Mtu Aero Engines Ag

Mtu Aero Engines Gmbh

Manfred Dopfer

Martin Pernleitner

Rudolf Stanka

Klaus Wittig

Dieter Freno

Wilfried Schuette

Marcus Woehler

Norman Cleesattel

Carsten Schoenhoff

Tino Hummel

Markus Schlemmer

Norbert Huebner

Bernd Kislinger

Christoph Lauer

Manfred Schill

Manuel Hein

Oliver Thiele

Daniel Theurich

Gottfried Schütz

Gottfried Schuetz

Christoph Bichlmaier

Norbert Hübner

Salome GASSMAnn

Johannes Linhard

Wilfrid Schuette

Carsten Schönhoff

Wilfried Schütte

Marcus Wöhler

The present invention relates to a rotating blade for a turbomachine, having a blade element and a shroud, wherein the shroud is profiled at a peripheral side, i.e., when observed in a tangential section, has a contact flank that is oriented axially, at least proportionally, and has a free flank that is also oriented axially, at least proportionally, and is in fact opposite the contact flank, wherein, when observed in the tangential section, an intermediate segment between the contact flank and the free flank has a first curvature in a first transition segment to the contact flank and has a second curvature in a second transition segment to the free flank, wherein the first curvature is greater than the second curvature.

Rotating blade for a turbomachine

A rotor for a turbomachine has at least one blade and has at least one rotor main part, which has at least one recess, in which a blade root of the least one blade is interlockingly received, wherein the blade root comprises at least one depression, in which at least one protrusion of the at least one rotor main part, which protrusion delimits the at least one recess in regions is received, wherein the at least one depression is delimited by a first delimiting face on the blade root side and the at least one protrusion is delimited by a second delimiting face on the rotor main part side. At least the first delimiting face has at least one elevation which narrows a gap at least in regions, which extends between the first delimiting face and the second delimiting face.

Rotor for a turbomachine and turbomachine

A rotor blade system includes a blade root, a blade neck adjoining the blade root, an airfoil adjoining the blade neck, a radially outer partition wall, an axially forward partition wall and an axially rearward partition wall, connected to the radially outer partition wall such that the partition walls surround the blade neck on three sides, the partition walls projecting beyond the blade neck in the circumferential direction, and further including a blade root protection plate designed to be arranged on the blade root. It is provided that the blade root protection plate have at least one sealing portion that extends in the axial direction from the forward partition wall to the rearward partition wall and whose radially outer side is disposed opposite the radially outer partition wall when the blade root protection plate is arranged on the blade root.

System with a rotor blade for a gas turbine with a blade root protective plate having a sealing section

An intermediate element is for a blade/rotor disk connection in a rotor of a fluid flow machine. The intermediate element is adapted to a shape of a blade root of a blade and to a blade root slot in a rotor disk for receiving the blade root such that, when arranged between the blade root and rotor disk, the intermediate element prevents contact between the blade root and rotor disk. The intermediate element has, on an outer surface that faces the rotor disk, at least one protrusion to reduce an air flow parallel to an axis of rotation of the rotor between the rotor disk and the intermediate element; and on an inner surface that faces the blade root, a recess that corresponds to the at least one protrusion.

Intermediate element for a blade/rotor disc connection in a rotor of a turbomachine, associated rotor for a turbomachine, and turbomachine

A rotor blade for a gas turbine, in particular an aircraft gas turbine, having a blade root element and a stream deflection portion adjoining the blade root element () in the longitudinal direction of the blade (RR); respective centroids () of blade cross-sectional areas of the stream deflection portion residing on a common stacking axis (). It is provided that, starting from a first centroid () of a first blade cross-sectional area adjoining the blade root element (), the stacking axis () extend within a cone () whose apex resides within the first centroid (), and whose cone height (KH) extends orthogonally to the plane of the blade cross-sectional area; the angle (&#x3b2;) of the cone () being greater than 0&#xb0; and smaller than or equal to 4&#xb0;; preferably greater than or equal to 0.5&#xb0; and smaller than or equal to 2&#xb0;.

Stacking of rotor blades on centroids

A blade () for a turbomachine is provided. The blade () includes an airfoil () having at least one thickening () on a pressure side () of the airfoil (). The thickening () has a region (B) of constant thickness (D). A turbomachine, a compressor and a turbine having stator vanes and/or rotor blades including at least one such blade () is also provided.

Blade for a turbomachine

The present invention relates to a blade or vane for a turbomachine, particularly for an aircraft engine, with a blade element for interacting with the flow medium, wherein the blade has different diffusion protective coatings in various regions on its surface for protection against corrosion and/or oxidation, wherein the diffusion protective coatings are produced by chromizing and/or aluminizing, wherein the blade element is divided into two regions along the longitudinal axis of the blade element, wherein the first region extends over 80 to 95% of the length of the blade element, and the second region extends over the remainder of the length of the blade element, and wherein in both regions, an AlCr diffusion protective coating is applied, and wherein in one of the regions, the AlCr diffusion protective coating has a higher Cr content.

Blade or vane for turbomachine with different diffusion protective coatings and method for manufacture thereof

A method for inspecting a rotor blade unit of a gas turbine, the rotor blade unit including a radially outwardly disposed outer shroud connected by a material-to-material bond to a radially inwardly adjoining rotor blade. The method includes: providing a rotor blade unit; processing the outer shroud at its axially forward or axially aft side, the side selected to be processed being the one that has a greater distance from the adjacent axial leading edge or axial trailing edge of the rotor blade, a contact surface being formed by processing the selected side; providing the rotor blade unit having the contact surface in a measuring device, in particular a measuring device for determining the torsion or twist of the rotor blade unit about an axis.

Method for inspecting a rotor blade unit

The invention relates to a wear-resistant shield for a rotating blade root of a rotating blade of a gas turbine, especially an aircraft gas turbine, having a base and two side walls connected to the base, wherein the side walls lie opposite each other and are shaped so as to be substantially complementary to an outer contour of a particular rotating blade root, and wherein the wear-resistant shield for this purpose is set up in such a way that, in an installed state, it is to be taken up between the respective rotating blade root and a rotating blade root mount of a rotor, especially between the respective rotating blade root and an axial securing element arranged in the rotating blade root mount.

Wear-resistant shield for a rotating blade of a gas turbine

A guide vane segmentfor a turbomachine includes a radially inner shroud platehaving a shroud plate surfacethat is adapted to be configured in the turbomachine to face a rotor bladeadjacent to the guide vane segment, and thereby essentially extend along an outer conical surface Kwhose cone axis coincides with the axis of rotation A of a rotor shaft. In a radially inner region, a rotor bladefor a turbomachine has a base platehaving a base plate surfacethat is adapted to be configured in the turbomachine to face a shroud of a guide vane rowadjacent to the rotor blade and thereby essentially extend along an outer conical surface Kwhose cone axis coincides with the axis of rotation A of a rotor shaft

Run-up surface for the guide-vane shroud plate and the rotor-blade base plate

A damper () for damping a blade movement of a turbomachine (), and to a method for producing the damper (). The damper () has at least one side surface (') which can be brought into frictional contact with a friction surface of the turbomachine () in order to damp a blade movement. The side surfaces (&#x2032;) are asymmetrically convex in shape.

Damping means for damping a blade movement of a turbomachine

A securing device () for axially securing a blade root () of a blade in a groove () of a turbine engine. The outer contour () of the securing device () that faces a groove wall, in particular a groove base (), is curved at least in some regions, the outer contour () having three different radii (R, R, R) in some regions.

Securing device for axially securing a blade root of a turbomachine blade

A blade for a turbomachine, in particular a jet engine, including a shroud, having two opposite lateral edges, for delimiting a main flow channel and including a blade which extends away from the shroud, a rounded transition area being provided which encompasses the blade on its root side and is guided beyond the one lateral edge, a section of the transition area protruding beyond the one lateral edge being severed and situated in the area of the other lateral edge as an elevation offset in the transverse direction, a blade arrangement having at least two of such blades as well as a turbomachine having a plurality of such blades.

Growing community of inventors

Unterschleissheim, Germany

Manfred Dopfer