Abstract: A method for producing a component of a turbomachine is disclosed. The method includes a) layer-by-layer deposition of a powder component material onto a component platform in a region of a buildup and joining zone, where the deposition takes place in accordance with layer information of the component to be produced; b) local layer-by-layer fusion or sintering of the powder component material by energy supplied in the region of the buildup and joining zone, where the buildup and joining zone is heated to a temperature just below a melting point of the powder component material; c) layer-by-layer lowering of the component platform by a predefined layer thickness; and d) repetition of steps a) to c) until the component is finished. A device for producing a component of a turbomachine is also disclosed.
Abstract: A device (2) having at least one connecting element (24, 26) for fastening a component (4) to, respectively for detaching it from a component carrier (6) using adhesive bonds; a mechanical peeling device (8) being provided for breaking the adhesive bonds. Also a method for fastening a component (4) to, respectively for detaching it from a component carrier (6).
Abstract: A gas turbine engine is utilized in combination with a gear reduction to reduce the speed of a fan relative to a low pressure turbine speed. The gas turbine engine is designed such that a blade count in the low pressure turbine multiplied by the speed of the low pressure turbine will result in operational noise that is above a sensitive range for human hearing. A method and turbine module are also disclosed.
Type:
Grant
Filed:
February 23, 2012
Date of Patent:
August 21, 2012
Assignees:
United Technologies Corporation, MTU Aero Engines GmbH
Abstract: A blade channel having a not-axially-symmetric end wall contour in a turbomachine is disclosed, the end wall contour having at least one individual contour in the form of an elevation, on the pressure side, and at least three individual contours in the form of two recesses and one elevation, on the suction side, the elevation being situated between the recesses in the flow direction; a turbomachine having a plurality of blade channels of this type is also disclosed.
Abstract: A ring element for a turbomachine, in particular for an aircraft gas turbine, is disclosed. The ring element has a ring element main body that has two adjacently arranged ring ends, the ring ends being connected to one another in a form-locking manner with respect to an axial plane. Also disclosed is a turbomachine having at least one such ring element.
Type:
Application
Filed:
February 1, 2012
Publication date:
August 9, 2012
Applicant:
MTU Aero Engines GmbH
Inventors:
Johannes RABE, Franz-Dieter SCHMIDT, Wilfried SCHUETTE
Abstract: The invention relates to a sealing arrangement for a non-hermetic seal of a gap (11) between a stator and a rotor with a brush seal (13), wherein the brush seal (13) has a bristle package (14) with several bristles (15), the bristles are wound about a core element (16) and fixed to the core element (16), wherein the brush seal (13) is disposed at least sectionally in an accommodating space, which is bounded by a supporting element (18) and a covering element (19), and wherein the bristles (15) have sealing bristle sections (21), which are at an angle to the bristle shafts (20) and the ends of which run against a sealing surface (22) of the rotor (12). According to the invention, the covering element (19) has a support surface (23) at an end adjacent to the sealing bristle sections (21), wherein the supporting surface (23) contacts the bristle shafts (20) in a defined, locally limited region, so that the radial mobility of the sealing bristle sections (21) is limited in the direction of the rotor (12).
Type:
Grant
Filed:
October 11, 2007
Date of Patent:
August 7, 2012
Assignee:
MTU Aero Engines GmbH
Inventors:
Alfons Gail, Wilhelm Gräbeldinger, Alexander Rauschmeier
Abstract: An apparatus and method for non-contact blade vibration measurement is disclosed. The apparatus and method includes sensors which are arranged around the circumference of a rotor which is formed with rotor blades, a signal detection unit, and an evaluation unit. Devices are provided to determine the rotor position and/or the housing deformation. This avoids the technical problems of the prior art, and provides an improved apparatus and an improved method for non-contact blade vibration measurement. In particular, the attainment according to the invention eliminates the effect of rotor radial movements and housing deformations, i.e., oval deformations, on the measurement data, thereby ensuring high amplitude resolution for the vibration analysis under all conditions.
Abstract: A method for producing a thin-walled structural component from a casting material. The casting material is supplied as a powder, and the powder is deposited on a support (1) by a kinetic cold gas spraying process so as to form the structural component (11, 11?). A structural component which is made of a casting material and in which the structure is formed from a plurality of particles (17) that are interlinked and deformed using a cold gas spraying process.
Type:
Application
Filed:
August 9, 2010
Publication date:
July 5, 2012
Applicant:
MTU Aero Engines GmbH
Inventors:
Andreas Jakimov, Erwin Bayer, Karl-Heinz Dusel, Carsten Butz
Abstract: A blading for a turbine, in particular a gas turbine, is disclosed. The blades of the blading in a section near the tip have a distribution ratio (t/l) of at least 0.70, in particular at least 0.9, and/or at most 0.97, in particular at most 0.95. A downstream flow angle (?) is at most 167°, in particular at most 165°, and at least 155°, in particular at least 160°. In addition, or alternatively, an acceleration ratio (w2/w1) is at least 1.4, in particular at least 1.5.
Abstract: A method for heat-treating gas turbine blades, namely for locally heat-treating at least one gas turbine blade in a blade section thereof; a blade root section, which is not to be heat-treated, of the gas turbine blade being positioned in a holding receptacle to prevent an unacceptable heating of the particular blade root section, which is not to be heat-treated, during the heat treatment of the particular blade section. The blade root section of the gas turbine blade is positioned in an interior space in a way that allows a remaining interior space of the holding receptacle, to be filled with a filler material; the holding receptacle, together with the gas turbine blade, being subsequently positioned in a heat treatment chamber to enable the gas turbine blade in the heat treatment chamber to undergo local heat treatment under vacuum.
Abstract: A method for replacing a blade (12) of a rotor (2) having integral blades, wherein a new blade (22) is joined to a base (34) arranged on a separation surface (28) on the main rotor body, and to a rotor repaired or mended according to said method.
Abstract: A blade segment (1) for a fluid flow machine (2), including: a first blade (3); a first shroud (7) attached to the tip of the first blade (3) and having a contact face (11); a second blade (3?); and a second shroud (7?) attached to the tip of the second blade (3?) and having a counter-contact face (12?) for engagement with the contact face (11) of the first shroud (7) in the radial direction (R) in such a way that a friction lock (RF) can be created between the contact face (11) and the counter-contact face (12?).
Abstract: An anti-wear coating, in particular an anti-erosion coating, which is applied to a surface of a component that is stressed under fluid technology and is to be protected, in particular a gas turbine part, is disclosed. The anti-wear coating includes one or more multilayer systems applied in a repeating order to the surface to be coated, and the/each multilayer system includes at least one relatively soft metallic layer and at least one relatively hard ceramic layer. All the layers of the/each multilayer system are based on chromium, and a diffusion barrier layer is applied between the surface to be protected and the multilayer system(s).
Type:
Application
Filed:
May 17, 2008
Publication date:
June 7, 2012
Applicant:
MTU Aero Engines GmbH
Inventors:
Wolfgang Eichmann, Falko Heutling, Thomas Uihlein
Abstract: A method and a measuring system for characterizing a deviation of an actual dimension of a component from a nominal dimension of the component, is disclosed. In an embodiment, the method includes a) determining at least one measured value characterizing the actual dimension at a position of the component by a measuring device; b) making a nominal value available with which the nominal dimension is characterized as a function of the position of the measured value; c) determining a spatial distance between the measured value and the nominal value; d) making a limiting criterion available with which a permissible deviation from the nominal value is characterized as a function of the position of the measured value; and e) determining a tolerance utilization value characterizing the deviation for the measured value as a function of the spatial distance and of the limiting criterion.
Abstract: A method for manufacturing integrally bladed rotors (preferably gas turbine rotors) is provided. The method includes the steps of a) providing a basic rotor body; b) placing the basic rotor body into an electrolyte; c) electrochemically machining the basic rotor body by simultaneously manufacturing a plurality of unmachined blades; and d) subsequently machining the unmachined blades to provide hydrodynamic surfaces, in particular a suction side and a pressure side, in the area of each unmachined blade.
Type:
Grant
Filed:
October 21, 2005
Date of Patent:
May 29, 2012
Assignee:
MTU Aero Engines GmbH
Inventors:
Erwin Bayer, Martin Bussmann, Thomas Kraenzler, Albin Platz, Juergen Steinwandel
Abstract: A method for hardfacing a metal component surface (14, 16), especially a shroud surface of a turbine blade made of a TiAl alloy, with at least one metal material (18, 20), in particular a Co—Cr alloy. The hardfacing coating is produced separately from the component surface and is then joined to the component surface in a high-temperature soldering process. A turbine blade including such a hardfacing coating, primarily in a shroud region (2).
Type:
Application
Filed:
July 8, 2010
Publication date:
May 24, 2012
Applicant:
MTU Aero Engines GmbH
Inventors:
Karl-Hermann Richter, Ulrich Knott, Piotr Kowalczyk
Abstract: A joining element, for example, an adapter, for connecting a rotor blade to a rotor disc or a rotor ring, or a rotor blade per se of an aircraft engine is disclosed. The joining element has a base body and a joining surface, with the joining surface formed by a fusion weldable coating made of a compressed powder that is connected in a positive fit to a contour of the base body. A method for the production of such a joining element, as well as an integrally bladed rotor, is also disclosed.
Abstract: The present invention concerns a device for measuring of layer thicknesses, especially for measuring of layer thicknesses of a structural part (18) during or after a coating process, with at least a first and a second path length measuring device (12, 14), wherein a first path length (a) to a surface (20) of a layer (22) being applied to the structural part (18) is measured by means of the first path length measuring device (12) and a second path length (b) to an uncoated surface (24) of the structural part (18) is measured by means of the second path length measuring device (14) continuously or at predetermined moments of time.
Type:
Grant
Filed:
October 30, 2007
Date of Patent:
May 15, 2012
Assignee:
MTU Aero Engines GmbH
Inventors:
Andreas Jakimov, Manuel Hertter, Stefan Schneiderbanger
Abstract: A turbine vane of a gas turbine, especially a gas turbine aircraft engine, is disclosed. The turbine vane includes a vane base body with an outer surface forming a suction side and a pressure side, the outer surface of the vane base body being at least partially coated with a thermal barrier coating. The thermal barrier coating extends continuously or uninterruptedly at least largely over the suction side and largely over the pressure side of the surface of the vane base body, with the layer thickness of the thermal barrier coating being variable or adjustable.
Type:
Grant
Filed:
October 6, 2007
Date of Patent:
May 8, 2012
Assignee:
MTU Aero Engines GmbH
Inventors:
Hernan Victor Arrieta, Rolf Kleinstueck, Paul Storm, Peter Wiedemann
Abstract: A multistage compressor for a gas turbine, having a compressor housing, guide vanes, impeller vanes, a hub area, discharge ports, and injection ports for stabilizing the compressor flow, wherein the discharge ports are situated at least one stage downstream from the injection ports, the injection ports being fashioned as nozzles, and the geometry of these nozzles is designed such that the flow speed of the injected air stream is significantly greater than the speed of the flow in a critical cross-section.