Abstract: A burner of a gas turbine including a reaction chamber and a plurality of jet nozzles opening into the reaction chamber is provided. Fluid is injected through an outlet into the reaction chamber by the jet nozzles using of a fluid stream wherein the fluid is burned into hot gas in the reaction chamber. An annular gap is disposed about the fluid stream for at least one jet nozzle so that a part of the hot gas is drawn out of the reaction chamber and flows opposite the fluid flow direction into the annular gap and is mixed with the fluid stream within the jet nozzle. The ring gap is formed by means of an insert tube, and wherein the insert rube includes a thickening at the upstream end. A method for stabilizing the flame of such a burner of a gas turbine is also provided.

Abstract: A fuel nozzle for a gas turbine combustor includes: an internal cylinder; a burner cylinder concentrically provided on the outer circumference of the internal cylinder in the vicinity of its distal end to surround the distal end portion of the internal cylinder; an air passage provided between the burner cylinder and the internal cylinder for delivering combustion air; a space provided at an axial center of the internal cylinder and extending in the axial direction of the internal cylinder; a plurality of inner passages provided at approximately equal intervals in the circumferential direction of the internal cylinder for injecting fluid through its distal end portion; and at least one communication portion provided on the upstream side of the distal end portion of the internal cylinder and extending from the outer wall of the internal cylinder toward the inner side in the radial direction to communicate with the space.

Abstract: According to one aspect of the invention, a fuel nozzle assembly for a turbine includes an inner conduit and a flange coupled to the inner conduit thereby forming a chamber for flow of a gas fuel. In addition, the flange includes a diaphragm member coupled to the inner conduit, the diaphragm member being configured to flex in response to relative movement between the inner conduit and the flange.

Abstract: According to various embodiments, a system includes a staggered multi-nozzle assembly. The staggered multi-nozzle assembly includes a first fuel nozzle having a first axis and a first flow path extending to a first downstream end portion, wherein the first fuel nozzle has a first non-circular perimeter at the first downstream end portion. The staggered multi-nozzle assembly also includes a second fuel nozzle having a second axis and a second flow path extending to a second downstream end portion, wherein the first and second downstream end portions are axially offset from one another relative to the first and second axes. The staggered multi-nozzle assembly further includes a cap member disposed circumferentially about at least the first and second fuel nozzles to assemble the staggered multi-nozzle assembly.

Abstract: A fuel injector is provided and includes a first tube connectable with a vessel and second and third tubes to deliver fuel to an interior of the vessel, the second and third tubes being supported within the first tube such that the second and third tubes are offset in first and second opposing directions from a first plane of the first tube, respectively, and in a third direction from a second plane of the first tube, the first tube defining an annulus about the second and third tubes through which gas is deliverable to the vessel interior.

Abstract: A fuel lance (7) for introducing fuel into a gas flow in a combustor (1) of a gas turbine engine includes a region of the lance (7) through which the fuel is introduced into the gas flow having a generally helical formation (12).

Abstract: Disclosed is a cross flow apparatus including a surface and at least one outlet located at the surface. The cross flow apparatus further includes at least one guide at the surface configured to direct an intersecting flow flowing across the surface and increase a velocity of a cross flow being expelled from the at least one outlet downstream from the at least one outlet.

Abstract: A combustor housing includes an inlet cover plate having a central inlet configured to receive a supply of one of a high BTU content fuel or air and at least one radially-spaced, peripheral fuel inlet configured to receive a supply of a low BTU content fuel. It also includes an outlet cover plate having at least one radially-spaced, peripheral fuel outlet. The combustor housing also includes a peripheral sidewall joining the inlet cover and the outlet cover and enclosing a plenum, the at least one peripheral fuel inlet opening through the inlet cover plate into the plenum and the at least one fuel outlet opening from the plenum through the outlet cover plate. The central inlet opens into at least one conduit which extends away from the central inlet and opens into at least one high BTU content fuel conduit or air supply conduit that is axially aligned with the at least one fuel outlet.

Abstract: A system is disclosed that includes a combustor having an end cover and at least one fuel nozzle assembly extending from an inner face of the end cover. A cartridge may extend through the end cover and into the fuel nozzle assembly. The cartridge may define an opening for receiving light emitted from within the combustor. Additionally, a fiber optic cable may be disposed within the cartridge and may be configured to capture at least a portion of the light received through the opening.

Abstract: A gas-turbine lean combustor includes a combustion chamber (2) and a fuel nozzle (1) which includes a pilot fuel injection (17) and a main fuel injection (18). The main fuel injection (18) includes central recesses (23) for a controlled inhomogeneous fuel injection, the number of said recesses on the circumference ranging from 8 to 40 and said recesses having an angle of inclination ?2 in circumferential direction of 10°??2?60° and an axial angle of inclination ?1 relative to the combustor axis (4) between ?10°??1?90°.

Abstract: A turbomachine including an annular combustion chamber fitted with fuel injectors and nozzle vanes arranged at the outlet from the chamber, the number of nozzle vanes being an integer multiple of the number of fuel injectors, and the head of each injector being situated angularly half-way between the leading edges of two consecutive nozzle vanes, these leading edges being in alignment with primary air holes and/or with dilution air holes.

Abstract: A system includes a fuel nozzle for a turbine engine that includes a tapered central body located at an interior base of the fuel nozzle, an air swirler, and a fuel port in the tapered central body, separate from the air swirler.

Abstract: A cooling circuit for a fuel nozzle in a gas turbine includes an end cap cavity receiving passive purge flow from a compressor of the turbine, and fuel nozzle swozzles disposed in a swozzle shroud that impart swirl to incoming fuel and air. Purge slots are formed in the swozzle shroud and through the fuel nozzle swozzles in fluid communication with the end cap cavity. The purge slots are positioned upstream of a quat fuel injection passage, and the passive purge flow enters fuel nozzle tip cavities of the fuel nozzle to provide tip cooling and tip purge volume without mixing the passive purge flow with quat fuel.

Abstract: An apparatus is disclosed including a main fuel supply fluidly coupled to a main fuel valve and an arcuate fuel passage receiving main fuel through the main fuel valve. The arcuate fuel passage includes a passage diameter and a radius of curvature which provides sufficient rotational acceleration to the main fuel such that a liquid portion of the main fuel is deposited on an outer wall of the arcuate fuel passage. The apparatus includes a fuel injector nozzle that receives the main fuel from the arcuate fuel passage, and injects the main fuel into a combustion chamber for a turbine engine. The apparatus further includes a pilot fuel supply fluidly coupled to a pilot fuel passage and a fuel selector structured to selectively provide the main fuel or the pilot fuel to the fuel injector nozzle.

Abstract: A protection system for a pre-mixing apparatus for a turbine engine, includes: a main body having an inlet portion, an outlet portion and an exterior wall that collectively establish a fuel delivery plenum; and a plurality of fuel mixing tubes that extend through at least a portion of the fuel delivery plenum, each of the plurality of fuel mixing tubes including at least one fuel feed opening fluidly connected to the fuel delivery plenum; at least one thermal fuse disposed on an exterior surface of at least one tube, the at least one thermal fuse including a material that will melt upon ignition of fuel within the at least one tube and cause a diversion of fuel from the fuel feed opening to at least one bypass opening. A method and a turbine engine in accordance with the protection system are also provided.

Abstract: A gas turbine engine fuel nozzle assembly includes a pilot fuel injector tip substantially centered about a centerline axis in an annular pilot inlet to a pilot mixer. A cross over arm extends across inlet to tip. Arm includes an arm fairing surrounding at least one fuel transfer tube to tip. A pilot nose cap is at an upstream end of tip. At least one of the cross over arm and the pilot nose cap includes a rounded forebody followed by a straight afterbody. Rounded forebody of pilot nose cap includes a substantially rounded dome extending forwardly from a rounded nose base and straight afterbody includes a substantially cylindrical nose afterbody parallel to a pilot nose centerline normal to nose base. An arm fairing surrounding fuel transfer tubes to tip includes rounded leading and trailing edges and a rectangular middle section therebetween having flat first and second sides.

Abstract: Disclosed is a swirler for use in a burner of a gas turbine engine, the swirler comprising a pluralkity of vanes arranged in a circle, flow slots being defined between adjacent vanes in a circle, each flow slot having an inlet end and an outlet end, in use of the swirler a flow of air and fuel traveling along each flow slot fom its inlet end to its outlet end such that the swirler provides a swirling mix of the air and fuel, at least one vane being configured to generate a flow vortex that extends from an edge of the vane adjacent an outlet end of a flow slot to within the swirling mix thereby to improve the mix of air and fuel in the swirling mix.

Abstract: A fuel nozzle sheath has a lateral opening for admitting air about a nozzle stern. The stress distribution along the perimeter of the window is smoothed out by increasing the corner radius of the window corner presenting the highest stress concentration. The different corner radii of the window opening in the sheath allows to reduce stresses resulting from the load transferred from the combustor liner to the sheath.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A fuel supply arrangement for a gas turbine burner is provided. The fuel distribution arrangement includes a rear section external to the burner located between a turbine wall and a fuel distribution system interface directing fuel into fuel supply circuits. The fuel distribution system interface has four fuel connections. An intermediate section is located between the turbine wall and the backside wall of a distribution chamber; and a front section in front of the second section located between said backside wall of the distribution chamber and a burner central backside block. The fuel distribution arrangement includes pipes for gaseous fuel, liquid fuel, as well as pilot gas, and liquid pilot fuel. In the rear section the pipes for gaseous fuel and for liquid fuel are arranged concentrically, and in at least one portion of the intermediate section the pipe for gaseous fuel is arranged non-concentrically with the liquid fuel pipe.

Abstract: The present application provides a fuel nozzle. The fuel nozzle may include a tube and a gas only insert tip positioned on the tube. The gas only insert tip may include a tip aperture. A plug may be positioned within the tip aperture.

Abstract: A thermally diluted exothermic reactor system is comprised of numerous orifices distributed within a combustor by distributed perforated contactor tubes or ducts. The perforated contactors deliver and mix diluent fluid and one or more reactant fluids with an oxidant fluid. Numerous micro-jets about the perforated tubes deliver, mix and control the composition of reactant fluid, oxidant fluid and diluent fluid. The reactor controls one or more of composition profiles, composition ratio profiles and temperature profiles in one or more of the axial direction and one or two transverse directions, reduces temperature gradients and improves power, efficiency and emissions.

Abstract: In one embodiment, a system includes a turbine engine fuel nozzle having an air path, a fuel path, and a surface along the air path. The fuel path may be directed toward the surface. The turbine engine fuel nozzle also may include a heating element configured to heat the surface.

Abstract: A gas turbine engine is piloted with a pilot flow of fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A second additional flow of the fuel is vaporized and delivered to the combustor as a vapor. A fuel injector may have passageways associated with each of the three flows.

Abstract: A swirler passage is provided for mixing fuel and compressor air with at least two side walls; at least first and second conduits arranged inside at least one of the at least two side walls, the first conduit forming a fuel gas conduit and the second conduit forming an air conduit, the fuel gas conduit connected to a gas fuel supply and the air conduit connected to an air supply; a tube connected to the fuel gas conduit and entirely traversing the air conduit inside the side wall; a fluid passage connected to the air conduit and surrounding the tube; at least one fuel outlet opening of the tube arranged on the side wall; and at least one air outlet opening of the fluid passage arranged on the side wall and surrounding the fuel outlet opening.

Abstract: The positioning arrangement comprises at least one pair of complimentary rounded recesses, each recess being provided on the mating face of the corresponding component. One ball is positioned between each pair of complementary recesses.

Abstract: A nozzle includes a center body that defines an axial centerline. A shroud circumferentially surrounds at least a portion of the center body to define an annular passage between the center body and the shroud. A plenum is inside the center body and substantially parallel to the axial centerline, and an igniter is inside the center body and generally adjacent to the plenum. A method for igniting a combustor includes flowing a fuel through a center body axially aligned in a nozzle and flowing a working fluid through an annular passage, wherein the annular passage is substantially parallel to and radially outward of the center body. The method further includes projecting at least one of a beam, spark, or flame from an igniter located inside the center body.

Abstract: Methods of managing fuel temperatures in fuel injectors for gas turbine engines include cooling fuel circuits by initiating fuel flow therethrough using opened, closed, and semi-open control techniques. A fuel injector for a gas turbine engine includes a feed arm having a fuel inlet fitting for delivering fuel to at least one fuel conduit extending through the feed arm. A nozzle body depends from the feed arm and has at least one fuel circuit extending therethrough. The fuel circuit is configured and adapted to receive fuel from the feed arm and to issue fuel from an exit orifice of the nozzle body. Sensing means are provided adjacent to the at least one fuel circuit. The sensing means are configured and adapted to provide temperature feedback in order to control fuel flow in the at least one fuel circuit to maintain fuel temperature within a predetermined range.

Abstract: The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole.

Abstract: A combustion dynamics control system for an aviation based or land based gas turbine engine employs an acoustic driver that is configured to drive pressure perturbations across a premixed fuel injection orifice to substantially zero in response to a control signal such that fuel flow perturbations across the fuel injection orifice are substantially zero.

Abstract: A combustor structure includes a combustor fuel-air mixing apparatus having a mainstream airflow region. One or more concavities are disposed within the mainstream airflow region of the fuel-air mixing apparatus. At least one fuel injection hole is disposed within an upstream base region of at least one concavity. Passing a stream of air through the main airflow region causes fuel injected into the upstream base region of at least one concavity to passively mix with a naturally oscillating vortex of air within the concavity.

Abstract: A mixer for a gas turbine engine combustor includes: (a) a primary fuel injector with upstream and downstream ends, the primary fuel injector including: (1) a hollow central portion configured to pass a first air stream; (2) an annular fuel passage at the downstream end, configured to discharge an annular first fuel stream surrounding the first air stream; (b) an annular splitter surrounding the primary fuel injector, configured to pass a second air stream surrounding the first fuel stream; (c) an annular first housing surrounding the splitter, and having an exit positioned axially downstream of the primary fuel injector and the splitter; and (d) fuel injection ports disposed in an array outside the first housing, and disposed in communication with a fuel supply and positioned to discharge a second fuel stream into a third air stream at a position axially upstream of the exit of the first housing.

Abstract: A late lean injection (LLI) manifold is provided and includes a central nozzle and first and second side nozzles positioned at circumferential locations defined around a vessel, a connector, a first leg, to which the connector is connected, formed to define a tube extending from the central nozzle to the first side nozzle such that fuel is communicable between the connector, the central nozzle and the first side nozzle and a second leg. The second leg is formed to define a tube extending from the central nozzle to the second side nozzle such that the fuel is communicable between the central nozzle and the second side nozzle.

Abstract: A fuel injector for a gas turbine engine is disclosed. The fuel injector includes an injector housing having a longitudinal axis. The injector housing includes one or more fuel inlets, one or more fuel galleries annularly disposed about the longitudinal axis, and an air inlet. The fuel injector also includes a premix barrel having a proximal end and a distal end circumferentially disposed about the longitudinal axis. The premix barrel is fluidly coupled to the fuel galleries and the air inlet at the proximal end, and configured to mechanically couple to a combustor of the gas turbine engine at the distal end. The fuel injector also includes a pilot assembly disposed radially inwards of the premix barrel. The pilot assembly may include a first end and a second end. The second end is removably coupled to the injector housing, and the first end is proximate the distal end of the premix barrel. The pilot assembly is fluidly coupled to the fuel galleries, the air inlet, and the combustor.

Abstract: A fuel-injection device in a turbomachine is disclosed. The device includes a high-pressure pump supplying a flow control valve, whose outlet is connected via a pressurization and cut-off valve to a fuel-injector feed pipe. The valve is connected to the inlet and to the outlet of the pump in order to define two fuel pressurization thresholds, one of which is used to start-up and restart the turbomachine and the other is used for operating the turbomachine from an idle speed and for commanding an equipment with variable geometry.

Abstract: A gas turbine combustor comprising a fuel nozzle for injecting mixed gas of fuel and air, a cylindrical liner for burning and reacting the mixed gas of fuel and air in a combustion chamber, a transition piece which is a flow path for leading combustion gas generated in the liner to turbine blades, and a transition piece flow sleeve for wrapping an outside surface of the transition piece, wherein a plurality of air introduction holes for introducing air into the transition piece flow sleeve are formed in regions of the transition piece flow sleeve excluding regions which are corner portions of the transition piece flow sleeve in a sectional direction thereof.

Abstract: A fuel injection nozzle for a gas turbine engine has a central fuel ejection nozzle and a plurality of airflow passages within the spray tip that include a first and second group of circumferentially spaced apart fuel-spray forming airflow passages disposed on opposite sides of a transverse axis and oriented towards each other such as to produce opposed fuel spray shaping air jets which generate a shaped final fuel spray.

Abstract: A fuel manifold for distributing fuel to a gas turbine engine includes a plurality of interconnected manifold segments. Each manifold segment extends between a pair of fuel injector inlet fittings. Each manifold segment includes a fuel liner defining an internal fuel passage therethrough fluidly connecting a pair of fuel injector inlet fittings. A wire braid layer surrounds the fuel liner, a fire sleeve surrounds the wire braid layer, and a flexible thermal shield surrounds the fire sleeve. An insulation space is defined between the fire sleeve and the thermal shield to thermally isolate the fire sleeve from conditions external to the thermal shield.

Abstract: A nozzle includes a nozzle body and a cavity defined at least in part by the nozzle body. A plenum extends through the nozzle body into the cavity. At least one passage through the plenum provides fluid communication between the plenum and the cavity. Orifices through the nozzle body and circumferentially spaced around the nozzle body provide fluid communication through the nozzle body. A method for cooling a face of a nozzle having a nozzle body that defines a cavity includes flowing a fuel through the cavity and inserting a plenum through the nozzle body into the cavity. The method further includes flowing a fluid through the plenum so that the fluid impinges on the face of the nozzle to remove heat.

Abstract: The present invention provides in one embodiment a unique gas turbine engine. Another embodiment is a unique gas turbine engine combustion system. Still another embodiment is a unique gas turbine engine combustor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine engine combustion systems and combustors.

Abstract: Disclosed herein are a fuel nozzle of a gas turbine combustor for dimethyl ether (DME, CH3OCH3) and a design method thereof. The fuel nozzle includes a pilot fuel injection hole formed at the center of the nozzle and a plurality of fuel injection ports formed at equiangular positions around the pilot fuel injection hole to inject DME. Each of the fuel injection ports includes a pair of upper and lower DME injection orifices communicating with a fuel-air mixture injection swirler. The upper DME injection orifice becomes gradually wider, whereas the lower DME injection orifice becomes gradually narrower. Thus, DME can be optimally combusted in the gas turbine combustor, thereby achieving cost reduction of power plants, enhancement in reliability of the power plants, and diversification of usable fuel.

Abstract: The lance of a burner includes a body that defines a first duct with first nozzles for injecting a liquid fuel and a second duct with second nozzles for injecting a gaseous fuel. Outlets of the first nozzles are spaced apart from outlets of the second nozzles. The body includes a third duct with third and fourth nozzles for injecting air. The third nozzles surround an axis of the first nozzles and the fourth nozzles surround an axis of the second nozzles.

Abstract: A reheat burner includes a channel with a lance projecting thereinto to inject a fuel over an injection plane perpendicular to a channel longitudinal axis. The channel and lance define a vortex generation zone upstream of the injection plane and a mixing zone downstream of the injection plane in the hot gas direction. The mixing zone has a cross section with diverging side walls in the hot gas direction. The diverging side walls define curved surfaces in the hot gas direction having a constant radius.

Abstract: A fuel delivery system for a turbine engine has at least one fuel injector having an upstream orifice arrangement that produces a first pressure drop of flowing fuel and a downstream orifice arrangement that produces a second pressure drop of the flowing fuel. The upstream orifice arrangement or the downstream orifice arrangement includes a nonecylindrical orifice.

Abstract: There are provided a fuel nozzle apparatus, a gas turbine, and a method of controlling a fuel nozzle apparatus, that prevent an occurrence of caulking while meeting the requirements of exhaust gas regulations. There are provided: a fuel pump that feeds fuel; a fuel flow passage formed with a circulation flow passage through which the fuel flows; a flow controller that controls a flow rate of the fuel flowing from the fuel pump into the circulation flow passage; a circulation pump that circulates the fuel within the circulation flow passage; a nozzle section formed with an injection nozzle that injects the fuel from the circulation flow passage to the outside; and an injection controller that controls an injection amount of the fuel to be injected from the injection nozzle.

Abstract: Premix nozzles and gas turbine engine systems involving such nozzles are provided. In this regard, a representative industrial gas turbine engine includes: a combustion section having a nozzle assembly operative to provide a fuel-air mixture for combustion, the nozzle assembly having an array of shuttered nozzles and non-shuttered nozzles; each of the shuttered nozzles being operative in an open position, in which air is directed through the shuttered nozzle for mixing with fuel, and a closed position, in which a reduced amount of air is directed through the shuttered nozzle; each of the shuttered nozzles being inoperative to independently alter an amount of air being directed therethrough.

Abstract: A fuel injector for a secondary fuel nozzle in a gas turbine includes axially oriented air slots and a plurality of fuel injection holes disposed between the air slots. The plurality of fuel injection holes include axially oriented injection holes and radially oriented injection holes such that fuel input through the plurality of fuel injection holes is injected in both a radial direction and an axial direction to mix with air flowing through the air slots.

Abstract: A fuel nozzle for turbine engine includes a cooling shroud located at the downstream end of the fuel nozzle to help cool the downstream end of the fuel nozzle. The cooling shroud surrounds the exterior circumference of the downstream end of the fuel nozzle. A flow of air is admitted into the cooling shroud and the flow of air travels in the downstream direction through a first passageway which covers the exterior of the fuel nozzle. The cooling air flow then turns 180° and travels in the upstream direction through a second passageway which is located concentrically outside the first passageway. The airflow then leaves the upstream end of the cooling shroud and enters the interior of the fuel nozzle.

Abstract: A gas turbine engine is provided and includes a combustor having a first interior in which a first fuel is combustible, a turbine into which products of at least the combustion of the first fuel are receivable, a transition zone, including a second interior in which a second fuel and the products of the combustion of the first fuel are combustible, a plurality of fuel injectors which are configured to supply the second fuel to the second interior in any one of a single axial stage, multiple axial stages, a single axial circumferential stage and multiple axial circumferential stages, a compressor, by which air is supplied to the first and second interiors for the combustion therein, and a control system configured to control relative amounts of the air to the first and second interiors and relative amounts of the first and second fuels supplied to the first and second interiors.

Abstract: A reverse flow annular combustor for a gas turbine includes a pre-vaporizer/pre-mixing region within a dome section, a liquid fuel injection system feeding the pre-mixing region, a combustion region downstream of the pre-vaporizer/pre-mixing region, a means for guaranteeing stable and sustained combustion in the combustion region, and a dilution region downstream of the combustion region.