Abstract: A geared turbofan engine includes a plurality of turbine stages, wherein for each stage (i) of the turbine, an inner radius Ri has a maximum deviation between +1.5% and ?3% as compared to the average inner radius of the inner blade platforms of the plurality of stages. The engine further includes a fan, the fan coupled to the turbine stages via a gear.

Abstract: A guide vane ring and/or a rotor blade ring for a low-pressure turbine of an aircraft gas turbine of a turbofan engine includes a plurality of airfoils disposed consecutively in a direction of rotation. The plurality of airfoils includes a first airfoil and a second airfoil adjacent to the first airfoil in the direction of rotation. The first and second airfoils are disposed relative to each other such that a ratio e/t, viewed in a conical section, results from a smallest free flow cross-section e taken on the pressure side of the first airfoil from its trailing edge to a suction side of the second airfoil, and a pitch t of the airfoils, such that e/t is less than 0.31, (e/t<0.31).

Abstract: The present invention relates to a method for designing a flow channel for a turbomachine, in particular a gas turbine that comprises a guide vane cascade having a plurality of guide vanes, which are distributed in the peripheral direction, and flow passages, each of which is bounded by two successive guide vanes, and a support rib arrangement having at least one support rib, wherein a design of one of the flow passages is adapted to this support rib, that it is situated downstream of, in order to reduce a pressure loss and/or a vibrational stimulation.

Abstract: The invention relates to a flow arrangement for placing in the hot gas duct of a turbomachine, having a first surrounding-flow structure and a second surrounding-flow structure, the surrounding-flow structures each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge, wherein the second surrounding-flow structure is provided as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure, and wherein, although the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure.

Abstract: A flow channel for a turbomachine, in particular a gas turbine, having a plurality of ribs that are disposed between a radially inner lateral surface and a radially outer lateral surface of the flow channel and are circumferentially distributed; a first rib of the ribs having a first rib thickness and a first rib length, and a second rib of the ribs having a second rib thickness and a second rib length, the second rib length being shorter than the first rib length, and/or the second rib thickness being smaller than the first rib thickness; a spacing in the circumferential direction between the first rib and the second rib adjacent thereto, and a spacing in the circumferential direction between at least two adjacent ribs of the ribs mutually deviating, and at least one of these ribs having a non-deflecting external profile and or an internal structure being disposed therein.

Abstract: A blade (10) for a turbomachine is provided. In at least one cross section disposed orthogonally to a longitudinal blade axis, an outer surface of the blade forms a curve (?) that runs continuously on the inflow side along an at least approximate ellipse (E) between a first separation point (P1) and a second separation point (P2). On the pressure side 12, the curve (?) has an inflection point W. A distance of the second separation point (P2) from the inflection point W along the curve is at least exactly equal, at least twice as great, at least three times as great, or even at least five times as great as a distance of the second separation point (P2) from the first separation point (P1) along the at least approximate ellipse (E). Also described is a blade ring having at least one blade of this kind, as well as a turbomachine.

Abstract: The present invention relates to a turbine module for a turbomachine, comprising a first flow bypass structure and a second flow bypass structure, said flow bypass structures being arranged in a hot gas duct, which is bounded by the turbine module and is designed to convey a hot gas and, namely, being arranged in succession in relation to a longitudinal axis of the turbine module in a direction of rotation, wherein, in relation to the bypass flow in the hot gas duct, the flow bypass structures each have a leading edge, and, downstream thereto, a trailing edge, and the second flow bypass structure is provided as a deflecting blade, wherein the second flow bypass structure has a smaller profile thickness than the first flow bypass structure, and wherein the hot gas duct is enclosed by a radial width.

Abstract: The present invention relates to a turbine module for a turbomachine, comprising a first flow bypass structure and a second flow bypass structure, said flow bypass structures being arranged in a hot gas duct, which is bounded by the turbine module and is designed to convey a hot gas and, namely, being arranged in succession in relation to a longitudinal axis of the turbine module in a direction of rotation, wherein, in relation to the bypass flow in the hot gas duct, the flow bypass structures each have a leading edge, and, downstream thereto, a trailing edge, and the second flow bypass structure is provided as a deflecting blade, wherein the second flow bypass structure has a smaller profile thickness than the first flow bypass structure, and wherein the hot gas duct is enclosed by a radial width.

Abstract: A flow channel for a turbomachine, in particular a gas turbine, having a plurality of ribs (11-23) that are disposed between a radially inner lateral surface (2) and a radially outer lateral surface (1) of the flow channel and are circumferentially distributed; a first rib (11) of the ribs having a first rib thickness (d11) and a first rib length (l11), and a second rib (12) of the ribs having a second rib thickness (d12) and a second rib length (l12), the second rib length being shorter than the first rib length, and/or the second rib thickness being smaller than the first rib thickness; and, disposed within the first rib is a first internal structure (31) which has a first structural thickness (d31), and disposed within the second rib is a second internal structure (32) which has a second structural thickness (d32) that is smaller than the first structural thickness; and/or a spacing (T12) in the circumferential direction between the first rib and the second rib adjacent thereto, and a spacing (T23) in the

Abstract: The present invention relates to a method for designing a flow channel for a turbomachine, in particular a gas turbine that comprises a guide vane cascade having a plurality of guide vanes, which are distributed in the peripheral direction, and flow passages, each of which is bounded by two successive guide vanes, and a support rib arrangement having at least one support rib, wherein a design of one of the flow passages is adapted to this support rib, that it is situated downstream of, in order to reduce a pressure loss and/or a vibrational stimulation.

Abstract: A blade (10) for a turbomachine is provided. In at least one cross section disposed orthogonally to a longitudinal blade axis, an outer surface of the blade forms a curve (?) that runs continuously on the inflow side along an at least approximate ellipse (E) between a first separation point (P1) and a second separation point (P2). On the pressure side 12, the curve (?) has an inflection point W. A distance of the second separation point (P2) from the inflection point W along the curve is at least exactly equal, at least twice as great, at least three times as great, or even at least five times as great as a distance of the second separation point (P2) from the first separation point (P1) along the at least approximate ellipse (E). Also described is a blade ring having at least one blade of this kind, as well as a turbomachine.

Abstract: The invention relates to a flow arrangement for placing in the hot gas duct of a turbomachine, having a first surrounding-flow structure and a second surrounding-flow structure, the surrounding-flow structures each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge, wherein the second surrounding-flow structure is provided as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure, and wherein, although the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure.

Abstract: A gas turbine (40) having a high-speed low-pressure turbine (24) and a turbine case (28) that bounds a flow path of a working fluid of the gas turbine (40) and an exit region (30), and extends between a rotor (32) of the high-speed low-pressure turbine (24) that is the most downstream in the through flow direction of the working fluid, and an exit opening (34) of the turbine case (28). The exit region (30) is designed to be free of exit guide vane assemblies.

Abstract: A turbomachine is described, which includes an annular space seal for at least reducing a fluid exchange between an annular space, through which a main stream flows, and at least one cavity situated radially on the inside and/or radially on the outside of the annular space, the annular space seal having a plurality of sheet-like elastic elements, which are oriented in the radial direction of the turbomachine and in the flow direction of a main stream flowing through the particular guide blade row, and thereby acting as a protective device for the rotor and/or stator during a relative axial displacement with respect to the stator in addition to providing an improved cavity coverage.

Abstract: The present invention relates to an airfoil for a fluid flow machine (100), having a suction side (5), a pressure side (7) and an airfoil trailing edge (200). The airfoil (100), at least in portions thereof, has a profile (9) in the region of the airfoil trailing edge (200), which profile extends over the suction side (5) and the pressure side (7) of the airfoil trailing edge (200). The present invention also relates to a blade and an integrally bladed rotor.

Abstract: A turbomachine is described, which includes an annular space seal for at least reducing a fluid exchange between an annular space, through which a main stream flows, and at least one cavity situated radially on the inside and/or radially on the outside of the annular space, the annular space seal having a plurality of sheet-like elastic elements, which are oriented in the radial direction of the turbomachine and in the flow direction of a main stream flowing through the particular guide blade row, and thereby acting as a protective device for the rotor and/or stator during a relative axial displacement with respect to the stator in addition to providing an improved cavity coverage.

Abstract: Disclosed is a blade for a gas turbine, in particular for an aircraft engine, wherein the blade comprises at least one internal and/or external supporting structure of a metallic and/or intermetallic material. Also disclosed are a blade cluster having at least one such blade for a gas turbine, a method for producing such a blade cluster and a gas turbine.

Abstract: The present invention relates to an airfoil for a fluid flow machine (100), having a suction side (5), a pressure side (7) and an airfoil trailing edge (200). The airfoil (100), at least in portions thereof, has a profile (9) in the region of the airfoil trailing edge (200), which profile extends over the suction side (5) and the pressure side (7) of the airfoil trailing edge (200). The present invention also relates to a blade and an integrally bladed rotor.

Abstract: A gas turbine is disclosed. The gas turbine includes a rotor which is driven by a turbine, a stator, struts that are fixed to the stator downstream from the turbine and that configure a guide ring for deflecting the rotational flow of hot gas, and a mixer arranged on the downstream end of the hot gas channel. The guide ring and the mixer are structurally and fluidically combined, the struts of the guide ring being connected to the wall structure of the mixer in the region of their radially outer ends.