Rolls-royce Deutschland Ltd & Co. Kg

Fraunhofer-Gesellschaft Zur Foerderung Der Angewandten Forschung E.V.

Oerlikon Surface Solutions Ag, Pfaffikon

Thomas Wunderlich

Volker Derflinger

Dan Roth-Fagaraseanu

Peter Broadhead

Harald Schiebold

Markus Weinert

Frank Brueckner

Susanne Schruefer

Yannick Cadoret

Mirko Riede

Hans-Frieder Vogt

Olaf Mistareck

Lin Shang

Dan Roth-fagaraseanu

Susanne SCHRÜFER

Sebastian Guimond

Toby Charles Middlemiss

Frank BRÜCKNER

A blade component of a compressor or turbine stage of a gas turbine, in particular of a gas turbine engine is provided. The blade component has at least two structural elements which can be connected together by means of a connection process, in particular sintering. The at least two structural elements can be coupled and/or connected to at least one means for internal cooling of the blade component, and/or the means for internal cooling of the blade component is arranged in at least one of the least two structural elements. The at least two structural elements, when assembled, can be coupled to a cooling insert inside the blade component, wherein during operation, cooling air flowing into the cooling insert can be guided in targeted fashion via impingement cooling openings onto the inside of the blade component, in particular as impingement flow cooling.

Blade component, method for manufacture of same, and gas turbine

A method for manufacturing a thermally deformable component for high thermal loads, includes: providing a first area of the component with a first metallic material by a generative laser process, or making the first area of the first metallic material; providing a second area of the component with a second metallic material by a generative laser process, or making the second area of the second metallic material; where at least one of the metallic materials is deposited by the generative laser process, and a ratio of a linear expansion coefficient &#x3b1;of the first metallic material and of a linear expansion coefficient &#x3b1;of the second metallic material is as: where x=0.5 to 1; T=mean operating temperature on a hot side; T=reference temperature; T=mean operating temperature on a cold side.

Method for the manufacture of a component for high thermal loads, a component producible by this method and an aircraft engine provided with the component

A component, especially contrived and designed for being used in a turbomachine, includes a high-temperature coating being arranged above a base of the component. The base has at least one structural element for connecting it to the high-temperature coating, with the cross-section of the at least one structural element having at least three different widths, i.e. a base width at the lower end of the at least one structural element, a center width above it, and a tip width above that, where on average the center width is greater than or equal to the base width, but less than four times the base width, in particular less than or equal to three times the base width.

Component and turbomachine having a component

A suspension arrangement for the casing shroud segments of the high-pressure turbine of a jet engine with inclination or graduation of the rotor blade tips () includes spaced apart attaching hooks () formed onto the casing shroud segments () which engage an axial retaining groove () on the turbine casing () on the upstream side and which are axially retained in the downstream direction by a retaining ring () arranged in a retaining ring groove () formed onto the turbine casing. In the retaining ring groove (), assembly pockets () are provided on the downstream side, with the assembly pockets () agreeing with the distance and dimensioning of the attaching hooks (), so that&#x2014;with the rotor installed&#x2014;axial travel (s) of the casing shroud segments () is reduced and handling clearance is increased, when the casing shroud segments () are installed.

Suspension arrangement for the casing shroud segments

With a door seal for a gas turbine engine, a gap () existing between an axially symmetrical component () and an assembly formed by several segments () is sealed by seal doors (), which are held via retaining holes () at retaining bolts () on one of the components, the seal doors () being pressed against sealing edges () of the two components, with the sealing edges () being disposed at a certain distance from each other. The retaining bolts and holes are located outside the pressure surface (A) of the seal doors, formed between the two sealing edges, thus preventing cooling air losses via a gap remaining between the retaining bolts and bores.

Door seal for a turbomachine

A passive internal control system for rotor blade tip clearance of a high-pressure turbine is based on inner rings that expand in a radial direction under the influence of heat and radially adjust the clearance delimited by liner segments () on the casing side. It comprises a radially expandable U-shaped downstream inner ring () mounted to inner platforms () of guide vanes () on a side where the rotor does not have a static bearing, said ring providing expansion compensation in the axial and peripheral directions and forming a torsion box () with struts () that absorbs rolling and tilting moments.

Arrangement for internal passive turbine blade tip clearance control in a high pressure turbine

On a two or multi-stage turbine, expansion rings are provided in all stages at the sides of the rotors for passive, continuous running gap control whose thermal expansion and contraction behavior corresponds to that of the rotors and which are connected to radially moveable upstream and downstream stator vanes (). The downstream stator vanes are assembled to the upstream stator vanes via a bridge () which is axially and circumferentially fixed and located flexibly in the radial direction on the outer casing () of the turbine. The stator vanes () arranged between the rotor disks are integrally connected to the bridge or, as separate components, held axially and circumferentially on the bridge to take up rolling and tilting moments.

Arrangement for the automatic running gap control on a two or multi-stage turbine

On an aircraft engine with a shroudless rotor wheel arranged in a flow duct, the flow duct includes a shroud segment () formed by a ceramic rubbing coating () which, while having good thermal conductivity, is highly-temperature resistant, attaches firmly to the metallic substrate () and is abradable by the tips of the rotor wheel to form a sealing gap as narrow as possible, this rubbing coating, owing to the lack of self-insulation, being coolable from a free side of the metallic substrate and, therefore, permitting for working gas temperatures occurring in high-pressure turbines and coating thicknesses sufficient for abrasion.

Shroud segment for a turbomachine

A gas turbine with several shroud segments () which enclose rotor blades () of a turbine wheel () as a seal, with the shroud segments () having at least a front and a rear attachment at the radially outward area of stator vane segments (), and with each of the stator vane segments () being located at their radially inner area on a control ring (), wherein the stator vane segments () are located on the control ring () in a radially adjustable manner.

Growing community of inventors

Rangsdorf, Germany

Thomas Wunderlich