Mtu Aero Engines Gmbh

Mtu Aero Engines Ag

Mtu Motoren-Und Turbinen-Union Muenchen, Gmbh

Martin Hoeger

Guenter Ramm

Franz Malzacher

Yavuz Guendogdu

Irene Raab

Inga Mahle

Fadi Maatouk

Jochen Gier

Nicolas Thouault

Marc Nagel

Uwe Schmidt-Eisenlohr

Günter Ramm

Yavuz Gündogdu

The invention relates to a flow arrangement for placing in the hot gas duct of a turbomachine, having a first surrounding-flow structure and a second surrounding-flow structure, the surrounding-flow structures each having, in reference to the surrounding flow in the hot gas duct, a leading edge and, downstream thereof, a trailing edge, wherein the second surrounding-flow structure is provided as a deflecting blade with a suction side and a pressure side and has a lesser profile thickness than the first surrounding-flow structure, which is arranged on the suction side of the second surrounding-flow structure, and wherein, although the second surrounding-flow structure has a partial axial overlap with the first surrounding-flow structure referred to a longitudinal axis of the turbomachine, the trailing edge of the second surrounding-flow structure is, at the same time, displaced axially downstream relative to the trailing edge of the first surrounding-flow structure.

Flow arrangement for placing in a hot gas duct of a turbomachine

The present invention relates to a turbine module for a turbomachine, comprising a first flow bypass structure and a second flow bypass structure, said flow bypass structures being arranged in a hot gas duct, which is bounded by the turbine module and is designed to convey a hot gas and, namely, being arranged in succession in relation to a longitudinal axis of the turbine module in a direction of rotation, wherein, in relation to the bypass flow in the hot gas duct, the flow bypass structures each have a leading edge, and, downstream thereto, a trailing edge, and the second flow bypass structure is provided as a deflecting blade, wherein the second flow bypass structure has a smaller profile thickness than the first flow bypass structure, and wherein the hot gas duct is enclosed by a radial width.

Turbine module for a turbomachine

An extruded profile for manufacturing a blade of an outlet guide vane of a turbine engine. A cross-sectional area has an axial length LAX and a thickness D/LAX relative to the axial length LAX. A cross-sectional area has an at least nearly axisymmetric leading edge region, a first transition region having a varying relative thickness D/LAX. A first constant region has a relative thickness D/LAX at least substantially constant and, relative to a leading edge of the extruded profile, begins at the closest at 10% LAX and ends at the furthest at 50% LAX. A second transition region has a varying relative thickness D/LAX and, relative to the leading edge of the extruded profile, begins at the closest at 30% LAX and ends at the furthest at 90% LAX. A second constant region has a relative thickness D/LAX at least substantially constant and an axial length X of 40% LAX at most; and an at least nearly axisymmetric trailing edge region.

Extruded profile for manufacturing a blade of an outlet guide vane

The present invention relates to a subassembly for a gas turbine, in particular a gas turbine aircraft engine, having a turbine casing (); a midframe (), which is adjacent downstream to the turbine casing and has a number of support ribs () spaced apart in the peripheral direction. The turbine casing and the midframe define a flow duct () for a working gas exiting a combustion chamber of the gas turbine, and a cavity, in particular a cooling air duct (), with an opening on the flow duct side is formed between the turbine casing and the midframe. An edge contour () of the opening on the turbine casing side varies along the periphery radially and/or axially.

Gas turbine subassembly

Intermediate housing (), in particular of turbines () of a gas turbine engine, having a radially inner bounding wall () and having a radially outer bounding wall ('), having a crossflow channel (), which is formed by the bounding walls (&#x2032;) and within which at least one supporting rib () is positioned that has a leading edge (), a trailing edge (), as well as side walls () extending between the leading edge () and the trailing edge () that direct a gas flow traversing the crossflow channel (); the radially outer bounding wall () having a contour that changes in the circumferential direction at least in one section upstream of the supporting rib ().

Intermediate housing of a gas turbine having an outer bounding wall having a contour that changes in the circumferential direction upstream of a supporting rib to reduce secondary flow losses

A turbo engine, particularly a gas turbine aircraft engine, has compressor components, turbine components, and at least one combustion chamber. At least one support rib is in flow channel between two turbine components, connected one behind the other. Each support rib diverts a flow through the flow channel. A preferably cylindrical guide element runs within each support rib. Each support rib has a suction side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. Each support rib has a pressure side with a greater thickness toward a radially inner flow channel wall as well as toward a radially outer flow channel wall, when viewed in the radial direction. The front edge and the rear edge of each support rib are inclined in the meridian direction.

Turbo engine

A tandem blade design for an axial turbomachine, comprising a front blade and a rear blade disposed with an offset thereto in the circumferential direction and in the axial direction. The rear blade is profiled and positioned with respect to the front blade such that it raises the speed level at the trailing edge of the front blade in a predetermined working range in interaction with the front blade.

Tandem blade design

A compressor, particularly a high-pressure compressor, of a gas turbine, particularly of an aircraft engine, includes at least one rotor and a number of blades (), which are assigned to the or to each rotor and which rotate together with the respective rotor. Each blade () is delimited, in essence, by a flow entry edge or leading edge (), a flow exit edge or trailing edge (), and by a blade surface (), which extends between the leading edge () and the trailing edge () while forming a suction side () and a pressure side. The leading edges () of the blades () are slanted at a sweep angle that changes with the height of the respective blade () in such a manner that the leading edges () comprise, in a radially external area () of the same, at least one forward sweep angle, a backward sweep angle or zero-sweep angle following in a radially external manner, and a forward sweep angle following, in a radially external manner, the backward sweep angle or the zero-sweep angle.

Compressor of a gas turbine and gas turbine

A flow structure of a gas turbine, in particular for an aircraft engine, in a transitional channel between two compressors or in a transitional channel between two turbines or in a transitional channel of a turbine outlet housing downstream from a low-pressure turbine is disclosed. Supporting ribs are positioned in the transitional channel and spaced a distance apart in the circumferential direction of the transitional channel. At least one guide vane and/or guide rib is positioned between two supporting ribs spaced a distance apart from one another. The flow outlet edge of the guide rib or each guide rib runs upstream from the flow outlet edges of the supporting ribs.

Flow structure for a gas turbine

A flow structure is for a gas turbine, e.g., for an aircraft engine, in a transition channel between two compressors or in a transition channel between two turbines or in a transition channel of a turbine outlet housing downstream from a low-pressure turbine, having support ribs positioned in the transition channel spaced apart from one another around the circumference of the transition channel. A channel wall which delimits the transition channel radially on the inside and/or a channel wall which delimits the transition channel radially on the outside may be drawn inwardly into the transition channel in the area of the outflow edges of the support ribs.

A rotary turbomachine, and particularly a turbojet engine, has a compressor stage that is especially adapted for operation in the transonic region. The compressor stage includes a rotor (1) having a hub (2) and a plurality of compressor blades (3) extending radially therefrom, and a stator having a plurality of compressor blades extending radially between a stator hub and a housing. The hub (2) of the rotor (1), and/or the hub of the stator, and/or the housing of the stator, have a circumferential surface contour that is not continuously rotationally symmetrical. Namely, a concave contour (K) is provided in the circumferential surface of the hub near the base of each blade (3) on the pressure side (PS) thereof, while the circumferential surface on the suction side (SS) of the base of each blade has a contour that is linear, slightly convex, convex/concave, or slightly concave to a lesser degree than the concave contour (K) on the pressure side of each blade. The hub contour slopes and transitions smoothly radially outwardly from the concave contour (K) on the pressure side to the flatter or convex contour on the suction side. In this manner, fluid flow velocities on both the pressure side and the suction side of each blade are reduced, especially under transonic operating conditions, and compression shock losses can be minimized.

Growing community of inventors

Erding, Germany

Martin Hoeger