Mtu Aero Engines Ag

Mtu Aero Engines Gmbh

Daimler Chrysler Ag

Rudolf Stanka

Erwin Bayer

Petra Kufner

Walter Gieg

Martin Pernleitner

Manfred Dopfer

Manfred Feldmann

Markus Schlemmer

Norman Cleesattel

Hermann Klingels

Werner Humhauser

Norbert Huebner

Bernd Kislinger

Karl Maar

Manuel Hein

Stefan Morgenstern

Marcus Woehler

Oliver Thiele

Thomas Hess

Inga Mahle

Markus Brettschneider

Karl-Heinz Dusel

Peter Geiger

Christoph Lauer

Erich Steinhardt

Hans-Peter Hackenberg

Stefan Hagen Weber

Manfred Schill

Wilfried Schuette

Markus Uecker

Klaus Wittig

Jochen Gier

Alois Eichinger

Klaus Pirker

Joachim Lorenz

Hans Zimmermann

Hans Stricker

Jens Wittmer

Carsten Schoenhoff

Martin Fischer

Lukasz Sosnowka

Gottfried Schütz

Martin Becker

Gottfried Schuetz

René Schneider

Norbert Hübner

Christoph Bichlmaier

Claus Riegler

Eckart Heinrich

Klaus Peter Rued

Marcin Regulski

Hugh Alun John Jackson

Wilfrid Schuette

Luitpold Frey

Carsten Schönhoff

Wilfried Schütte

Marcus Wöhler

Sven-Jürgen Hiller

Marcus Dr Wöhler

Inga Dr Mahle

A gas turbine includes a housing and an outer sealing ring. In an assembled state of the housing and the outer sealing ring, a first planar structure of the outer sealing ring is axially moveable with respect to a second planar structure of the housing against a direction of flow through the gas turbine.

Gas turbine with axially moveable outer sealing ring with respect to housing against a direction of flow in an assembled state

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 rotor, for a turbomachine, in particular a gas turbine, having a first flange and a second flange with recesses that are distributed in a direction of distribution, in particular in the peripheral direction, wherein the second flange is fastened to the first flange, in particular detachably, by at least one fastener, which engage in the first of these recesses of the first and second flanges, wherein second ones of these recesses of the first flange, which are free of a fastener, are covered by the second flange.

Turbomachine component arrangement

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

The invention relates to a guide vane segment for an aircraft engine, having at least one guide vane element, having at least one flange formed in a radial extension direction (R) of the guide vane element and at least one positioning member projecting from the flange in the radial extension direction (R), and at least one seal carrier, which is arranged at the flange and is aligned relative to the guide vane element by way of the positioning member, wherein the flange has a supporting surface on which the seal carrier is supported, wherein a partial surface of the positioning member is adjacent to the flange at a transition from the flange to the positioning member, and respective surface normal lines of the supporting surface and of the partial surface enclose between them an angle (&#x3b1;) that is different from a zero angle.

Guide vane segment for a turbomachine

Described is a turbine center frame for a gas turbine, in particular an aircraft gas turbine, the turbine center frame having a radially inner wall and a radially outer wall bounding an annular space through which hot gas flows and each having a contour facing the annular space, the contours describing an inner annular space curve along the inner wall and an outer annular space curve along the outer wall. At least one vane element extends in the radial direction through the annular space and has an axial leading edge and an axial trailing edge, the vane element having an outer axial width. The outer annular space curve or/and the inner annular space curve have at least one curve section having an inflection point of the respective annular space curve or/and a point of maximum slope of the respective annular space curve, the curve section being located in the region of the leading edge or trailing edge as considered with respect to the outer axial width or/and the inner axial width and having a length projected parallel to the axial direction which length is up to 20% of the respective axial width, and the curve section intersecting the piercing point where the leading edge or the trailing edge pierces the outer wall or the inner wall.

Turbine center frame having a specifically designed annular space contour

The aircraft-engine gas turbine includes an outer sealing ring for sealing an array of rotor blades that can be attached to a housing by a clamping mechanism () in a friction fit, and a plurality of ring segments (), whereina free axial path length (a) of a sealing ring segment counter to the direction of through-flow is at least as large as an axial engagement (a) of a rotation locking member () of the outer sealing ring (a&#x2265;a), which is free of form fit counter to the direction of through-flow, and/or an axial overhang (a) of a radial mounting rail () of the outer sealing ring (a&#x2265;a), and/or an axial offset (a, a) of a sealing fin (); and/ora quotient of a specific clearance sum of the outer sealing ring attached to the housing in a friction fitand pi is at least as large as a difference between a maximum outer diameter of the outer sealing ring and a minimum inner diameter of the flow channel inlet of the housing.

Gas turbine

The invention relates to a seal arrangement for a turbomachine, especially for an aircraft engine, for sealing a radial gap between a rotor and a stator, comprising at least one seal carrier for the supporting and/or fastening of at least one seal element, wherein the seal carrier comprises a radial crosspiece extending in a radial direction of extension and an axial crosspiece, formed as a single piece with the latter and extending in an axial direction of extension, wherein the seal element is arranged at a radially inner-lying bearing surface of the axial crosspiece, and a front ring or a front ring segment, viewed in the flow direction, and/or a rear ring or a rear ring segment, each with a radially running crosspiece.

Seal arrangement for a turbomachine, method for manufacturing a seal arrangement and turbomachine

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

The invention relates to a module for a gas turbine comprising a rotationally symmetrical outer casing of a first material having a first linear thermal expansion coefficient, a rotationally symmetrical component for guiding hot gas of a second material having a linear thermal expansion coefficient which is lower than the first coefficient, at least one annular structure arranged radially within the component, and at least three struts having a radially inner end secured on the outer casing and a radially outer end secured on the annular structure. The component also comprises at least one first fixing which is free in radial direction and fixed in axial direction and arranged between the component and the outer casing and/or the annular structure, as well as at least three second fixings having an elastic effect in radial, axial and/or circumferential direction and being arranged between the component and the strut and/or the annular structure.

Module for a gas turbine

The invention relates to a run-in coating () for an outer air seal of a turbomachine, especially of an aero engine, comprising a substrate () of segmented form which carries a honeycomb structure () with an run-in surface () to face rotor blades of a rotor of the turbomachine, wherein a radial thickness of the honeycomb structure () varies in the circumferential direction. The invention furthermore relates to an outer air seal for a turbomachine and also to a turbomachine with an outer air seal.

Run-in coating for an outer air seal of a turbomachine

The invention relates to a guide vane segment for a gas turbine comprising at least one radially outer shroud and one radially inner shroud, which extend along a respective arc and together form an annular segment, wherein, in the radial direction between the outer shroud and the inner shroud, a plurality of guide vanes are arranged adjacent to one another in the peripheral direction and are joined to the inner shroud and to the outer shroud in a material-bonded manner and, in particular, are joined in a one-piece manner, wherein, in an axial lengthwise direction, the outer shroud comprises an axially front end wall element and an axially rear end wall element in such a way that the outer shroud and the two end walls form a tub-like profile in lengthwise section.

Guide vane segment with radical securing elements

A guide vane arrangement for a turbomachine, particularly a gas turbine, is disclosed. The guide vane arrangement includes an external shroud band, on which at least two guide vanes protrude radially inward. A pair of axial ribs spaced apart from each other in the peripheral direction and disposed between two adjoining guide vanes in the peripheral direction protrude radially outward from a circumferential surface of the external shroud band and/or axially from a radial flange of the external shroud band.

Guide vane arrangement

A method for disassembly of a rotor, in particular the front rotor of a gas turbine with a housing and a channel which diverges in a direction of flow and in which the rotor is arranged, is disclosed. An embodiment of the method includes axially displacing an outer sealing ring that is radially opposite the rotor, and whose minimum inside diameter is smaller than the maximum outside diameter of the rotor, against the direction of flow. Then, the rotor is axially displaced against the direction of flow, in particular out of the housing. A method for assembly of such a rotor, as well as a tool for the same, is also disclosed.

Method for disassembly and assembly of a rotor of a gas turbine

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

Growing community of inventors

Rattenkirchen, Germany

Rudolf Stanka