Siemens Aktiengesellschaft

Siemens Aktiengesell Schaft

Deutsches Zentrum Fur Luft-und Raumfahrt E.v.

Christian Cornelius

Peter Tiemann

Eckart Schumann

Iris Oltmanns, Legal Representative

Ralf Müsgen

Christian Lerner

Ulrich Waltke

Malte Blomeyer

Thorsten Mattheis

Bernhard W Küsters

Uwe Sieber

Silvio-Ulrich Martin

Torsten Matthias

Rostislav Teteruk

Georg Kröger

Andreas Peters

Bernd Van Den Toorn

Christoph Starke

Thomas Hofbauer

Stefan Braun

Bernhard Küsters

Olaf Hein

Eberhard Nicke

Marcel Aulich

Stephan Mais

Reinhard Mönig

Romain Bayere

Achim Schirrmacher

Bernhard Kuesters

Andreas Heilos

Lutz Stephan

Annika Emde

Adam Zimmermann

A compressor rotor blade for an axial-type compressor has a blade profile having a transonic section and a profile section which extends in the transonic section and has concave and convex suction-side regions on the suction side, the convex suction-side region arranged downstream of the concave suction-side region, and has convex and concave pressure-side regions on the pressure side, the concave pressure-side region arranged downstream of the convex pressure-side region. Curvature progressions on the pressure side and on the suction side are both applied in a continuous manner over a profile chord of the profile section, the positions of the minimum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord, and the positions of the maximum values of the curvature progressions deviate from each other by not more than 10% of the length of the profile chord.

Compressor rotor blade, compressor, and method for profiling the compressor rotor blade

A method for profiling a replacement blade for an axial turbomachine is provided. The method includes: measuring hub contour geometry and housing contour geometry of the flow channel of the old blade and the axial position of the centre of gravity of the blade aerofoil of the old blade; laying out geometry of the blade aerofoil of the replacement blade, the blade aerofoil having a rearward sweep on its leading edge; defining a region of the blade aerofoil near the mounting and in which positive influence of the rearward sweep on the degree of stage efficiency is rated as low; and axially displacing the replacement blade aerofoil section arranged outside this region in the upstream direction, until the axial position of the centre of gravity of the blade aerofoil of the replacement blade coincides with the old blade. In the region from the mounting-side section to the displaced blade aerofoil section of the replacement blade, the leading edge is inclined downstream towards the mounting such that the blade aerofoil has a transition sweep in this region.

Method for profiling a replacement blade as a replacement part for an old blade for an axial-flow turbomachine

The invention relates to a compressor blade of a compressor which, along a main axis, comprises a blade base, a platform area and an adjacent blade profile having a profile tip. The blade profile is configured by a convex wall at the suction end and a concave wall at the pressure end opposite the wall at the suction end. These surfaces extend, with respect to a flow medium, from a leading edge to a trailing edge, a profile center line extending in the center between the two. A front face is arranged on the profile tip at an angle to the main axis, a sealing lip at least partially extending from the leading edge to the trailing edge and the blade profile including the sealing lip having a blade profile height extending in the direction of the main axis. In order to allow for an inexpensive compressor blade having improved aerodynamic properties and a modified sealing lip while having the same sealing properties, the height of the sealing lip is less than 2 percent of the height of the blade profile.

Compressor blade and production and use of a compressor blade

A compressor blade for an axially permeated compressor, preferably of a stationary gas turbine is provided. The camber line of the blade tip side profile of the blade of the compressor blade includes at least two inflection points for reducing radial gap losses. By means of two inflection points, there is a concavely designed suction side contour segment in a segment from 35% to 50% for the suction side contour, and a convexly implemented pressure side contour segment for the pressure side contour. It is possible by means of the geometry to generate low-loss gap vortices, in order to increase the overall efficiency of an axial compressor including the compressor blades.

Compressor blade for an axial compressor

A ring diffuser for an axial turbomachine is provided. The ring diffuser includes an outer wall and an inner wall coaxial thereto, between which a diffuser channel extends in a ring shape along an axial extension from an inflow-side end diverging to an outflow-side end, wherein the inner wall and the outer wall each include a wall surface bounding the diffuser channel. In order to specify a diffuser channel that is tailored to an inflow that is uneven along the circumference and that facilitates a particularly efficient conversion of kinetic energy into static energy, a rotationally asymmetric wall surface bounding the diffuser channel on the inlet side is provided.

Ring diffuser for an axial turbomachine

An arrangement with a ring diffuser for an axial turbomachine, for example for a gas turbine, is provided. The ring diffuser includes an outer wall and with an inner wall coaxial thereto, between which a diffuser duct extends in ring form along an axial extent so as to diverge from an inflow-side end to an outflow-side end, the inner wall and the outer wall in each case including a wall surface delimiting the diffuser duct. In order to specify a diffuser duct which is adapted to an inflow uneven along the circumference and by means of which a specially efficient conversion of kinetic energy into static energy is possible, the wall surface which delimits the diffuser duct on the inlet side is rotationally asymmetrical.

An axial turbomachine including a rotor blade cascade is provided. The rotor blade cascade includes rotor blades each including front edge, a radially outer free blade tip and an annular enclosure enclosing the rotor blade cascade with an annulus inner side by means of which the annular enclosure is arranged directly adjacent to the blade tips to give a radial gap between the enveloping ends of the blade tips and the annulus inner side. The rotor blades include a radial projection in the region of the front edge on the blade tip and the annular enclosure has an annular radial recess in the annulus inner side, arranged at a radial distance from the ends of the blade tips, such that in the main flow direction of the axial turbomachine the line of the radial projections on the side facing the radial gap matches the line of the radial recess.

Axial turbo engine with low gap losses

A gas turbine engine having an axial flow compressor, an annular combustion chamber, a turbine, and a diffuser. The diffuser includes a flow-dividing element formed by an inner deflecting flank and an outer deflecting flank that divides a compressed gas flow into two partial flows at a branching point. The two deflecting flanks define: an angle of less than 90&#xb0; along at least a portion of the deflecting flanks, and an angle between 15&#xb0; and 90&#xb0; between the deflecting flank angle bisector and the turbine longitudinal axis. The deflector also includes a main deflecting region arranged upstream of the branching point and directed at an acute angle from the turbine longitudinal axis toward an inner combustion chamber shell.

Diffuser arranged between the compressor and the combustion chamber of a gas turbine

Disclosed is an annular flow duct for a turbomachine, the flow duct arranged concentrically about a machine axis running in the axial direction and defined by a boundary wall of circular cross section for directing a main flow, wherein the boundary wall has a plurality of return flow passages distributed over its circumference and through which in each case a partial flow is detached from the main flow at a bleed position and returned to the main flow at a feed position situated upstream of the bleed position, and having aerofoils, arranged radically in the flow duct, of a blade ring, the aerofoil tips of which lie opposite the boundary wall, with a gap formed in each case, wherein the aerofoils are movable in a predetermined rotation direction along the circumference of the boundary wall. Furthermore, a compressor which is insensitive to pumping and flow separations, as viewed in rotation direction, the bleed position of each return flow passage lies upstream of the corresponding feed position.

Annular flow duct for a turbomachine through which a main flow can flow in the axial direction

Growing community of inventors

Sprockhövel, Germany

Christian Cornelius