Abstract: An annular combustor includes an annular outer shell that includes a first flange that defines an inner diameter IDOS and an annular inner shell that includes a second flange that defines an outer diameter ODIS. An annular hood includes a radially outer hood flange and a radially inner hood flange. A bulkhead includes a radially outer bulkhead flange that defines an outer diameter ODB and a radially inner bulkhead flange that defines an inner diameter IDB. The first flange is secured at a radially outer joint between the radially outer hood flange and the radially outer bulkhead flange. The second flange is secured at a radially inner joint between the radially inner hood flange and the radially inner bulkhead flange. The IDOS and the ODB define a ratio R1 of IDOS/ODB that is 0.998622-1.001129, and the IDB and the ODIS define a ratio R2 of IDB/ODIS that is 0.998812-1.001388.

Abstract: A mid-turbine frame is incorporated into a turbine section of a gas turbine engine intermediate a high pressure turbine and a low pressure turbine. The high pressure and low pressure turbines rotate in opposite directions. The mid-turbine frame carries a plurality of vanes to redirect the flow downstream of the high pressure turbine as it approaches the low pressure turbine. In another feature, a power density is defined as the thrust divided by the volume of a turbine section, and the power density is of about 1.5 lbf per in3.

Abstract: A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5.

Abstract: The problem to be solved of the present invention is to provide a precipitation hardening martensitic stainless steel having excellent tissue stability, strength, toughness, and corrosion-resistance, requiring no sub-zero treatment, and having excellent productivity; and also a steam turbine long blade using the same. The problem is solved by providing a precipitation hardening martensitic stainless steel containing, by mass, 0.1% or less of C; 0.1% or less of N; 9.0% or more and 14.0% or less of Cr; 9.0% or more and 14.0% or less of Ni; 0.5% or more and 2.5% or less of Mo; 0.5% or less of Si; 1.0% or less of Mn; 0.25% or more and 1.75% or less of Ti; 0.25% or more and 1.75% or less of Al, and the rest is Fe and inevitable impurities; and a steam turbine long blade using the precipitation hardening martensitic stainless steel.

Abstract: An article which includes a structure of a ceramic material that has a composition SiOxMzCy, where Si is silicon, O is oxygen, M is at least one metal and C is carbon and wherein x<2, y>0 and z<1 and x and z are non-zero.

Abstract: A turbomachine comprises a turbine shaft, first and second rotors, first and second propulsion stages, and a magnetic stator. The first rotor is rotationally coupled to the turbine shaft, and coaxially arranged along an axis. The first propulsion stage is rotationally coupled to the first rotor, opposite the turbine shaft. The second rotor is coaxially arranged about the first rotor, and the second propulsion stage is rotationally coupled to second rotor, opposite the turbine shaft and adjacent the first propulsion stage. The magnetic stator is coaxially arranged between the first rotor and the second rotor, forming a magnetic coupling between the and second rotors to drive the second propulsion stage in contra-rotation with respect to the first propulsion stage.

Abstract: An airfoil includes, among other possible things, a main body extending between a leading edge and a trailing edge. Channels are formed into the main body, with a plurality of ribs extending intermediate the channels. A cover skin is attached to the main body. The cover skin is welded to the main body at outer edges. An adhesive is placed between inner surfaces of the cover skin and the main body. The adhesive is deposited inwardly of the outer edges of the cover skin.

Abstract: According to one aspect of the invention, a turbine airfoil includes a first cavity inside the turbine airfoil configured to receive a fluid and a second cavity inside the turbine airfoil. The turbine airfoil also includes a passage inside the turbine airfoil that provides fluid communication between the first and second cavities, wherein the passage includes a curved portion configured to separate particulates from the fluid as the fluid flows through the passage.

Abstract: A combustor system for use in a turbine engine is provided. The turbine engine includes turbine assembly that includes a fluid inlet, a fluid outlet, and a combustion gas path defined therebetween. The combustor system includes a first combustor assembly and a second combustor assembly. The first combustor assembly is coupled to the turbine assembly for channeling a first flow of combustion gases through the turbine assembly. The first combustor assembly is oriented adjacent to the turbine assembly inlet to channel the first flow of combustion gases to the turbine assembly through the turbine assembly inlet. The second combustor assembly is coupled to the turbine assembly along the combustion gas path for channeling a second flow of combustion gases through the turbine assembly.

Abstract: It is an objective of the invention to provide an Ni-based forged alloy having good large ingot formability and good hot formability as well as high mechanical strength at high temperature. There is provided an Ni-based forged alloy comprising: 0.001 to 0.1 mass % of C; 0.001 to 0.01 mass % of B; 16 to 22 mass % of Cr; 0.5 to 1.5 mass % of Al; 0.1 to 6.0 mass % of W; 3.5 to 5.5 mass % of Nb; 0.8 to 3.0 mass % of Ti; 16 to 20 mass % of Fe; 2.0 mass % or less of Mo; and the balance including Ni and unavoidable impurities, in which: a segregation parameter Ps defined by a formula of “Ps (mass %)=1.05[Al concentration (mass %)]+0.6[Ti concentration (mass %)]?0.8[Nb concentration (mass %)]?0.3[Mo concentration (mass %)]” satisfies a relationship of “Ps??3.0 mass %”; and total amount of W and Mo is 1.75 atomic % or less.

Abstract: A power plant incorporating attributes of a gas turbine engine, flywheel, and electrical generator (hereafter turbine/flywheel or TF) in a single compact unit, having a compressor arrayed with magnets which weight the periphery of the TF. Intermittent combustion periods accelerate the TF to a first rotational velocity, then combustion ceases, and the inlet/outlet of the TF are sealed, causing it to self-evacuate. Conductive coils surround the TF. Magnetic flux between the magnets and coils acts as a motor/generator, electrically powering a load, and absorbing electrical power therefrom via regenerative braking; power out decelerating the TF (now a flywheel), power in accelerating it. A pressure accumulator accepts the TF exhaust, and is pressurized by the combustion periods. Between combustion periods, exhaust in the accumulator expands in a small pump/motor that drives a generator, routing electricity to the TF to raise its rotational velocity.

Abstract: A can annular industrial gas turbine engine, including: a single-piece rotor shaft spanning a compressor section (82), a combustion section (84), a turbine section (86); and a combustion section casing (10) having a section (28) configured as a full hoop. When the combustion section casing is detached from the engine and moved to a maintenance position to allow access to an interior of the engine, a positioning jig (98) is used to support the compressor section casing (83) and turbine section casing (87).

Abstract: A device (30) for optimizing the cooling in a gas turbine of the type comprising at least one compressor, equipped with a combustion chamber (10) and an outer casing (16) and inner casing (32), at least one turbine wheel (14), equipped with a series of blades 12), and at least one high-pressure rotor (38), equipped with one or more supporting bearings (34), the compressor being capable of generating cooling air sent to the turbine wheel (14) through a suitable channel (20). On the outer surface of the device (30), there are one or more grooves (26) which allow the passage of additional flow-rates of air from the compressor towards the channel (20), in order to increase the overall cooling air flow towards said turbine wheel (14). The device (30) can be easily installed without the necessity of dismantling the outer casing (16), and is assembled on pre-existing machines in substitution of the vent tube (22) of the seals (36) of the supporting bearings (34).

Abstract: System and methods for hydrogen assisted oxy-fuel combustion are provided. The system includes a combustor, an air separation unit, a fuel stream source, and a condenser. The combustor includes an oxygen input port, a fuel stream input port, a carbon dioxide input port, and an exhaust output port. The air separation unit is in fluid communication with the combustor via the oxygen input port of the combustor. The fuel stream source is in fluid communication with the combustor via the fuel stream input port and includes a fuel source and a hydrogen source. The condenser is disposed to receive an exhaust from the combustor via the exhaust output port and to return an output stream to the combustor via the carbon dioxide input port. The method includes combusting a fuel stream in a combustor in the presence of an oxidizer to generate an exhaust gas.

Abstract: Provided is a combustor connection structure in which a cross-sectional area (Dout) of a transition piece outlet is set within a range of 0.79?Dout/Din?0.9, relative to a cross-sectional area (Din) of a transition piece inlet of a combustor, and in a first stage turbine nozzle of a turbine connected to the transition piece outlet, a size (Hin) in a radial direction between upstream ends of the inner shroud and the outer shroud has a same dimension as a size (a) in a radial direction of the transition piece outlet. In a combustor transition piece that is included in the combustor having a center line (S) arranged with an angle to an axial center (R) of a rotor of a gas turbine and that guides combustion gas to the turbine, a cross-sectional area is monotonously reduced from the transition piece inlet in which the combustion gas flows to the transition piece outlet from which the combustion gas flows out.

Abstract: Motive power supply equipment for a natural gas liquefaction plant includes gas turbine equipment having a compressor for compressing air, a combustor for generating combustion gases by burning a fuel and the compressed air from the compressor, and a turbine rotated by the combustion gases from the combustor. A low-temperature heat source supply system is provided which, during rated load operation of the gas turbine equipment, supplies carbon dioxide as a cooling medium to the gas turbine equipment in order to cool the gas turbine equipment. The carbon dioxide supplied is generated by separation from the natural gas in the natural gas purification equipment.

Abstract: A turbine of a turbomachine is provided and includes opposing endwalls defining a pathway into which a fluid flow is receivable to flow through the pathway; and a nozzle stage at which adjacent nozzles extend across the pathway between the opposing endwalls to aerodynamically interact with the fluid flow. The adjacent nozzles are configured to define a throat distribution exhibiting endwall throat decambering and pitchline throat overcambering.

Abstract: A turbomachine includes an inner casing component having a first end that extends to a second end and a seal member. An outer casing component is coupled to the inner casing component. The annular outer casing component includes a first end portion that extends to a second end portion and a seal element that aligns with the seal member of the annular inner casing component to form a seal passage. A seal is arranged in the seal passage. The seal includes a first end section that extends to a second end section through an intermediate zone. The first end section includes a recessed portion and the second end section includes a connecting portion. The connecting portion is configured and disposed to nest within the recessed portion to form a substantially continuous seal.

Abstract: A gas turbine engine comprises an outer engine case, a combustor, a high pressure spool and a low pressure spool. The combustor is disposed within the outer engine case. The high pressure spool is configured for rotation coaxially with the combustor. The low pressure spool is configured for rotation coaxially with the high pressure spool. The low pressure spool is configured to rotate at speeds faster than that of the high pressure spool during operation of the gas turbine engine. In one embodiment, the low pressure spool includes a gear system configured to rotate a low pressure compressor faster than a low pressure turbine. In another embodiment, the high pressure spool rotates outward of the combustor within the outer engine case. In another embodiment, the high pressure spool includes a drum having radially inward projecting blades and vanes and radially outward projecting fan blades.

Abstract: In a method for operating a gas turbine (10), a CO2-containing gas is compressed in a compressor (13), the compressed gas is used to burn a fuel in at least one subsequent combustion chamber (14, 18), and the hot combustion gases are used to drive at least one turbine (17, 21). Improved control and performance can be achieved by measuring the species concentration of the gas mixture flowing through the gas turbine (10) at several points within the gas turbine (10) by a distributed plurality of species concentration sensors (22; 22a-1; 23), and utilizing the measured concentration values to control the gas turbine (10) and/or optimize the combustion performance of the gas turbine (10).

Abstract: An exhaust gas diffuser is provided and includes a peripheral body, a center body, formed to define an interior and disposed within the peripheral body to define an annulus between the peripheral body and the center body through which a first fluid flows along a main flow direction, a plurality of first members, each of which is respectively coupled to the peripheral body and the center body, to support the center body within the peripheral body and a plurality of second members, each of which extends across the annulus from the peripheral body to the center body downstream from the plurality of the first members relative to the main flow direction, to transport a second fluid to the center body interior. The plurality of the second members is circumferentially clocked relative to the plurality of the first members.

Abstract: An air cycle machine includes two turbines and a compressor mounted on an integral shaft. The integral shaft includes a plurality of shaft sections that are welded together and machined in a single set-up process into a desired shaft shape to provide highly aligned bearing surfaces. The shaft includes three stops that cooperate with three fasteners to secure the two turbines and the compressor on the shaft.

Abstract: A turbomachine including an improved backflow margin includes a combustor portion fluidly connected to a turbine portion. The turbine portion includes a gas path, a first stage having a first plurality of airfoil members arranged along the gas path, and a second stage having a second plurality of airfoil members arranged along the gas path downstream from the first stage. The first plurality of airfoil members are configured to intercept combustion gases from the combustor portion at a first momentum and create a wake zone having a second momentum that is lower than the first momentum. The turbomachine includes a gas flow aeromechanics system configured and disposed to improve gas flow aeromechanics along the gas path by circumferentially clocking the second plurality of airfoil members relative to the first plurality of airfoil members to intercept the wake zone at the second momentum.

Abstract: A rotor stack assembly for a gas turbine engine includes a first rotor assembly and a second rotor assembly axially downstream from the first rotor assembly. The first rotor assembly and the second rotor assembly include a rim, a bore and a web that extends between the rim and the bore. A tie shaft is positioned radially inward of the bores. The tie shaft maintains a compressive load on the first rotor assembly and the second rotor assembly. The compressive load is communicated through a first load path of the first rotor assembly and a second load path of the second rotor assembly. At least one of the first load path and the second load path is radially inboard of the rims.

Abstract: An aero-derivative can annular gas turbine engine having: an aero gas turbine engine core including an aero high pressure compressor (65) interconnected with an aero high pressure turbine (73) by an aero high pressure shaft (142) in a geometric arrangement appropriate for association with an aero annular combustor (84), but with the aero annular combustor (84) and a first row of turbine vanes (38) of the aero high pressure turbine (73) absent; and a can annular combustor assembly (122) assembled with the aero gas turbine engine core and configured to receive compressed air from the aero high pressure compressor (65) and to accelerate and orient combustion gasses directly onto a first row of blades of the aero high pressure turbine (73).

Abstract: A combustor (14) is placed next to a turbine (16), on the side opposite a compressor (12). A heat insulation device (20) for reducing the transmission of heat from the high-temperature side to the low-temperature side is provided between the combustor/turbine and the compressor. A connection shaft (18) has an axial hole (18a) open on the inlet side of the compressor and axially extending to near a turbine impeller, and also has a radial hole (18b) open near the turbine impeller to the outside of the connection shaft and radially extending to be in communication with the axial hole.

Abstract: A mid-turbine frame buffer system for a gas turbine engine includes a mid-turbine frame that supports a shaft by a bearing. An air compartment and a bearing compartment are arranged radially inward of the mid-turbine frame. The bearing compartment is arranged within the air compartment and includes first and second contact seals arranged on either side of the bearing. The air compartment includes multiple air seals. A high pressure compressor is fluidly connected to the air compartment and is configured to provide high pressure air to the air compartment. A method of providing pressurized air to a buffer system includes sealing a bearing compartment with contact seals, surrounding the bearing compartment with an air compartment, and supplying high pressure air to the air compartment.

Abstract: An internal combustion engine includes in one aspect a source of a pressurized working medium and an expander. The expander has a housing and a piston, movably mounted within and with respect to the housing, to perform one of rotation and reciprocation, each complete rotation or reciprocation defining at least a part of a cycle of the engine. The expander also includes a septum, mounted within the housing and movable with respect to the housing and the piston so as to define in conjunction therewith, over first and second angular ranges of the cycle, a working chamber that is isolated from an intake port and an exhaust port. Combustion occurs at least over the first angular range of the cycle to provide heat to the working medium and so as to increase its pressure.

Abstract: A gas turbine engine includes a compressor, a combustor section, and a turbine. The turbine includes an integrated case/stator segment that is comprised of a ceramic matrix composite material.

Abstract: A tandem fan-turbine rotor assembly for a tip turbine engine includes a hollow fan blade rotor and a solid fan blade rotor. The hollow fan blade rotor includes a first fan blade rotor hub, hollow fan blades, and an inducer. The solid fan blade rotor includes a second fan blade rotor hub, solid fan blades, an exo-ring, and tip turbine blades. The exo-ring connects the solid fan blades at their maximum outward radial extent. The tip turbine blades are connected to the exo-ring and extend radially outward from the exo-ring. The hollow fan blade rotor and the solid fan blade rotor rigidly attach to each other at the first and second fan blade rotor hubs.

Abstract: An example turbomachine assembly includes a first combustor configured to combust fuel and compressed air and a second combustor configured to combust fuel and compressed air. Flow moves through the first combustor in a first direction and flow moves though the second combustor in a second direction different than the first direction. The first combustor is axially spaced from the second combustor.

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 combustor (14) is placed next to a turbine (16), on the side opposite a compressor (12). A heat insulation device (20) for reducing the transmission of heat from the high-temperature side to the low-temperature side is provided between the combustor/turbine and the compressor. A connection shaft (18) has an axial hole (18a) open on the inlet side of the compressor and axially extending to near a turbine impeller, and also has a radial hole (18b) open near the turbine impeller to the outside of the connection shaft and radially extending to be in communication with the axial hole.

Abstract: A rotationally symmetrical wall for an aircraft turbomachine combustion chamber, including primary holes, dilution holes positioned downstream from the primary holes, and a plug installation aperture upstream from the primary holes, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane (D) in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane (D), wherein the other primary holes are distributed equidistantly, at a predefined interval P, and the distance between the two primary holes is greater than the interval P separating two of the other adjacent primary holes.

Abstract: The system according to the invention comprises a rotary shaft (50), a compressor (54) mounted integral with the rotary shaft (50), a power turbine (52) capable of rotating the rotary shaft (50), a cold expansion turbine (56) rotated by the rotary shaft (50) and supplied with a compressed gas from the compressor (54) toward the cold turbine (56). The system (22) comprises an upstream assembly (42) supplying outside air to the aircraft not having passed through a propulsion engine (16A, 16B) of the aircraft, the upstream supply assembly (42) being connected to an inlet of the compressor (54).

Abstract: Various systems and apparatuses are provided for a flow delivery system for an engine. In one example, a system includes a first turbine providing an exhaust flow and a second turbine having an inlet and being fluidically coupled to the first turbine. The second turbine further includes a plurality of nozzle vanes positioned within the inlet of the turbine. A transition conduit is curved about an axis and coupled to the inlet and to the first turbine. The transition conduit is configured to impart an angular momentum component to at least a portion of the exhaust flow, and includes a slot that delivers at least a portion of the exhaust flow to the plurality of nozzle vanes.

Abstract: A compressor-less micro gas turbine has a compressed working medium container for maintaining a gas turbine under pressure. A combustion chamber is in fluid communication with the compressed gas container for receiving a gas from the gas container and heating the gas within the combustion chamber to create an expanded gas. A heater heats the combustion chamber to expand the working gas therein. A turbine in fluid communication with the combustion chamber for receiving the expanded gas. The expanded gas drives the turbine. A generator is operatively coupled to the turbine. The turbine provides a mechanical input to the generator causing the generator to produce electricity.

Abstract: A compressor assembly for a gas turbine engine includes rotatable impeller with forward and aft ends, including: an annular hub defining a generally concave-curved cross-sectional inner flowpath surface at its radially outer periphery, the inner flowpath surface extending between an inlet and an exit; an annular array of airfoil-shaped inducer blades extending radially outward from the inner flowpath surface near the forward end; and an annular array of exducer blades extending outward from the inner flowpath surface, the array of exducer blades axially spaced apart from the array of inducer blades; a non-rotating shroud assembly surrounding the impeller and including a convex-curved outer flowpath surface, wherein the inner and outer flowpath surfaces cooperate to define a primary flowpath past the blades and the stator vanes; and an annular array of airfoil-shaped stator vanes extending radially inward from the outer flowpath surface into a space between the inducer and exducer blades.

Abstract: A head part of an annular gas turbine combustion chamber including: an end wall with a passage opening for accommodating a burner, the end wall including a set back portion adjacent the passage opening; a heat shield covering a back side of the end wall which faces towards the combustion chamber, the heat shield including a protruding portion shaped to cooperate with the set back portion of the end wall; and a burner collar adapted to fit within the passage opening and receive a burner, the burner collar including a protruding portion radially protruding from an outer surface of the burner collar; wherein the head part of the annular gas turbine combustion chamber is configured such that in an installed configuration the protruding portion of the burner collar is held between the protruding portion of the heat shield and the set back portion of the end wall.

Abstract: A composite annular shroud supported by a support assembly including at least two single piece full 360 degree rings and at least partially disposed within an innermost one of the rings. The shroud is biased against and in sealing engagement with an inner flange of the innermost ring. A three ring assembly includes the inner ring disposed radially inwardly of a middle ring disposed radially inwardly of an outer ring and the shroud at least partially disposed within the inner ring. At least three clocking pins extend radially inwardly from the middle ring through slots in the inner ring into notches in the shroud. The middle ring may be an aft end of a support ring fixedly connected to an engine backbone. Mounting pins may be press fitted into pin holes in the middle ring and extend radially outwardly from the middle ring through radial holes in the outer ring.

Abstract: In a communicating structure between combustors that generates combustion gas inside pipe pieces and a turbine portion that generates a rotational driving force by making the combustion gas sequentially pass through a turbine stage formed of turbine stator vanes and turbine rotor blades, at least some of the first-stage turbine stator vanes closest to the combustor among the turbine stator vanes are disposed downstream of sidewalls of one pipe piece and another pipe piece that are adjacent to each other, and the distance from leading edges of the first-stage turbine stator vanes disposed downstream of the sidewalls of the pipe pieces to end portions of the sidewalls closer to the turbine portion is equal to or less than a spacing between an internal surface of the sidewall of the one pipe piece and an internal surface of the sidewall of the other pipe piece.

Abstract: A burner assembly having a fuel distribution ring, number of fuel nozzles which are mounted on the fuel distribution ring in the direction of low is provided. The fuel distribution ring has a ring-shaped surface in the direction of flow and wherein the fuel distribution ring center and an opposite outer outer side and wherein there is at least one slot on the surface between the fuel nozzles and the at least one slot extends on the surface from the outside to the inside. There is at least one recess arranged on the surface and the at least one recess also partially includes the outside of the fuel distributor.

Abstract: An igniter arranged to ignite combustion in a primary flow including a fuel and air mixture, the igniter including one or more geometric features arranged to: induce a shockwave flow structure at least partially disposed in the primary flow; and ignite the fuel and air mixture by virtue of the shockwave flow structure. The present disclosure also relates to a method of igniting combustion in a primary flow including a fuel and air mixture, the method including: providing one or more geometric features arranged to induce a shockwave flow structure; inducing the shockwave flow structure at least partially in the primary flow; and igniting the fuel and air mixture by virtue of the shockwave flow structure.

Abstract: A micro-turbine engine, consisting of at least a compressor, combustor, and turbine, is a complicated fluid flow device that controls the flow rate and thermodynamic properties of a working fluid in order to generate shaft power. Existing micro-turbines are costly to manufacture because they are designed with sophisticated contours and exotic materials. The present invention discloses a method for designing a micro-turbine with stacked layers of structure, each of which is designed with vertically simple geometry such that it can be manufactured using conventional machining technology. The resulting micro-turbine is low cost compared to existing alternatives in the target range of power outputs and applications. The present invention also describes a method for connecting the micro-turbine to an electrical generator to generate power. Lastly, the method for designing the micro-turbine is applied to heat exchangers, Rankine engines, fluid mixers, and other fluid flow devices.

Abstract: A method for producing the electrical power required for equipment on board an airplane is disclosed. Auxiliary power is taken off by a shaft driven by the high pressure turbine and, when idling, the efficiency of the low pressure turbine is degraded so as to enable the high pressure turbine to operate at a speed that is sufficient for delivering the required auxiliary power.

Abstract: A ring element for a turbomachine, in particular for an aircraft gas turbine, is disclosed. The ring element has a ring element main body that has two adjacently arranged ring ends, the ring ends being connected to one another in a form-locking manner with respect to an axial plane. Also disclosed is a turbomachine having at least one such ring element.

Abstract: A combustor (14) is placed next to a turbine (16), on the side opposite a compressor (12). A heat insulation device (20) for reducing the transmission of heat from the high-temperature side to the low-temperature side is provided between the combustor/turbine and the compressor. A connection shaft (18) has an axial hole (18a) open on the inlet side of the compressor and axially extending to near a turbine impeller, and also has a radial hole (18b) open near the turbine impeller to the outside of the connection shaft and radially extending to be in communication with the axial hole.

Abstract: There is described a gas turbine with a channel wall that limits a flow, said channel wall comprising first and second wall sections, one of the wall sections being part of a combustion chamber of the gas turbine and the other wall section being part of an annular hot gas channel of a turbine unit, and the two adjacent wall sections facing each other with a gap there between. In order to reduce the leakage quantities of blocking gas that protect the gap from the entry of hot gas, it is proposed that blocking element projects from the side of the first wall section into the gap and can be displaced transversely thereto, said blocking element comprising a piston face disposed in the first wall section and a sealing structure lying opposite the piston face, it being possible to press said sealing means against an end sealing section of the second wall section by means of a pressure medium that acts on the piston face.

Abstract: A method for balancing a rotating assembly of a gas turbine engine includes removing a stator vane from a section of the gas turbine engine. Removing the stator vane provides access to a rotating assembly of the gas turbine engine. The method further includes at least one of adding, removing, and repositioning a weight with respect to the rotating assembly via access to the rotating assembly provided by removing the stator vane.

Abstract: A cogeneration plant is provided that includes a gas turbine, a heat recovery steam generator, a steam turbine and a cooler/condenser. A division module is provided at a division point, via which downstream the heat recovery steam generator the combustion gas is cooled and dehumidified in the cooler/condenser and then divided into a first combustion gas flow and a second combustion gas flow. A second condenser is provided for receiving the second combustion gas flow to separate contained carbon dioxide from contained water by condensation of the water. The cogeneration plant further includes a heater and a compressor for receiving the first combustion gas flow, which is heated, compressed and partly extracted to by-pass the combustor for cooling of the gas turbine before it enters the combustor and mix with the flow of oxygen and fuel to be burned in the gas turbine.