Abstract: An industrial gas turbine engine (10), including: a can annular combustion assembly (80), having a plurality of discrete flow ducts configured to receive combustion gas from respective combustors (82) and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for delivery directly onto the first row (56) of turbine blades (62); and a compressor diffuser (32) having a redirecting surface (130, 140) configured to receive an axial flow of compressed air and redirect the axial flow of compressed air radially outward.

Abstract: An industrial gas turbine engine (10) rated for at least 75 MW maximum output, including: a can annular combustion assembly (80); and a single rotor shaft (114); wherein a combustion section length (112) between a trailing edge (28) of a last row of compressor airfoils (20, 22) and a leading edge (54) of first row of turbine blades (56) is less than 20% of an engine length (154) between a leading edge (26) of a first row of compressor airfoils and a trailing edge (66) of a last row of turbine airfoils (60, 62).

Abstract: A system and method for reducing combustion dynamics includes first and second combustors arranged about an axis, and each combustor includes a cap assembly that extends radially across at least a portion of the combustor and a combustion chamber downstream from the cap assembly. Each cap assembly includes a plurality of tubes that extend axially through the cap assembly to provide fluid communication through the cap assembly to the combustion chamber and a fuel injector that extends through each tube to provide fluid communication into each tube. Each cap assembly has an axial length, and the axial length of the cap assembly in the first combustor is different than the axial length of the cap assembly in the second combustor.

Abstract: A support assembly for a turbine system is disclosed. The support assembly includes a transition duct extending between a fuel nozzle and a turbine section. The transition duct has an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis. The support assembly further includes an annular contact member extending from the transition duct. The contact member is configured to allow movement of the transition duct about at least one axis.

Abstract: The present application and the resultant patent provide a method of improving durability of a first stage bucket of a turbine of a gas turbine engine. The method may include the steps of generating a first combustion flow in a first can combustor and a second combustion flow in a second can combustor, wherein the first can combustor and the second can combustor meet at a joint comprising a flow disruption surface; passing the first combustion flow and the second combustion flow over the flow disruption surface and to a mixing region; substantially mixing the first combustion flow and the second combustion flow in the mixing region to form a mixed combustion flow; and passing the mixed combustion flow to a first stage bucket of a turbine.

Abstract: A gas turbine combustor includes a combustor liner enclosing a combustion chamber; at least one fuel nozzle arranged to provide fuel to the combustion chamber; a flow sleeve surrounding the combustor liner forming a passage radially between the combustor liner and the flow sleeve for supplying air to the combustion chamber, the flow sleeve configured to permit air to flow substantially axially into the passage via a substantially annular flow sleeve inlet. A downstream end of the flow sleeve is formed to include an annular manifold provided with plural outlets about a circumference of the downstream end of the flow sleeve and adapted to supply supplemental air from an external variable air source substantially radially into the passage to thereby maintain axial air flow boundary layer attachment at the flow sleeve inlet.

Abstract: The present invention provides methods and system for power generation using a high efficiency combustor in combination with a CO2 circulating fluid. The methods and systems advantageously can make use of a low pressure ratio power turbine and an economizer heat exchanger in specific embodiments. Additional low grade heat from an external source can be used to provide part of an amount of heat needed for heating the recycle CO2 circulating fluid. Fuel derived CO2 can be captured and delivered at pipeline pressure. Other impurities can be captured.

Abstract: A paint drying booth includes a plurality of walls, a ceiling, and an access door for egress and ingress. A turbine generates heated compressed air which is delivered to the interior of the booth via a duct system. The duct is in fluid communication with an air knife having a slit which delivers a stream or sheet of air into the interior of the booth. More than one air knife can be provided, and the orientation of the plurality of air knives are coordinated to provide a flow of heated turbulent air within the booth to accelerate the drying of painted items within the booth. The turbine comprises a multi-stage compressor including a plurality of rotors and stators for compressing air passing therethrough. Optionally a pneumatic spray gun can be provided for painting the items within the booth as well.

Abstract: A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a trailing edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Abstract: A method is disclosed for controlling gas turbine operation in response to lean blowout of a combustion can. The gas turbine comprises a pair of combustion cans. The method includes sensing that a first combustion can is extinguished during a full load operation of the gas turbine, adjusting a fuel ratio between the fuel nozzles in each can, delivering a richer fuel mixture to the fuel nozzles nearest to the cross-fire tubes, generating a cross-fire from the second combustion can to the first combustion can, detecting a recovery of the turbine load, and adjusting the fuel ratio to the normal balanced fuel distribution between the fuel nozzles in each can.

Abstract: A gas turbine engine (10), including: a turbine having radial inflow impellor blades (38); and an array of advanced transition combustor assemblies arranged circumferentially about the radial inflow impellor blades (38) and having inner surfaces (34) that are adjacent to combustion gases (40). The inner surfaces (34) of the array are configured to accelerate and orient, for delivery directly onto the radial inflow impellor blades (38), a plurality of discrete flows of the combustion gases (40). The array inner surfaces (34) define respective combustion gas flow axes (20). Each combustion gas flow axis (20) is straight from a point of ignition until no longer bound by the array inner surfaces (34), and each combustion gas flow axis (20) intersects a unique location on a circumference defined by a sweep of the radial inflow impellor blades (38).

Abstract: A combustor for a gas turbine engine comprises an inner annular liner and an outer annular liner. A first and a second combustion stages are defined between the liners, each said combustion stage having a plurality of fuel injection bores distributed in a liner wall defining the respective stage. Valves at the fuel injection bores of one of the combustion stages, the valves each defining an air passage from an exterior to an interior of the combustion stage, the valves each having an actuatable member for adjusting a size of a respective air passage for air staging the combustor.

Abstract: A fuel delivery system for a gas turbine engine includes one or more subsets of combustors, and at least two fuel manifolds including a fuel manifold coupled to each combustor subset and configured to deliver fuel to only the subset of combustors during a predetermined gas turbine engine operating mode. A gas turbine engine assembly including a fuel delivery system and a method of operating a gas turbine engine is also described herein.

Abstract: In order to reduce a manufacturing cost and a running cost, casings each housing a combustion oscillation detecting device are attached in an alternating manner to cross-flame tubes by which a plurality of combustors arranged annularly in a combustor casing are connected to each other.

Abstract: A section of a gas turbine engine including a radial spoke is provided. The spoke includes an aerodynamic shape with a leading side and a trailing side and, extending from the leading side to the trailing side, a first side and a second side, opposite the first side. The spoke also includes at least one flow guiding element arranged on at least the first side.

Abstract: A gas turbine combustor (26) disposed in a combustion air plenum (65) and a transition piece (28) for the combustor disposed in a separate cooling air plenum (58, 67). The combustion air plenum may receive combustion air (50) from a high-pressure compressor stage (22A). The cooling air plenum may receive cooling air (52) from an intermediate-pressure compressor stage (22B) at lower temperature and pressure than the combustion air. This cools the transition piece using less air than prior systems, thus making the gas turbine engine (20) more efficient and less expensive, because less expensive materials are needed and/or higher combustion temperatures are allowed. The cooling air may exit the cooling air plenum through holes (62) in a downstream portion (61) of the transition piece. An outer wall (72) on the transition piece may provide forced convection along the transition piece.

Abstract: A can-annular gas turbine engine including a casing defining an annulus for receiving air flow from a compressor. A combustor basket is supported within the casing, and a swirler assembly is located within the basket and includes an air inlet and an outlet located in a combustion chamber. The swirler assembly is configured to discharge a fuel/air mixture from the outlet downstream into the combustion chamber. An acoustic filter is provided comprising a Herschel-Quincke tube having a first open end at the annulus and a second open end adjacent to the swirler assembly within the combustion chamber. Components of acoustic pressure oscillations travel along two paths of different lengths from one end of the Herschel-Quincke tube to the opposite end of the tube where out of phase components of the pressure oscillations interact to attenuate the pressure oscillations.

Abstract: A power generation system is provided and includes a compressor to increase a pressure of inlet ambient air and a turbine downstream from the compressor having a first stage and additional stages downstream from the first stage, the first stage structurally incorporating a combustor in which the compressed ambient air is mixed with fuel and in which the mixture is combusted to generate high temperature exhaust gas from which mechanical energy is derived in the first and additional stages.

Abstract: A transition duct for a gas turbine engine of the kind with can combustors, the transition duct is formed from an inner section and an outer section that form a spiral shaped cooling channel from an inlet end to an outlet end. The spiral channel includes a number of fins that extend into the cooling air passage. an annular plate with an annular arrangement of exit holes closes an aft end of the spiral cooling channel and forms an expansion chamber for the exit holes.

Abstract: A gas turbine includes a combustor, a combustor casing formed in a cylindrical shape to house the combustor therein, and a combustor insertion hole, which is a hole formed on a side periphery of the combustor casing, where a size thereof in a direction orthogonal to a circumferential direction of the combustor casing is formed larger than that in the circumferential direction of the combustor casing, and when the combustor inserted into the hole, a member moved together with the combustor, and the hole are viewed from a moving direction of the combustor, the hole includes an entire external form of the combustor and the member moved together with the combustor.

Abstract: A bullhorn or support for a turbine combustor is provided. The support maintains proper positioning of the transition element relative to the combustor while also allowing for certain movements of the support to alleviate mechanical stresses during turbine operation. The support is constructed from at least two pieces having an interlocking and releasable connection.

Abstract: A splitter is provided for improving a pressure recovery of a compressor discharge casing (CDC) by substantially aligning a fluid vortex about a longitudinal axis of a gas turbine engine. The CDC includes an annular wall that extends generally axially along at least a portion of the CDC. The CDC is configured to house a first combustor and a second combustor. The splitter extends from the annular wall between the first and second combustors. The splitter includes a first surface that is configured to direct fluid towards the first combustor and a second surface that is configured to direct fluid towards the second combustor.

Abstract: An arrangement (10) for conveying combustion gas from a plurality of can annular combustors to a turbine first stage blade section of a gas turbine engine, the arrangement (10) including a plurality of interconnected integrated exit piece (IEP) sections (16) defining an annular chamber (18) oriented concentric to a gas turbine engine longitudinal axis (20) upstream of the turbine first stage blade section. Each respective IEP (16) includes a first flow path section (40) receiving and fully bounding a first flow from a respective can annular combustor along a respective common axis (22) there between, and delivering a partially bounded first flow to a downstream adjacent IEP section (42). Each respective IEP further includes a second flow path section (112) receiving a partially bounded second flow from an upstream adjacent IEP (66) and delivering at least part of the second flow to the turbine first stage blade section.

Abstract: A transition duct used in a gas turbine engine to direct the hot gas flow from the combustor into the turbine section of the gas turbine engine. The transition duct includes a plurality of guide vane integral with the duct. The transition duct includes a circular shaped inlet end for connection to a can combustor and a rectangular and arched shaped outlet end for connection to a first stage turbine section. the guide vanes extend within the flow path between inner and outer projections each having a curved opening in the shape of the airfoil each airfoil includes inner and outer airfoil ends with retainer slots formed between the airfoil ends and the duct projections that form shear pin retainer slots. Shear pin retainers are secured within the slots to secure the guide vanes to the duct in a thermally uncoupled manner to reduce thermal stresses. The guide vanes can be made from a single crystal material for higher gas flow temperatures.

Abstract: A combustion duct assembly has a transition duct and a combustion liner having a hula seal at a downstream end that is forced within an inner wall of the transition duct. The combustion liner is held within the transition duct by the hula seal, but allowed to move relative to the transition duct. The combustion liner is formed with heat transfer columns adjacent the downstream end, and radially inwardly of the hula seal.

Abstract: An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.

Abstract: A crossfire tube assembly is configured for connecting adjacent combustion cans in a gas turbine, and includes a first tube segment having a first end and an opposite female end. A second tube segment has a first end and an opposite male end fitted concentrically within the female end with an overlap region between the female and male ends. Each of the first ends of the tube segments is configured for securing to a liner of a respective combustion can. Oppositely oriented first and second impingement sleeves extend from the female end of the first tube segment to the respective first ends of the tube segments. Combustion cooling air is directed through metering holes in the impingement sleeves and flows axially along concentric cavities defined between the impingement sleeves and the first and second tube segments. The combustion cooling air vents from the cavities into an axial combustion air flow stream between the respective combustion can liners and sleeves.

Abstract: A combustor includes a combustion chamber and an interior wall circumferentially surrounding at least a portion of the combustion chamber and defining an exterior surface. A plurality of turbulators are on the exterior surface. The combustor further includes means for preferentially directing fluid flow across a predetermined position of the turbulators. A method for cooling a combustion chamber includes locating a plurality of turbulators to an exterior surface of the combustion chamber and preferentially directing fluid flow across a predetermined position of the plurality of turbulators.

Abstract: A turbomachine including an annular combustion chamber, the combustion chamber presenting a connection flange at its downstream end for connection to an outer casing. The connection flange bears axially against the outer casing and is blocked in an axial direction by the upstream end of an inner casing of a high pressure turbine.

Abstract: A gas turbine transition duct apparatus is provided comprising first and second turbine transition ducts and a strip seal. The strip seal may comprise a sealing element and a spring structure.

Abstract: An engine assembly includes a combustor having a combustion chamber in which an air and fuel mixture is combusted to produce combustion gases. The engine assembly further includes a transition scroll coupled to the combustor for receiving the combustion gases. The transition scroll includes an interior surface, an exterior surface, and effusion cooling holes for providing cooling air to the interior surface. The engine assembly further includes a turbine coupled to the transition scroll for receiving and extracting energy from the combustion gases.

Abstract: Methods and systems for delivering fuel in a gas turbine engine are provided. The system includes a plurality of can annular combustors that includes at least a first set of combustors of the plurality of can annular combustors and at least a second set of combustors of the plurality of can annular combustors wherein each set of combustors is supplied by a separately controllable respective fuel delivery system. The method includes supplying fuel at a first fuel schedule to the first set of combustors and supplying fuel at a second fuel schedule to the second set of combustors during a first mode of operation wherein the second fuel schedule is different than the first fuel schedule, and supplying fuel at the second fuel schedule to the first and second sets of combustors during a second mode of operation.

Abstract: A combustor liner (230) for a gas turbine engine combustor (200) comprises an inner wall (232), an outer wall (238), a cooling air flow channel (244) formed there between, and a flow control ring (246). The flow control ring (246) is sealingly attached to the downstream ends of the inner wall (232) and the outer wall (238), and comprises a plurality of holes (250) that, during gas turbine engine operation, may regulate a flow of cooling air that passes through the cooling air flow channel (244). One or more surfaces may be coated with a thermal barrier coating (237) to provide additional protection from thermal damage.

Abstract: A combustor cap is cooperable with an impingement plate in a turbine fuel nozzle and includes a plurality of cap segments independently securable to the impingement plate. The plurality of cap segments are radially and tangentially movable relative to the impingement plate.

Abstract: A method of operating a turbine engine, wherein the turbine engine includes a compressor, a combustor, a turbine, a plurality of successive axially stacked stages that include a row of circumferentially spaced stator blades and circumferentially spaced rotor blades, and a plurality of circumferentially spaced injection ports disposed upstream of a first row of stator blades in the turbine; the injection ports comprising a port through which cooling air is injected into the hot-gas path of the turbine, the method comprising: configuring the stator blades in the first row of stator blades such that the circumferential position of a leading edge of one of the stator blades is located within +/?15% pitch of the first row of stator blades of the circumferential location of the injection port midpoint of at least a plurality of the injection ports.

Abstract: A gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

Abstract: The present invention improves the reliability of a combusting system including multiple combustors and improves an environmental performance thereof. The invention includes a device that delays the start time of fuel injection from a fuel nozzle of a combustor including a ignitor from the start time of fuel injection from a fuel nozzle of the combustor not including the ignitor.

Abstract: Disclosed is a combustor seal including a seal support locatable at a first combustor component and having a plurality of through impingement holes. A wave-shaped seal located at the seal support and defining at least one seal cavity between the wave-shaped seal and the seal support. A peak of the wave-shaped seal is locatable at a second combustor component. The wave-shaped seal includes at least one through passageway located upstream of the peak capable of flowing cooling fluid therethrough into the at least one seal cavity and through the plurality of impingement holes thereby cooling the first combustor component. Further disclosed is a combustor including a combustor seal and a method for cooling a first combustor component.

Abstract: A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have an internal passage extending between an inlet to an outlet. An axis of the transition duct body may be generally linear such that gases expelled from the transition duct body flow in a proper direction into the downstream turbine blades. The linear transition duct may include an outlet with exhaust mouths that are configured such that sides of the transition duct are coplanar with adjacent transition ducts, thereby eliminating destructive turbulence between adjacent, linear transition ducts.

Abstract: A diffuser for a gas turbine system having a longitudinal axis is disclosed. The diffuser includes a plurality of diffuser ducts. Each of the plurality of diffuser ducts is disposed annularly about the longitudinal axis and has an inlet, an outlet, and a passage extending between the inlet and the outlet. The outlet of each of the plurality of diffuser ducts is tangentially offset from the inlet of the respective diffuser duct. Each of the plurality of diffuser ducts is configured to flow a gas flow therethrough, reducing the gas flow velocity.

Abstract: A combustion section for a gas turbine includes a casing defining a chamber, a plurality of combustor cans disposed in the casing and oriented in an annular pattern, and a plurality of transition pieces each coupled with one of the combustor cans. The transition pieces direct products of combustion from the combustor cans into contact with rotating buckets of the gas turbine. Each of the transition pieces is angled in two planes to effect turning of the products of combustion and to shorten the gas turbine.

Abstract: A gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

Abstract: A combustor assembly includes a convergent segment followed by a divergent segment to advantageously improve combustion. The combustor assembly includes a first segment beginning at a forward end that transitions to a second segment past a transition segment in a direction along a combustor axis toward an aft end. The reduction in cross-sectional area within the first segment provides desirable fuel and air mixing properties. The convergent first segment in combination with the divergent second segment decreases residence time of fuel-air mixture within the combustor chamber that decreases production of undesirable emissions from the combustor assembly.

Abstract: A transition duct for conveying hot combustion gas from a combustor to a turbine in a gas turbine engine. The transition duct includes a panel including a middle subpanel, an inner subpanel spaced from an inner side of the middle subpanel to form an inner plenum, and an outer subpanel spaced from an outer side of the middle subpanel to form an outer plenum. The outer subpanel includes a plurality of outer diffusion holes to meter cooling air into the outer plenum. The middle subpanel includes a plurality of effusion holes to allow cooling air to flow from the outer plenum to the inner plenum. The inner subpanel includes a plurality of film holes for passing a flow of cooling air from the inner plenum through the film holes into an axial gas flow path adjacent to the inner side of the inner subpanel.

Abstract: A cover assembly disposed about a rotor in a gas turbine engine. The cover assembly comprises a first cover, a second cover, and securing structure. The first cover is disposed about the rotor and comprises a forward end and an opposed aft end. The first cover is associated with a case mounting structure that is fixed to an engine casing. The second cover is disposed about the rotor and comprises a forward end and an opposed aft end. At least a portion of the first cover is disposed radially outwardly from the second cover. The securing structure couples the first cover to the second cover and permits relative radial movement between the first and second covers.

Abstract: A gas turbine combustion system for a turbine engine is disclosed. The gas turbine combustion system may include a plurality of combustor baskets positioned circumferentially about a longitudinal axis of the turbine engine, at least one firing nozzle extending into each of the plurality of combustor baskets and a fuel delivery system in communication with each of the at least one firing nozzles and in communication with at least one fuel source. Each combustor basket may include a fuel flow control device of the fuel delivery system inline with each of the at least one firing nozzles in that combustor basket such that fuel flow to some of the combustor baskets can be shut off while fuel flow to firing nozzles in other combustor baskets remains unobstructed, thereby permitting some combustor baskets to remain non-fired while other combustor baskets are firing to reduce CO emissions.

Abstract: In exemplary embodiments, a gas turbine system is provided. The gas turbine system can include a compressor configured to compress air and combustor cans in flow communication with the compressor, the combustor cans being configured to receive compressed air from the compressor and to combust a fuel stream. The gas turbine system can also include a multi-circuit manifold coupled to the combustor cans and configured to provide a split fuel stream from the fuel stream to the combustor cans.

Abstract: A system comprises a gas turbine combustor having a plurality of combustor cans, crossfire tubes for connecting combustor cans, and a tubular connection system connecting the combustor cans to control combustion dynamics. The tubular connection system comprises tubes for connecting at least a pair of the combustor cans.

Abstract: Embodiments of a transition (400) of the present invention comprise a cooling channel (20, 30, 410) defined in part by an outer wall (23) and an inner wall (24). The cooling channel (20, 30, 410) also comprises lateral side walls (33, 34). Two or more subdomains (AA, BB, A, B1, B2, C1, C2, D) of impingement jets (25, 35) are provided through the outer wall (23), and one or more metering outlets (26, 36, 37, 38) is provided through the inner wall (24), all communicating with a respective channel (20, 30, 410). The impingement jets of each respective subdomain are designed to supply cooling fluid to one of the one or more metering outlets.

Abstract: A method for assembling a gas turbine combustor system is provided. The method includes providing a combustion liner including a center axis, an outer wall, a first end, and a second end. The outer wall is orientated substantially parallel to the center axis. The method also includes coupling a transition piece to the liner second end. The transition piece includes an outer wall. The method further includes coupling a plurality of lean-direct injectors along at least one of the liner outer wall and the transition piece outer wall such that the injectors are spaced axially apart along the wall.