Abstract: A secondary combustion system for a stage one turbine nozzle. The secondary combustion system may include a supply tube extending into the stage one nozzle, a number of injectors extending from the supply tube to an outer surface of the stage one nozzle, and an air gap surrounding each of the number of injectors.

Abstract: A method for fabricating a unitary distributor is disclosed, the method comprising the steps of determining three-dimensional information of the unitary distributor having at least one flow passage having an arcuate portion located within a distributor ring body having an axis, converting the three-dimensional information into a plurality of slices that each define a cross-sectional layer of the unitary distributor and successively forming each layer of the unitary distributor by fusing a metallic powder using laser energy. A distributor is disclosed comprising a flow passage with an arcuate portion located in a distributor ring body wherein the distributor is made by using a rapid manufacturing process. A unitary fuel distributor comprising a fuel conduit and a distributor is disclosed wherein the unitary fuel distributor is made by using a rapid manufacturing process.

Abstract: A fuel injector for a gas turbine engine includes a stem extending along a longitudinal axis from a proximal end to a distal end. The stem includes one or more fuel tubes configured to deliver a fuel toward the distal end of the stem, and a pilot assembly coupled to the distal end of the stem. The pilot assembly includes one or more components configured to inject fuel into a combustor of the engine. The fuel injector also includes a substantially tubular premix barrel extending along the longitudinal axis from a first end to a second end. The premix barrel is circumferentially disposed about the stem and configured to couple with the combustor at the second end. The fuel injector also includes an annular premix duct formed between the premix barrel and the stem. The premix duct includes an air inlet port at the first end, and a plurality of first orifices located downstream of the air inlet port. The first orifices are configured to inject fuel from the fuel tube into the premix duct.

Abstract: Injection system of a turbojet engine. It comprises a fixed part and a sliding crossmember. The sliding crossmember comprises a flange capable of sliding on the sliding surface of the fixed part. The fixed part comprises antirotation means and the sliding crossmember of the complementary antirotation means to limit the rotation of the sliding crossmember relative to the fixed part. These means consist of at least one cutout formed in the flange and of at least one lug protruding into said at least one cutout of the flange of the sliding crossmember.

Abstract: Disclosed is a fuel nozzle for a combustor which includes a tip section and a tube section. The tube section, in which a first end is connected to the tip section includes an outer tube having a tube axis and connected to the tip section and an inner tube disposed at least partially inside of the outer tube and connected to the tip section. The inner tube includes at least one flexible portion capable of reducing stresses in the fuel nozzle. Also disclosed is a fuel nozzle which includes a tip section, a base section, and a tube section disposed between the tip section and the base section and connected to the tip section at a first end and connected to the base section at a second end. The tube section includes an outer tube having a tube axis and an inner tube disposed at least partially inside of the outer tube and connected to the tip section. At least one rib is disposed on the outer tube and extends radially outwardly from an outer surface of the outer tube.

Abstract: An air-blast fuel injector is disclosed which includes an outer air circuit having an exit portion, an inner air circuit having an outlet configured to direct air toward the exit portion of the outer air circuit, and a fuel circuit radially outboard of the inner air circuit and having an exit communicating with the outer air circuit upstream from the exit portion of the outer air circuit.

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: An axial-flow pyrospin combustor comprises inner and outer combustor liners and a plurality of pyrospin effusion holes. The inner liner is coaxially mounted inside the outer liner, about a central combustor axis. The pyrospin effusion holes are formed in at least one of the outer combustor liner and the inner combustor liner. Each of the pyrospin effusion holes has a down angle and a back angle, which control a global swirl flow about the central axis, and promote film cooling without detachment.

Abstract: A fuel injector for an aero-mechanical injection system for injecting an air/fuel mixture into a turbomachine combustion chamber, the injector comprising a main tubular structure of axis XX? opening out at a downstream end for delivering the air/fuel mixture, a tubular fuel duct that is disposed inside the main structure and that opens out into the main structure via a fuel atomizer plug so as to introduce fuel into the main structure at a pressure PC into the main structure, at least one air feed channel that opens out into the main structure so as to introduce air at a pressure PA therein, and means for injecting into the fuel duct a gas at a pressure PG that is greater than PA and greater than or equal to PC so as to create effervescence in the fuel on being introduced into the main structure.

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: An assembly that includes two components joined by a pre-compressed braze where the compression in the braze is progressively relieved upon relative thermal expansion of the two components. Also disclosed is a process for producing a pre-compressed braze.

Abstract: A fuel manifold of gas turbine engines comprises a tube for fluid communication with a fuel source and for connection with at least one fuel nozzle tip to thereby define a fuel conveying passage. A redundant seal is provided around the tube and covers substantially an entire length thereof to reduce a risk of fuel leakage of the fuel conveying passage.

Abstract: This injector with two valves (7, 10) which open at different fuel pressures so as to establish a primary flow regime and then a secondary flow regime is characterized in that the primary regime is enriched with fuel via an orifice (21) which adds a branch to the primary flow path of the fuel, but which is closed when the secondary valve is displaced so as to prevent enrichment of the fuel in the secondary regime. This invention applies in particular to certain aircraft engines.

Abstract: A lean, low NOx gas-fired can combustor having liquid fuel operation capability includes a housing, a combustor liner disposed within the housing and defining a combustion zone for combusting fuel with air, and a head assembly for joining the longitudinal housing and liner ends and for supporting swirler vanes that admit about 45-55% of a total combustor air flow to the combustion zone. The head assembly includes multiple gaseous fuel injection ports upstream of the swirler vanes, and a single liquid fuel injector positioned along the liner axis for directing atomized liquid fuel into the combustion zone. The head assembly further includes a heat shield disposed between the injector and the combustion zone, and the head assembly is configured to provide a flow of air for cooling the heat shield during all operations of the can combustor.

Abstract: A device for guiding an element in an orifice in a wall of a turbomachine combustion chamber comprises a coaxial ring and bushing mounted one inside the other, the bushing comprising two coaxial annular parts that are fastened to each other by welding and that define a groove for guiding a rim of the ring, welding zones between the parts of the bushing being situated away from the path of the rim of the ring when moved along a privileged transverse direction of ring movement relative to the bushing when the combustion chamber is in operation.

Abstract: The invention relates to a turbomachine combustion chamber comprising a chamber bottom which comprises at least one opening designed to receive a combustion bowl in the axis of which an air and fuel injection device is mounted, said flared bowl comprising downstream a divergent consisting of a double partition delimiting an annular cavity, the first outer partition comprising inlet orifices arranged to cool the second inner partition by impact; the second inner partition comprising outlet orifices; a chamber characterized by the fact that the inlet orifices, distributed in at least two circular rows on the periphery of the divergent, are in staggered rows with the outlet orifices.

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: An outer liner of a combustor of a gas turbine engine includes a plurality of liner regions arranged adjacent one another and separated by boundary lines; and a plurality of air admission holes formed in the liner regions. The air admission holes within each liner region form a V-pattern.

Abstract: An apparatus and method of providing a gas turbine combustor having increased combustion stability and reducing pressure drop across a gas turbine combustor is disclosed. A plurality of vanes is fixed to a flow sleeve radially between the flow sleeve and a combustion liner. The plurality of vanes serve to direct a flow of air entering the region between the flow sleeve and combustion liner in a substantially axial direction, such that components of tangential velocity are removed thereby providing a more uniform flow of air the combustion chamber and reducing the amount of pressure lost due attempting to straighten the airflow by pressure drop alone.

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 fuel injector for a gas turbine engine is disclosed. The fuel injector includes an injector housing extending from a first end to a second end along a longitudinal axis. The second end of the housing is fluidly coupled to a combustor of the turbine engine and the housing includes a liquid fuel gallery annularly disposed about the longitudinal axis. The fuel injector also includes a stem extending longitudinally from the first end of the housing to a third end. The stem includes a liquid tube configured to deliver liquid fuel to the fuel injector. The fuel injector also includes an annular shell extending along the longitudinal axis from the first end to the third end and circumferentially disposed about the stem. The fuel injector further includes an insulating air shroud formed inside the shell. The air shroud includes a layer of air between the shell and the stem.

Abstract: A gas turbine engine fuel nozzle includes an axis of symmetry extending therethrough, the nozzle body including a first passage extending coaxially therethrough, a second passage, and a third passage, the second passage circumscribing the first passage, the third passage formed radially outward of the second passage, and a nozzle tip coupled to the nozzle body, the nozzle tip including at least one primary discharge opening in flow communication with the first passage, at least one secondary discharge opening in flow communication with the second passage, and at least one tertiary discharge opening in flow communication with the third passage.

Abstract: A fuel system includes a fuel deoxygenator for removing oxygen from a liquid fuel. A vaporizer is in fluid communication with the fuel deoxygenator. The vaporizer vaporizes at least a portion of the liquid fuel to produce vaporized fuel. At least a portion of the vaporized fuel pre-mixes with oxidizer to reduce formation of undesirable emissions.

Abstract: A fuel injection system comprises a fuel conveying member and a nozzle tip assembly threadably engaged thereto. A pair of sealing elements is engaged in a fuel passage between the fuel conveying member and the nozzle tip for sealing the junction therebetween. The pair of sealing elements includes a first and a second sealing element located proximate each other and having different cross-sectional shape.

Abstract: In order to allow dual fuel operation, that is to say gas and liquid fuel, a fuel injection arrangement and an assembly are provided in which an injection conduit provides constriction to an airflow towards an outlet. In an assembly a plurality of arrangements including liquid fuel jets are arranged to such that through high angle incidents by the airflow appropriate vaporization and droplet formation is achieved for a fuel/air mixture combustible in a combustor of a turbine engine. At low liquid fuel flow rates below a pre-determined level the liquid fuel is atomised and entrained by the airflow. At high liquid fuel flow rates the fuel splash impinges upon a wash surface to develop a wash film which is again vaporised and atomised about an edge for a fuel/air mixture capable of combustion.

Abstract: A fuel conveying member in a fuel system of a gas turbine engine includes a heat pipe disposed in heat transfer communication with a fuel flow passage of the fuel conveying member, such that heat is transferred by the heat pipe from a high temperature region to a lower temperature region of the fuel flow passage. The heat pipe has a constant mass or working fluid enclosed within a sealed chamber in heat transfer communication with fuel in said fuel flow passage, said working fluid being self-circulating in a fluid flow having a liquid portion and a vapor portion, said working fluid passively boiling, flowing and condensing to transfer heat from said high temperature region to said lower temperature region.

Abstract: A joint for a gas turbine fuel injector, the joint comprises a housing for supporting a fuel injector, a first tube terminating within the housing and a second tube. The housing at least partially encloses an adaptor having a bore, with one end of the first tube and one end of the second tube being located within opposing ends of the bore. First sealing means are provided between the outside of the first tube and the adaptor bore, second sealing means are provided between the outside of the second tube and the adaptor bore and third sealing means are provided between the outside of the adaptor and the housing. The joint is particularly suitable for conduits carrying a relatively cold fluid mounted in a relatively hot housing.

Abstract: A method for operating a combustor is provided. The method includes supplying a predetermined amount of a first gaseous fuel to the combustor, wherein the first gaseous fuel has a first Modified Wobbe Index (MWI) and a first fuel reactivity, and supplying a predetermined amount of a second gaseous fuel to the combustor, wherein the second gaseous fuel has a second MWI that is lower than the first MWI and a second fuel reactivity that is higher than the first fuel reactivity. The method also includes mixing the first and second gaseous fuels together to form a blended gaseous fuel, and injecting the blended gaseous fuel into the combustor.

Abstract: A lean premix burner for a gas-turbine engine includes an annular center body (3) with a conically flaring fuel film applicator (11) supplied with fuel via an annular distributor chamber (14) and fuel channels (12) as well as air ducts (7, 9) with swirler elements (8, 10) provided on the outer and inner circumference. A fuel prefilmer lip (15) is attached to the fuel film applicator (11), with a flow area of a portion of the annular air ducts (7, 9) downstream of the swirler elements (8, 10) decreasing to accelerate the air swirled in correspondence with the airflow direction. By this, the fuel is transported positively without interim separation and without the occurrence of compressive oscillation—to a defined flow break-away edge (16), providing for a good mixture, high combustion efficiency and reduced formation of nitrogen oxide.

Abstract: A method of operating a gas turbine engine is provided. The method includes discharging pilot fuel into a combustion chamber from a nozzle through at least one pilot fuel outlet defined in a tip of the nozzle, discharging steam from the nozzle through a plurality of steam outlets that are spaced circumferentially about the plurality of pilot fuel outlets, and discharging primary fuel from the nozzle through a plurality of primary fuel outlets that are circumferentially aligned with the plurality of steam outlets. To facilitate mixing the primary fuel with the steam, the primary fuel is discharged from the nozzle tip at an oblique angle with respect to a centerline extending through the nozzle tip.

Abstract: A fuel injector for a gas turbine engine includes a support member and a nozzle tip disposed on the support member with a deformable, removable metallic or non-metallic seal member disposed between facing sealing surfaces of the support member and a shroud member of the nozzle tip. The deformable seal member includes one or more fuel passages through which fuel in one or more fuel circuits flows with the seal member sealingly separating the first and second fuel circuits and preventing external fuel leakage from the nozzle tip. The nozzle tip can include an inner fuel swirler that includes one or more swirler slots wherein each slot is so angled in an axial direction toward the discharge orifice of the nozzle tip assembly that operation of the nozzle tip assembly is unaffected by changes in fuel viscosity.

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: A fuel conveying member for an engine, the fuel conveying member having heating means integrated therein to permit pyrolysis of carbonaceous deposits.

Abstract: A gas fuel injector includes a first header plate, a second header plate, and a plurality of venturi tubes. The second header plate is spaced downstream from the first header plate. The plurality of venturi tubes is sealably secured to the first and second header plates. Each venturi tube of the plurality of venturi tubes is defined by a plurality of fixable components.

Abstract: An aerodynamic injection system for injecting an air/fuel mixture into a turbomachine combustion chamber comprises a tubular structure of axis XX? that opens out at a downstream end for delivering the air/fuel mixture, at least one air feed channel that opens out into the structure so as to introduce air at a pressure PA therein, an annular fuel passage that is formed in the structure around its axis XX?, that is connected to at least one fuel feed channel in which there flows fuel at a pressure PC, and that opens out at a downstream end into the structure, and means for injecting gas into the at least one fuel feed channel, said gas being at a pressure PG that is greater than PA and greater than or equal to PC so as to create effervescence in the fuel on being introduced into the structure.

Abstract: A staged fuel injector is disclosed that includes a main fuel circuit for delivering fuel to a main fuel atomizer and a pilot fuel circuit for delivering fuel to a pilot fuel atomizer located radially inward of the main fuel atomizer. The pilot fuel circuit is in close proximity to the main fuel circuit enroute to the pilot fuel atomizer so that the pilot fuel flow cools stagnant fuel located within the main fuel circuit during low engine power operation to prevent coking.

Abstract: A tip turbine engine (40) provides a peripheral combustor (30) with a more efficient combustion path through the combustor and through the tip turbine blades (34). In the combustor, the core airflow is received generally axially from compressor chambers in hollow fan blades (28) and then turned radially outwardly into a combustion chamber (112), where it is then mixed with the fuel and ignited. The combustor has a combustion path extending axially from a forward end of its combustion chamber through a combustion chamber outlet (122) and through a turbine (32) mounted to the fan. Thus, when the core airflow begins to expand in a high-energy gas stream, it has a substantially axial path from the combustion chamber through the turbine.

Abstract: A bi-propellant rocket engine may include a primary propellant flowing in a central passageway, a secondary propellant flowing in a secondary passageway generally coaxial with central passageway and a pintle tip having a central chamber sidewall coaxial with the primary passageway and surrounding a central chamber, the central chamber sidewall having a first plurality of apertures there through so that some of the primary propellant exits the central chamber transverse to the flow of the secondary propellant in the secondary passageway. The pintle tip may have a secondary chamber sidewall, substantially thicker than the primary chamber sidewall, surrounding a secondary chamber downstream of and in fluid communication with the primary chamber, the secondary chamber sidewall having a second plurality of apertures there through so that some of the primary propellant exits the secondary chamber transverse to the flow of the secondary propellant in the secondary passageway.

Abstract: Turbomachine with an annular combustion chamber (46) comprising at its upstream end an annular chamber bottom (28), and attached at its downstream end by annular flanges (24, 26) to a radially inner casing (14) and outer casing (18), wherein the upstream end of the chamber (46) is connected to the inner casing (14) or outer casing (18) by elastically deformable means (48, 49) of variable stiffness.

Abstract: A method and a pre-shaped workpiece for making at least a portion of a gas turbine fuel nozzle are disclosed having opposed gripping tabs which can be secured to machine tool holders during manufacturing by a machine tool.

Abstract: A fuel injector for a gas turbine engine of an aircraft, and more particularly a novel and unique heatshield structure for a fuel nozzle wherein a labyrinth seal is uniquely provided in the nozzle to isolate a portion of an insulating gap from an interface whereat fuel may enter the insulating gap, and the insulating gap is provided with a positive purge flow for forcing vapors out of the insulating gap.

Abstract: A fuel injector stick for a gas turbine includes: a body of a length, L, the body including an annular shape forming a fuel channel and adapted for insertion into a premixer of the gas turbine; a mounting section for mounting the injector stick to the gas turbine; and a nozzle for injecting fuel into the premixer of the gas turbine. A gas turbine combustor and a method for changing fuel injector sticks are provided.

Abstract: A fuel injector for a gas turbine combustor is disclosed which includes a feed arm having a flange for mounting the injector within the combustor and a fuel nozzle depending from the feed arm for injecting fuel into the combustor for combustion. An optical sensor array is operatively associated with the fuel nozzle for observing combustor flame characteristics. The optical sensor array includes a plurality of sapphire rods positioned to be close enough to the combustor flame to oxidize soot deposits thereon.

Abstract: A combustor assembly including an injector element able to inject and substantially mix a fuel with a vitiated air stream, so that the substantially all of the fuel injected into the vitiated air stream reaches an auto ignition temperature before it combusts. Generally, the injector element expels the fuel from the injector orifice at a rate, a velocity, a trajectory, or a flow geometry able to mix the fuel with the vitiated air. Generally, the fuel is able to mix with the vitiated air in such a manner that the fuel is substantially mixed and no substantially high or substantially low concentrations of fuel exist in the air stream at the time that the fuel reaches the auto ignition temperature.

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 injector nozzle for dispensing fuel in the combustion chamber of a gas turbine engine, comprises a fluid feed conduit having at least one internal channel for the passage of fluid from an inlet end to an outlet end of the fluid feed conduit. The fluid feed conduit has a first annular segment receiving fluid from the inlet end and a second annular segment fluidly connected to receive fluid from the first annular segment at a junction having a circumferential length less than the circumferential lengths of the first and second annular segments. The second annular segment includes fluid dispensing openings to dispense fluid from the conduit, and the first and second annular segments are coaxial and axially separated for relative movement over a major portion of the second segment to accommodate differential thermal expansion.

Abstract: A mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer, and a fuel manifold positioned between the pilot mixer and main mixer. The pilot mixer includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing.

Abstract: An end cap for a fuel injector of a turbine engine is disclosed. The end cap includes an annular first surface including a plurality of perforations. The annular first surface is exposed to a combustion chamber of the turbine engine. The end cap is coupled to an end face of the fuel injector to define an enclosed cavity. The enclosed cavity and the plurality of perforations form an array of Helmholtz resonators.

Abstract: An inlet tube supplies fuel to a manifold of a combustor in a gas turbine engine, the manifold defining a manifold plane. The inlet tube comprises a tube body having a channel providing fluid flow communication between first and second ends of the tube body. At least a portion of the channel defines a tube body plane which is spaced apart from the manifold plane.

Abstract: A turbine fuel supply system is disclosed as including a first sub-system having: a first nozzle for injecting fuel into a combustor of a turbine engine; a first valve controllable to communicate fuel to the first nozzle; and a first fuel manifold for communicating fuel to the first valve from a fuel source; the system further including a controller assembly for raising the pressure of fuel in the first fuel manifold; and the system being characterized in that the first valve is adapted to open in response to a predetermined pressure difference between the first fuel manifold and the pressure in the combustor, thereby allowing fuel to be communicated from the first fuel manifold to the first nozzle. The system may include a first recirculating conduit in fluid communication with the first fuel manifold and the fuel source, the first recirculating conduit allowing fuel not communicated by the first valve to be returned to the fuel source.