Abstract: A gas turbine with a sequential combustor arrangement as disclosed includes a first combustor with a first burner for admitting a first fuel into a combustor inlet gas during operation and a first combustion chamber for burning the first fuel, a dilution gas admixer for admixing a dilution gas to the first combustor combustion products leaving the first combustion chamber, a second burner for admixing a second fuel and a second combustion chamber. To assure a temperature profile after the dilution gas admixer and to increase the gas turbine's power and efficiency a vane and/or blade of the turbine has a closed loop cooling. The outlet of the closed loop cooling is connected to the dilution gas admixer for admixing the heated cooling gas leaving the vane and/or blade into the first combustor combustion products.

Abstract: A sealing arrangement is provided at the interface between a combustor and a turbine of a gas turbine. The turbine includes guiding vanes at its inlet. The guiding vanes are each mounted within the turbine at their outer diameter by means of rear outer diameter vane hook and are each at their inner diameter in sealing engagement by means of a front inner diameter vane tooth with a honeycomb seal arranged at the corresponding inner diameter part of the outlet of said combustor. The rear outer diameter vane hook allows a relative movement of the guiding vane in form of a rotation around the rear outer diameter vane hook. The sealing adapts the front inner diameter vane tooth and the corresponding honeycomb seal in their configuration to the rotating relative movement of the guiding vane, such that the compression of said honeycomb seal through transients of the gas turbine is reduced.

Abstract: The invention relates to an axial flow turbine and method of operating thereof. The turbine comprises a last stage of rotating blades located towards a downstream end of the turbine having a distal region at an end of the airfoil of the blades. A monitoring control system has at least one sensor in the distal region of at least one last stage blade for measuring at least one physical property of the airfoil and a control element that is capable of influencing at least one physical property of the distal region. The control system further includes a controller that adjusts the control element based on at least measured physical property so by controlling the at least one physical property.

Abstract: Disclosed is a method for operating a power plant. The method includes providing a power target setpoint to a power plant controller; dependent on said power target setpoint, determining a setpoint of at least one first power plant operation parameter; operating the power plant at a set power output; and and with the at least one first operation parameter adjusted to the setpoint of the operation parameter, applying a measurement to determine an actual value of at least one second operation parameter. The method further includes applying a theoretical power plant model to determine an expected value of the at least one second operation parameter; comparing the expected value of the at least one second operation parameter and the measured value of the at least one second operation parameter; and applying a calibration loop for the theoretical power plant model.

Abstract: A gas turbine with a sequential combustor arrangement as disclosed includes a first combustor with a first burner for admitting a first fuel into a combustor inlet gas during operation and a first combustion chamber for burning the first fuel, a dilution gas admixer for admixing a dilution gas to the first combustor combustion products leaving the first combustion chamber, a second burner for admixing a second fuel and a second combustion chamber. To assure a temperature profile after the dilution gas admixer and to increase the gas turbine's power and efficiency a vane and/or blade of the turbine has a closed loop cooling. The outlet of the closed loop cooling is connected to the dilution gas admixer for admixing the heated cooling gas leaving the vane and/or blade into the first combustor combustion products.

Abstract: A sealing arrangement is provided at the interface between a combustor and a turbine of a gas turbine. The turbine includes guiding vanes at its inlet. The guiding vanes are each mounted within the turbine at their outer diameter by means of rear outer diameter vane hook and are each at their inner diameter in sealing engagement by means of a front inner diameter vane tooth with a honeycomb seal arranged at the corresponding inner diameter part of the outlet of said combustor. The rear outer diameter vane hook allows a relative movement of the guiding vane in form of a rotation around the rear outer diameter vane hook. The sealing adapts the front inner diameter vane tooth and the corresponding honeycomb seal in their configuration to the rotating relative movement of the guiding vane, such that the compression of said honeycomb seal through transients of the gas turbine is reduced.

Abstract: The invention relates to a method for temporary increasing the power of a gas turbine, which is part of a combined-cycle power plant, wherein the exhaust gas of the gas turbine is used in a heat recovery steam generator to generate steam for a water/steam cycle including a steam turbine with a high pressure steam turbine, an intermediate pressure steam turbine and a low pressure steam turbine, whereby the method comprises the step of injecting steam from the water/steam cycle at a predetermined injection pressure into the gas turbine. The efficiency is improved by taking steam of a lower pressure than the predetermined injection pressure from the water/steam cycle, increasing the pressure of the steam taken from the water/steam cycle to the predetermined injection pressure by subjecting it in separate compressing means to a compression step, and injecting the compressed steam into the gas turbine.

Abstract: An improved gas turbine engine is adapted to include an inlet flow control arrangement disposed in an inlet passage at least between a filter member and a row of Variable Inlet Guide Vanes (VIGV), and adjacent the filter member, to enable pressure loss in the inlet passage and supply reduced amount of inlet air to a compressor to enable the reduction of power output of the gas turbine engine at a low load condition.

Abstract: The disclosure relates to a turbo-machine comprising a stator and a rotor arranged rotatable inside the stator as well as at least one electric heating device which is arranged on the surface of at least part of the stator for active clearance control. Besides the turbo-machine, a method for operating the active clearance control comprising electric heating devices is disclosed.

Abstract: The invention relates to an axial flow turbine and method of operating thereof. The turbine comprises a last stage of rotating blades located towards a downstream end of the turbine having a distal region at an end of the airfoil of the blades. A monitoring control system has at least one sensor in the distal region of at least one last stage blade for measuring at least one physical property of the airfoil and a control element that is capable of influencing at least one physical property of the distal region. The control system further includes a controller that adjusts the control element based on at least measured physical property so by controlling the at least one physical property.

Abstract: A single shaft steam turbine plant is disclosed with a first and a second low pressure steam turbine located at either end of the shaft. A generator and at least one high-pressure steam turbine are located on the shaft between low pressure steam turbines.

Abstract: A single shaft steam turbine plant is disclosed with a first and a second low pressure steam turbine located at either end of the shaft. A generator and at least one high-pressure steam turbine are located on the shaft between low pressure steam turbines.

Abstract: A labyrinth seal has a plurality of arcuate labyrinth seal segments 14 in which the confronting ends of adjacent arcuate labyrinth seal segments 14 have planar faces 26 which are oriented radially and axially but which are canted away from a strictly axial orientation 28 by a cant angle. A plurality of arcuate brush seal segments 16 may be retained in the plurality of arcuate labyrinth seal segments 14, thereby forming a combined labyrinth and brush seal.

Abstract: A gas turbine includes a compressor for compressing the combustion air, at least one combustion chamber, in which a fuel is burned while compressed combustion air is supplied, and at least one first turbine arranged downstream of the combustion chamber and in which the hot combustion gases from the combustion chamber are expanded to perform work. Part of the compressed combustion air is branched off, cooled in a cooling air cooler, brought to a lower cooling air pressure by a second turbine and supplied to the gas turbine for cooling purposes.

Abstract: A gas turbine system comprises a compressor that takes in suction air on the inlet side and compresses it to compressor end air that is available on the outlet side, a combustor in which a fuel is burned by using the compressor end air while resulting in the formation of hot gas, as well as a turbine in which the hot gas is expanded while providing work output. In a method for cooling this gas turbine system, compressed air is removed from the compressor, is fed as cooling air for cooling inside an internal cooling channel through thermally loaded components of the combustor and/or the turbine, is then compressed and added to the compressor end air. An improved cooling without disadvantage for the efficiency of the system is achieved in that, in the manner of a targeted leakage, a small part of the cooling air is fed for film cooling into the turbine stream through drilled film cooling openings provided on the components.

Abstract: A gas turbine includes a compressor for compressing the combustion air, at least one combustion chamber, in which a fuel is burned while compressed combustion air is supplied, and at least one first turbine arranged downstream of the combustion chamber and in which the hot combustion gases from the combustion chamber are expanded to perform work. Part of the compressed combustion air is branched off, cooled in a cooling air cooler, brought to a lower cooling air pressure by a second turbine and supplied to the gas turbine for cooling purposes.

Abstract: A gas turbine system comprises a compressor that takes in suction air on the inlet side and compresses it to compressor end air that is available on the outlet side, a combustor in which a fuel is burned by using the compressor end air while resulting in the formation of hot gas, as well as a turbine in which the hot gas is expanded while providing work output. In a method for cooling this gas turbine system, compressed air is removed from the compressor, is fed as cooling air for cooling inside an internal cooling channel through thermally loaded components of the combustor and/or the turbine, is then compressed and added to the compressor end air. An improved cooling without disadvantage for the efficiency of the system is achieved in that, in the manner of a targeted leakage, a small part of the cooling air is fed for film cooling into the turbine stream through drilled film cooling openings provided on the components.

Abstract: A gas turbine system comprises a compressor that takes in suction air on the inlet side and compresses it to compressor end air that is available on the outlet side, a combustor in which a fuel is burned by using the compressor end air while resulting in the formation of hot gas, as well as a turbine in which the hot gas is expanded while providing work output. In a method for cooling this gas turbine system, compressed air is removed from the compressor, is fed as cooling air for cooling inside an internal cooling channel through thermally loaded components of the combustor and/or the turbine, is then recooled and subsequently compressed and added to the compressor end air. The influence of the cooling on the efficiency of the system is minimized by the fact that at least part of the compressor end air is used to recool the cooling air.