Rolls-royce Deutschland Ltd & Co. Kg

Erik Johann

Frank Heinichen

Rainer Mielke

Jens Ortmanns

Cetin Eryilki

A rotor blade airfoil of a turbomachine, which rotor blade airfoil has: a leading edge, a trailing edge, and a profile chord length which is dependent on the height of the blade airfoil. In a side view of the blade airfoil, a maximum projected chord length his defined as the axial spacing between the axially foremost point of the leading edge and the axially rearmost point of the trailing edge of the blade airfoil in the side view under consideration. Here, the axial position of the leading edge varies in a manner dependent on the height of the blade airfoil above a front axial region. Provision is made whereby, furthermore, with respect to the side view under consideration, the axial position of the trailing edge of the blade airfoil varies in a manner dependent on the height of the blade airfoil above a rear axial region, wherein the variation of the axial position of the trailing edge in the rear axial region amounts to at least 10% of the maximum projected chord length, the trailing edge of the blade airfoil assumes the axially rearmost point at a height of the blade airfoil that lies in the range between 20% and 50% of the total height of the blade airfoil at the trailing edge, and the leading edge of the blade airfoil assumes the axially foremost point at a height of the blade airfoil that lies in the range between 15% and 35% of the total height of the blade airfoil at the leading edge.

Rotor blade of a turbomachine

A compressor rotor of a turbomachine includes a rotor disk; a rotor hub forming or connected to a radially outer edge of the rotor disk; and a plurality of rotor blades on the rotor hub extending radially outwards. The rotor hub includes an axially frontal leading edge, an axially rear trailing edge, a top side, a frontal bottom side extending on a bottom side of the rotor hub from the leading edge in a direction of the rotor disk and transitioning into same, and a rear bottom side extending on the bottom side from the trailing edge in the direction of the rotor disk and transitioning into same. The frontal bottom side and/or the rear bottom side of the hub is contoured in a circumferential direction of the rotor hub to form respectively one indentation in an area below a rotor blade.

Compressor rotor of a fluid flow machine

A method for the manufacture of a double-row blade wheel for a fluid-flow machine that includes a first grid row with a plurality of first blades and a second grid row with a plurality of second blades is provided. The method includes the following steps: manufacturing a first grid row or segments of the first grid row; manufacturing a second grid row or segments of the second grid row, independently of the first row; and after the respective manufacture, connecting the first grid row to the second grid row or connecting segments of the first grid row and/or of the second grid row to one another or to one of the grid rows.

Method for producing a tandem blade wheel for a jet engine and tandem blade wheel

The present invention relates to a stator of an axial compressor stage of a turbomachine featuring a radially outer blade ring forming an outer ring surface, a radially inner blade ring forming an inner ring surface, and several stator blades connected to the blade rings. It is provided that the outer ring surface and/or the inner ring surface has at least in a partial area a changing radius relative to a central axis of the stator both in the axial direction and in the circumferential direction.

Stator of an axial compressor stage of a turbomachine

A fluid-flow machine includes at least one rotor having a rotary element with a plurality of rotor blades arranged on the rotary element, and a circumferential casing having a central axis and surrounding the rotor. The circumferential casing or a part connected thereto has an annular space surface on the inside, which delimits a flow duct of the fluid-flow machine radially outwards. The annular space surface has a structuring at least in one area adjoining a rotor on the circumferential side. At least one structuring of the annular space surface has, relative to the central axis of the circumferential casing, a circumferentially asymmetrical design.

Fluid-flow machine

The present invention relates to a rotor of an axial compressor stage of a turbomachine featuring a rotor assembly with a rotary axis, forming on its circumference a blade ring with a radially outer ring surface, and several rotor blades arranged on the blade ring. It is provided that the ring surface between two adjacent rotor blades has at least in a partial area a changing radius relative to the rotary axis of the rotor assembly both in the axial direction and in the circumferential direction.

Rotor of an axial compressor stage of a turbomachine

A fluid flow machine includes a flow path formed by a casingand a rotating shaft, in which rows of bladesare provided, with at least one annular groove-type recessbeing arranged in a bladetip area in an annulus duct wall of the shaft. A cross-section and the position of the recessare defined as follows: the axial arrangement is selected such that a partial length A of the recessextends by max. 30 percent of the chord length of a bladefrom the bladeleading edge against the direction of flow, and up to max. 50 percent of the chord length of the bladefrom the bladeleading edge in the direction of flow.

Casing treatment for axial compressors in a hub area

In the repair of a blisk or blisk drum, the damaged airfoils are completely cut off a low-stress nodal line of the respective blade, except for a blade stump (), and are each replaced by a new blade repair element () of a same shape and size which is accurately clamped onto the blade root by a ledge () formed onto its mating end, with this ledge () also serving for fixation of the processing fixtures and stabilization of the thin airfoil, and is then welded to the blade root and subsequently subjected to a cutting or non-cutting operation for removal of the ledge (). This highly efficient repair method can similarly also be applied to the new manufacture of blisks.

Method and blade repair element for blisk repair or new blisk manufacture

The compressor blades of an aircraft engine are, in at least one natural-vibration critical area, designed such that at the blade leading edge (), the leading edge shock wave () attaches to the leading edge (), as a result of which a laminar boundary layer flow () on the suction side () quickly transitions into a turbulent boundary layer flow () which is kept constant and prevented from re-lamination by the further, continuous curvature of the suction side. Therefore, the transition, whose periodic movement is also suppressed, cannot communicate with a suction-side compression shock (), preventing the compression shock from augmenting the natural vibrations of the blade occurring under certain flight conditions. The blade leading edge can, for example, be designed as an ellipse with a semi-axis ratio equal to or smaller than 1:4.

Compressor blade with reduced aerodynamic blade excitation

On a compressor with compressor blades, a flow transition fixation mechanism () is provided on the suction side (), approximately parallel to the leading edge () and upstream of the compression shocks acting upon the blade, which prevents the transition point from the laminar to the turbulent boundary layer flow from oscillating, thus suppressing oscillation of the compression shocks and their coupling effect with the natural frequencies of the compressor blades.

Growing community of inventors

Berlin, Germany

Erik Johann