Abstract: A guide vane element for a gas turbine, the guide vane element having a first guide vane, a second guide vane distanced by one division in the peripheral direction, and at least one band joining these guide vanes, in particular, a radially inner band and/or a radially outer band, wherein at least one band joining these guide vanes has a vane-side surface having a contouring and a first front side in the peripheral direction having a groove, which is particularly straight in the axial direction, for the uptake of a sealing element.

Abstract: A low pressure turbine engine component for use in an engine, for propelling a vehicle such as an aircraft is formed by a disk portion and a plurality of low pressure turbine blades extending outwardly from the disk portion. Each of the low pressure turbine blades has an airfoil portion with an axial chord length and a trailing edge. The low pressure turbine blades are spaced apart so that there is a pitch-to-chord ratio greater than 1.4, wherein the pitch is a distance between the trailing edges of adjacent ones of the low pressure turbine blades and the chord is the axial chord length of the blades.

Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

Abstract: 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 (1) for deflecting a flow of working fluid and a first platform (3) connected thereto, in particular integrally connected thereto, to radially bound a flow duct for the working fluid, the airfoil having a suction side and a pressure side (1.1) which are connected at a leading edge (1.2) and at a trailing edge (1.3). The trailing edge has a first minimum wall thickness in a first region (A) of a radial longitudinal extent (R) of the airfoil proximal to the first platform (3), and a maximum wall thickness that is smaller than the first minimum wall thickness in a platform-distal region (C) of the radial longitudinal extent.

Abstract: A blade for a gas turbine has a leading edge and a trailing edge, which are connected by a pressure side and an intake side. In at least one segment of a stacking axis, the blade has a cross section with a profile with a common profile tangent at a leading edge region and a trailing edge region, a leading edge tangent at the leading edge, which is perpendicular to the common profile tangent, a trailing edge tangent at the trailing edge, which is perpendicular to the common profile tangent, and a camber line, which extends, at an equal distance from the pressure side and the intake side, from a center point of a leading edge circle, up to a center point of a trailing edge circle.

Abstract: 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 (?) 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 (w2/w1) is at least 1.4, in particular at least 1.5.

Abstract: The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. The mean outer radius of the last stage of the second turbine divided by the length of the second turbine is at least 1.4.

Abstract: The invention relates to a turbofan aircraft engine that comprises a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled to a fan for feeding a secondary duct of the turbofan aircraft engine. The bypass ratio of the inlet area of the secondary duct to the inlet area of the primary duct is at least 7 and the second turbine comprises at least two stages. For the first stage the mean radius r of a stator vane expressed in inch divided by the number of stator vanes is at least 0.18.

Abstract: 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.

Abstract: 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.

Abstract: The present invention relates to a guide vane element for a gas turbine, having a first guide vane (1), a second guide vane (2) distanced by one division (T) in the peripheral direction, and at least one band (3) joining these guide vanes, in particular, a radially inner band and/or a radially outer band, wherein at least one band (3) joining these guide vanes (1, 2) has a vane-side surface having a contouring and a first front side (4) in the peripheral direction having a groove (5), which is particularly straight in the axial direction, for the uptake of a sealing element (6) with other details in accordance with the present invention.

Abstract: 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.

Abstract: 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 ([(p1/p2)?1]) of the total pressure ratio minus one, and the total pressure ratio is greater than 4.5, and the turbine has at least two and no more than five turbine stages; and/or a product (An2) of an exit area (AL) of the second turbine and a square of a rotational speed of the second turbine at the design point is at least 4.5·1010 [in2·rpm2], and a blade tip velocity (uTIP) of at least one turbine stage of the second turbine at the design point is at least 400 meters per second. A jet and method are also provided.

Abstract: A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W1) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W2) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (DF) of the fan to a maximum blade diameter (DL) of the second turbine is at least 3.5, in particular at least 4.

Abstract: The present invention relates to a blade (1) for a gas turbine, having a leading edge and a trailing edge, which are connected by a pressure side and an intake side, wherein, in at least one segment of a stacking axis, the blade has a cross section with a profile with a common profile tangent (P) at a leading edge region and a trailing edge region, a leading edge tangent (V) at the leading edge, which is perpendicular to the common profile tangent, a trailing edge tangent (H) at the trailing edge, which is perpendicular to the common profile tangent, and a camber line (S), which extends, at an equal distance from the pressure side and the intake side, from a center point (MV) of a leading edge circle (KV), up to a center point (MH) of a trailing edge circle.

Abstract: The invention refers to polynucleotides selected from the group consisting of a) polynucleotides encoding for the polypeptide RBM20 comprising a P638L mutation for a human polypeptide RBM20, or a P641L mutation for a rat polypeptide RBM20, b) polynucleotides with a reverse complementary sequence of the polynucleotide of a) above, and c) polynucleotides with an identity at least 50% to a polynucleotide of a) or b) above.

Abstract: 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 (1) for deflecting a flow of working fluid and a first platform (3) connected thereto, in particular integrally connected thereto, to radially bound a flow duct for the working fluid, the airfoil having a suction side and a pressure side (1.1) which are connected at a leading edge (1.2) and at a trailing edge (1.3). The trailing edge has a first minimum wall thickness in a first region (A) of a radial longitudinal extent (R) of the airfoil proximal to the first platform (3), and a maximum wall thickness that is smaller than the first minimum wall thickness in a platform-distal region (C) of the radial longitudinal extent.

Abstract: 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.

Abstract: 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.

Abstract: 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.