Abstract: The invention relates to a turbine blade, including a blade having a front edge and a rear edge, which blade transitions by means of a shaft into a blade root designed for fastening the turbine blade, and including a platform, which is arranged at the lower end of the blade in order to bound a flow channel. The platform is designed as a separate component and can be connected to the blade in a form-fit manner. Flexible application is achieved in that the platform is composed of several individual platform elements, which enclose the blade in the assembled state.

Abstract: Aspects of the present invention relate to systems and methods of a vane design utilizing welding techniques. The present invention concerns a method for preventing cracking within a vane assembly utilizing full penetration welding. Additional embodiments concern a vane design that, when assembled with another vane, comprises an axial slot that prevents cracking within a vane assembly.

Abstract: A rotor blade arrangement (20), especially for a gas turbine, which can be fastened on a blade carrier (19) and includes in each case a blade aerofoil element (10) and a platform element (14), wherein the platform elements (14) of a blade row form a continuous inner shroud. With such a blade arrangement, a mechanical decoupling, which extends the service life, is achieved by the blade aerofoil element (10) and the platform element (14) being formed as separate elements and by being able to be fastened in each case separately on the blade carrier (19).

Abstract: Disclosed is a turboengine blading member, having at least one platform member and at least one airfoil member. A receiver through opening extends through the platform member. The connector post is received within the receiver through opening. Each of the connector post and the receiver through opening have a retainer flute provided on the circumferential wall and extending along at least a part of the circumferential extent. The retainer flutes are arranged juxtaposed each other with the open sides facing each other such as to jointly form a joint retainer cavity. First and second retainer members are bonded to each other to provide a common retainer member extending into both corresponding retainer flutes, thereby retaining the connector post within the receiver through opening and interlocking the airfoil member and the platform member.

Abstract: The disclosure relates to a sandwich arrangement having at least two peripheral disposed ceramic panels and a ceramic felt which is inserted between a first and second ceramic panel, The material of the first ceramic panel is equal or different to the material of the second panel, wherein the ceramic felt is formed by a textile structure with a regularly or quasi-regularly structured woven fibres. The fibres are made of at least one material and/or composition, wherein at least one adhesive is provided between an underside of the panels and adjacent fibres.

Abstract: In a method for manufacturing an article (1), particularly a prototype of a product or component, a tool prototype or spare part, by using selected laser melting, for the application onto the article (1) of a layer (13) or portion of a second metallic material, which is different from the material of the first metallic powder (4), a tape (12), sheet (14), foil or three-dimensional pre-form (18) of a second material is applied to the article (1) and is heated by a focused laser or electron beam (6) to a specified temperature such that the tape (12), sheet (14), foil or pre-form, respectively, are made molten by the electron laser beam (6), wherein the focused beam (6) is applied to a given area corresponding to a selected cross-sectional area of the model of the article (1) under formation of a new layer or part made of second material integral with the article (1).

Abstract: The invention relates to a damper for reducing pulsations in a gas turbine, which includes an enclosure, a main neck extending from the enclosure, a spacer plate disposed in the enclosure to separate the enclosure into a first cavity and a second cavity and an inner neck with a first end and a second end, extending through the spacer plate to interconnect the first cavity and the second cavity. The first end of the inner neck remains in the first cavity and the second end remains in the second cavity. A flow deflecting member is disposed proximate the second end of the inner neck to deflect a flow passing through the inner neck. With the solution of the present invention, as a damper according to embodiments of the present invention operates, flow field hence damping characteristic in the second cavity constant regardless the adjustment of the spacer plate in the enclosure.

Abstract: The present invention discloses a novel and improved device and method for securing a combustion liner within a gas turbine combustor. The improved configuration comprises a plurality of equally spaced support tab assemblies secured about the combustion liner and positioned radially outward of the combustion zone. The support tab assemblies comprise parallel liner tabs which provide increased flexibility, greater structural support, and lower operating stresses.

Abstract: An auto-tune controller and tuning process implemented thereby for measuring and tuning the combustion dynamics and emissions of a GT engine, relative to predetermined upper limits, are provided. Initially, the tuning process includes monitoring the combustion dynamics of a plurality of combustors and emissions for a plurality of conditions. Upon determination that one or more of the conditions exceeds a predetermined upper limit, a fuel flow split to a fuel circuit on all of the combustors on the engine is adjusted by a predetermined amount. The control system continues to monitor the combustion dynamics and to recursively adjust the fuel flow split by the predetermined amount until the combustion dynamics and/or emissions are operating within a prescribed range of the GT engine. Additionally, a method of automated extended turndown of a GT engine to find a minimum load is provided.

Abstract: An internally cooled airfoil for a rotary machine, for example, a gas turbine engine includes a suction and pressure side wall each extending in an axial direction, i.e. from a leading to a trailing edge of the airfoil. A suction wall sided cooling channel and a pressure wall sided cooling channel extend in the axial direction. A feed chamber is defined between a first and second inner wall for feeding the suction wall and pressure wall sided cooling channel each by at least one through hole inside of the first and second inner wall. The suction wall sided cooling channel and the pressure wall sided cooling channel extend into the trailing edge region separately. The suction wall sided cooling channel and the pressure wall sided cooling channel join before discharging at the trailing edge.

Abstract: Disclosed is a method for manufacturing a mechanical component, by applying additive manufacturing, wherein the method includes depositing a powder material and locally melting and resolidifying the powder material, thereby providing a solid body, the method including choosing a powder material of a specified chemical composition.

Abstract: A method for repairing a blade tip of a turbine blade based on a dual fusion weld process for portions of a tip plate and a mechanical locking bead over the tip plate. A damaged portion of blade is removed. The damaged portion is removed at an initial cut level. An undamaged portion of the turbine blade exists as a base below the initial cut level. A replacement section is rebuilt via laser cladding. The replacement section is machined to form a tip pocket. A tip plate is placed in the tip pocket. The tip plate comprises at least a first portion and a second portion. The tip plate is welded using a dual fusion weld process comprising a first weld process and a second weld process. A squealer tip is built using a welding process that forms a mechanical locking bead over the tip plate.

Abstract: The invention refers to a method for producing a near-surface cooling passage in a thermally highly stressed component, which includes: a) providing a component which has a surface on a hot side in a region which is to be cooled; b) letting a channel into the surface; c) inserting a cooling tube into the channel; d) filling the channel, with the cooling tube inserted, with a temperature-resistant filling material in such a way that the inserted cooling tube is embedded into the filling material, leaving free an inlet and an outlet; and e) covering the channel, with the cooling tube embedded, with an anti-oxidation, temperature-stable cover layer. The method is inexpensive and can be used in a flexible manner in the most diverse situations in order to save cooling medium or to reduce the thermal load.

Abstract: A method for controlling a gas turbine assembly includes: a compressor in which compression of the outside air occurs for producing a flow of compressed air; a sequential combustor including a first combustor, in which combustion of a mixture of fuel and compressed air arriving from the compressor occurs for producing a flow of hot gasses, and a second combustor which is located downstream of the first combustor and in which combustion of a mixture of fuel and hot gasses arriving from the first combustor occurs; an intermediate turbine in which a partial expansion of the hot gasses arriving from the first combustor occurs; and a second combustor in which combustion of a mixture of fuel and hot gasses arriving from the intermediate turbine occurs; the method further includes, on a start-up transient operating phase of the gas turbine assembly, the step of controlling the fuel mass flow-rate supplied to the first and/or the second combustor on the basis of the flame temperature inside the first combustor.

Abstract: An improved intake arrangement, for a compressor having a compressor blading, includes a manifold divided by a barrier into two sections, to convey, from one section, a flue gas stream, and, from other section, an air stream. Further, the intake arrangement includes a converging section configured to the manifold and extends convergingly to the compressor defining an inlet to the compressor blading. The converging section includes inner and outer ring members disposed coaxially to each other, between which there extends, coaxially and convergingly, the barrier to at least up to a certain distance within the converging section, defining a converging nozzle therebetween. The converging nozzle includes a mixing feature adapted to enhance mixing of the flue gas and air streams.

Abstract: The invention refers to a method for producing means with thermal resist for applying at a surface of a heat exposed component. The method includes providing at least one mixture of at least two different inorganic powder materials: with different chemical composition and thermal expansion coefficient, where at least one of the powder materials experiences volume changes due to crystallographic phase change at a given temperature, and/or\ where at least one of the two powder materials experiences volume changes due to chemical composition change, forming the at least one mixture into a green body shape, sintering the green body to obtain a ceramic body and cooling the ceramic body and causing micro-cracking during heating and or/sintering and/or cooling to obtain the means with thermal resist. The forming is performed such that a first mixture forming a first layer onto which a second mixture is applied forming a second layer differs in porosity from that first layer to obtain a layered green body.

Abstract: A combustion device (1) for a gas turbine includes portions (12) having an inner and an outer wall (13, 14) with an interposed noise absorption plate (15) having a plurality of holes (16). The combustion device (1) further has first passages (17) connecting zones between the inner wall (13) and the plate (15) to the inside of the combustion device (1) and second passages (21) for cooling the inner wall (13). The portions (12) also have an inner layer (22) between the inner wall (13) and the plate (15) defining inner chambers (23), each connected to at least a first passage (17), and an outer layer (24) between the outer wall (14) and the plate (15) defining outer chambers (25) connected to the inner chambers (23) via the holes (16) of the plate (15).

Abstract: The present invention discloses a novel and improved apparatus and method for reducing the emissions of a gas turbine combustion system. More specifically, a combustion system is provided having a first combustion chamber and a premixer positioned proximate an outlet end of a combustion liner for mixing a second fuel/air mixture with hot combustion gases and burning the subsequent mixture to achieve reduced emissions levels. The premixer is positioned generally about the combustion liner and includes a plurality of channels and fuel injectors for introducing a fuel/air mixture, induced with a swirl, into a second, axially staged combustor.

Abstract: The invention relates to a coating system for a component of a turbomachine, which includes at least two different base powders. Each of the at least two different base powders has an individual predetermined distribution within the coating system. Further, each of the at least two different base powders is responsible for a specific property of the coating system.

Abstract: A robotic system includes a main drive unit, a non-actuated extendable arm unit an axial drive unit, an arm guidance member, and a head articulation unit. The unit includes a mounting structure, and an arm storage unit coupled to the mounting structure. The arm unit is coupled to the arm storage unit to be moved axially in a lateral plane from the arm storage unit. The arm unit is flexible and thin in the lateral plane, and rigid and wide in a vertical plane. The unit may be coupled to the arm storage unit to enable axial extension and retraction of the unit. The head articulation unit is coupled to the arm unit to actuate thereto.