Abstract: A mixing region of a fuel nozzle assembly for a combustor in a gas turbine, the assembly including: a gaseous fuel nozzle having a center axis and extending along the center axis, the fuel nozzle including a gaseous fuel passage and a fuel nozzle at a distal end of the passage; an air tube concentric with the fuel nozzle and defining an air passage between the air tube and the fuel nozzle, wherein the air tube includes a distal section extending axially beyond the fuel injection nozzle; a first fuel-air mixing zone defined by and inside the distal section of the air tube, wherein said first fuel-air mixing zone is downstream of the fuel injection nozzle; a flame holder comprising a porous structure with thermal barrier coating and micro swirlers and defining a downstream end of the first fuel-air mixing zone, wherein fuel and air from the first fuel-air mixing zone pass through the porous structure of the flame holder and into a combustion zone of the combustor.

Abstract: A turbofan gas turbine engine augmenter includes a fuel/air swirler disposed between an axially extending bypass flowpath and an axially extending exhaust flowpath. A swirler inlet is axially open to and positioned substantially normal to the bypass flowpath and a swirler outlet is open to and positioned substantially parallel to the exhaust flowpath. A swirl chamber within the fuel/air swirler is between the swirler inlet and the swirler outlet. A swirl axis of the fuel/air swirler extends through the swirler outlet substantially normal to the exhaust flowpath. An air swirler may be centered about the swirl axis within the fuel/air swirler. The air swirler may be a louvered or have a plurality of swirling vanes. The swirler inlet may be radially offset with respect to the swirl axis. An air scoop may lead from the swirler inlet to a rounded swirler housing within which the air swirler is disposed.

Abstract: A lance (3) for introducing a fuel into a mixing path (2) of a burner (1) which is provided with a swirler for producing a swirled combustion air flow, especially a burner of a turbine plant, includes a lance head (5), a fuel feed (8) and an oxidator feed (10). The lance head (5) projects into the mixing path (2), has at least one fuel nozzle (6) for injection of the fuel into the mixing path (2), and also has at least one oxidator outlet orifice (7) for introducing an oxidator into the mixing path (2). The fuel feed (8) extends in the lance (3) and is connected to the at least one fuel nozzle (6). The oxidator feed (10) extends in the lance head (5) and is connected to the at least one oxidator outlet orifice (7).

Abstract: A combustor for a turbine engine is disclosed. The combustor may have a can-like dilution zone, a primary combustion zone, and a secondary combustion zone. The primary combustion zone may be disposed radially about the can-like dilution zone. The secondary combustion zone may be disposed at an end of the can-like dilution zone to fluidly communicate the primary combustion zone and the can-like dilution zone.

Abstract: An apparatus for producing power from a source of liquid fuel. The apparatus includes at least one capillary flow passage, the at least one capillary flow passage having an inlet end and an outlet end, the inlet end in fluid communication with the source of liquid fuel, a heat source arranged along the at least one capillary flow passage, the heat source operable to heat the liquid fuel in the at least one capillary flow passage to a level sufficient to change at least a portion thereof from a liquid state to a vapor state and deliver a stream of substantially vaporized fuel from the outlet end of the at least one capillary flow passage, a combustion chamber for combusting the stream of substantially vaporized fuel and air, the combustion chamber in communication with the outlet end of the at least one capillary flow passage and a conversion device operable to convert heat released by combustion in the combustion chamber into mechanical or electrical power.

Abstract: To meet emissions standards, many gas turbine engines use some form of lean, pre-mixed combustion systems. These systems may lead to combustion oscillations or other instabilities. Jet combustion techniques provide a stable alternative lean, pre-mixed combustion system. The present invention presents a jet combustion system that includes a first cylinder having a first array of orifices. A second cylinder is positioned coaxially with the first cylinder. The second cylinder has a second array of orifices offset from the first array of orifices.

Abstract: The invention relates to a combustion unit for combusting a liquid fuel. The combustion unit has a fuel inlet, an air inlet and a flue gas outlet which are connected to a combustion chamber for combusting the fuel, wherein the fuel inlet is connected to at least one explosion atomizing unit which is disposed and adapted such that atomized fuel fragments due to gas formation in the atomized fuel. The explosion atomizing unit is preferably an explosion swirl atomizing unit to a system for generating power having at least one gas turbine, at least one compression device driven by the gas turbine and at least one such combustion unit.

Abstract: In a combustion apparatus (10), in particular for driving gas turbines, in which combustion apparatus (10) a gaseous fuel in a burner (11) is sprayed through a plurality of separate fuel-injection devices (15, 16) into a gas flow containing combustion air, and the resulting mixture flows into a combustion chamber (12) for combustion and burns there, the acoustic impedance or stiffness of the fuel-injection devices (15, 16) is selected to be different in order to avoid thermoacoustic combustion instabilities.

Abstract: A fuel air mixer for a gas turbine engine combustor having a longitudinal axis therethrough, wherein the fuel air mixer is configured for use in a dome oriented substantially radial to the longitudinal axis. The fuel air mixer includes a fuel injection assembly having a first end, a second end, a fuel passage extending therethrough, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The fuel air mixer also includes a mixer assembly having a first end, a second end, a cavity formed in a central portion thereof, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The mixer assembly is configured to receive the fuel injection assembly in the cavity so that the fuel injection assembly and the mixer assembly are able to be connected to an outer casing of the combustor by means of the respective flange portions.

Abstract: This invention relates to an apparatus and method for increasing the reactivity of a fuel/air mixture prior to homogenous combustion of the mixture. More specifically, this invention is a pilot for a gas turbine combustor which utilizes the heat of combustion within the pilot to increase the reactivity of a portion of the fuel/air mixture utilized by the pilot.

Abstract: Lean premixed combustion of a hydrocarbon fuel and air is combined with lean direct injection of hydrocarbon fuel and carrier fluid such as air or inert gas or a mixture of air and inert gas into a combustor downstream of the premixed reaction zone in order to achieve extremely low levels of emissions of oxides of nitrogen at the high combustor exit temperatures required by advanced heavy duty industrial gas turbines. One or more premixing fuel nozzles are used to supply a lean mixture of hydrocarbon fuel and air to the main or primary reaction zone of a gas turbine combustor. This lean fuel/air mixture has an adiabatic flame temperature below the temperature that would result in substantial thermal NOx formation.

Abstract: This invention relates to an apparatus and method for increasing the reactivity of a fuel/air mixture prior to homogenous combustion of the mixture. More specifically, this invention is a pilot for a gas turbine combustor which utilizes the heat of combustion within the pilot to increase the reactivity of a portion of the fuel/air mixture utilized by the pilot.

Abstract: An ultra-low NOx gas turbine combustor having a dual fuel capability. The combustor has a plurality of venturis in flow communication with a corresponding number of fuel lances such that fuel injected from each lance into a corresponding venturi is mixed with high velocity air also flowing into the venturis. A vee-gutter flame holder is positioned downstream of the array of venturis for holding a combustion flame within the combustor at stable conditions. Due to the high velocities of the air-fuel mixture in the vicinity of the flame-holder, as a result of the fuel-air mixing in the venturis, destructive flash back conditions are obviated. A stable flame within the combustor provides for minimum acoustic disturbance and low overall pressure losses within the combustor provide for high operating efficiency levels of the turbine apparatus. Burning of the fuel in air at lean mixtures allows for minimal operation of a conventional pilot nozzle assembly so as to reduce the undesirable formation of NOx.

Abstract: A gas turbine combustor includes a diffusion combustor arranged at an axis portion of a combustion chamber for effecting diffusion combustion, a plurality of first premixing combustors arranged at an outer periphery of the diffusion combustor for effecting premixed combustion, each of the plurality of first premixing combustors being formed so that a mixture outlet end thereof projects more downstream than a fuel outlet end of the diffusion combustor, and a plurality of second premixing combustors each formed so that a mixture outlet thereof projects more downstream than the mixture outlet end of the first combustor, wherein the first premixing combustors and the second premixing combustors are arranged alternately at the outer periphery of the diffusion combustor, and wherein a swirler is provided on the first premixing combustor in the vicinity of the mixture outlet end thereof for swirling mixture.

Abstract: A combustor for a gas turbine includes a diffusion flame combustion zone, a catalytic combustion zone and a post-catalytic combustion zone. At start-up or low-load levels, fuel and compressor discharge air are supplied to the diffusion flame combustion zone to provide combustion products for the turbine. At mid-range operating conditions, the products of combustion from the diffusion flame combustion zone are mixed with additional hydrocarbon fuel for combustion in the presence of a catalyst in the catalytic combustion zone. Because the fuel/air mixture in the catalytic reactor bed is lean, the combustion reaction temperature is too low to produce thermal NO.sub.x. Under high-load conditions, a lean direct injection of fuel/air is provided in a post-catalytic combustion zone where auto ignition occurs with the reactions going to completion in the transition between the combustor and turbine section.

Abstract: This invention pertains to small air breathing gas turbine engines (1). Engines of the invention employ improved combustion systems, including improved fuel injection systems and improved combustor design. These improvements enable the engine to idle at speeds as low as one-tenth the maximum operating speed of the engine, which gives the engine an increased range of operating speeds. The improved designs generally provide for pre-mix assemblies intimate with the fuel injectors, a typically flame-free primary "A" mixing zone in the combustion chamber, and a primary "B" combustion zone, in addition to the secondary combustion zone and the dilution zone. In preferred embodiments, the bearing assembly is located between the compressor and the turbine wheel. A cooling cavity is provided between the turbine wheel and the bearing assembly, thus protecting the bearings from the turbine heat.

Abstract: Combustion-induced instabilities are minimized in gas turbine combustors by incorporating one or more Helmholtz resonators into the combustor. First and second plates located in the head end of the combustor casing define one cavity, and a sleeve located between the casing and the liner defines another cavity. Each of the two cavities is connected to the combustion chamber by one or more throats, thus forming Helmholtz resonators. The throats of each resonator can be tubes of different lengths and/or different cross-sectional areas to provide dynamics suppression over a broad band of frequencies. The throats can also be arranged such that each throat is associated with a different portion of its respective cavity, each cavity portion having a different volume.

Abstract: In the case of a combustion chamber which is respectively arranged upstream of and downstream of a fluid-flow machine and which essentially comprises an inflow duct (5) and a downstream premixing and combustion zone (5a), a fuel (11) is injected into the combustion air (4) coming from the fluid-flow machine acting upstream after the combustion air (4) flows through vortex generators (200). The injection (7a, 7b) of the fuel (11) into the premixing and combustion zone (5a) is effected in varying direction and quantity. The hot gases from the combustion of the aforesaid mixture form a temperature-graduated front (8), the minimum temperature of which corresponds fluidically with the base of the blades to be acted upon of the fluid-flow machine (2) . arranged downstream. The fuel (11) can be assisted with assisting air (12).

Abstract: In a combustion chamber (14) of a gas-turbine group which is arranged downstream of a first turbine (11) and upstream of a second turbine (12) and has a burnerless combustion space, extending between the outflow plane of the first turbine (11) and the oncoming-flow plane of the second turbine (12), and a fuel injector (2) which comprises a fuel-nozzle holder (8), extending transversely to the main flow (1) approximately up to the center axis, and a fuel nozzle (9) arranged in the direction of flow in the center axis and connected to the fuel-nozzle holder (8), the combustion chamber (14) working with premixing combustion, and the inlet temperature of the exhaust gas of the first turbine (11) into the combustion chamber (14) being above the self-ignition temperature of the fuel (13), the fuel injector (2) has means for simultaneously evening out the velocity profile and for lowering the temperature of the flow in the wake region.

Abstract: An annular combustor with an air fuel mixer is disclosed including a plurality of mixing tubes having a forward end and an exit end, wherein a longitudinal axis through the exit end for each mixing tube is at both an axial angle and a radial angle to the centerline axis of the mixer. A fuel injector for providing fuel to the forward end of each mixing tube, a fuel manifold in flow communication with each of the fuel injectors, and a fuel supply and control means are provided. High pressure air from a compressor is injected into the mixing tubes and fuel is injected into the mixing tubes from the fuel injectors, wherein the high pressure air and the fuel is uniformly mixed therein so as to produce minimal formation of pollutants when the fuel/air mixture is exhausted out the downstream end of the mixing tubes into the combustor and combusted. The mixing tubes preferably impart a swirl to the fuel/air mixture as it exits into the combustor.

Abstract: A fuel/oxidizer pre-mixing combustion chamber for a turbojet engine is disclosed in which the pre-mixing device is incorporated into the upstream end wall structure of the combustion chamber. The upstream end portion of the combustion chamber, which is in communication with the oxidizer, has a plurality of generally radially extending, "V"-shaped members oriented such that the vertex of the "V" shape faces in an upstream direction. These members extend generally radially between the inner and outer walls which define the boundaries of the combustion chamber. The members define upstream edges and downstream edges which are axially the upstream edges. Fuel injectors extend radially inwardly between these members and are axially positioned between the locations of the upstream and downstream edges so as to spray fuel onto the adjacent sides of each of the members.

Abstract: In a combustor including a premixing type combustion burner which has an atomizer for ejecting a liquid fuel together with combustion air to atomize the liquid fuel, the atomizer is comprised of an inner shell to an inner peripheral surface of which the liquid fuel is supplied, an outer shell defining a passage for the combustion air running substantially straightly between the outer shell itself and an outer peripheral surface of the inner shell, and a swirling-flow guide plate for swirling the combustion air passed into the inner shell, while directing it in a downstream direction. The combustor further includes a resistor abruptly decreased in sectional area downstream and provided substantially downstream of the center of the swirling flow and in the vicinity of an outlet of the premixing type combustion burner for providing a resistance to a premixture ejected from the premixing type combustion burner.

Abstract: A staged combustion apparatus for a gas turbine to lower the concentration of NO.sub.x and CO in the exhaust gases. Combustion occurs in a first and second stage, the first stage being sub-stoichiometric combustion and the second stage being above-stoichiometric combustion. Cyclonic combustion occurs in one of the first and second combustion stages and a swirling pattern of combustion occurs in the other of the combustion stages. Steam injection is provided to further lower the concentration of NO.sub.x in the exhaust emissions from the gas turbine.

Abstract: In order to enhance fuel injection efficiency, and particularly to atomize low fuel flows at high velocity with low fuel pressure, a fuel injection system 24 for a combustor 22 of a turbine engine 20 has an air blast tube 48 and a fuel supply tube 76. The air blast tube 48 is mounted at an acute angle relative to a wall 26 of the combustor 22 and has a first end 50 in communication with the combustor 22 and a second, enlarged end 52 disposed at an acute angle to an axis 54 of the air blast tube 48 in communication with a source of compressed air externally of the wall 26 of the combustor 22. The air blast tube 48 is operable to deliver compressed air from the source into the combustor 22.

Abstract: A burner for a gas turbine combustor has a perforated plate facing the combustor and a plurality of gaseous fuel premixing tubes conveying a lean premixed fuel through the plate. A gas pilot extending through the plate projects jets of fuel parallel to the plate between the tube locations nearest the pilot. Ignition stability at turndown is maintained.

Abstract: Difficulties in achieving reliable starts in gas turbine engines operating at high altitude are avoided in a gas turbine engine including a rotary compressor (10), a rotary turbine wheel (12) coupled to the compressor (10) to drive the same, and a nozzle (42) for directing gases of combustion against the turbine wheel (12). An annular combustor (34) has an outlet (40) connected to the nozzle (42) and an opposed dome (38) that is axially spaced from the outlet (40). At least three sets of air injection openings (80, 92, 100) are axially spaced from one another with one set (82) in close proximity to the dome (38). Fuel injectors (86, 94) are associated with two of the sets (80, 92, 100), including the one set (80) and another of the sets (92) that is nearest the one set (80).

Abstract: Difficulties with fuel atomization at low fuel flow rates in gas turbine fuel injection systems may be overcome by use of an impingement injector (52) including a cylindrical orifice (64, 94, 114) that is smooth, straight and uninterrupted by burrs or other disturbances having an exit opening (68, 96, 118) transverse to the axis of the orifice (64, 94, 114) and facing an impingement surface (72, 101, 126) that is planar and spaced from the exit opening (68, 96, 118).

Abstract: In order to reduce emissions of NO.sub.x, CO and unburned hydrocarbons, the flame temperature must be kept between 2500.degree.-2800.degree. F. Premixing combustor tubes are used to premix the fuel and air before the mixture is burned in the combustion chamber. The tubes are flashback resistant due to the flow of the fuel and air mixture through them. If it is desired to reduce the load operation of the combustor to substantially below 100% of its maximum load operation, as is the case during off-peak hours, the fuel and air flow rate can be regulated to reduce the load operation while still keeping the low emissions and the low flame temperature.

Abstract: In order to efficiently atomize a fuel/air mixture to reduce smoke, a fuel injection system (26) for a combustor (12) of a turbine engine (10) includes an air blast tube (16), a fuel supply passage (42), and an impingement surface (30). The air blast tube (16) is mounted in relation to an opening (18) in a wall (20) of the combustor (12) such that a first end (22) of the tube (16) is in communication with the combustor (12) and a second end (24) of the tube (16) is in communication with a source of compressed air externally of the combustor (12) to facilitate the delivery of compressed air from the source into the combustor (12).

Abstract: In order to facilitate operation in a wide variety of operating conditions, and to enhance ultra high altitude starting capability while eliminating start injectors, a radial turbine engine (10) includes a pair of fuel injection zones (24 and 26). The radial turbine engine (10) also includes a turbine wheel (12) coupled to a rotary compressor (14) for axially driven movement, an annular nozzle (16) for directing gases of combustion radially at the turbine wheel (12), and an annular combustor (17). The annular combustor (17) defines an annular combustion space disposed about the turbine wheel (12) and in fluid communication with both the compressor (14) and the nozzle (16), and it receives fuel from a source and air from the compressor (14) which are combusted in the combustion space to generate the gases of combustion.

Abstract: In order to provide a simple, low-cost fuel injection system, particularly for a dual zoned annular combustor (10), a single fuel manifold (42) is utilized having first and second manifold portions (42a and 42b) each with a respective plurality of openings (44a and 44b) through which fuel may flow into first and second fuel injection zones (34 and 36). The annular combustor (10) typically includes a combustor housing (12) comprising an axially extending sleeve (14) disposed about a longitudinally extending axis (16) and an annular liner (18) defining a combustion chamber (26) and being spaced from the housing (12) and the sleeve (14) to form a compressed air flow path (28) therebetween.

Abstract: A combustor for a gas turbine engine includes divergent mixing cones disposed substantially within the combustion chamber proper to provide a flow restriction which separates the combustion chamber into primary and secondary combustion zones. Placement of the mixing cones within the chamber enhances vaporization of the fuel and permits combustion to take place in the primary zone at flame temperatures below the stoichiometric temperature thereby reducing formation of nitrous oxides. The mixing cones have external cooling shrouds to prevent autoignition, and the mixing cones for the primary zone provide tangential swirl of the vaporized fuel/air charge in a direction opposite that of the secondary mixing cones. The mixing cones together with an associate fuel nozzle sub-assembly form an integrated unit separable from the combustor for calibration and setting of the fuel/air ratio.

Abstract: The high cost of a fuel injection system in a turbine engine may be substantially reduced by utilizing a combination manifold/fuel injector (80) that is formed of a flattened tube (82) of annular shape and disposed within the case (58) of the engine in substantially surrounding relation to an annular combustor (44). The flattened tube includes fuel injecting apertures (88, 100) which need not be precision formed but which are aligned with the flared ends (74) of inlet tubes (70) extending into the combustor (44) from the plenum defined by the space between the combustor (44) and the case (58).

Abstract: A gas turbine engine 10 having a rotor 12 with turbine blades 14 and a nozzle 16 adjacent the turbine blades 14 for directing hot gases of combustion at the turbine blades 14. The engine 10 also includes an annular combustor 18 about the rotor 12 and an annular combustor housing 30 substantially surrounding the combustor 18 in generally concentric spaced relation. A fuel injection system is operatively associated with the combustor 18 for injecting fuel into a combustion space 28 to be combusted with air from a compressor 34. The fuel injection system may include a plurality of circumferentially spaced fuel injectors 68 disposed in an outer wall 20 of the combustor 18 together with a generally oval shaped manifold 70 in fluid communication with a primary fuel source and the fuel injectors 68. A turbine shrould 36 is provided to extend radially outward from the rotor to the outer wall of the combustor on the side of the nozzle 16 opposite the combustion space 28.

Abstract: Poor atomization of fuel at low fuel flows in a gas turbine engine may be overcome in an engine including a rotary compressor (14), a turbine wheel (22) coupled to the compressor (14) to drive the same and a nozzle of (26) for directing gases of combustion at the turbine wheel (22). An annular combustor (34) defines a combustion annulus about the turbine wheel (22) and has an outlet (30) connected to the nozzle (26). A plurality of air blast tubes (50) include exit openings (52) circumferentially spaced about the combustor (32) and each has an axis (54) that is generally circumferentially directed with respect to the combustion annulus. A fuel discharge orifice (64) is located within each air blast tube (50) and a flange (66) is aligned with each orifice (64) in the path of fuel being discharged therefrom and at a substantial acute angle (.theta.) to the axis (54).

Abstract: A fuel vaporizer is formed from two concentric tubes. A mixture of fuel and air having a high proportion of fuel passes along the annular space between the two tubes while air alone passes along the inner tube.

Abstract: A gas turbine engine fuel burner suitable for burning diesel fuel oil comprises an annular body and a center body coaxially located within the annular body so that an annular flow path is defined between them. At the upstream end of the burner an annular fuel manifold cooperates with the annular body and the center body to define two coaxial passages which direct air into the annular flow path with minimal turbulence. The low level of turbulence within the fuel burner reduces the possibility of spontaneous combustion occurring within the burner. The center body is hollow and is provided at its downstream end with an end cap which cooperates with the downstream end of the annular body to define a radial fuel/air mixture outlet.

Abstract: An improved gas turbine combustor is provided which reduces nitric oxide emissions through premixing of the fuel gas and air and feeding the mixture through a venturi, providing an air cooled passage around the venturi, and extending the passage downstream toward the combustion zone to optimize the stability of the combustion and reduce NO.sub.x and CO emissions.

Abstract: The expense of fabricating an annular combustor (10) for a gas turbine is minimized by providing a combustor housing (12) including an axially extending sleeve (14) and an annular liner (18) disposed within the housing (12) and about the sleeve (14). The annular liner (18) has concentric inner and outer axially elongated walls (20, 22) spaced from the sleeve (14) and the housing (12), respectively, and also has a radially extending wall (24) spaced from the housing (12) and interconnecting the inner and outer walls (20, 22) at one end to define a combustion chamber (26). The liner (18) to spaced from the housing (12) the sleeve (14) to define a compressed air flow path (28) extending from a radially outer compressed air inlet (30) to a radially inner compressed air outlet (32) in communication with the combustion chamber (26) axially remote from the radially extending wall (24).

Abstract: A combustor for a gas turbine engine includes divergent mixing cones disposed within the combustion chamber proper to provide a flow restriction which separates the combustion chamber into primary and secondary combustion zones. Placement of the mixing cones within the chamber enhances vaporization of the fuel and permits combustion to take place in the primary zone at flame temperature below the stoichiometric temperature thereby reducing formation of nitrous oxides.

Abstract: Improved fuel atomization for turbine engines operating at low fuel flows and at high altitudes is accomplished in an engine 10 having an annular combustor 12 by utilizing fuel injectors 18 provided with fuel supply tubes 24 disposed within air tubes 20 wherein the fuel supply tubes 24 each have an exit orifice 26 internally of the corresponding air tube 20 and upstream of an exit orifice 22 thereof with a fuel impingement surface 28 being provided within the air tube 20 in confronting relation to the exit orifice 26 of the fuel supply tube 24 to produce a conical fuel film 36 subjected to pressurized air to enhance fuel atomization.

Abstract: A gas turbine combustor including a first-stage combustion chamber being arranged to serve as a premix chamber for a second-stage combustion chamber and an auxiliary burner provided in the first-stage combustion chamber and arranged to effect, when fired, combustion and holding of a flame in the first-stage combustion chamber and to effect, when extinguished, feed a flame from the first-stage combustion chamber to the second-stage combustion chamber; and a method of driving the gas turbine combustor including a step of preparing the auxiliary burner in the first stage combustion chamber, the auxiliary burner being sparked when said combustion is fired, and causing the first-stage combustion camber to serve as the premixing chamber for the second-stage combustion chamber by extinguished the auxiliary burner.

Abstract: A fuel injector 46 for a turbine engine includes an elongated, generally cylindrical metal casting 76 having a cylindrical surface 78 terminating in an end 84 having a frustoconical surface 86. A bore 88 having an axis 89 normal to the frustoconical surface 86 extends through the casting 76 to provide an air inlet 90 in the cylindrical surface 78 and a fuel and air outlet 94 in the frustoconical surface 86. The bore 88 is narrowed at the outlet to provide a constriction 110 thereat. A curved tube 96 of substantially lesser diameter than the bore 88 is located within the bore 88 to serve as a fuel injecting tube and has an open end 100 on and normal to the axis 89 and located in close adjacency to the constriction 110 without increasing the resistance to air flow through the constriction 110.

Abstract: A gas trubine combustor includes a main body with a combustion portion into which mixture of fuel and air is supplied. The gas mixture is burned through a catalytic reaction at a catalyst body arranged in the main body. The gas mixture passed through the catalyst body is introduced into a gas phase combustion portion provided in the main body. Between the catalyst body and the gas phase combustion portion is arranged a dividing unit which has a plurality of branch passages. The gas mixture passed through the catalyst body is divided by the dividing unit into a plurality of gas streams flowing through the branch passages. Each of the gas streams is mixed with fuel supplied from a fuel supply tube.

Abstract: Excellent fuel atomization in a turbine engine may be obtained with fuel injectors 46, 47 including elongated, laminar discharge orifices 72, 100 and impingement surfaces 76, 102 disposed in the path of fuel 78, 106 being discharged through the discharge orifices 72, 100. Preferably, the fuel 78, 106 is discharged as a flat spray generally tangentially to the annular combustion space 40 of an annular combustor 26.

Abstract: A fuel injector 46 for a turbine engine includes an elongated, generally cylindrical metal housing 76 having a cylindrical surface 78 terminating in an end 84 having a frustoconical surface 86. A bore 88 having an axis 89 normal to the frustoconical surface 86 extends through the housing 76 to provide an air inlet 90 in the cylindrical surface 78 and a fuel and air outlet 94 in the frustoconical surface 86. The bore 88 is narrowed at the outlet to provide a constriction 110 thereat. A curved tube 96 of substantially lesser diameter than the bore 88 is located within the bore 88 to serve as a fuel injecting tube and has an open end 100 on and normal to the axis 89 and located in close adjacency to the constriction 110 without increasing the resistance to air flow through the constriction 110.

Abstract: A domestic or industrial fuel spray combustion device which includes a cylinder-shaped swirl chamber into which air streams are introduced; a fuel injection atomizer for spraying fuel to the swirl chamber; an ignition plug; a first combustion cylinder connected through a choke to the swirl chamber, the first combustion cylinder being coaxial with the swirl chamber; and being larger than an equal to the swirl chamber in diameter and a second combustion cylinder connected through a choke to the first combustion cylinder, the second combustion cylinder being coaxial with the first combustion chamber and being equal to or smaller than the first combustion cylinder in diameter. This structure provides a compact combustion device which enables stable combustion even with a small amount of fuel.

Abstract: A fuel spraybar (16) having an elongated, closed housing (26) discharges fuel from a plurality of openings (32) into an exhaust gas stream (12). A first fuel conduit (34) conducts liquid fuel through the housing (26) and into an annular volume (38) formed therebetween. The fuel, heated by contact with the housing (26) is vaporized in the annulus (38) for preferentially discharging liquid fuel (18) in an axial pilot zone (40) within the gas stream (12).

Abstract: A fuel vaporizer for a gas turbine engine combustion chamber comprises a hollow cylindrical stem open at one end and branching into two radial branch ducts at the other end which terminate in outlets facing towards the inlet end. In order to ensure more even cooling of the vaporizer stem, the fuel injector which lies inside the stem has two outlets which direct fuel onto those parts of the interior wall of the stem which lie between the branch passages, and a baffle divides the interior of the stem to ensure that each branch passage is fed with fuel from only one fuel injector, and each fuel injector supplies only one branch passage.

Abstract: An exhaust purification apparatus includes a regenerative burner for repetitively regenerating a trap element located in the exhaust system of an internal combustion engine. The regenerative burner has a combustion chamber adapted to permit flow of exhaust gas from the engine exhaust conduit to the trap element. A hollow member is disposed within the combustion chamber to define an evaporation chamber therein. The hollow member has a side wall formed with a number of flame holes to permit fluid flow from the evaporation chamber into the combustion chamber. The evaporation chamber has a supply of air-fuel mixture through a mixture conduit having an end terminating in a discharge outlet opening into the evaporation chamber. The mixture conduit has a portion extending through the combustion chamber to promote fuel evaporation during a trap regeneration operation.