Mtu Aero Engines Ag

Markus Brettschneider

Inga Mahle

Fadi Maatouk

Martin Pernleitner

Markus Schlemmer

Detlef Korte

Bernd Kislinger

Franz Malzacher

Nina Wolfrum

Rudolf Stanka

Guenter Ramm

Klaus Wittig

Christoph Lauer

Bjoern Grueber

Christoph Mueller-Schindewolffs

Inga Dr Mahle

A guide vane ring and/or a rotor blade ring for a low-pressure turbine of an aircraft gas turbine of a turbofan engine includes a plurality of airfoils disposed consecutively in a direction of rotation. The plurality of airfoils includes a first airfoil and a second airfoil adjacent to the first airfoil in the direction of rotation. The first and second airfoils are disposed relative to each other such that a ratio e/t, viewed in a conical section, results from

Guide vane ring and rotor blade ring for a turbofan engine

A stator vane for a turbomachine, including an airfoil having a leading edge and a trailing edge, which are interconnected by a suction side and a pressure side, the airfoil extending substantially in a radial direction between an inner platform and an outer platform, wherein, when viewed in an axial direction of the turbomachine toward the trailing edge, the trailing edge extends so as to be inclined at a first angle to the pressure side radially on an outside in relation to the radial direction at a circumferential position at which the trailing edge meets the outer platform, and wherein the first angle between the trailing edge and a tangent to the outer platform at a transition from the outer platform to the trailing edge is between 72° and 84°.

Stator vane for a turbomachine

The invention relates to a blade or vane, particularly of a turbine stage of a gas turbine, in particular of an aircraft gas turbine, having a blade or vane root and a blade or vane element joined to the blade or vane root, wherein the blade or vane element has a pressure side and a suction side, and wherein the blade or vane root has at least one raised region on its radial outer side facing the blade or vane element. It is proposed according to the invention that the blade or vane has a first raised region on the pressure side and a second raised region on the suction side, wherein the highest point of the first raised region is disposed essentially directly adjacent to the pressure side, and the highest point of the second raised region is disposed essentially directly adjacent to the suction side.

Blade or guide vane with raised areas

A guide vane airfoil for placement in a flow path portion of a turbomachine is provided, which, relative to a flow pattern in flow path portion, has a leading edge and, downstream thereof, a trailing edge, as well as a suction side and a pressure side; relative to a longitudinal axis of the turbomachine, viewed in the axial direction, in a radially inner portion, forming a first angle &#x3b1; with a circular arc about the longitudinal axis, and, in a radially outer portion, a second angle &#x3b3; with a circular arc about the longitudinal axis. The guide vane airfoil is inclined in the outer portion, thus &#x3b3;&#x2212;90&#xb0;, in terms of absolute value, being &#x3e;0&#xb0; (|&#x3b3;&#x2212;90&#xb0;|&#x3e;0&#xb0;), and the guide vane airfoil being more highly inclined in the outer portion than in the inner portion, thus &#x3b3;&#x2212;90&#xb0;, in terms of absolute value, being &#x3e;&#x3b1;&#x2212;90&#xb0; (|&#x3b3;&#x2212;90&#xb0;|&#x3e;&#x3b1;&#x2212;90&#xb0;).

Guide vane airfoil for the hot gas flow path of a turbomachine

Described are an airfoil array segment () having at least two airfoils () and a platform () that features an axis asymmetrical platform surface (). This platform surface features an elevation () that extends from the pressure side () of the first to the suction side () of the second airfoil (). A highest point () of the elevation is more proximate to the suction sideof the second airfoil () than to the pressure side () of the first airfoil (). Also described are an airfoil, a platform, an airfoil passage and a turbomachine.

Endwall contouring 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

An airfoil array for a turbomachine, in particular a turbine or compressor stage of a gas turbine. The airfoil array includes at least two airfoils and at least one contoured circumferential surface which connects a pressure side of one airfoil to a suction side of the other airfoil and includes an upstream first section and a downstream second section which adjoins the first section along an elevation contour line; the first section being depressed relative to a rotationally symmetric reference surface containing this elevation contour line away from the airfoils, and the second section not being depressed relative to this reference surface away from the airfoils; this elevation contour line lying in an axial area which terminates at most 30% of an axial chord length of one of the airfoils downstream of its leading edge; and an axial distance of this elevation contour line increasing toward the pressure side and toward the suction side, starting at a point between the pressure and suction sides that is closest to the leading edge; and the first section extending over at least 90% of the space between the pressure side and the suction side.

Turbomachine airfoil array

A blade/vane cascade segment of a blade/vane cascade for a turbomachine is disclosed, which comprises a stage with a stage surface as well as a first blade/vane element and a second blade/vane element. The stage surface comprises a first elevation reaching as far as the pressure side of the first blade/vane element and a second elevation reaching as far as the suction side of the second blade/vane element. A furthest downstream point of a boundary of the second elevation has an axial position which differs from the axial position of at least one highest point of the first elevation by a maximum of 10% of an axial cascade span. Furthermore, a blade/vane cascade, a blade/vane channel, a stage, a turbomachine and an aircraft engine are disclosed.

Contouring a blade/vane cascade stage

Disclosed is an airfoil array segment () of an airfoil array for a turbomachine, the airfoil array segment including a platform () having a platform surface, as well as at least two airfoils (). The platform surface has a depression () which extends up to the first airfoil and contacts the pressure side () of the first airfoil downstream of 80% of the axial chord (g) downstream of the leading edges () and which contacts the pressure side () of the first airfoil downstream up to no more than 80% of the axial chord (g) downstream of the leading edges (). At least one lowest point () of the depression is located at least 90% of the axial chord (g) downstream of the leading edges (). Also disclosed are an airfoil array, an airfoil passage, a platform, a turbomachine, and an aircraft engine.

Contouring of an airfoil array platform

A blade/vane cascade segment for a blade/vane cascade of a turbomachine includes a stage and at least two blade/vane elements that define a blade/vane intermediate strip with an axial cascade span on the stage surface by their leading and trailing edges. A stage edge on the inflow side has a contour with a depression. In the axial direction, this depression extends at most by 10% of the cascade span into the blade/vane intermediate strip. Also provided is a corresponding stage, a blade/vane cascade, a blade/vane channel, and a turbomachine.

Disclosed is a blade/vane cascade segment of a blade/vane cascade of a turbomachine, which comprises at least two blade/vane elements defining an axial cascade width, and a stage with a stage surface, as well as a stage edge on the inflow side. The stage surface has an elevation extending up to the pressure side of a first blade/vane element and a depression extending up to the suction side of the other blade/vane. At least one highest point of the elevation and at least one lowest point of the depression each lie at least 30% and at most 60% of the axial cascade width downstream of the inflow edges of the blade/vane elements. In this case, the elevation and the depression each come close to the stage edge on the inflow side.

Disclosed is a turbomachine including a stator, a rotor rotatable about an axis of rotation, and an annulus for carrying a core flow, the annulus having a side wall on the stator and a side wall on the rotor, at least one airfoil array having a plurality of airfoils being disposed in the annulus. In a departure from an ideal aerodynamic annulus contour, a radial annulus enlargement begins upstream of the airfoils and extends downstream up to an aft portion of the airfoil array that follows the ideal aerodynamic annulus contour. Also disclosed is an airfoil for such a turbomachine.

Turbomachine having an annulus enlargment and airfoil

A blade channel of a turbomachine that is delimited in the circumferential direction of the turbomachine by a pressure side of an airfoil and by an opposite suction side of an adjacent airfoil that, in the radial direction of the turbomachine, is delimited by two opposing side walls, and whose extent in the axial direction of the turbomachine is delimited by leading edges and by trailing edges of airfoils; at least one of the side walls being provided with localized contours, of which at least two are formed as elevations and at least two as depressions; a saddle surface being formed between the contours that, in one rotation, alternately merges into an elevation and a depression; a blade cascade having such blade channels, as well as a turbomachine having such a blade cascade.

Blade channel, blade cascade and turbomachine

The invention relates to a guide vane cluster for a turbomachine, in particular for a turbojet engine, having at least two vanes that are joined together by way of at least one common platform for the radial delimiting of a flow channel of the turbomachine, wherein each vane has a vane element with a suction side and a pressure side that are joined together by a leading edge and a trailing edge, wherein the at least two vanes have different trailing edge wall thicknesses at least in a corresponding radial lengthwise extension region.

Guide vane segment for a turbomachine

A blade cascade for a continuous-flow machine having a non-axisymmetrical side wall contour, whereby the side wall contour has at least one pressure-side elevation and one suction-side depression, whose highest section and lowest section are located over an area 30% to 60% of an extension of the blades in the axial direction, and the axial positions of the outermost sections differ from each other at the maximum by 10% in the axial direction.

Growing community of inventors

Karlsfeld, Germany

Markus Brettschneider