Mtu Aero Engines Ag

Mtu Aero Engines Gmbh

Giovanni Brignole

Georg Zotz

Vitalis Mairhanser

Harsimar Sahota

Carsten Zscherp

Tobias Mayenberger

A gas turbine compressor has a flow duct wall disposed radially opposite to an airfoil tip and has a circumferential groove having an upstream groove edge and a downstream groove edge, the circumferential groove having a web having a radial cutback. In at least one meridional section through an airfoil-tip-side end face of the web, an axial distance between an upstream beginning of the cutback and the upstream leading edge of the airfoil tip is at least 1% and/or no more than 40% of a chord length and/or an axial distance between the upstream leading edge of the airfoil tip and the downstream groove edge is at least 5% and/or no more than 40% of the chord and/or an axial distance between the upstream leading edge of the airfoil tip and a kink in an airfoil-tip-side upper edge of the web in the cutback is no more than 10% of the chord length and/or a radial distance between the airfoil tip and an airfoil-tip-side upper edge of the web in the cutback is at least 50% and/or no more than 1500% of a radial distance between the airfoil tip and the downstream groove edge radially opposite thereto.

Gas turbine compressor

The present invention relates to a guide vane cascade for a turbomachine, which has guide vanes that are mounted adjustably about an axis of rotation, so as to change an inflow angle each time, wherein a first guide vane and a second guide vane arranged on the pressure side thereof, referred to a longitudinal axis of the turbomachine, are arranged with an axial offset, namely, the axis of rotation of the second guide vane is offset axially toward the back, wherein the first guide vane and the second guide vane are provided such that, in an adjusted state, they form together a tandem configuration in a radially outer region but together they delimit a divergent channel in a radially inner region.

Guide vane cascade for a turbomachine

A gas turbine compressor having at least one airfoil tip () and a flow duct wall () which is disposed radially opposite thereto and has a circumferential groove (-) therein in which is disposed at least one web () having a radial cutback () is provided. An upstream beginning () of the cutback is located axially downstream of an upstream groove edge () between this groove edge and an upstream leading edge () of the airfoil tip, and a downstream end () of the cutback is located in an airfoil-tip-proximal half () of a radial height () of the circumferential groove.

A turbomachine having at least one circulation structure is disclosed. The circulation structure has an annular space with baffle elements surrounding a main flow path and is open to the main flow path. A housing of the turbomachine to receive the circulation structure is divided in an axial plane into a front housing region and a rear housing region and the circulation structure is divided into a front structure region and a rear structure region in the axial plane of separation. A circulation structure divided into two parts in the axial direction and a method for the same is also disclosed.

Turbomachine, circulation structure and method

A circulation structure for a turbo compressor, in particular for a compressor of a gas turbine, is disclosed. The circulation structure includes at least one annular chamber that can be traversed in a circumferential direction, is concentric with a shaft of the turbo compressor in the region of the free blade ends of a blade ring, and radially borders a main flow channel. Several chambers that can be traversed in an axial direction are situated upstream of the or each annular chamber, when viewed from the main flow direction of the main flow channel.

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Munich, Germany

Giovanni Brignole