Mtu Aero Engines Ag

Fadi Maatouk

Markus Brettschneider

Inga Mahle

Martin Hoeger

Guenter Ramm

Yavuz Guendogdu

Irene Raab

Yavuz Gündogdu

Günter Ramm

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.

Flow arrangement for placing in a hot gas duct of a turbomachine

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.

Blade or guide vane with raised areas

Described are an airfoil array segment () having at least two airfoils () and a platform () that features an axis asymmetrical platform surface (). This platform surface features an elevation () that extends from the pressure side () of the first to the suction side () of the second airfoil (). A highest point () of the elevation is more proximate to the suction sideof the second airfoil () than to the pressure side () of the first airfoil (). Also described are an airfoil, a platform, an airfoil passage and a turbomachine.

Endwall contouring for a turbomachine

An airfoil array for a turbomachine, in particular a turbine or compressor stage of a gas turbine. The airfoil array includes at least two airfoils and at least one contoured circumferential surface which connects a pressure side of one airfoil to a suction side of the other airfoil and includes an upstream first section and a downstream second section which adjoins the first section along an elevation contour line; the first section being depressed relative to a rotationally symmetric reference surface containing this elevation contour line away from the airfoils, and the second section not being depressed relative to this reference surface away from the airfoils; this elevation contour line lying in an axial area which terminates at most 30% of an axial chord length of one of the airfoils downstream of its leading edge; and an axial distance of this elevation contour line increasing toward the pressure side and toward the suction side, starting at a point between the pressure and suction sides that is closest to the leading edge; and the first section extending over at least 90% of the space between the pressure side and the suction side.

Turbomachine airfoil array

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.

Contouring a blade/vane cascade stage

Disclosed is an airfoil array segment () of an airfoil array for a turbomachine, the airfoil array segment including a platform () having a platform surface, as well as at least two airfoils (). The platform surface has a depression () which extends up to the first airfoil and contacts the pressure side () of the first airfoil downstream of 80% of the axial chord (g) downstream of the leading edges () and which contacts the pressure side () of the first airfoil downstream up to no more than 80% of the axial chord (g) downstream of the leading edges (). At least one lowest point () of the depression is located at least 90% of the axial chord (g) downstream of the leading edges (). Also disclosed are an airfoil array, an airfoil passage, a platform, a turbomachine, and an aircraft engine.

Contouring of an airfoil array platform

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.

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Munich, Germany

Fadi Maatouk