Mtu Aero Engines Ag

Mtu Aero Engines Gmbh

Jochen Gier

Karl Engel

Inga Mahle

Yavuz Guendogdu

Manfred Feldmann

Hermann Klingels

Werner Humhauser

Rudolf Stanka

Martin Hoeger

Norbert Huebner

Hans-Peter Hackenberg

Erich Steinhardt

Karl Maar

Stefan Hagen Weber

Andreas Fiala

Kai Koerber

Claus Riegler

Klaus Peter Rued

Eckart Heinrich

Kai Körber

Yavuz Gündogdu

Engel Karl

The present invention relates to a rotor blade shroud for a turbomachine, comprising a sealing tip and a support structure that abuts the sealing tip. The support structure has at least one intermediate region in which a structural segment is arranged, wherein the radially outwardly arranged surface of the support structure and of the structural segment forms an essentially planar surface. The present invention further relates to a rotor blade for a turbomachine, comprising a rotor blade shroud as well as two methods of manufacturing a rotor blade shroud and a method of manufacturing a rotor blade.

Rotor blade shroud for a turbomachine, rotor blade, method of making a rotor blade shroud and a rotor blade

A blade cascade of a turbomachine whose at least one side wall is configured to be circumferentially undulated and which has at least two elevations and at least one depression or at least two depressions having at least one elevation, as well as a turbomachine having a blade cascade of this kind.

Blade cascade and turbomachine

A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (D) of the fan to a maximum blade diameter (D) of the second turbine is at least 3.5, in particular at least 4.

Turbofan aircraft engine

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 blade cascade of a continuous-flow machine is disclosed. The blade cascade has at least one side wall that is constructed in a wave-like manner in the peripheral direction and has at least one elevation and at least one depression. At least a profile-like rib is integrated in or combined with the side wall, which has a blade profile with a pressure side and with an opposite suction side. A continuous-flow machine with such a blade cascade is also disclosed.

Blade cascade 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 construction for a turbo device, preferably a blade array of a low-pressure turbine, is disclosed. On the inlet-side profile surface of the blade a vortex generator is arranged downstream from the speed maximum. The vortex generator is formed by a surface undulation with at least one wave, the wave tail of which runs in the form of a wave trough and/or a wave peak in the blade vertical direction.

Blade for a turbo device with a vortex-generator

Growing community of inventors

Karlsfeld, Germany

Jochen Gier