Disclosed is a cooled airfoil having a hub end and tip, an airfoil height being defined between the hub end and the tip. The airfoil has a leading edge, trailing edge, suction side and pressure side. The airfoil has a first airfoil height section adjacent the hub end and extending towards the tip, wherein, in a meridional view, the leading edge and trailing edge are straight along the first airfoil height section. The airfoil has a second airfoil height section adjacent the tip and extending towards the hub end, wherein, in a meridional view, the airfoil is concavely shaped at the leading edge and is convexly shaped at the trailing edge along the second airfoil height section. At least one cooling channel has a length principally extending along the airfoil height, extends straight in a first cooling channel length section, and is bent in a second cooling channel length section.
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
B22D 25/02 - Special casting characterised by the nature of the product by its peculiarity of shapeSpecial casting characterised by the nature of the product of works of art
3.
Method for manufacturing a burner assembly for a gas turbine combustor and burner assembly for a gas turbine combustor
A method for manufacturing a burner assembly for a gas turbine combustor, having a pilot burner extending along a longitudinal axis, and a premix burner surrounding the pilot burner, wherein the method includes manufacturing at least one portion of the pilot burner by an additive manufacturing technique which includes manufacturing at least one first thermal bridge connecting parts of the pilot burner having a temperature difference not greater than a threshold value.
A heat exchanger is disclosed which includes a pressure vessel with an inlet for air, an outlet for air, and a bundle of pipes housed within the pressure vessel for a thermo-vector fluid. A gas/solid separator is provided within the pressure vessel for separating particles drawn by the air.
F28F 19/01 - Preventing the formation of deposits or corrosion, e.g. by using filters by using means for separating solid materials from heat-exchange fluids, e.g. filters
F02C 7/052 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with dust-separation devices
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F28D 7/16 - Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall the conduits being arranged in parallel spaced relation
F02C 7/141 - Cooling of plants of fluids in the plant of working fluid
5.
Method for operating a supply assembly for supplying fuel gas and inert media to a gas turbine combustor, such supply assembly and a gas turbine comprising such supply assembly
A method for operating a supply assembly configured for supplying fuel gas and an inert purge media to a gas turbine combustor, the method including supplying fuel gas in a fuel gas circuit with an upper flow rate; reducing the fuel gas flow rate in the fuel gas circuit from the upper flow rate to a lower flow rate; stopping the supply of the fuel gas in the fuel gas circuit; and starting the supply of the inert purge media in the inert purge media circuit, wherein the starting is performed before the stopping to have a temporary parallel supply of fuel gas and of inert purge media to a fuel distribution system.
A gas turbine having a compressor section provided with a plurality of compressor blades and vanes for compressing air; a combustion section; a turbine section a rotor having an axis and extending from a compressor section to the turbine section for supporting the compressor and turbine blades, the rotor having a last rotor disk having a downstream portion for supporting a bearing cover and an upstream portion for supporting a last turbine blade; and a cooling duct system configured for supplying cooling air from a downstream end of the turbine section to the last turbine blade passing inside the last rotor disk, and having an axial bore in the downstream portion of the last rotor disk along the rotor axis; and a first plurality of inclined radial bores in the upstream portion of the last rotor disk off the axis.
F01D 5/08 - Heating, heat-insulating, or cooling means
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
The invention refers to a blade of a rotary flow machine including an airfoil having a suction surface and a pressure surface joining each other along a trailing and a leading edge. A radially outward directed airfoil tip and a radially inward directed end joining an inner platform connect the airfoil to a shank at a radial end of the airfoil and providing, at least one shank pocket radially encircled by an axially extending portion of the platform. At least one radially extending rim extends from the trailing edge side of the shank and has an essentially radially orientated first slot for receiving a seal. A mount extends radially inwardly from said shank pocket. The first slot has a first aperture on a shank surface orientated in an axial direction.
A vortex generating device is provided as a generally airfoil-shaped lobed body. The flow deflection varies along a spanwise extent, such that the body exhibits a corrugated geometry in a trailing edge region, and the trailing edge exhibits an undulating shape. The undulating shape includes at least one corner. The trailing edge may include or may consist of straight trailing edge sections. The trailing edge may exhibit a polygonial waveform shape, in particular a trapezoidal or rectangular waveform shape. The vortex generating device may be provided as a fuel discharge device which is suited to discharge a fuel into a vortex flow generated by the vortex generating device. To this extent, a fuel supply plenum may be provided inside the body and at least one fuel discharge opening which is in fluid communication with the fuel supply plenum may be provided on the trailing edge.
Disclosed is a vortex generating device having a body, extending between a leading edge and a trailing edge. The body, in profile cross sections taken across the spanwise direction, exhibits an airfoil-shaped geometry. Each airfoil-shaped profile cross section has a camber line extending from the leading edge to the trailing edge, at least two of the camber lines exhibiting different camber angles, such that the body exhibits at least two different flow deflection angles along the spanwise extent. An imaginary trailing edge diagonal extends straight from a first spanwise end of the trailing edge to a second spanwise end of the trailing edge. When seen from the downstream viewpoint, the trailing edge crosses the imaginary trailing edge diagonal exactly once at one diagonal crossing point.
An additive manufactured article for a gas turbine having a body with two lateral surfaces elongated in a first direction; at least a nested duct housed within the lateral surfaces and elongated in the first direction; a structure so that the nested duct is structurally connected to an attachment within the lateral surfaces, wherein at least the body, the duct and the structure are manufactured by an additive manufacturing process and the structure includes an array of ribs attached to the duct in order to compensate differential elongation along the first direction of the duct with respect to the attachment of the ribs by flexural deformation.
wherein the blade tip is configured to have a cylindrical shape along the axial direction in a hot running condition starting from a conical shape along the axial direction in a cold starting condition.
F01D 11/18 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
F01D 5/20 - Specially-shaped blade tips to seal space between tips and stator
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
12.
Stator heat shield segment for a gas turbine power plant
A stator heat shield segment for a gas turbine includes an inner surface facing a hot gas main flow of the gas turbine; an outer surface opposite to the inner surface and at least partially exposed to cooling air; a first impingement cavity provided with a plurality of first impingement holes fluidly connected with the cooling air supplied outside the outer surface; a feeding cavity isolated from the cooling air supplied outside the outer surface and fluidly connected with first impingement cavity; and a second impingement cavity provided with a plurality of second impingement holes fluidly connected to the feeding cavity.
A combustor device for a gas turbines engines includes first and a second tubular members telescopically fitted in axially sliding manner to one another with interposition of annular centering and sealing which include at least a centering annular shoulder and a sealing ring arranged coaxial to one another. The sealing ring is axially spaced apart from the centering annular shoulder so that an axial distance between the centering annular shoulder and the sealing ring is greater than a maximum axial movement allowed between the first and said second tubular members.
F02C 3/20 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
A dual fuel concentric nozzle such as for a dual fuel injector of a sequential burner of a sequential gas turbine, the dual fuel concentric nozzle having a nozzle downstream end, a liquid fuel duct, a gas fuel duct concentrically surrounding the liquid fuel duct and defining a gas fuel passage between the outer surface of the liquid fuel duct and the inner surface of the gas fuel duct, and a lateral surface concentrically surrounding the gas fuel duct and defining a carrier air passage, a downstream end edge of the gas fuel duct being recessed inside the nozzle downstream end with respect to a downstream end edge of the liquid fuel duct.
A sealing arrangement at an interface between a combustor and a turbine of a gas turbine. The turbine can include deflecting vanes at its inlet, which deflecting vanes are each mounted within the turbine so as to define an inner or outer diameter platform and are in sealing engagement via an inner or outer diameter vane tooth with a seal arranged at the corresponding inner or outer diameter part of the outlet of the combustor. The seal is movable and is pressed on the inner or outer diameter vane tooth by a differential pressure such that the pressure of the mainstream hot gas flow is a lower pressure.
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.
F02C 7/228 - Dividing fuel between various burners
F23R 3/42 - Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
17.
Annular Helmholtz damper for a gas turbine can combustor
An annular Helmholtz damper for a gas turbine can combustor, the annular Helmholtz damper having an axis; an inner wall and an outer wall concentrically arranged with respect to the axis to define an annular damping volume arranged around a can combustor; a front and rear circumferential plates for closing the annular damping volume upstream and downstream; at least one intermediate circumferential plate arranged between the front and the rear plates for dividing the annular damping volume in a main and a secondary volume; and a plurality of intermediate drain holes passing the intermediate circumferential plate and configured for draining collected liquid from the main volume to the secondary volume.
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23M 20/00 - Details of combustion chambers, not otherwise provided for
F23R 3/50 - Combustion chambers comprising an annular flame tube within an annular casing
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
18.
Combustor device for a gas turbine engine and gas turbine engine incorporating said combustor device
Combustor device having a twin-shell tubular casing including: an inner tubular member which extends roughly coaxial a longitudinal axis of the combustor device, delimits the combustion chamber and surrounds the burner; and an outer tubular housing which extends roughly coaxial outside of the inner tubular member. An intermediate supporting structure includes an outer annular supporting member, an inner annular supporting member, and a series of three or more oblong connecting beams angularly staggered about the longitudinal axis of combustor device in cantilever manner to stably connect the inner and outer annular supporting members.
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
F23R 3/60 - Support structuresAttaching or mounting means
F23R 3/50 - Combustion chambers comprising an annular flame tube within an annular casing
F23R 3/42 - Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
F01D 25/26 - Double casingsMeasures against temperature strain in casings
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
F02C 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
A temperature probe and method for determining a temperature in a gas flow are disclosed. The probe includes a probe body. A free flow temperature sensor a free flow temperature of the gas flow and a total temperature sensor measures a total temperature of the gas flow. The method includes measuring a flow temperature in a free gas flow, providing a static gas volume in which essentially all kinetic energy of the flowing gas is recovered and converted into thermal energy, and measuring a total temperature in the static gas volume. An accurate determination of the total temperature of a gas flow, which is representative of a specific total enthalpy, can thereby be achieved while detecting fast and transient temperature changes.
A gas turbine transition duct includes: an inner tubular body, defining a transition channel and having a first upstream end and a first downstream end for coupling to a can combustor and to a turbine section of a gas turbine assembly, respectively; an outer tubular body, arranged around the inner tubular body and having a second upstream end at the first upstream end of the inner tubular body and a second downstream end at the first downstream end of the inner tubular body; wherein a convective cooling channel is defined between the inner tubular body and the outer tubular body, the convective cooling channel having an inlet between the first downstream end and the second downstream end; and wherein the outer tubular body is continuous between the second upstream end and the second downstream end.
F23R 3/42 - Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
A temperature detecting device for a gas turbine power plant is provided with at least one optical probe configured to detect a parameter indicative of a temperature and with at least one capsule configured to define a camera inside which the optical probe is housed.
A gas turbine unit having a combustor having a liner, a turbine, arranged downstream of the liner along a main flow gas direction and including a plurality of first stage vanes, a rotor cover support located inwardly of the vanes, and a sealing arrangement at a combustor to turbine interface, wherein the sealing arrangement includes a first dogbone seal extending between the rotor cover support and an inner downstream end of the liner or between the rotor cover support and a bulkhead located at the inner downstream end of the liner.
F01D 3/04 - Machines or engines with axial-thrust balancing effected by working fluid axial thrust being compensated by thrust-balancing dummy piston or the like
An airfoil for a gas turbine engine, the airfoil comprising a wall having a first surface, a second surface, and a passageway extending through the wall from a first opening in the first surface to a second opening in the second surface, the passageway having one or more sections between the first opening and the second opening, the one or more sections in fluid communication with each other, the plurality of sections comprising a first diffuser section providing a first change in cross-sectional area within the passageway, a second diffuser section providing a second change in cross-sectional area within the passageway, a flow conditioning section, and an edge section having two surfaces set opposite each other across the passageway, the two surfaces extending along the passageway substantially in parallel to one another, the edge section being located adjacent to the second opening.
An assembly of at least two members. One of the members supports the other member, the assembly defining an axial direction, a radial direction, and a circumferential direction, an inner member of the at least two members being received radially inside an outer member of the at least two members, wherein the inner member and the outer member am attached to each other by a support arrangement, the support arrangement including at least one floating support assembly as a displaceable coupling between an inner member support point provided at the inner member and an outer member support point provided at the outer member. A displacement of support points in a radial direction results in an interrelated relative displacement of the support points in an axial direction end vice versa.
A blade includes an airfoil and a root having diverging walls. The diverging walls are made of a ceramic matrix composite material. A reinforcement element is provided between the diverging walls.
A method for determining fatigue lifetime consumption of an engine component, by defining a reference thermal load cycle, the reference thermal load cycle being characterized by a reference load cycle amplitude and a reference load cycle time, and determining a reference load cycle lifetime consumption. The method includes measuring a temperature of the engine component, determining a thermal load cycle based upon the temperature measurement, determining a load cycle amplitude, determining a load cycle time, relating the load cycle time to the reference load cycle time, thereby determining a load cycle time factor, relating the load cycle amplitude to the reference load cycle amplitude, thereby determining a load cycle amplitude factor, combining the load cycle time factor and the load cycle amplitude factor into a combined load cycle factor for determining a load cycle lifetime consumption.
A turboengine as disclosed includes an outer wall structure and an inner wall structure, wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures has a surface, the surfaces being arranged facing each other in the radial direction. At least one guide vane member includes at least one airfoil, a radially inner end and a radially outer end. The inner wall structure and the outer wall structure are jointly provided as a vane carrier unit, wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member extending between the inner wall structure and the outer wall structure.
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
F02C 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
F02C 7/20 - Mounting or supporting of plantAccommodating heat expansion or creep
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
An airfoil for a working fluid path of a turboengine extends along a spanwidth direction from a base to a tip. An aerodynamic body thereof includes a suction side surface, a pressure side surface, a leading edge, a trailing edge and a tip, the tip of the aerodynamic body having a tip cross section and a cross-sectional contour circumscribing the tip cross section. A rim extends to the tip of the airfoil and follows the cross-sectional contour on the pressure side, the suction side and extends over the leading edge of the airfoil, the rim delimiting a tip cavity which is open at the tip of the airfoil. The rim is further open at the trailing edge of the airfoil such that the tip cavity is open at the trailing edge of the airfoil.
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.
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F02C 9/32 - Control of fuel supply characterised by throttling of fuel
F02C 9/34 - Joint control of separate flows to main and auxiliary burners
A cooled wall of a turbine component includes a first layer of channels for a coolant arranged along a side of the wall facing to a flow of hot gas, the first layer of channels having a serpentine shape, each channel of the first layer having an inlet and an outlet; a second layer of channels for the coolant disposed further from the flow of hot gas than the first layer, each channel of the second layer having an inlet and an outlet, the outlet of each of the channels of the first layer being in fluid communication with corresponding inlet of associated channel of the second layer creating a bend for changing a direction of the coolant leaving the channel of the first layer when entering the channel of the second layer.
A stator heat shield for a gas turbine having a hot gas flow path, is disclosed. The stator heat shield includes a first surface configured to face the hot gas flow path of the gas turbine; a second surface opposite to the first surface; cooling channels for directing cooling fluid from the second surface towards the first surface; and cavities arranged at the first surface for receiving the cooling fluid from at least a part of the cooling channels; wherein at least a part of the cavities each have at least two corresponding cooling channels open thereto, the at least two corresponding cooling channels being inclined towards each other. In use, a vortex is created in the cavity.
A method is disclosed for repairing a precision casted turbine blade having a squealer and at least one letterbox which was closed by brazing during new-part manufacturing, wherein the letterbox area was damaged during operation in the first life cycle of the blade.
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.
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F02C 9/32 - Control of fuel supply characterised by throttling of fuel
F02C 9/34 - Joint control of separate flows to main and auxiliary burners
A hula seal as disclosed extends in a circumferential direction, a radial direction and an axial direction relative to a central axis.The hula seal includes a first leaf extending from the first edge region to the second edge region and a second leaf extending from the first edge region to the second edge region. The first leaf is the same distance as the second leaf from the central axis in the radial direction, and is adjacent to and overlaps the second leaf in the circumferential direction, when installed in a gas turbine. The first leaf is attached to the second leaf such that the first leaf can move relative to the second leaf in the circumferential direction when installed in a gas turbine.
A boroscope sheath is disclosed for providing a boroscope with temperature protection during a boroscope inspection of a machine such as a gas turbine or a steam turbine. The boroscope sheath includes an elongate tube having a wall extending from a front end to a back end and around a central bore configured and arranged to allow removable insertion of a boroscope cooling channels extend in the wall. The boroscope sheath can be held in a first position relative to the machine when the machine is in use and moved to a second position relative to machine for inspection. Part of the boroscope sheath remain in the machine during use of the machine.
A fuel nozzle assembly for a gas turbine and a method of reconditioning fuel nozzle assemblies is provided. The method is performed on a fuel nozzle assembly (100) of a gas turbine, and comprises providing a pre-assembled fuel nozzle assembly (100) having a base (102), a body (104) extending from the base (102) to a fuel nozzle tip (106), an inner assembly (110), and an outer assembly (112). The method further comprises removing at least a portion of the fuel nozzle tip (106) and the inner assembly (110), coupling and joining a replacement inner assembly (152) to the base (102), and coupling and joining a replacement fuel nozzle tip (154) to the replacement inner assembly (152) and to the outer assembly (112) to provide a reconditioned fuel nozzle.
A method of reconditioning and fabricating turbine components is provided. In one embodiment, the method is performed on a fuel nozzle assembly of a gas turbine, and comprises providing a pre-assembled fuel nozzle assembly having a base, a body extending from the base to a fuel nozzle tip, an inner assembly, and an outer assembly. The method further comprises removing at least a portion of the fuel nozzle tip and the inner assembly, coupling and joining a replacement inner assembly to the base, and coupling and joining a replacement fuel nozzle tip to the replacement inner assembly and to the outer assembly to provide a reconditioned fuel nozzle.
Methods and systems for determining that a sensor, such as a pressure sensor, that provides feedback on one or more conditions of a gas turbine is deficient are provided. The amplitude of measurements from the sensor may be monitored in different frequency ranges in order to detect certain abnormal conditions of the gas turbine that require attention by the control system in one frequency range, and also, concurrently and/or separately, detect a sensor deficiency in another frequency range prior to actual failure of the sensor, at which time the failure may otherwise be noticeable in the first frequency range. This permits better detection of deficient sensors during operation of the gas turbine.
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for
G01L 27/00 - Testing or calibrating of apparatus for measuring fluid pressure
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
F01D 21/14 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to other specific conditions
39.
Method for controlling the operation of a gas turbine with an averaged turbine outlet temperature
A method is disclosed for operating a gas turbine having a compressor, a combustor, a turbine downstream of the combustor, and a total number of turbine outlet temperature measurements. The method includes locally measuring the turbine outlet temperature of the turbine with the turbine outlet temperature measurements of the respective turbine, and averaging measured temperatures of the selected turbine outlet temperature measurements to obtain an average turbine outlet temperature. The gas turbine operation is controlled depending on the determined average turbine outlet temperature.
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
F01D 21/12 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to temperature
F01D 17/08 - Arrangement of sensing elements responsive to condition of working fluid, e.g. pressure
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
The present disclosure relates to gas turbines and to a damper assembly for a combustion chamber of a gas turbine. A damper assembly as disclosed herein may be adjusted to different frequencies during operation and/or deactivated for different operation regimes.
F01N 1/02 - Silencing apparatus characterised by method of silencing by using resonance
F23R 3/42 - Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
G10K 11/16 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general
F23M 20/00 - Details of combustion chambers, not otherwise provided for
41.
IMPROVED METHOD OF ELECTRON BEAM WELDING AND ELECTRON BEAM WELDED TURBINE COMPONENT
This disclosure describes an improved method of electron beam ("EB") welding utilizing a collection pocket (118). The method includes providing a first surface (110,146) and a second surface (100,132,140,144), forming a collection pocket (118) in at least one of the first surface (110,146) and the second surface (100,132,140,144), coupling the first surface to the second surface at a joining location (136,138), and EB welding the first surface (110,146) and the second surface (100,132,140,144) to each other at the joining location (136,138). The collection pocket (118) captures and contains excess weld material to prevent the excess material from escaping the joining location (136,138), and also reduces an amount of wall thickness required for EB welding. A method of reconditioning gas turbine components is also disclosed.
The present disclosure generally relates to a guide vane for a gas turbine, and provides for example an innovative guide vane with improved flexibility leading to a reduction of stresses at the interface between the vane platform and the vane carrier. Exemplary embodiments provide only circumferential line contact or point contact between the guide vane and the guide vane carrier.
A gas turbine rotor cover includes a body, the body having an inner side facing towards a central axis and an outer side facing away from the central axis, wherein the gas turbine rotor cover is configured and arranged to extend in an axial direction and a circumferential direction relative to the central axis. An inner layer of insulation is attached to the inner side of the body and extends along at least part of the length of the body in the axial direction and/or an outer layer of insulation is attached to the outer side of the body, and extends along at least part of the length of the body in the axial direction. A gas turbine containing the gas turbine rotor cover is also disclosed.
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F01D 11/18 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
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.
A gas turbine membrane seal is disclosed. The gas turbine membrane seal includes a membrane, the membrane configured and arranged to extend from a first gas turbine component to a second gas turbine component and to separate two volumes, and an anti-fretting part configured and arranged to be attached to the first gas turbine component. A face of the anti-fretting part is adjacent to the membrane, and the face of the anti-fretting part is convex. Further embodiments of the gas turbine membrane seal are also described, along with a gas turbine having the gas turbine membrane seal and a retrofitting method.
A gas turbine having a compressor, a combustor downstream from the compressor in a gas flow direction, and a turbine downstream from the combustor in the gas flow direction is described herein. The turbine includes a rotating part and a stationary part arranged around the rotating part. A gap between the rotating part and the stationary part, extends in a substantially radial direction relative to the rotation axis of the rotating part. A cooling fluid flows from the compressor to the gap, wherein at least a part of the cooling path extends in the stationary part, and wherein a pre-swirl nozzle is arranged adjacent to the gap and within the cooling path in the stationary part.
A first stage vane arrangement having an array of first stage vanes and an array of frame segments and method for cooling frame segments of the first vane arrangement of a gas turbine are disclosed. The frame segments are designed for axially receiving aft ends of a combustor transition pieces. The first stage vanes include extended vanes, each vane having a leading section, a trailing edge, and an airfoil extending between an outer platform and an inner platform. The frame segments having an I-beam with an upper horizontal element, a lower horizontal element, and a vertical web. The vertical web having a downstream face facing towards a first stage of a turbine when installed in a gas turbine. The downstream face of the vertical web of at least one of the frame segments overlaps, at least partially, the leading section of at least one of the extended vanes.
A locking pin arrangement with a basic element for receiving an insert and a fall-proof locking pin for fixing the insert in the basic element. The locking pin is at least partially inserted into a borehole of the basic element and is pushed towards an opening of the borehole by a spring. The locking pin has a guiding surface for axial guidance in the borehole, an outer thread and a constriction positioned between the outer thread and the guiding surface, wherein the diameter of the outer thread is smaller than the diameter of the guiding surface and the diameter of the locking pin at the constriction is smaller than the diameter of the outer thread. Further, the arrangement comprises an inner thread going from the wall of the borehole into the space formed by the constriction.
The present invention generally relates to a sequential combustor for a gas turbine having second and/or subsequent stages of a re-heat, sequential or axially-staged combustion system. A variation in Mach number along the flow path can be used to control static temperature variation, which in turn influences the progress of auto-ignition reactions that eventually lead to the onset of combustion.
F23R 3/18 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
F23R 3/34 - Feeding into different combustion zones
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
F23R 3/50 - Combustion chambers comprising an annular flame tube within an annular casing
F23R 3/12 - Air inlet arrangements for primary air inducing a vortex
F02C 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
The present invention relates to a gas turbine implemented for example at the interface between the combustor and the vane platform. An efficiency of a cooling film associated to the vane platform can be increased, hence reducing the quantity of the air needed.
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
The application describes a method of operating a gas turbine during a cool-down phase. The gas turbine provides a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor, with the turbine providing a turbine vane carrier. The method includes feeding a flow of cooling air from the compressor to the turbine vane carrier, measuring a temperature of the flow of cooling air and measuring a temperature of the turbine vane carrier. In the method, the flow of cooling air is fed at a first flow rate when the temperature of the turbine vane carrier is lower than the temperature of the cooling air, and the flow of cooling air is fed at a second flow rate when the temperature of the turbine vane carrier is higher than the temperature of the cooling air, wherein the first flow rate is higher than the second flow rate.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 21/12 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to temperature
A gas turbine blade includes a blade root and a blade aerofoil, a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip, a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes, and a pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate. The blade root impingement plate can direct the cooling fluid from the blade root to the pipe.
A method for repairing a blade tip of a turbine blade based on welding below and above a designated depth recommended for repair of turbine blades. A damaged portion of the turbine blade is inspected to identify a standard cut portion and an angled cut portion. The standard cut portion is damaged above a standard cut line and the angled cut portion is damaged below the standard cut line. The damaged portion of the turbine blade is removed. The standard cut portion is removed using a first removal process and the angled cut portion is removed using a second removal process. The angled cut portion is built up with a first weld repair process. The angled cut portion is built up to the standard cut portion. The standard cut portion and the angled cut portion are built up with a second weld repair process.
B23K 26/34 - Laser welding for purposes other than joining
B24B 19/14 - Single purpose machines or devices for particular grinding operations not covered by any other main group for grinding turbine blades, propeller blades or the like
F01D 5/18 - Hollow bladesHeating, heat-insulating, or cooling means on blades
B23K 101/00 - Articles made by soldering, welding or cutting
54.
Method for cooling a gas turbine and gas turbine for conducting said method
A method is disclosed for cooling a gas turbine having a turbine, wherein a rotor, which rotates about a machine axis, carries a plurality of rotating blades, which are mounted by blade roots and extend with their airfoils into a hot gas path of the gas turbine. The rotor is concentrically surrounded by a turbine vane carrier carrying a plurality of stationary vanes, whereby the rotating blades and the stationary vanes are arranged in alternating rows in axial direction. An extended lifetime with external cooling is achieved by providing first and second cooling systems for the turbine.
This application describes a mixing system for a gas turbine combustor arrangement, the mixing system including a lobed mixer and a wall, enclosing a fluid flow path, wherein the lobed mixer is arranged in the wall, between a first part of the wall and a second part of the wall, and wherein the first part of the wall and the second part of the wall are spaced apart in the direction of a lobed mixer axis. Details of the lobed mixer and a method of mixing two flows in a mixing system are also described.
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/06 - Arrangement of apertures along the flame tube
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F02C 7/04 - Air intakes for gas-turbine plants or jet-propulsion plants
A blade includes a leading edge, a trailing edge, a pressure surface, a suction surface, a root end, a tip end, a tip shroud attached to the tip end, the tip shroud comprising a platform and a fin, wherein the fin having a leading edge side facing towards the leading edge of the blade, a trailing edge side facing towards the trailing edge of the blade, a back end and a front end, the leading edge side and the trailing edge side extending between the back end and the front end, the fin extending across the tip end of the blade at an angle to the chord of the blade at the tip end of the blade. A first platform portion extends from the leading edge side of the fin to the suction surface. A second platform portion extends from the trailing edge side of the fin to the pressure surface at the tip end of the blade.
A method is disclosed of disassembling a gas turbine which includes a combustor, a liner, a turbine vane carrier (TVC) and an outer housing, the method including disengaging the liner from the TVC and removing the TVC from the gas turbine. The liner is disengaged from the TVC before the TVC is removed, the liner remains inside the outer housing and the combustor remains in the outer housing.
The invention concerns a turbine for a gas turbine comprising a blade, a vane and an abradable lip attached to the blade or to the vane, wherein the blade and the vane are separated by a gap and the abradable lip extends part of the distance across the gap. Embodiments include the addition of an abrasive layer attached on the other side of the gap from the abradable lip. A method of manufacturing is also described.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 11/02 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
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.
The invention proposes a vortex generating arrangement, especially for a pre-mixing burner of a gas turbine, having an air conducting channel of predetermined height extending between two essentially parallel channel walls and having a predetermined direction of air flow, and having a plurality of vortex generators arranged in the channel. An improved mixing is achieved by the vortex generators each having the shape of a triangular plate, which is arranged essentially perpendicular to the channel walls and oriented relative to the direction of air flow with a predetermined off-axis angle ≠0°, and a first side of the triangular plate being oriented perpendicular to the channel walls.
A fuel injector device includes a body having a leading edge and a trailing edge and defining a streamwise direction from the leading edge to the trailing edge. The fuel injector device body includes a first surface and a second surface opposite the first surface, each surface extending between and including the leading edge and the trailing edge, and the surfaces conjoining each other at the leading edge and the trailing edge. The trailing edge, when seen in the streamwise direction, undulates along a trailing edge mean line and, along its extent, deviates in opposite directions from the mean line and includes at least one inflection point along its extent.
F23D 14/24 - Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other at least one of the fluids being submitted to a swirling motion
Disclosed is a fuel injector device having a body with a leading edge and a trailing edge and defining a streamwise direction from the leading edge to the trailing edge, the fuel injector device body having a first wall and a second wall opposite the first wall, each wall extending between and including the leading edge and the trailing edge and the walls conjoining each other at the leading edge and the trailing edge, each wall having a streamwise extent and a crosswise extent, the walls further enclosing an internal space, at least one fluid plenum being provided within the internal space, the fluid plenum at least at one of an upstream end and/or a downstream end being delimited by an internal wall structure, wherein at least one surface of the wall structure is an inclined surface which forms an angle with the streamwise direction which is smaller than or equal to a maximum angle, wherein the maximum angle is 60°.
The present invention discloses a novel apparatus and way for directing a supply of compressed air into a fuel nozzle assembly for mixing with a fuel source. The apparatus comprises a fuel nozzle assembly having a plurality of coaxial tubes and radially-extending swirler vanes for directing a supply of fuel to a mixing tube. Compressed air is directed to flow in a primarily axial direction by passing through a hemispherically-shaped dome portion at an air inlet region of the fuel nozzle assembly. The hemispherically-shaped dome includes a plurality of openings for directing air into the fuel nozzle assembly in a direction having a radial and axial component.
The present invention relates to dual fuel delivery system for a gas turbine. A dual fuel delivery system for a gas turbine includes: a main fuel line having a main fuel oil conduit and a main fuel gas conduit, wherein the main fuel gas conduit encloses, at least partially, the main fuel oil conduit; and a fuel ring connected to the main fuel line, the fuel ring having a fuel gas ring connected to the main fuel gas conduit and a fuel oil ring connected to the main fuel oil conduit, wherein the fuel gas ring encloses, at least partially, the fuel oil ring.
The present invention discloses a novel apparatus and way for directing a supply of compressed air into a fuel nozzle assembly for mixing with a fuel source. The apparatus comprises a fuel nozzle assembly having a plurality of coaxial tubes and radially-extending swirler vanes for directing a supply of fuel to a mixing tube. Compressed air is directed to flow in a primarily axial direction by passing through a hemispherically-shaped dome portion at an air inlet region of the fuel nozzle assembly. The hemispherically-shaped dome includes a plurality of openings for directing air into the fuel nozzle assembly in a direction having a radial and axial component.
An apparatus for providing improved cooling to a combustion liner (308) of a gas turbine combustor is provided. A plurality of flow deflectors (302) is secured to a flow sleeve (304) in order to improve the flow of impingement air from a flow sleeve to the combustion liner outer surface, such that the amount of impingement air being swept away by a cross flow of cooling air is reduced.
The present invention discloses a fuel nozzle assembly and method for axially staging fuel injection. The fuel nozzle assembly comprises a plurality of vanes connected to a core, with an annular ring connectedly surrounding the plurality of vanes. Fuel is directed through the core and then through the vanes and is injected at different axial planes generally perpendicular to an oncoming air stream to obtain a broad spectrum of gas residence time between the point of fuel injection and the flame.
An apparatus for providing improved cooling to a combustion liner of a gas turbine combustor is provided. A plurality of flow deflectors is secured to a flow sleeve in order to improve the flow of impingement air from a flow sleeve to the combustion liner outer surface, such that the amount of impingement air being swept away by a cross flow of cooling air is reduced.
The present invention relates to dual fuel delivery system for a gas turbine. A dual fuel delivery system for a gas turbine includes a main fuel line having a main fuel oil conduit and a main fuel gas conduit, wherein the main fuel gas conduit encloses, at least partially, the main fuel oil conduit; and a first fuel divider having a first fuel oil divider connected to the main fuel oil conduit and a first fuel gas divider connected to the main fuel gas conduit, wherein the first fuel gas divider encloses, at least partially, the first fuel oil divider.
F02M 43/00 - Fuel-injection apparatus operating simultaneously on two or more fuels, or on a liquid fuel and another liquid, e.g. the other liquid being an anti-knock additive
F02C 9/40 - Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
The invention concerns a gas turbine having a picture frame, a first vane, and a sealing arrangement to seal a gap between the picture frame and the first vane, the sealing arrangement including two seals arranged in series between the picture frame and the first vane. In exemplary embodiments, one of the seals is a honeycomb seal, a dogbone seal, a hula seal or a piston seal and the other seal is a honeycomb seal, a dogbone seal, a hula seal or a piston seal. A method of supplying cooling fluid to the gap between the picture frame and the first vane is also disclosed.
The invention concerns a conical hula seal wherein an inner part of the conical hula seal is closer to a conical hula seal longitudinal axis than an outer part of the conical hula seal, and wherein the conical hula seal extends in a circumferential direction relative to the conical hula seal longitudinal axis. A gas turbine which includes the conical hula seal is also disclosed, along with a method of operating a gas turbine which includes a conical hula seal that seals a gap between a first vane and a picture frame, the method including purging the gap with a cooling fluid.
The present invention discloses a fuel nozzle assembly and method for axially staging fuel injection. The fuel nozzle assembly comprises a plurality of vanes connected to a core, with an annular ring connectedly surrounding the plurality of vanes. Fuel is directed through the core and then then through the vanes and is injected at different axial planes generally perpendicular to an oncoming air stream to obtain a broad spectrum of gas residence time between the point of fuel injection and the flame.
An exhaust gas liner for a gas turbine includes an annular inner shell and an annular outer shell, which are arranged concentrically around a machine axis of the gas turbine to define an annular exhaust gas channel in between. The inner shell and/or said outer shell are composed of a plurality of liner segments, which are attached to a support structure. To compensate thermal expansion and achieving resistance against dynamic loads, the liner segments are fixed to the support structure at certain fixation spots, which are distributed over the area of said liner segments, such that said liner segments are clamped to said support structure through a whole engine thermal cycle without hindering thermal expansion.
A combustion chamber includes a duct wall for guiding a hot gas flow in a hot gas flow path during operation. The duct wall is a double-walled construction including an inner face, an outer face, and a wall cavity. A sleeve at least partly encloses the duct wall for guiding a cooling gas in a cooling channel between the sleeve and the duct wall along the outer surface of the duct wall to an exit end, and the cavity opens to the cooling channel. A gas turbine is disclosing as having such a combustion chamber.
A turbine part replacing apparatus for moving a combustor of an industrial gas turbine from and to a mounting port of the gas turbine is disclosed, the apparatus including a mobile frame with a moving mechanism for substantially horizontal translation motion on the ground, a vertically extendable arm attached to and extending from said mobile frame, and a carriage attached to the extendable arm such that the carriage may be moved vertically with respect to the mobile frame. Moreover, the turbine part replacing apparatus is adapted for receiving the combustor in the carriage and for vertically moving the received combustor by the extendable arm from and to the mounting port.
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
B23P 19/04 - Machines for simply fitting together or separating metal parts or objects, or metal and non-metal parts, whether or not involving some deformationTools or devices therefor so far as not provided for in other classes for assembling or disassembling parts
F23R 3/60 - Support structuresAttaching or mounting means
76.
Mounting and dismounting device for a liner of a gas turbine and a related method
A mounting and dismounting device for a liner of a gas turbine includes two inner rails attached at the turbine housing and each one having a first straight portion, two straight outer rails releasably attached at the turbine housing, wherein the adjacent free ends of the inner and outer rails can be positioned in full alignment one to the other. The rails are adapted to support the liner to be moveable in its axial direction. The inner rails include a second straight portion connected to the first straight portion through a curved portion wherein the axis of the second straight portion is parallel to the axis of the combustion gas passageway.
The invention concerns a sequential liner for a gas turbine combustor, having a sequential liner outer wall spaced apart from a sequential liner inner wall to define a sequential liner cooling channel between the sequential liner outer wall and the sequential liner inner wall. The sequential liner outer wall includes a first face, a first adjacent face and a second adjacent face, the first and second adjacent faces each being adjacent to the first face, the first face of the sequential liner outer wall having a first convective cooling hole adjacent to the first adjacent face and a second convective cooling hole adjacent to the second adjacent face, each convective cooling hole being arranged to direct a convective cooling flow into the sequential liner cooling channel adjacent to each adjacent face. The invention also concerns a method of cooling using the sequential liner and a method of retrofitting a gas turbine.
The invention concerns a method for operating a gas turbine arrangement, wherein the gas turbine arrangement can be actively connected to a grid system and includes a separation of compressor and turbine shaft to operate both components individually as unit. A first unit can include at least one turbine and at least one generator and a second unit can include at least one compressor and least one motor. Various switches are situated along power lines and are actively connected to a frequency converter and/or the grid system, wherein the compressed air duct operating downstream of the compressor includes a flap.
A turbine blade having an airfoil and a fir tree root is disclosed. The fir tree root has a lengthwise direction, a crosswise direction extending between two lateral sides of the fir tree root, and a span direction extending from a root base towards an airfoil tip. The fir tree root includes a web interposed between each pair of at least two neighboring channels. For each web, a web-to-channel ratio with each of the two neighboring channels is chosen to be larger than or equal to 0.5 and is smaller than or equal to 0.85 at least at a position where the root width is a minimum load bearing root width in a load bearing section of the fir tree root. In another aspect, overall web-to-channel ratio is defined as a ratio between the sum of all web lengths and the sum of all channel lengths.
A fuel injection device for a gas turbine can include streamlined bodies arranged adjacent to one another in a circumferential direction. Each streamlined body can have a fuel nozzle on its trailing edge. The streamlined bodies can be split into at least a first group of streamlined bodies and a second group of streamlined bodies. All the streamlined bodies in the first group being the same and being different from the streamlined bodies in the second group. The first group of streamlined bodies includes at least two streamlined bodies adjacent to one another. Alternatively, the streamlined bodies are clustered in sets of two adjacent streamlined bodies. As yet another alternative, the streamlined bodies can be arranged such that the fuel injection device has a maximum of four-fold rotational symmetry in the plane perpendicular to the central axis.
A centering arrangement includes a ring-like inner part and a ring-like outer part, whereby the outer part surrounds the inner part in a concentric arrangement and with an interspace established between the outer part and the inner part, and whereby the outer part and the inner part are subject to a differential radial expansion. A centering contact element, procures a centering mechanical contact between the outer part and the inner part at a plurality of circumferentially distributed contact points, and is able to deform in order to compensate for differential radial expansion.
F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
F01D 25/26 - Double casingsMeasures against temperature strain in casings
F16D 1/08 - Couplings for rigidly connecting two coaxial shafts or other movable machine elements for attachment of a member on a shaft or on a shaft-end with clamping hubCouplings for rigidly connecting two coaxial shafts or other movable machine elements for attachment of a member on a shaft or on a shaft-end with hub and longitudinal key
A sequential combustor arrangement and method are disclosed which can include a first burner, a first combustion chamber, a mixer for admixing a dilution gas 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 mixer can include at least one injection opening in the mixer wall for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber. Further, the mixer can include a damper with a damper volume and a neck connecting the damper volume to the mixer, for modulating and damping pressure pulsations inside the mixer.
F02C 3/20 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F23R 3/06 - Arrangement of apertures along the flame tube
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
F23R 3/34 - Feeding into different combustion zones
F23M 20/00 - Details of combustion chambers, not otherwise provided for
A wall for a hot gas channel in a gas turbine is described, the wall having a back side and a front side and an impingement sheet having impingement cooling holes, the wall being for exposure to a hot fluid at the front side, and the wall having an array of pins attached to the back side and extending between the back side and the impingement sheet, the wall additionally having a plurality of ribs attached to the back side, each rib extending between two pins to delineate an array of cells on the back side, and/or at least one compartment wall attached to the back side to delineate compartments on the back side. Embodiments include an impingement sheet with impingement cooling holes and cooling exit holes. A gas turbine including the wall is also described.
A method of manufacturing a hot gas wall for a gas turbine is described. The method is carried out on a hot gas wall having a wall part with a front side and a back side, the wall part being for exposure to a hot fluid on the front side, and the hot gas wall also having a turbulator structure. In an exemplary embodiment, a turbulator structure is attached to the wall by placing a braze foil on the back side of the wall part, placing a turbulator structure on the braze foil, and brazing to attach the turbulator structure to the wall part. In another embodiment, the turbulator structure is attached by passing a current through the turbulator structure part and the wall part to resistance weld the turbulator structure part to the wall part.
It is disclosed an arrangement of a rotor and at least a blade. The blade includes a root, a platform and an airfoil. The rotor includes a seat for the root. The root has side walls which complement side walls of the seat and axial walls between the side walls. A chamber is provided between the root and the rotor. A shank cavity is provided between the root and the platform. A lock plate facing at least an axial wall is connected to the rotor and the blade. The lock plate has at least a slot on a side facing the root.
A combustor arrangement of a gas turbine engine or power plant is disclosed, having at least one combustion chamber, at least one mixer arrangement for admixing air or gas to the hot gas flow leaving the combustion chamber. The mixer arrangement is configured to guide combustion gases in a hot gas flow path extending downstream of the combustion chamber, wherein the mixer includes a plurality of injection pipes pointing inwards from the side walls of the mixer arrangement for admixing air portions to cool at least the hot gas flow leaving combustion chamber. The mixer arrangement is applied to at least one volume of dilution air flowing from a first plenum and at least one volume of cooling air flowing from a second plenum.
A method is disclosed for operating an axially staged mixer arrangement with dilution air or dilution gas injection in connection with a combustion arrangement of a gas turbine engine or turbo machinery, having at least one combustion chamber, at least one mixer arrangement for admixing air or gas portions to the hot gas flow leaving the combustion chamber. The mixer arrangement includes a plurality of injectors for admixing air portions to cool at least the hot gas flow leaving combustion chamber. The spacing between the last injection location of acting injectors and at least one subsequently arranged dilution air injection, inside of the mixer arrangement in the hot gas flow, where there is no node between the injection locations, corresponds to a distance equal or approximating to half of convective wave length.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F23R 3/34 - Feeding into different combustion zones
F23R 3/00 - Continuous combustion chambers using liquid or gaseous fuel
F23R 3/06 - Arrangement of apertures along the flame tube
F23R 3/12 - Air inlet arrangements for primary air inducing a vortex
F23C 6/04 - Combustion apparatus characterised by the combination of two or more combustion chambers in series connection
F23C 9/08 - Combustion apparatus characterised by arrangements for returning combustion products or flue gases to the combustion chamber for reducing temperature in combustion chamber, e.g. for protecting walls of combustion chamber
88.
Mixer for admixing a dilution air to the hot gas flow
The invention refers to a combustor arrangement of a gas turbine engine or power plant, having at least one combustion chamber, at least one mixer for admixing a dilution medium or air to the hot gas flow leaving the combustion chamber. The mixer is configured to guide combustion gases in a hot gas flow path extending downstream of the combustion chamber, wherein the mixer includes a plurality of injection pipes pointing inwards from the side walls of the mixer for admixing the dilution medium or air to cool the hot gas flow leaving combustion chamber. The mixer includes at least one dilution air plenum having at least one pressure-controlled compartment which is directly or indirectly connected to at least one injection pipe.
F02C 3/20 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
F23C 9/08 - Combustion apparatus characterised by arrangements for returning combustion products or flue gases to the combustion chamber for reducing temperature in combustion chamber, e.g. for protecting walls of combustion chamber
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
The present invention generally relates to a vane for a gas turbine, and more in particular it provides an innovative vane with improved flexibility leading to a reduction of stresses at the transition from the vane trailing edge to the vane platform, without interfering into the cooling scheme of such component. The present invention can increase flexibility of the vane platform by introducing on the vane platform a material cutback confined in the proximity of the trailing edge portion of the vane airfoil.
The invention concerns a gas turbine fuel pipe, having a fuel line, the fuel line having a fuel line volume, a fuel line outer wall and an opening in the fuel line outer wall, a damper having a damper volume and a damper outer wall and attached in fluid communication with the fuel line, wherein the damper covers the opening in the fuel line outer wall, and a perforated lining extending across at least part of the opening in the fuel line outer wall.
The present invention relates to dampers for gas turbines and, for example, to a compensation assembly for a damper of a gas turbine for reducing the pulsations occurring in the combustion chamber. The damper can include a resonator cavity with a neck tube in flow communication with the interior of the combustion chamber, wherein the compensation assembly includes a spherical joint associated to the neck tube and configured to allow relative rotation between the combustion chamber and the resonator cavity, and having a bulb portion disposed around the neck tube and a spherical socket configured to internally host the bulb portion, wherein the spherical socket can have a top collar portion and a bottom collar portion connected to each other.
A rotating blade for a gas turbine includes an airfoil extending in a longitudinal direction and having a leading edge and a trailing edge, whereby the airfoil is bordered at its outer end by a tip shroud, whereby the airfoil includes two or more internal passages, which run in longitudinal direction and are separated by solid webs, and whereby a plurality of shroud fins is arranged on top of the tip shroud to improve gas sealing against a corresponding stator heat shield. The stability and life time of the blade can be enhanced by selecting a position of each of the shroud fins to be exclusively above one of the webs and/or a leading edge wall.
A rotating gas turbine blade is disclosed which includes an airfoil with a suction side and a pressure side, the airfoil extending in a radial direction from a blade root to a blade tip. The blade tip includes a tip shroud, the airfoil having internal cooling passages for a cooling medium, which extend through the tip shroud. Outlet ports are provided above a selected internal airfoil cooling passage for the cooling medium to be ejected above the tip shroud in a direction of the blade's pressure side. Dust accumulation is avoided at the tip end of the selected internal cooling passage.
A sequential burner for an axial gas turbine comprises: a burner body, which is designed as an axially extending hot gas channel and further comprises a fuel injection device, which extends into said burner body perpendicular to the axial direction. The manufacturing of the burner body is simplified and the fuel injection is stabilized by designing said fuel injection device as a mechanically stiff component, and fixing said fuel injection device to said burner body in order to keep it aligned with said burner body and to stiffen said burner body against creep.
3 and the abrasive particles are coated with a first particle coating layer disposed on the abrasive particles and an optional second particle coating layer disposed on the first particle coating layer, wherein the matrix material contains or consists of the compound MCrAlY, wherein M is at least one element selected from the group consisting of Ni, Co and Fe. A method for manufacturing such a coated substrate is also disclosed.
B23K 26/00 - Working by laser beam, e.g. welding, cutting or boring
C09D 1/00 - Coating compositions, e.g. paints, varnishes or lacquers, based on inorganic substances
C23C 28/02 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of metallic material
B23K 26/03 - Observing, e.g. monitoring, the workpiece
B23K 26/34 - Laser welding for purposes other than joining
B32B 15/01 - Layered products essentially comprising metal all layers being exclusively metallic
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
C23C 24/10 - Coating starting from inorganic powder by application of heat or pressure and heat with intermediate formation of a liquid phase in the layer
C09D 7/62 - Additives non-macromolecular inorganic modified by treatment with other compounds
C08K 9/02 - Ingredients treated with inorganic substances
A damper for a gas turbine combustion chamber as shown in FIG. 1 includes a damper volume wall and a main neck. The damper volume wall defines a damper volume inside the damper volume wall. The main neck includes a main neck wall defining a main neck volume inside the main neck wall. The main neck is associated with the damper volume for fluid communication between the damper volume and the gas turbine combustion chamber. In addition, the damper includes a gap between the main neck wall and the damper volume wall. The main neck defines a main neck axis. For example, the gap is a second neck, and in further embodiments, multiple damper volumes are provided.
The present disclosure generally relates to a rotor assembly, and in particular relates to an improved rotor heat shield which provides an innovative configuration for securing the same to the rotor assembly. The rotor heat shield element is secured to the rotor assembly in correspondence of the groove in which it is inserted. Embodiments of the present disclosure can allow the removal of current fixation features on heat shields and blades. Furthermore, since the heat shield is no longer connected to a blade but directly to the rotor assembly, there is more freedom in selecting the number of heat shield elements to be provided to form the circumferential heat shield.
The disclosure relates Frame segment for a transition piece-turbine interface having a picture frame receptacle for axially receiving an aft end of a combustor transition piece. The frame segment can include an I-beam with an upper horizontal element, a lower horizontal element, and a vertical web, wherein the upper horizontal element has mounting face for fixation to a vane carrier. The vertical web has a downstream face, facing towards a first stage of a turbine when installed in a gas turbine. The vertical web includes a cooling gas duct for supplying cooling gas to the downstream face of the vertical web.
The disclosure relates to a vane arrangement with a vane carrier, an array of rocking first stage vanes, and an array of frame segments for axially receiving aft ends of a combustor transition pieces. The frame segments comprise an I-beam with an upper horizontal element, a lower horizontal element, a vertical web, a radially outer fixation to the vane carrier, and an arm extending from the lower horizontal element in axial direction below the inner rim segment for supporting the inner platform of the vane and for sealing a gap between the inner platform and the lower horizontal element. Besides the vane arrangement the disclosure relates to a method for assembly of such an arrangement as well as to a gas turbine comprising such a vane arrangement.
A lobe lance is disclosed for a gas turbine combustor which includes a plurality of N (N≥4) lobe fingers, each configured as a streamlined body with two lateral surfaces. A plurality of nozzles for injecting a gaseous and/or liquid fuel mixed with air are provided whereby lobes running between the nozzles are provided for improving the mixing quality and reducing pressure loss in said combustor. The lobes of each lobe finger have one of two opposite orientations with respect to said flow direction, and the lobes of all lobe fingers follow a predetermined pattern of orientation across the lobe fingers at least one pair of neighboring lobe fingers has the same lobe orientation resulting in a grouped lobe arrangement ( . . . LL . . . or . . . RR . . . ) such that at least two of the vortices generated by the lobe shape downstream of the lobe fingers combine.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F23C 7/00 - Combustion apparatus characterised by arrangements for air supply
F23R 3/12 - Air inlet arrangements for primary air inducing a vortex
F23R 3/18 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
F23R 3/20 - Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
F23R 3/16 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
patents
