Abstract: Disclosed is an airfoil array segment (110, 120, 130, 140, 150) of an airfoil array for a turbomachine, the airfoil array segment including a platform (10) having a platform surface, as well as at least two airfoils (20, 30). The platform surface has a depression (111, 121, 131, 141, 151) which extends up to the first airfoil and contacts the pressure side (21) of the first airfoil downstream of 80% of the axial chord (g) downstream of the leading edges (23, 33) and which contacts the pressure side (21) of the first airfoil downstream up to no more than 80% of the axial chord (g) downstream of the leading edges (23, 33). At least one lowest point (112, 122) of the depression is located at least 90% of the axial chord (g) downstream of the leading edges (23, 33). Also disclosed are an airfoil array, an airfoil passage, a platform, a turbomachine, and an aircraft engine.

Abstract: A blade/vane cascade segment of a blade/vane cascade for a turbomachine is disclosed, which comprises a stage with a stage surface as well as a first blade/vane element and a second blade/vane element. The stage surface comprises a first elevation reaching as far as the pressure side of the first blade/vane element and a second elevation reaching as far as the suction side of the second blade/vane element. A furthest downstream point of a boundary of the second elevation has an axial position which differs from the axial position of at least one highest point of the first elevation by a maximum of 10% of an axial cascade span. Furthermore, a blade/vane cascade, a blade/vane channel, a stage, a turbomachine and an aircraft engine are disclosed.

Abstract: A blade/vane cascade segment for a blade/vane cascade of a turbomachine includes a stage and at least two blade/vane elements that define a blade/vane intermediate strip with an axial cascade span on the stage surface by their leading and trailing edges. A stage edge on the inflow side has a contour with a depression. In the axial direction, this depression extends at most by 10% of the cascade span into the blade/vane intermediate strip. Also provided is a corresponding stage, a blade/vane cascade, a blade/vane channel, and a turbomachine.

Abstract: Disclosed is a blade/vane cascade segment of a blade/vane cascade of a turbomachine, which comprises at least two blade/vane elements defining an axial cascade width, and a stage with a stage surface, as well as a stage edge on the inflow side. The stage surface has an elevation extending up to the pressure side of a first blade/vane element and a depression extending up to the suction side of the other blade/vane. At least one highest point of the elevation and at least one lowest point of the depression each lie at least 30% and at most 60% of the axial cascade width downstream of the inflow edges of the blade/vane elements. In this case, the elevation and the depression each come close to the stage edge on the inflow side. Also disclosed are a blade/vane cascade, a stage for a blade/vane cascade segment, a blade/vane channel, and a turbomachine.

Abstract: The invention relates to a blade or vane, particularly of a turbine stage of a gas turbine, in particular of an aircraft gas turbine, having a blade or vane root and a blade or vane element joined to the blade or vane root, wherein the blade or vane element has a pressure side and a suction side, and wherein the blade or vane root has at least one raised region on its radial outer side facing the blade or vane element. It is proposed according to the invention that the blade or vane has a first raised region on the pressure side and a second raised region on the suction side, wherein the highest point of the first raised region is disposed essentially directly adjacent to the pressure side, and the highest point of the second raised region is disposed essentially directly adjacent to the suction side.

Abstract: A blade cascade of a turbomachine having at least one shape variation of a blade situated on the blade side in the proximity of a side wall and extending downstream, and at least one side wall contouring of the side wall or at least one second shape variation of an adjacent blade near the side wall, as well as a turbomachine, are disclosed.

Abstract: The invention relates to a blade or vane ring for a gas turbine having a plurality of blades or vanes that are arranged next to one another in the peripheral direction (UR), wherein blades or vanes have a flow segment extending essentially in the radial direction, these blades or vanes having a convex suction side, a concave pressure side, a leading edge and a trailing edge, wherein the suction side and the pressure side are joined together by the leading edge and the trailing edge, wherein the blades or vanes transition into an annular segment of the blade or vane ring radially inside and/or radially outside, wherein annular segment connects a first blade or vane (12) and a second blade or vane, which are adjacent to one another, between the suction side of the first blade or vane and the pressure side of the second blade or vane.

Abstract: The invention relates to a guide vane cluster (10) for a turbomachine, in particular for a turbojet engine, having at least two vanes (14) that are joined together by way of at least one common platform (12, 16) for the radial delimiting of a flow channel of the turbomachine, wherein each vane (14) has a vane element (17) with a suction side (18) and a pressure side (20) that are joined together by a leading edge (22) and a trailing edge (24), wherein the at least two vanes (14) have different trailing edge wall thicknesses at least in a corresponding radial lengthwise extension region.

Abstract: A blade channel of a turbomachine that is delimited in the circumferential direction of the turbomachine by a pressure side of an airfoil and by an opposite suction side of an adjacent airfoil that, in the radial direction of the turbomachine, is delimited by two opposing side walls, and whose extent in the axial direction of the turbomachine is delimited by leading edges and by trailing edges of airfoils; at least one of the side walls being provided with localized contours, of which at least two are formed as elevations and at least two as depressions; a saddle surface being formed between the contours that, in one rotation, alternately merges into an elevation and a depression; a blade cascade having such blade channels, as well as a turbomachine having such a blade cascade.

Abstract: A blade cascade for a continuous-flow machine having a non-axisymmetrical side wall contour, whereby the side wall contour has at least one pressure-side elevation and one suction-side depression, whose highest section and lowest section are located over an area 30% to 60% of an extension of the blades in the axial direction, and the axial positions of the outermost sections differ from each other at the maximum by 10% in the axial direction.

Abstract: Disclosed is a turbomachine including a stator, a rotor rotatable about an axis of rotation, and an annulus for carrying a core flow, the annulus having a side wall on the stator and a side wall on the rotor, at least one airfoil array having a plurality of airfoils being disposed in the annulus. In a departure from an ideal aerodynamic annulus contour, a radial annulus enlargement begins upstream of the airfoils and extends downstream up to an aft portion of the airfoil array that follows the ideal aerodynamic annulus contour. Also disclosed is an airfoil for such a turbomachine.

Abstract: A blade cascade of a turbomachine having at least one shape variation of a blade situated on the blade side in the proximity of a side wall and extending downstream, and at least one side wall contouring of the side wall or at least one second shape variation of an adjacent blade near the side wall, as well as a turbomachine, are disclosed.

Abstract: A blade cascade for a continuous-flow machine having a non-axisymmetrical side wall contour, whereby the side wall contour has at least one pressure-side elevation and one suction-side depression, whose highest section and lowest section are located over an area 30% to 60% of an extension of the blades in the axial direction, and the axial positions of the outermost sections differ from each other at the maximum by 10% in the axial direction.