Mtu Aero Engines Ag

Mtu Aero Engines Gmbh

Inga Mahle

Markus Brettschneider

Fadi Maatouk

Martin Pernleitner

Markus Schlemmer

Norbert Huebner

Nina Wolfrum

Kai Koerber

Alexander Boeck

Rudolf Stanka

Petra Kufner

Martin Hoeger

Walter Gieg

Jochen Gier

Oliver Thiele

Norman Cleesattel

Karl Engel

Gottfried Schuetz

Ulrich Haid

Michaela Niebler

Eckart Henrich

Kai Körber

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.

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.

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.

Blade channel, blade cascade and turbomachine

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

Guide vane segment for a turbomachine

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.

Sealing system for a turbomachine

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.

Gas turbine blade

A turbomachine stage includes guide vanes and an airfoil platform forming a guide vane cascade, and rotor blades and an airfoil platform forming a rotor blade cascade. Airfoil platforms have cascade regions extending between circumferentially adjacent airfoils, and gap regions which radially and/or axially bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade. A contour of at least one of these gap regions varies in the radial and/or axial direction around the circumference. A maximum extent of this contour in the radial direction toward the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the radial direction being no more than 50% of the cascade pitch and/or a maximum extent in the axial direction away from the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the axial direction being no more than 50% of the cascade pitch.

Turbomachine and turbomachine stage

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.

Gas turbine stage

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.

Blade cascade with side wall contours and continuous-flow machine

Intermediate housing (), in particular of turbines () of a gas turbine engine, having a radially inner bounding wall () and having a radially outer bounding wall ('), having a crossflow channel (), which is formed by the bounding walls (&#x2032;) and within which at least one supporting rib () is positioned that has a leading edge (), a trailing edge (), as well as side walls () extending between the leading edge () and the trailing edge () that direct a gas flow traversing the crossflow channel (); the radially outer bounding wall () having a contour that changes in the circumferential direction at least in one section upstream of the supporting rib ().

Intermediate housing of a gas turbine having an outer bounding wall having a contour that changes in the circumferential direction upstream of a supporting rib to reduce secondary flow losses

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.

Growing community of inventors

Munich, Germany

Inga Mahle