Ansaldo Energia Switzerland Ag

General Electric Technology Gmbh

General Electric Company

Alstom Technology Limited

Ansaldo Energia IP UK Limited

Ewald Freitag

Adnan Eroglu

Andrea Ciani

Douglas Pennell

Urs Benz

Andre Theuer

John Philip Wood

Peter Stuttaford

Michael Thomas Maurer

Stefano Bernero

Fulvio Magni

Jaan Hellat

Mirko Ruben Bothien

Bettina Paikert

Klaus Doebbeling

Franklin Marie Genin

Robert Marmilic

Andreas Huber

Ennio Pasqualotto

Michael Duesing

Uwe Rüdel

Nicolas Tran

Marcel Rieker

Rohit Madhukar Kulkarni

Martin Zajadatz

Simone Roberto Walter Camponovo

Douglas Anthony Pennell

Rohit Kulkarni

A burner for a combustor includes (a) a swirl generator enclosing a burner interior on an inlet side and including a tangential air inlet relative to a longitudinal center axis; (b) a mixing chamber enclosing the burner interior on an outlet side and defining a burner outlet fluidly connecting the burner interior with a combustion chamber; and (c) a lance arranged coaxially with the longitudinal center axis. The lance introduces fuel through a nozzle tip at or near the burner outlet into the combustion chamber. The nozzle tip includes a cartridge defining a center fuel passage; fuel swirl vanes within the center fuel passage at an outlet end of the nozzle tip; a first tube surrounding the center fuel passage and defining a first fluid passage; a second tube surrounding the first tube and defining a second fluid passage; and air swirl vanes in the second fluid passage.

Pilot nozzle tips for extended lance of combustor burner

A burner for a combustor of a turbomachine includes a pilot nozzle with inline premixing. The pilot nozzle is formed in an aft end of the burner. An air inlet is formed in a forward end of the burner in fluid communication with the pilot nozzle. A mixing channel extends along the axial direction between the air inlet and the pilot nozzle such that the air inlet is in fluid communication with the pilot nozzle via the mixing channel. An annular fuel plenum extends along the circumferential direction. A fuel port is in fluid communication with the annular fuel plenum and the mixing channel, the fuel port includes an outlet configured to inject fuel into the mixing channel such that a shear flow is induced.

Pilot nozzle with inline premixing

The invention relates to a gas turbine assembly which substantially includes at least one compressor, at least one first burner, at least one second burner that is connected downstream of the first burner, and at least one turbine that is connected downstream of the second burner. At least the first and second burner form a component of a tubular or quasi-tubular combustion chamber element in the flow direction of the combustion path of the burners. The combustion chamber element being closed or quasi-closed and extending between the compressor and the turbine. The combustion chamber elements are arranged around the rotor of the gas turbine assembly in the shape of a ring.

Gas turbine assembly and corresponding operating method

The invention refers to a sequential combustor arrangement comprising a first burner, a first combustion chamber, a mixer for admixing a dilution gas via a dilution gas inlet to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The sequential combustor arrangement further includes four cooling zones with a cooling channel. During operation a cooling gas flows through the cooling channels. The disclosure further refers to a method for operating a gas turbine with such a sequential combustor arrangement.

Sequential combustion with dilution gas mixer

A burner for a combustion chamber of a gas turbine with a mixing and injection device, which includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies. Each streamlined body extends in a first transverse direction into the gas-flow channel, and has two lateral surfaces that are arranged essentially parallel to the main-flow direction. The lateral surfaces are joined to one another at their upstream and downstream sides to form leading and trailing edges of the body, respectively. At least one of the streamlined bodies includes a mixing structure and at least one fuel nozzle at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel. At least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor.

Burner for a can combustor

The present invention relation to a burner for a combustion chamber of a gas turbine with a mixing and injection device. The mixing and injection device includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies, each extending in a first transverse direction into the gas-flow channel. Each streamlined body has two lateral surfaces that are arranged essentially parallel to the main-flow direction, the lateral surfaces being joined to one another at their upstream side to form a leading edge of the body and joined at their downstream side to form a trailing edge of the body. Each streamlined body has a cross-section perpendicular to the first transverse direction that is shaped as a streamlined profile. At least one of the streamlined bodies is provided with a mixing structure and with at least one fuel nozzle located at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel, wherein at least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor. The invention also relates to a method of using said burner in a gas turbine.

Burner for a combustion chamber of a gas turbine having a mixing and injection device

A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The combustor arrangement includes acentral lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and into the second burner, wherein the lance body includes a fuel duct for providing fuel for the first burner and/or for the second burner.

Combustor arrangement for a gas turbine

The invention provides a new path of combustion technology for gas turbine operation with multifuel capability, low emissions of NOand CO and high thermal efficiency. Further to the present invention providing a method for operating a combustor for a gas turbine and a combustor for a gas turbine are disclosed. The combustor includes a first combustion chamber with a wide operating range, a subsequent deflection unit for deflecting the hot gas flow of the first combustion chamber at least in circumferential direction and components for injecting and mixing additional air and/or fuel, and a sequential combustion chamber with a short residence time, where the temperature of the hot gases reaches its maximum.

Combustor for a sequential gas turbine having a deflection unit between first and second combustion chambers

The invention refers to a sequential combustor arrangement including 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, and a second burner for admixing a second fuel and a second combustion chamber. To assure good mixing over a wide operating range, the ratio of the pressure loss of the first combustor to the pressure loss of the dilution gas admixer is in the range of 1 to 6. The invention further refers to a gas turbine including such a sequential combustor arrangement as well as method for operating a gas turbine with such a sequential combustor arrangement.

Sequential combustion arrangement with dilution gas

A reheat burner arrangement including a center body, an annular duct with a cross-section area, an intermediate fuel injection plane located along the center body and being actively connected to the cross section area of the annular duct, wherein the center body is located upstream of a combustion chamber, wherein the structure of the reheat burner arrangement is defined by various parameters and the structure of the reheat burner arrangement is defined by various dependencies.

Reheat burner arrangement having an increasing flow path cross-section

The invention refers to burner arrangement for producing hot gases to be expanded in a gas turbine, including a burner inside a plenum, where the burner has means for fuel injection, means for air supply and means for generating an ignitable fuel/air mixture inside the burner, and a combustion chamber following downstream said burner having an outlet being fluidly connected to the gas turbine. The invention is characterized in that the means for air supply includes at least two separate flow passages, and that the one of the two flow passages is fed by a first supply pressure and the other flow passage is fed by a second supply pressure.

Burner arrangement including an air supply with two flow passages

The invention relates a method for mixing a dilution air with a hot main flow in a sequential combustion system of a gas turbine, wherein the gas turbine essentially comprises at least one compressor, a first combustor which is connected downstream to the compressor, and a second combustor. The hot gases of the first combustor are admitted to at least one intermediate turbine or directly or indirectly to at least one second combustor, wherein the hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery device. The method comprises a coaxial injection of first combustor liner cooling air with second combustor liner cooling air, the first combustor liner cooling air having a sufficient excess pressure margin with respect to the second combustor liner cooling air.

Method for mixing a dilution air in a sequential combustion system of a gas turbine

The invention discloses a method for operating a gas turbine with sequential combustion, which gas turbine includes a compressor, a first combustor with a first combustion chamber and first burners, which receives compressed air from the compressor, a second combustor with a second combustion chamber and second burners, which receives hot gas from the first combustor with a predetermined second combustor inlet temperature, and a turbine, which receives hot gas from the second combustor. The CO emission for part-load operation is reduced by reducing the second combustor inlet temperature for base-load operation of the gas turbine, and increasing the second combustor inlet temperature when decreasing the gas turbine load (RL) from base-load to part-load.

Growing community of inventors

Baden, Switzerland

Ewald Freitag