Abstract: A system and method is described that allows a communication session to be transferred between endpoints. A user may provide an indication using a first endpoint of the communication session that they desire to transfer the communication session to a second endpoint. Communication session endpoints associated with the user that are available to transfer the communication session can be determined and the user may select one of the available endpoints and then the communication session can be transferred from the first endpoint to the selected endpoint.

Abstract: A system and method is described that allows a communication session to be transferred between endpoints. A user may provide an indication using a first endpoint of the communication session that they desire to transfer the communication session to a second endpoint. Communication session endpoints associated with the user that are available to transfer the communication session can be determined and the user may select one of the available endpoints and then the communication session can be transferred from the first endpoint to the selected endpoint.

Abstract: In the case of a device and a method for monitoring the service life of an engine or of a turbine having a compressor blisk and/or a turbine blisk, momentary stresses of blisk substructures, such as the blade (1), the disk (2), and the join region (3) between the blade (1) and the disk (2), are calculated on the basis of operating parameters that are measured in the course of engine or turbine operation; and accumulated damage to the individual substructures, that was caused by the momentary stresses, is estimated.

Abstract: An integrally bladed rotor disk (20) for a turbine, including rotor blades (40), which are joined in a substance-to-substance bond to a disk element (30), and a sealing device (60) for preventing or reducing the extent to which cooling air is able to flow from a high-pressure side (12) of the rotor disk (20) through openings (24) on the rotor disk (20) to a low-pressure side (14) of the rotor disk (20).

Abstract: A shroud for the rotating blades of a turbo machine, particularly a gas turbine, in which shroud is arranged along the circumference of a row of several rotating blades disposed on a rotor and has at least one separation gap along its circumference. The separation gap is formed in zigzag shape and has at least three damping gaps that are distanced from one another and extend at an angle relative to an axis of rotation of the rotor, and adjacent to these, has connection gaps each connecting damping gaps or extending the latter in the direction of shroud edges, wherein, when the rotor rotates, the gap width of damping gaps is reduced until the gap walls forming the damping gaps come to rest against each other.

Abstract: A graphite-metal carbide-rhenium article of manufacture is provided, which is suitable for use as a component in the hot zone of a rocket motor at operating temperatures in excess of approximately 3,000 degrees Celsius. One side of the metal carbide is chemically bonded to the surface of the graphite, and the rhenium containing protective coating is mechanically bonded to the other side of the metal carbide. Rhenium forms a solid solution with carbon at elevated temperatures. The metal carbide interlayer serves as a diffusion barrier to prevent carbon from migrating into contact with the rhenium containing protective coating. The metal carbide is formed by a conversion process wherein a refractory metal carbide former is allowed to react with carbon in the surface of the graphite. This structure is lighter and less expensive than corresponding solid rhenium components.

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.

Abstract: The English-language Abstract from the international application is to be retained and is therefore not duplicated in the specification.

Abstract: An integrally bladed rotor disk (20) for a turbine, including rotor blades (40), which are joined in a substance-to-substance bond to a disk element (30), and a sealing device (60) for preventing or reducing the extent to which cooling air is able to flow from a high-pressure side (12) of the rotor disk (20) through openings (24) on the rotor disk (20) to a low-pressure side (14) of the rotor disk (20).

Abstract: In the case of a device and a method for monitoring the service life of an engine or of a turbine having a compressor blisk and/or a turbine blisk, momentary stresses of blisk substructures, such as the blade (1), the disk (2), and the join region (3) between the blade (1) and the disk (2), are calculated on the basis of operating parameters that are measured in the course of engine or turbine operation; and accumulated damage to the individual substructures, that was caused by the momentary stresses, is estimated.

Abstract: The invention relates to a shroud for rotating blades of a turbo machine, in particular a gas turbine. Said shroud (10) is arranged along the circumference of a row of several rotating blades (14, 50) disposed on a rotor (12) and has at least one separation gap (16, 16?, 16?) along the circumference. The separation gap (16, 16?, 16?) is serrated and has at least three damping gaps (20, 22, 24, 26, 28) that are located at a distance from one another and extend at an angle from the axis of rotation (18) of the rotor (12), and adjacent connection gaps (30, 32, 34, 36, 38, 40) which connect the damping gaps (20, 22, 24, 26, 28) or extend the same in the direction of the shroud edges (46, 48). When the rotor (12) rotates, the width of the damping gaps (20, 22, 24, 26, 28) is reduced to the point where the gap walls (42, 44) forming the damping gaps (20, 22, 24, 26, 28) rest against each other.

Abstract: An apparatus and method for attaching a rotor blade to a rotor, particularly for a gas turbine, is disclosed. The rotor blade has a blade vane and a blade root, such that the blade vane can be attached to or anchored in the rotor by the blade root. The blade root is configured asymmetrically, such that the front region of the blade root as seen in the direction of rotation of the rotor blade is provided with more anchoring teeth than a rear region of the blade root as seen in the direction of rotation of the rotor blade.