Abstract: A system includes an injector having a scheduling valve assembly and a nozzle in fluid communication with the valve assembly. The scheduling valve assembly is configured for regulation of flow from an inlet of the injector to the nozzle. The injector includes two fluid circuits between the inlet of the injector and two respective outlets of the nozzle for staged flow output from the nozzle. A first one of the two fluid circuits is a primary circuit, and a second one of the two fluid circuits is a secondary circuit. A solenoid valve is connected in fluid communication with the scheduling valve assembly, wherein the solenoid valve is configured to adjust position of a hydromechanical valve spool of the valve assembly.

Abstract: A single-crystal turbine blade and a method of making such single-crystal turbine blade are disclosed. During manufacturing, a secondary crystallographic orientation of the material of the single-crystal turbine blade is controlled based on a parameter of a root fillet between an airfoil of the single-crystal turbine blade and a platform of the single-crystal turbine blade. The parameter can be a location of peak stress in the root fillet expected during use of the turbine blade.

Abstract: A gas turbine engine including a compressor section, a combustor for combusting a fuel, and a turbine. Compressed air flows through a combustion liner of the combustor in a bulk airflow direction. The combustor includes a primary fuel nozzle and a secondary fuel nozzle. The secondary fuel nozzle is downstream of the primary fuel nozzle in the bulk airflow direction. The primary fuel nozzle is configured to inject a primary portion of the fuel into a primary combustion zone, and the secondary fuel nozzle is configured to inject a secondary portion of the fuel into a secondary combustion zone. The secondary combustion zone is located downstream of the primary combustion zone in the bulk airflow direction. The fuel may be one of diatomic hydrogen fuel and a hydrogen enriched fuel.

Abstract: A burner arrangement has a plurality of mixing channels which extend parallel to the main axis of the burner arrangement and are arranged in at least two concentric circles, in which mixing channels fuel and discharge air from the compressor are mixed during the operation of the burner arrangement. The mixing channels are grouped together into fuel stages so as to produce an irregular staging in the peripheral direction of the burner arrangement.

Abstract: A fuel injector includes an outer body and an inner body. The outer body extends about an axis and has a radially inner surface and a retention groove defined in the inner surface of the outer body. The inner body is positioned within the outer body has an outer surface and a retention tab. The retention tab retains the inner body relative to the outer body by engagement of the retention tab within the retention groove. A is axially offset from the retention tab and fixes the inner body within the outer body.

Abstract: A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution array having a plurality of dilution passages, each dilution passage of the plurality of dilution passages having a concatenated geometry repeating in a predetermined pattern and extending circumferentially around the liner body. The dilution passage integrates a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow and injects the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor. The dilution array is repeated along an axial length of the liner body.

Abstract: A gas turbine arrangement for dual fuel operation has a first manifold that delivers a first fuel or compressor bleed fluid and is connected to a bleed port and a first passage for ejecting fuel or fluid into a combustor space. A second manifold delivers a second fuel and is connected to a second passage for ejecting the second fuel into the combustor space. A control system, when operated with the second fuel, provides the second fuel to the second manifold and continuously opens the bleed valve to provide bleed fluid into the first manifold to replace the first fuel. The control system controls the bleed valve over time by throttling a mass flow of the bleed fluid provided to the first passage or by increasing a mass flow of the bleed fluid provided to the first passage to adapt to fuel properties of the second fuel.

Abstract: A gas turbine engine with a compressor section, a turbine section, and a combustion section located downstream from the compressor section and upstream from the turbine section, the combustion section including: a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a primary fuel injector fluidly coupled to the dome inlet, and a second fuel injector fluidly coupled to the combustion chamber.

Abstract: A method of operating a combustor of a turbomachine on a total fuel input that contains a concentration of hydrogen that is greater than about 80% is provided. The method includes injecting a first mixture of air and a first fuel containing a first amount of hydrogen into the primary combustion zone of the combustor to generate a first flow of combustion gases having a first temperature. The method further includes injecting, with one or more premix injectors disposed downstream of the fuel nozzles, a second mixture of air and a second fuel containing a second amount of hydrogen into the secondary combustion zone of the combustor to generate a second flow of combustion gases having a second temperature. The method further includes separately injecting a third fuel into secondary combustion zone to generate a third flow of combustion gases having a third temperature.

Abstract: A combustor assembly of a gas turbine engine having a trapped vortex feature to reduce emissions where the trapped vortex is formed using ammonia injected into an annular cavity located in a wall surrounding a combustion chamber of the combustor assembly. The annular cavity, and therefore the trapped vortex, is positioned such that when the combustion occurs within the combustion chamber the position of the annular cavity, and therefore of the trapped vortex, is downstream of a flame front. The emissions resulting from combustion travel through the combustion chamber and pass by the annular cavity before exiting the combustion chamber. The trapped vortex in the combustion chamber supplies NH2 radicals, resulting from the ammonia of the trapped vortex, to the passing by emissions and converts NOx and/or N2O in the emissions to non-polluting products, mainly water and nitrogen.

Abstract: A combustor and a gas turbine capable of uniformly supplying air into a burner are provided. The combustor may include a burner including a tubular nozzle casing, a head plate coupled to an end of the nozzle casing, and a plurality of nozzles to inject fuel and air, and a duct assembly coupled to the burner, a mixture of the fuel and the air being burned in the duct assembly to produce combustion gas. Each of the nozzles may include outer nozzles and an inner nozzle installed inside the outer nozzles, each of the outer nozzles may include a nozzle tube configured to provide a channel through which air and fuel flow and a nozzle shroud configured to surround the nozzle tube, and a flow distribution member may be installed between the head plate and the nozzle shroud to distribute a flow rate of air introduced into the outer nozzle.

Abstract: Fuel injectors, combustors, and methods of fabricating a fuel injector are provided. A fuel injector includes a forward end wall and an aft end wall disposed oppositely from one another. The fuel injector also includes side walls that extend between the forward end wall and the aft end wall. The forward end wall and the aft end wall are arcuate. The forward end wall, the aft end wall, and the side walls collectively define an opening for passage of air. The fuel injector further includes at least one fuel injection member disposed within the opening and extending between the forward end wall and the aft end wall.

Abstract: A combustor may include a combustor liner and flow sleeve. A high pressure air cools an outer surface of the combustor liner via openings in the flow sleeve, creating a lower pressure air in an annulus between the combustor liner and the flow sleeve. A first fuel nozzle is positioned at a primary combustion zone, and a second fuel nozzle is positioned at a secondary combustion zone of the liner. A fuel source is configured to deliver a fuel to the fuel nozzles. The fuel nozzles produce a premixture of high pressure air and the fuel, and produce a mixture of the premixture and the lower pressure air, prior to introducing the mixture to a respective primary or secondary combustion zone of the combustor. The combustor provides improved fuel premixing and is fuel flexible, and reduces pressure drop requirements. The combustor is usable in a can, annular, or segmented annular combustor assembly.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a liner defining a combustion chamber therewithin and a pressure plenum surrounding the liner. The liner defines an opening and includes a walled chute disposed at least partially through the opening. A plurality of flow openings is defined through the walled chute.

Abstract: An injection assembly for a gas turbine combustor having a liner defining a combustion zone and a secondary combustion zone and a forward casing circumferentially surrounding at least a portion of the liner is provided. The injection assembly includes a thimble assembly and an injector unit. The thimble assembly, which is mounted to the liner, includes a thimble that extends through a thimble aperture in the liner. The injector unit, which is mounted to and extends through the forward casing, includes an injector blade that extends into the thimble. The injection assembly introduces a flow of fuel into a flow of air flowing through the thimble, such that fuel and air are injected into the secondary combustion zone in a direction transverse to a flow of combustion products from the primary combustion zone.

Abstract: A multipoint fuel injection system comprises an injection system segment including a circumferentially extending outer support defining a fuel manifold with a plurality of manifold passages extending circumferentially therethrough. A first connector is included at a first circumferential end of the outer support and a second connector is included at a second circumferential end of the outer support opposite the first circumferential end. The first and second connectors are each configured to connect each manifold passage with a manifold passages of a respective outer support of a circumferentially adjacent injection system segment. The system includes a circumferentially extending inner support and a plurality of circumferentially spaced apart feed arms extending radially between the inner support and the outer support. A plurality of outlet openings extend in an axial direction from each feed arm for feeding respective injection nozzles.

Abstract: A fuel swirler, for a gas turbine engine, has a primary cone housing defining an interior chamber. The interior chamber has an inlet in communication with a source of pressurized fuel. The interior chamber has a transition portion and a socket portion with an axisymmetric interior surface. A swirler core is disposed within the interior chamber. The swirler core has a downstream end and an upstream shank portion having an exterior surface mating the axisymmetric interior surface of the socket portion. The shank portion has a plurality of axially extending grooves. The grooves are disposed axisymmetrically about the exterior surface of the shank portion.

Abstract: A fuel nozzle for a gas turbine engine is generally provided. The fuel nozzle includes an outer sleeve extended circumferentially around a fuel nozzle centerline and extended along a longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve defines a plurality of first radially oriented air inlet ports through the outer sleeve in circumferential arrangement relative to the fuel nozzle centerline. A centerbody is positioned radially inward of the outer sleeve. The centerbody includes an outer annular wall extended circumferentially around the fuel nozzle centerline and extended along the longitudinal direction co-directional to the fuel nozzle centerline. The outer sleeve further includes a first radial wall extended from a downstream end of the outer annular wall. The centerbody defines a plenum radially inward thereof. The outer annular wall and the outer sleeve together define a first fuel air mixing passage extended along the longitudinal direction therebetween.

Abstract: A fuel injector includes an outer body and an inner body. The outer body extends about an axis and has a radially inner surface and a retention groove defined in the inner surface of the outer body. The inner body is positioned within the outer body has an outer surface and a retention tab. The retention tab retains the inner body relative to the outer body by engagement of the retention tab within the retention groove. A is axially offset from the retention tab and fixes the inner body within the outer body.

Abstract: A fuel injector including a fuel supply tube (8), a plurality of premix tubes (2), a support plate (3) which supports the fuel supply tube and the plurality of premix tubes, a substrate which supports downstream end portions of the plurality of premix tubes, an outer wall which is cylindrical, which forms a plenum (P) inside, a partition plate which partitions the plenum (P) into a fuel plenum (PF) and a cooling air plenum (PA), a baffle which partitions the cooling air plenum (PA) into an upstream cooling air plenum (PA1) and a downstream cooling air plenum (PA2) and has a plurality of cooling holes formed therein, and a cooling air supply tube configured to supply cooling air to the upstream cooling air plenum (PA1), in which the end portion on the downstream side of the fuel supply tube opens in the fuel plenum (PF), and a fuel introduction hole.

Abstract: The present disclosure is directed to a method for operating a turbine engine, the method including arranging a fluid conduit through a fuel nozzle in a first direction toward a downstream end and in a second direction toward an upstream end, the fluid conduit in fluid communication with a premix passage via a fluid injection port; flowing an oxidizer into the premix passage via a radially oriented first inlet port and a radially oriented second inlet port; flowing a first fuel to the premix passage through the fluid conduit and the fluid injection port, wherein the first fuel is provided to the premix passage axially downstream of the first inlet port; and generating a premixed flame from a mixture of the oxidizer and the first fuel.

Abstract: A System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

Abstract: A state quantity acquiring unit is configured to acquire a state quantity of a turbine. A time calculation unit is configured to calculate an operable time at an over firing operation of the turbine based on a design life of the turbine and a state quantity. A distance calculation unit configured to calculate a Mahalanobis distance based on the state quantity. A determination unit configured to determine whether or not the over firing operation of the turbine is possible by the Mahalanobis distance and the operable time.

Abstract: A gas turbine engine combustor has a dome and a shell extending axially from the dome. The dome and the shell cooperates to define a combustion chamber. A dome heat shield is mounted to the dome inside the combustion chamber. A front heat shield is mounted to the shell inside the combustion chamber. The dome heat shield and the front heat shield have axially overlapping portions cooperating to define a flow guiding channel. The flow guiding channel has a length (L) and a height (h). The length (L) is at least equal to the height (h).

Abstract: An annular combustor is disclosed in which free-vortices are generated which: enhance fuel air mixing, recirculate the air, provide cooling for the combustor walls, and provide low emissions and a substantially uniform exit temperature profile. The combustor is provided fuel and air through a plurality of fuel injectors which atomizes the fuel and promote individual vortices emanating from each injector. The combustor includes an annular prechamber, an annular main chamber, and an annular dilution zone. A first swirler is provided on the inner side of the annular prechamber and a second swirler on the outer side of the annular prechamber, the swirlers causing flow to rotate in mutually opposite directions to intensify the vortices imposed by the injectors. The vortices cause a pressure depression to promote some backflow that enhance mixing to: promote complete combustion, produce low emissions, and provide cooling.

Abstract: Disclosed are a fuel nozzle assembly, and a fuel nozzle module and gas turbine having the same. The fuel nozzle assembly includes a fuel nozzle, a shroud spaced apart from the fuel nozzle and defining a flow path between an inner wall and the fuel nozzle, a rim formed around the shroud to guide air, and a turning guide spaced apart from the rim to distribute the air. The turning guide includes a turning separator spaced apart from the rim to extend in a circumferential direction of the rim, and inner separators extending in a radial direction of the rim to interconnect opposite circumferential ends of the turning separator and an outer surface of the fuel nozzle. The fuel nozzle assembly prevents air pockets, and ensures uniform supply of air, thereby preventing a local increase in combustion temperature inside the fuel nozzle and reducing generation of NOx.

Abstract: A combustor or augmentor of a gas turbine engine is provided. The combustor or augmentor includes a wall defining apertures and an interior having an inlet, an outlet and a mixing region between the inlet and the outlet and fuel injectors. The fuel injectors are respectively arrayed along the wall at corresponding apertures to inject fuel into the mixing region. Each fuel injector includes a fuel injector body defining an injection outlet through which fuel exits the fuel injector body toward the mixing region and a pintle. The pintle includes a first end, a second end and a pintle body extending between the first and second ends. At least one of the first and second ends is attachable to the fuel injector body to position the pintle body between the injection outlet and the mixing region.

Abstract: A combustor system for a GT system may include: a plurality of burners, each burner including an inflow region for receiving a combustion air flow and a mixing zone disposed downstream of the inflow region for receiving the air flow and a fuel flow; a combustion chamber disposed downstream of the mixing zone; a fuel flow valve system disposed to control the fuel flow to each of the plurality of burners; a combustion sensor configured to determine a combustion parameter; and an exhaust sensor configured to determine an exhaust parameter. A control system may be connected to the combustion sensor, the exhaust sensor and fuel flow valve system. The control system, in response to the gas turbine system operating at a low partial load, redistributes the fuel flow to at least one burner of the plurality of burners as a function of a predetermined emission limit.

Abstract: A combustor dome system includes an annular combustor dome defining a main axis therethrough. The combustor dome includes opposed upstream and downstream faces, wherein the upstream face is configured to face upstream toward a compressor discharge space, wherein the downstream face is configured to face downstream toward a combustor space. The downstream face has a curved cross-sectional profile. A plurality of nozzles extends at least partially through the combustor dome from the upstream face to the downstream face for injection of fuel into the combustor space. A fuel manifold is in fluid communication with the plurality of nozzles.

Abstract: A fuel nozzle assembly includes an inner fuel nozzle; a plurality of outer fuel nozzles disposed radially around the inner fuel nozzle, each outer fuel nozzle including a central body for fuel injection, a shroud spaced apart from and surrounding the central body, the shrouds forming an outer periphery of the fuel nozzle assembly, and an inlet formed at one end of the shroud; and a peripheral rim formed at the inlets and disposed to cover at least a portion of the outer periphery. The fuel nozzle assembly is connected to an end plate of a combustion chamber. The one end of a corresponding shroud is disposed at a set distance from an end plate of a combustion chamber, the set distance depending on at least one of relative positions of the inner fuel nozzle and the plurality of outer fuel nozzles and an inlet radius of the corresponding shroud.

Abstract: A gas turbine may include a rotatable shaft, a compressor disposed about the rotatable shaft and configured to output compressed air, and a combustor disposed about the rotatable shaft. The combustor may be configured to receive the compressed air and output high temperature compressed gas. The gas turbine may further include a power turbine disposed about the rotatable shaft and configured to receive the high temperature compressed gas, and a first liner defining a plurality of holes and disposed around the combustor. The power turbine may be configured to expand the high temperature compressed gas and rotate the rotatable shaft. The first liner may have a first end and a longitudinally opposite second end. The first end may be coupled to an inner surface of the casing at or adjacent an upstream end of the combustor and the second end may be substantially free from any connection with the casing.

Abstract: A combustor panel for use in a gas turbine engine includes a body portion having an outwardly curved edge defining a channel between the body portion and the outwardly curved edge. The outwardly curved edge includes a plurality of slots. The panel further includes a fastening member having a base disposed within the channel and a plurality of fasteners extending from the base and disposed within the slots.

Abstract: An assembly for a turbine engine includes a combustor bulkhead, a swirler and a mounting strap. The swirler includes a mounting lug and a tubular swirler body that extends along an axis. The mounting lug projects radially out from the swirler body. The mounting strap is attached to the combustor bulkhead. The mounting lug is axially restrained between the combustor bulkhead and the mounting strap.

Abstract: A fuel injection nozzle (31) of the present invention comprises a pre-mixing passage (34) which mixes fuel (103) and air (102) to generate an air-fuel mixture; and an injection port (33) which is located downstream of the pre-mixing passage (34) and injects the air-fuel mixture into a combustion chamber (26). The fuel injection port (33) has a slit shape and a width that is less than a dimension that is twice as large as a quenching distance.

Abstract: A burner arrangement having a combustion chamber, a multiplicity of mixing ducts discharging into the combustion chamber, in which ducts combustion air and fuel introduced during proper operation are mixed, and at least one resonator which has a defined resonator volume and resonator openings. The mixing ducts have mixing tubes which extend axially through an annular space defined between a tubular outer wall, a tubular inner wall arranged with the radial separation from the outer wall, an annular end plate arranged upstream and an annular end plate arranged downstream, wherein the end plates are provided with through openings which receive and/or prolong the mixing tubes. The resonator openings of the at least one resonator are designed as air ducts that extend through the downstream end plate, and the resonator volume of the at least one resonator is formed by at least one part of the annular space.

Abstract: A collar configured to couple a burner to a partition plate is provided including a body having a diameter configured to couple to a diameter of the burner. A flange extends outwardly from the body. The collar is formed from a heat resistance material such that heat transfer between the burner and the partition plate is limited.

Abstract: A system for dissipating fuel egress includes a fuel injector and a fuel supply assembly. The fuel supply assembly has a fuel supply line that provides fuel and a purge air conduit arranged co-axially around the fuel supply line, thereby defining an annular channel around the fuel supply line. The fuel injector includes a perimeter body, a fuel injection port associated with the perimeter body, and an internal mixing chamber downstream of the fuel injection port in a direction of fuel flow. The fuel supply line provides fuel to the fuel injection port. The perimeter body of the fuel injector defines a fuel egress passage therethrough, which has an inlet in fluid communication with the annular channel and an outlet in fluid communication with the mixing chamber. A gas turbine combustor having an axial fuel staging (AFS) system with the present fuel egress dissipation system is also provided.

Abstract: A system is provided with a turbine combustor having a first diffusion fuel nozzle, wherein the first diffusion fuel nozzle is configured to produce a diffusion flame. The system includes a turbine driven by combustion products from the diffusion flame in the turbine combustor. The system also includes an exhaust gas compressor, wherein the exhaust gas compressor is configured to compress and route an exhaust gas from the turbine to the turbine combustor along an exhaust recirculation path.

Abstract: Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Abstract: A gas turbine engine and combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two axially spaced apart annular trapped vortex cavities are located on the combustor liner and staged axially and radially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities. A plurality of injectors are configured tangentially relative to circular radially outer wall extending between an aft wall and a forward wall of each cavity to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid.

Abstract: A fuel nozzle apparatus for a gas turbine engine includes: an annular outer body extending parallel to a centerline axis and having an exterior surface, and having a ring of forward openings passing through the exterior surface, and a ring of aft openings passing through the exterior surface, the aft openings positioned axially aft of the forward openings; an annular main injection ring disposed inside the outer body and including: a forward main fuel gallery extending in a circumferential direction; an aft main fuel gallery extending in a circumferential direction; a ring of forward main fuel orifices communicating with the forward main fuel gallery and each aligned with one of the forward openings; a ring of aft main fuel orifices, communicating with the aft main fuel gallery and each aligned with one of the aft openings; and a pilot fuel injector disposed along the centerline axis.

Abstract: A combustor for a gas turbine engine is disclosed. The combustor is described as comprising a dome plate coupled to a liner thereof, with at least one heat shield comprised of a ceramic matrix composite coupled at the aft end of the dome plate. Also described is a method for assembling a combustor for a gas turbine engine, including releasing a metal alloy heat shield from a dome plate and providing a ceramic matrix composite heat shield as replacement.

Abstract: A combustor assembly includes a first wall, a second wall, a bulkhead and a plurality of fuel injectors. The bulkhead forms a combustion chamber with the first and the second walls. The fuel injectors are configured with the first wall in a unique and/or a fluctuating pattern.

Abstract: A fuel injection system for use in a combustor of a turbine assembly includes a substantially annular fuel injection housing at least partially defining an annular cavity. The fuel injection system also includes a fuel manifold and an air manifold. The fuel manifold includes a plurality of fuel injection ports extending through the fuel injection housing. The air manifold includes a first plurality of air injection ports and a second plurality of air injection ports that each extends through the fuel injection housing.

Abstract: A fuel delivery system for a gas turbine combustor is provided. The fuel delivery system includes a fuel tube and an attachment assembly. The fuel tube extends from a downstream injector (e.g., a late lean injector) towards a mounting ring of the combustor and is moveably attached to the mounting ring using the attachment assembly. Accordingly, the exemplary fuel delivery system may provide fuel to a downstream injector while accommodating a thermal expansion or contraction of the fuel tube along an axial direction of the combustor.

Abstract: A fuel injector for a gas turbine engine includes a monolithic nozzle body that defines within its interior one or more fuel circuits. Each fuel circuit includes an inlet, an outlet orifice, a main passage fluidly coupling the inlet with the outlet orifice, and a branch passage connected to the main passage. The branch passage connects to the main passage downstream of the inlet and upstream of the outlet orifice to form an effective metering flow area that is smaller than the flow area of the outlet orifice.

Abstract: The present invention discloses a novel apparatus and way for controlling a velocity of a fuel-air mixture entering a gas turbine combustion system. The apparatus comprises a hemispherical dome assembly which directs a fuel-air mixture along a portion of the outer wall of a combustion liner and turns the fuel-air mixture to enter the combustion liner in a manner coaxial to the combustor axis and radially outward of a pilot fuel nozzle so as to regulate the velocity of the fuel-air mixture.

Abstract: A gas turbine engine that includes: a combustor coupled to a turbine and a downstream injection system that includes two injection stages, a first stage and a second stage, positioned within an interior flowpath, wherein the first stage comprises an axial position that is aft of the primary air and fuel injection system and the second stage comprising an axial position that is aft of the first stage. Each of the first stage and the second stage include a plurality of circumferentially spaced injectors, each injector of which is configured to inject air and fuel into a flow through the interior flowpath. The first stage and the second stage have a configuration that limits a fuel injected at the second stage to less than 50% of a fuel injected at the first stage.

Abstract: An annular combustion chamber for a turbine engine, the chamber including an inner wall and an outer wall forming surfaces of revolution, the walls being connected together by a chamber end wall fitted with a fuel injection mechanism, each of the inner and outer walls including primary holes and dilution holes situated downstream from the primary holes in a gas flow direction, the primary and dilution holes being regularly distributed around the circumference of the inner and outer walls. The outer wall has a greater number of dilution holes than the inner wall.

Abstract: A gas turbine that includes: a combustor coupled to a turbine that together define an interior flowpath, the interior flowpath extending aftward about a longitudinal axis from a primary air and fuel injection system that defines a forward end, through an interface at which the combustor connects to the turbine, and through a row of stator blades in the turbine that defines an aft end; and a downstream injection system that includes two injection stages, a first stage and a second stage, that are axially spaced along the longitudinal axis of the interior flowpath. The first stage and the second stage each includes multiple injectors configured to inject an air and fuel mixture into the interior flowpath.