Abstract: A locking element for a borescope opening of a gas turbine is disclosed. The borescope opening includes a radially inner opening edge facing a gas-carrying annular channel of the gas turbine where the locking element includes a main section and a locking section connected to the main section. The locking section is configured at least partially in a rotationally-symmetric manner with regard to a longitudinal axis of the locking element. The locking element is dimensioned such that the locking section protrudes partially over the radially inner opening edge of the borescope opening in the inserted state into the borescope opening.

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 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: Intermediate housing (14), in particular of turbines (11, 13) of a gas turbine engine, having a radially inner bounding wall (23) and having a radially outer bounding wall (24, 24?), having a crossflow channel (33), which is formed by the bounding walls (23, 24, 24?) and within which at least one supporting rib (15) is positioned that has a leading edge (16), a trailing edge (17), as well as side walls (18) extending between the leading edge (16) and the trailing edge (17) that direct a gas flow traversing the crossflow channel (33); the radially outer bounding wall (24) having a contour that changes in the circumferential direction at least in one section upstream of the supporting rib (15).

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: A blade channel having a not-axially-symmetric end wall contour in a turbomachine is disclosed, the end wall contour having at least one individual contour in the form of an elevation, on the pressure side, and at least three individual contours in the form of two recesses and one elevation, on the suction side, the elevation being situated between the recesses in the flow direction; a turbomachine having a plurality of blade channels of this type is also disclosed.

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 row for a turbomachine is disclosed. The blade row has an inner lateral wall and an outer lateral wall for bordering a hot gas channel in the turbomachine through which hot gas passes. At least one of the lateral walls has a rounded lateral wall front edge, which is provided with a circumferentially asymmetrical edge contouring that is elliptical or is comprised of one or more segments of a circle.

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 sealing system for a turbomachine, in particular for a gas turbine, the sealing system being formed in an annular space between a flow-limiting wall of the turbomachine and at least one rotor blade tip of a rotor blade or an outer shroud arranged on the rotor blade tip, and comprising at least one sealing point. The sealing point comprises at least one run-in coating arranged on the rotor blade tip or on the outer shroud in the direction of the flow-limiting wall of the turbomachine. The invention furthermore encompasses a gas turbine comprising the sealing system.

Abstract: A sealing system for a turbomachine, in particular, for a gas turbine, having an annular space between a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades. The sealing system comprises at least one first sealing fin disposed on an end of the rotor blade facing the wall, at least one second abradable lining disposed following the first sealing fin in the flow direction on the end of the rotor blade facing the wall, at least one first abradable lining disposed on the inside of the wall and opposite the first sealing fin, and at least one second sealing fin disposed in the flow direction following the first abradable lining on the inside of the wall and opposite the second abradable lining. The invention furthermore relates to a gas turbine comprising at least one sealing system.

Abstract: Intermediate housing (14), in particular of turbines (11, 13) of a gas turbine engine, having a radially inner bounding wall (23) and having a radially outer bounding wall (24, 24?), having a crossflow channel (33), which is formed by the bounding walls (23, 24, 24?) and within which at least one supporting rib (15) is positioned that has a leading edge (16), a trailing edge (17), as well as side walls (18) extending between the leading edge (16) and the trailing edge (17) that direct a gas flow traversing the crossflow channel (33); the radially outer bounding wall (24) having a contour that changes in the circumferential direction at least in one section upstream of the supporting rib (15).

Abstract: A blade channel having a not-axially-symmetric end wall contour in a turbomachine is disclosed, the end wall contour having at least one individual contour in the form of an elevation, on the pressure side, and at least three individual contours in the form of two recesses and one elevation, on the suction side, the elevation being situated between the recesses in the flow direction; a turbomachine having a plurality of blade channels of this type is also disclosed.