addition, or alternatively, an acceleration ratio (w | 06-21-2012 |
20120275922 | HIGH AREA RATIO TURBINE VANE - A vane for a turbine engine comprises an airfoil section, an inner platform and an outer platform. The airfoil section comprises pressure and suction surfaces extending from a leading edge to a trailing edge. The inner platform is attached to the airfoil section along an inner flow boundary, where the inner flow boundary extends from an upstream inlet region of the vane to a downstream outlet region of the vane. The outer platform is attached to the airfoil section along an outer flow boundary, where the outer flow boundary extends from the inlet region to the outlet region. An area ratio of the outlet region to the inlet region is greater than 2.4. | 11-01-2012 |
20130272884 | BLADE FOR A TURBOMACHINE, BLADE ARRANGEMENT, AND TURBOMACHINE - A blade for a turbomachine, in particular a jet engine, including a shroud, having two opposite lateral edges, for delimiting a main flow channel and including a blade which extends away from the shroud, a rounded transition area being provided which encompasses the blade on its root side and is guided beyond the one lateral edge, a section of the transition area protruding beyond the one lateral edge being severed and situated in the area of the other lateral edge as an elevation offset in the transverse direction, a blade arrangement having at least two of such blades as well as a turbomachine having a plurality of such blades. | 10-17-2013 |
20130287579 | BLADE OF A TURBOMACHINE, HAVING PASSIVE BOUNDARY LAYER CONTROL - A blade of a turbomachine is disclosed. A contour variation is provided on the suction side of the blade, where the contour variation has a negative step as viewed in a direction of flow. The step has a stepped surface extending perpendicularly to a contour of the suction side and the contour variation has a tangential surface which leads upstream tangentially on the contour of the suction side starting from a step edge. | 10-31-2013 |
20150016985 | GAS TURBINE STAGE - A gas turbine stage including a rotor blade array having a plurality of rotor blades and an adjacent stator vane array having a plurality of stator vanes which have leading edges facing the rotor blade array. In a first radial position of a rear face of the rotor blade array, a minimum axial gap is formed between this rear face and an opposite first contact region a stator vane leading edges, and in a second radial position of the rear face different from the first position, the minimum axial gap is formed between the rear face and an opposite second contact region. Between the first and second contact regions, this stator vane leading edge has an axial offset of no more than 0.6% of a radial height of the stator vane leading edge. | 01-15-2015 |
20150017011 | BLADE FOR A GAS TURBOMACHINE - A blade, in particular a rotor blade or a stator vane, for a gas turbomachine, in particular a turbojet engine, the blade having an airfoil ( | 01-15-2015 |
20150285080 | Unknown - The present invention relates to a blade ( | 10-08-2015 |
20160032826 | TURBOFAN AIRCRAFT ENGINE - A turbofan aircraft engine has at least one stage pressure ratio is at least 1.5, and a quotient of the total blade count divided by 110 is less than a difference ([(p | 02-04-2016 |