Abstract: A low NOx combustor and method improve dynamic stability of a combustion flame fed by a fuel and air mixture. The combustor includes a chamber having a dome at one end thereof to which are joined a plurality of premixers. Each premixer includes a duct with a swirler therein for swirling air, and a plurality of fuel injectors for injecting fuel into the swirled air for flow into the combustion chamber to generate a combustion flame therein. The fuel injectors are axially staged at different axial distances from the dome to uncouple the fuel from combustion to reduce dynamic pressure amplitude of the combustion flame.

Abstract: A gas turbine engine carburetor includes a fuel injector having a plurality of circumferentially spaced apart fuel outlets. A plurality of air swirlers (24) are circumferentially spaced apart around the fuel injector (22), with each swirler having a fuel inlet (24b) disposed in flow communication with a respective one of the fuel outlets for receiving fuel therefrom. Each swirler includes a plurality of swirl vanes to swirl air for mixing with the fuel from the fuel inlets. The multi-swirler carburetor having a common fuel injector increases injection points for reducing NO.sub.x emissions.

Abstract: In a method of operating a plant with staged combustion, the first combustion stage (1a) is operated with a fuel/air mixture (3) whose air coefficient is larger than the overall air coefficient of the combustion system. The hot combustion gases (5) from the first combustion stage (1a) are mixed with an additional fuel/air mixture (4) whose air coefficient is smaller than the overall air coefficient of the combustion system, before the further combustion in the second stage (2a) takes place. Since hot-gas backmixing is no longer required in the second stage (2a) for the flame stabilization, this combined mixture burns without the formation of further NOx emissions.

Abstract: In a pilot burner fuel nozzle for a premixed type combustor, diffusion pilot fuel injection ports (fuel holes 101) are disposed unevenly in the circumferential direction, whereby a nonuniform pilot flame is formed in the circumferential direction in the premixed type combustor. Thereupon, the heat is distributed in the axial direction, so that oscillated combustion does not occur, by which stable combustion can be maintained.

Abstract: A gas turbine combustor having a fuel injection nozzle for diffusion combustion disposed in a central portion of the combustor and an annular premixed nozzle disposed in an outer peripheral portion of the fuel injection nozzle for injecting a mixed gas of fuel and air. The annular premixed nozzle includes an annular premixing chamber and a plurality of premixed fuel nozzles for injecting fuel into the annular premixing chamber. The annular premix nozzle is divided by a partition (108) to form a plurality of premixing chambers each adjacent to the fuel injection nozzle for diffusion combustion. Each of the premixing chambers includes at least one of the premixed fuel nozzles. A mechanism is provided for supplying fuel to a predetermined number of the premixing chambers in a low load operation of the gas turbine and for supplying fuel to all of the premixing chambers in a high load operation.

Abstract: In a combustion chamber of a gas-turbine group which essentially comprises a mixing section (2) for premixing an air/fuel mixture (16) and a downstream combustion space (3), a jump (5) in cross section is provided at the transition between the two said flow sections (2/3). This jump (5) in cross section induces the cross section of flow of the combustion space (3) and at the same time forms outer recirculation zones (10) in the combustion space (3). Flow passages (4) branch off in the end phase of the mixing section (2), which flow passages (4) then lead into the outer recirculation zones (10). A portion (9) of the air/fuel mixture flows out of the mixing section (2) through these flow passages (4) and into the outer recirculation zones (10), the portion being enriched here with an additional fuel (6). This fuel (6) is introduced via a circular line (19) provided with bores (18).

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: A dual fuel injector for injecting liquid and/or gaseous fuel into a gas turbine engine includes a plurality of hollow spoke members for injecting gaseous fuel that are located upstream of a plurality of main air swirling vanes that are in turn upstream of a plurality of air-blast atomizers for injecting main liquid fuel. A pilot fueling arrangement is provided that is capable of starting the gas turbine engine using either gaseous or liquid fuel. The injector includes a labyrinth-shaped cooling passage capable of providing cooling air to cylindrical walls of an injector centerbody.

Abstract: A process and a device for quickly switching over from the premixing operation to the diffusion operation in combustion chambers of a gas turbine upon the extinction of the premixing flame or upon the occurrence of a disturbance that may lead to the extinction of the premixing flame. A pilot gas line, into which a pilot gas control valve is inserted, branches off from the fuel gas supply line before the fuel gas valve. The fuel gas supply line branches into fuel supply lines for both combustion chambers after the fuel gas control valve with displacement measurement. The fuel supply to the diffusion gas line and to the premixing gas line is controlled by installing a quick switch-over valve each with servo valves in the branching fuel supply lines. The combustion chambers are monitored in the area of the flame tube by pressure-measuring devices and by an optical flame-monitoring device.

Abstract: A gas turbine combustion system includes a cylindrical combustor, a plurality of combustion sections in an arrangement spaced apart in an axial direction of the combustor, a plurality of fuel supply lines independently connected to the combustion sections, respectively, premixed fuel supply sections respectively provided for the fuel supply lines for supplying a premixed fuel, a diffusion combustion fuel supply section for supplying a diffusion combustion fuel to the combustion sections, and a control switching over the fuel supply sections to selectively supply either one of the premixed fuel and the diffusion combustion fuel. The premixed fuel at a first combustion stage is burned while the premixed fuel of subsequent stage is ignited by a high-temperature gas generated from combustion of the premixed fuel of a preceding combustion stage.

Abstract: A gas turbine combustion chamber which has primary, secondary and tertiary combustion zones in flow series has a secondary mixing duct and a tertiary mixing duct. The secondary and tertiary mixing ducts reduce in cross-sectional area from their intakes to their outlet apertures to provide an accelerating flow through the mixing ducts to prevent the formation of recirculating zones. Fuel injectors have fuel discharge apertures downstream of any recirculating zones formed at the intakes. The fuel injectors extend across a major portion of the width of the ducts to effectively subdivide the ducts over at least part of the streamwise length of the ducts. The portions of the fuel injectors within the ducts are oval shaped in cross-section and the portions outside the ducts are aerofoil shaped in cross-section. The fuel injectors reduce in dimension perpendicular to the widthwise direction of the duct.

Abstract: The present invention relates generally to a low emission can-annular combustion system for an industrial gas turbine engine to satisfy increasingly stringent environmental requirements. The combustion system of the present invention employs a dual mode combustion technique to meet engine operability requirements and high power emission targets without the use of combustor diluent injection or post combustor exhaust treatment. A lean premix combustion mode is utilized to minimize primary zone combustion temperatures and limit the oxide of nitrogen production during high power engine operation. A pilot-starting auxiliary fueling system is utilized to augment the main premix fueling system. The lean premix combustion mode is enabled by a lean premix dome having a fixed axial swirler with radial fuel pathways connecting to a circumferential main fuel manifold for distributing the fuel more uniformly across the flow path.

Abstract: An air fuel mixer is disclosed having a mixing duct, a shroud surrounding the upstream end of the mixing duct in which a fuel manifold is provided in flow communication with a fuel supply and control means, a set of inner and outer counter-rotating swirlers adjacent the upstream end of the mixing duct for imparting swirl to an air stream, a hub separating the inner and outer annular swirlers to allow independent rotation of the air stream, and a centerbody located axially along and substantially the full length of the mixing duct. In order to inject one type of fuel into the mixing duct, fuel is supplied to the outer annular swirlers which include hollow vanes with internal cavities, wherein the internal cavities of the outer swirler vanes are in fluid communication with the fuel manifold in the shroud. The outer swirler vanes further include a plurality of fuel passages therethrough in flow communication with the internal cavities.

Abstract: A gas turbine engine fuel injection assembly includes a combustor case surrounding a combustor having a dome. An arcuate manifold is disposed inside the combustor case adjacent to the dome, and includes a plurality of fuel injectors disposed in flow communication therewith for receiving fuel for injection into the combustor. An inlet stem is joined in flow communication with the manifold for channeling fuel thereto, and includes an inner fitting disposed in a mounting port in the combustor case. A mounting adaptor is joined to the combustor case at the mounting port, and circumferentially engages the inner fitting to restrain torsional movement thereof. An outer fitting extends in the adaptor and engages the inner fitting in flow communication therewith for supplying fuel thereto.

Abstract: A combustor is disclosed capable of averting engine flameout. This combustor comprises a combustion chamber, which comprises an outlet for discharging heated combustion products. The combustor further comprises a primary injector for injecting a fuel and compressed air into a primary zone of the combustion chamber such that the fuel is induced into a recirculation zone at a corner of the primary zone. Combustion means for combusting the fuel and the compressed air in the combustion chamber resulting in heated combustion products are also incorporated in the combustor. Moreover, the combustor comprises sensing means for monitoring the combustor to substantially detect a pre-flameout condition. To realize the advantages of the present invention, the combustor additionally comprises a pilot injector for injecting a propellant into the recirculation zone in response to detecting said flameout condition by the sensing means, such that the propellant combusts as a diffusion flame.

Abstract: A high performance annular combustor for a gas turbine engine powering aircraft consisting of a primary combustion zone with the fuel nozzles and air swirlers in the dome and having typical combustion/dilution air holes in the liner and a secondary zone downstream of the primary zone having fuel nozzles and air swirlers operational solely when the primary zone is at a predetermined stoichiometric condition and issuing fuel which is at parity with the fuel/air ratio of the primary zone.

Abstract: The invention relates to a combustion chamber (6) for a gas turbine (6, 18, 19) through which a flow (2) of compressed air is allowed to pass, flowing along an axis (3) from a compressor part (18) to a turbine part (19), said flow (2) having a swirl relative to the axis (3). In this arrangement, an annular passage (4) is provided at the inlet of the combustion chamber (6) having an inflow part (8) for separating a partial flow (1) from the flow (2), which comprises means (11) for removing the swirl from the partial flow (1), and which communicates with the cooling passages (14) for cooling the combustion chamber (6) as well as with pilot burners (7) for stabilizing combustion in the combustion chamber (6). The combustion chamber (6) is configured in an especially favorable manner with regard to avoiding thermodynamic losses and permits the separation of a partial flow (1) which is outstandingly suitable both for cooling purposes and for stabilizing purposes.

Abstract: In a gas turbine having a plurality of combustors, each having an annular array of outer fuel nozzles arranged about a center fuel nozzle. A method for operating the combustors wherein the annular array is supplied with fuel from a diffusion manifold and at least one pre-mix manifold, and further wherein the center nozzle is supplied with fuel from a center premix manifold includes the steps of:a) at start-up, supplying some of the annular array nozzles with diffusion fuel;b) at part speed supplying pre-mix fuel to one of the nozzles in the annular array;c) at full speed, part load, transferring the annular array nozzles receiving diffusion fuel to pre-mix fuel;d) as load is increased supplying pre-mix fuel to the center nozzle; and thene) supplying additional pre-mix to all nozzles as load increases.

Abstract: In a gas turbine having a plurality of combustors, each having an annular array of outer fuel nozzles arranged about a center axis, an improved construction is provided wherein each combustor has a center nozzle located on the center axis. Various methods of operating the improved combustor are also disclosed.

Abstract: In a gas-operated, flame-stabilizing premixing burner for a combustion chamber, the combustion air can be introduced at least approximately tangentially into a premixing space (130). The fuel is injected via a plurality of nozzles (117) lined up in the longitudinal direction of the premixing space and is intensively mixed with the combustion air prior to ignition. The nozzles (117) are subdivided into at least two groups having a separate fuel feed (120, 121) in each case.

Abstract: A combustion chamber with two-stage combustion has primary burners (110) of the premixing type of construction, in which the fuel injected via nozzles (117) is intensively mixed with the combustion air inside a premixing space (130) prior to ignition. The primary burners are of flame-stabilizing design, i.e. they are designed without a mechanical flame retention baffle. They are provided with tangential inflow of the combustion air into the premixing space (130). Arranged downstream of a precombustion chamber (61) are secondary burners (150) which are designed as premixing burners which do not operate by themselves.

Abstract: A heat shield for a gas turbine engine combustor of the type composed of multiple annular stages. The heat shield is particularly adapted for use in a middle stage disposed between radially inward and outward stages of a multistage combustor, and configured to form centerbodies that isolate the middle stage from the other stages of the combustor. The heat shield is composed of an annular array of single crystal superalloy segments. To achieve acceptable levels of durability for the superalloy segments, their primary and secondary crystal orientations are controlled to achieve a balance between the ability to withstand thermally-induced stresses and those stresses induced by high levels of acoustic energy, particularly those high levels attained as a result of a combustor operating with lean fuel/air ratios.

Abstract: An air fuel mixer is disclosed having a mixing duct, a shroud surrounding the upstream end of the mixing duct in which a fuel manifold is provided in flow communication with a fuel supply and control means, a set of inner and outer counter-rotating swirlers adjacent the upstream end of the mixing duct for imparting swirl to an air stream, a hub separating the inner and outer annular swirlers to allow independent rotation of the air stream, and a centerbody located axially along and substantially the full length of the mixing duct. In order to inject one type of fuel into the mixing duct, fuel is supplied to the outer annular swirlers which include hollow vanes with internal cavities, wherein the internal cavities of the outer swirler vanes are in fluid communication with the fuel manifold in the shroud. The outer swirler vanes further include a plurality of fuel passages therethrough in flow communication with the internal cavities.

Abstract: A fuel combustion chamber, and a method of and a nozzle for mixing liquid fuel and air in the fuel combustion chamber in lean direct injection combustion for advanced gas turbine engines, including aircraft engines. Liquid fuel in a form of jet is injected directly into a cylindrical combustion chamber from the combustion chamber wall surface in a direction opposite to the direction of the swirling air at an angle of from about 50.degree. to about 60.degree. with respect to a tangential line of the cylindrical combustion chamber and at a fuel-lean condition, with a liquid droplet momentum to air momentum ratio in the range of from about 0.05 to about 0.12. Advanced gas turbines benefit from lean direct wall injection combustion. The lean direct wall injection technique of the present invention provides fast, uniform, well-stirred mixing of fuel and air.

Abstract: An air fuel mixer is disclosed having a mixing duct, a shroud surrounding the upstream end of the mixing duct in which a fuel manifold is provided in flow communication with a fuel supply and control means, a set of inner and outer counter-rotating swirlers adjacent the upstream end of the mixing duct for imparting swirl to an air stream, a hub separating the inner and outer annular swirlers to allow independent rotation of the air stream, and a centerbody located axially along and substantially the full length of the mixing duct. In order to inject one type of fuel into the mixing duct, fuel is supplied to the outer annular swirlers which include hollow vanes with internal cavities, wherein the internal cavities of the outer swirler vanes are in fluid communication with the fuel manifold in the shroud. The outer swirler vanes further include a plurality of fuel passages therethrough in flow communication with the internal cavities.

Abstract: A method and device for operating a combined burner for liquid and gaseous fuels for the purpose of generating hot gases functions to raise the lean stability limit of the gas flame without impairing the atomization of the liquid fuel and improve the regulating range of the burner. According to the invention, this is achieved when the inflow rate and/or swirl of the blast air (5) into the inner burner space (16) is controlled. To this end, the blast air (5), during operation with gaseous fuel (6), is throttled back by injection of pilot fuel into the blast air, and additionally swirled by swirl generators in the burner. In addition, active regulation of the blast air inflow rate is effected at the burner inlet during the use of both gaseous fuel and liquid fuel.

Abstract: Combustion system with low polluting emissions for gas turbines, of the pre-mixing type, wherein a series of parallel burners, autonomously fed with additional fuel, is circumferentially arranged around the conjunction choke of the pre-mixing chamber with the combustion chamber, to create in the combustion zone immediately before the choke a corresponding series of additional flames for the stabilization of the main flame, the combustion air for the burners deriving from the cooling air of the tapered head of the combustion chamber, which is sent to the burners through twirled blades to give a substantially helicoidal movement to the air.

Abstract: An annular combustor structure particularly suited to staged combustion has a faired centrebody incorporating a diffuser in the head of the combustor. The structure carries fuel injector mountings for pilot and main stage combustion which receive fuel from respective manifolds running around the hollow interior of the structure. There are two annular chambers each containing a fuel manifold, and cooling air passes through the chambers before entering the combustion volume. The manifolds are connected to an external source of fuel by conduits which pass radially through the combustion chamber outer casing walls.

Abstract: A fuel injection assembly for a combustion chamber of a gas turbine engine is disclosed in which a plurality of premixing chambers are in communication with the combustion chamber and adjacent pairs of the premixing chambers are connected to a single prevapoization chamber by prevaporization conduits. Fuel is injected into the prevaporization chamber via a fuel injector and is mixed with oxidizer passing into the prevaporization chamber such that the fuel/oxidizer mixture has a richness exceeding the stoichiometric ratio. The fuel/oxidizer mixture passes into the prevaporization conduits through a diaphragm opening and, upon entering the premixing chamber, is mixed with additional oxidizer such that the fuel/oxidizer mixture in each premixing chamber has a richness less than the stoichiometric ratio.

Abstract: A combustor for a gas turbine having primary and secondary combustion zones. The combustor has primary gas fuel spray pegs for supplying a lean mixture of gaseous fuel to the primary combustion zone via a first annular pre-mixing passage and secondary fuel spray bars for supplying a lean mixture of fuel to the secondary combustion zone via a second annular pre-mixing passage. The fuel spray bars are aerodynamically shaped and a row of fuel discharge ports are formed on opposing sides of the spray bar. A pair of mixing fins project outwardly from the spray bar sides. The fins create turbulence in the air flow that ensures adequate mixing of the fuel and air. The fins have sufficient height and are displaced sufficiently far from the fuel discharge ports so that although the turbulence has not dissipated by the time the air flow reaches the fuel discharge ports, the zone of recirculation located downstream from the fins does not extend to the fuel discharge ports.

Abstract: A combustion chamber which has a primary combustion zone and a secondary combustion zone is provided with a plurality of secondary fuel and air mixing ducts arranged around the primary combustion zone. The secondary fuel and air mixing ducts are defined by a pair of annular walls and by a plurality of walls extending radially between the annular walls. Each secondary fuel and air mixing duct has an aperture to direct a fuel and air mixture into the secondary combustion zone. The apertures have the same flow area. Each secondary fuel and air mixing duct has one or more fuel injectors to inject fuel into the upstream end of the secondary fuel and air mixing duct. This arrangement ensures that the fuel/air ratio emitted from each aperture is within 3.

Abstract: A gas turbine engine combustion chamber which has primary, secondary and tertiary combustion zones in flow series has a secondary fuel and air mixing duct and a tertiary fuel and air mixing duct. The secondary mixing duct has a secondary air intake at its upstream end and the tertiary mixing duct has a tertiary air intake at its upstream end. The tertiary air intake is arranged adjacent to the secondary air intake. A combined secondary and tertiary fuel system is provided to supply fuel to the secondary and tertiary mixing ducts. The fuel system comprises a manifold arranged adjacent to the secondary mixing duct but is spaced from the tertiary mixing duct by the secondary mixing duct. The manifold has apertures to direct fuel towards the tertiary air intake across the secondary air intake. Variations in fuel pressure cause the fuel to flow into the secondary air intake or the tertiary air intake.

Abstract: A combustor for a gas turbine engine is disclosed which is able to operate efficiently at inlet air flows having a high subsonic Mach Number. The combustor has an outer liner and an inner liner spaced therefrom, wherein a combustion chamber is defined therebetween. A dome inlet module is provided which is in flow communication with compressed air flow from a compressor upstream thereof. The dome inlet module has an outer member fixed to the outer liner and an inner member fixed to the inner liner, wherein a flow passage is defined by the outer and inner members for the compressed air flow to flow freely from a compressor to the combustion chamber. Fuel atomizers are then utilized to inject fuel into the flow passage so that the fuel is mixed with the compressed air stream to form a fuel/air mixture which enters the combustion chamber.

Abstract: A fuel and air mixing apparatus for a combustor and gas turbine generator. A primary portion of the fuel is injected into the mixing air at long distances from the combustor prechamber. The primary portion of the fuel is almost completely mixed with the mixing air. A secondary portion of fuel is injected into the mixing air in the boundary layer at a short distance form the combustor prechamber. This minimally mixed second portion provides some rich portions of fuel-air in the prechamber to improve stability and reduce the chances of blowout.

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 system for the destruction of volatile organic compounds while generating power. In a preferred embodiment the system comprises a combustor and a reaction chamber connected to an exit of the combustor. A primary inlet to the combustor supplies a primary fuel to the combustor. A secondary fuel, comprising air and an amount of one or more volatile organic compounds, is supplied to a compressor, which compresses the secondary fuel and directs the secondary fuel to the combustor and the reaction chamber. The system is suitably configured to enable the stoichiometric reaction of the two fuels in a manner sufficient to destroy the volatile organic compounds contained in the secondary fuel and power a turbine engine connected to an exit of the reaction chamber.

Abstract: In the case of a heat generator which essentially consists of a premix burner (100) and a flame tube (1), the hot gases (10) from the combustion in the premix burner (100) are fed into the flame tube (1), and there undergo staged post-combustion. This post-combustion takes place by means of a first post-combustion stage (11) and a second post-combustion stage (12). The air/fuel mixture (11a, 12a) is provided for each post-combustion stage (11, 12) in individual mixers (200, 300). These mixers are arranged axially with respect to the flame tube (1) and work in such a way that injection of the corresponding mixture (11a, 12a) makes it possible to obtain different combustion zones which extend in a staged sequence over the flame tube (1). By virtue of this staged post-combustion mode NO.sub.x emissions can be reduced by a factor of 5 compared to conventional techniques.

Abstract: Sleeves (28) are circumferentially spaced from one another about the liner of a combustor body (12) of a dry low NO.sub.x combustor. The sleeves carry dilution air into the dilution zone. Cooling air is supplied a venturi (20) to cool the venturi and the cooling air flows into the reaction volume. The dilution air sleeves penetrate sufficiently to thoroughly mix the dilution air with the core of hot gases of combustion and, by vorticity effects caused by the flow past the sleeves, thoroughly mix the generally annular flow of cooling air from the venturi with the hot gases of combustion. The thorough mixing of both the cooling air and dilution air inhibits or minimizes the formation of cold areas or streaks within the reaction volume such that CO to CO.sub.2 reactions are not quenched, affording reduced CO emissions.

Abstract: A gas turbine fuel nozzle and method of fuel flow provide resistance to coking of the fuel by means of a multiple passage heat transfer cooling circuit. Fuel streams to primary and secondary sprays of pilot and main nozzle tips are arranged to transfer heat between the pilot primary fuel stream and each of the main secondary fuel stream and the pilot secondary fuel stream. This protects the fuel in the streams from coking during both low flow, lower engine heat conditions and high flow, high engine heat conditions. This nozzle and flow can protect engines with both single tip and dual tip nozzles.

Abstract: The present invention relates to a method and apparatus for supplying fuel to and cooling a fuel injector of a gas turbine engine having a dual head combustion chamber. The method comprises the steps of supplying a total fuel flow to the fuel injector such that at least a portion of the total fuel flow circulates through the high power injector nozzle during all operational modes of the gas turbine engine, including those modes in which the high power injector is not supplying fuel to the combustion chamber, and evacuating unused fuel from the high power injector.

Abstract: A fixed geometry combustor for a gas turbine comprising at least one combustion chamber, a dilution chamber disposed downstream of said combustion chamber and in communication with said combustion chamber, at least one primary inspirator for introducing a fuel/air mixture into said combustion chamber in communication with said combustion chamber, and means for introducing dilution air into the dilution chamber.

Abstract: Fuel flow is scheduled to a plurality of combustion stages by putting a plurality of engine fuel flow governors in parallel with each other and letting the governor which outputs the lowest fuel demand signal be the one which schedules the total fuel flow to the engine. Fuel flows to the plurality of combustion stages are then individually schedulled to achieve the lowest fuel flows consistent with predefined temperature limits for the corresponding stages of combustion.

Abstract: In the operation of gas turbine engines, it is an ever increasing goal to reduce the amount of harmful elements contained within the emissions of the engine. In particular, it is of primary importance to reduce the amounts of nitrogen oxides contained within the emissions. The use of a catalytic combustor is one known means by which the formation of nitrogen oxides may be controlled. However, the use of a catalytic combustor in a gas turbine engine has been known to be effective primarily in the upper ranges of engine operation. The present invention provides a combustor arrangement for a gas turbine engine that employs a first combustor means that utilizes a catalytic bed to react a mixture of fuel and air during an upper range of engine operation and a second combustor means that utilizes a plurality of premixed fuel injectors to ignite a mixture of fuel and air during a lower range of engine operation.

Abstract: A fuel and air mixing apparatus for a combustor and gas turbine generator. A primary portion of the fuel is injected into the mixing air at long distances from the combustor prechamber. The primary portion of the fuel is almost completely mixed with the mixing air. A secondary portion of fuel is injected into the mixing air in the boundary layer at a short distance form the combustor prechamber. This minimally mixed second portion provides some rich portions of fuel-air in the prechamber to improve stability and reduce the chances of blowout.

Abstract: A fuel injection system for a gas turbine engine combustion chamber is disclosed having a sleeve defining a passageway extending along a longitudinal axis of a combustion zone, a first fuel injector orifice to inject fuel into the passageway, and second fuel injector holes arranged in a plurality of substantially linear arrays extending radially from the sleeve and angularly equidistantly spaced from each other, the second fuel injection holes located in a plane extending substantially perpendicular to the longitudinal axis of the combustion chamber such that the plane is located axially between an upstream end wall of the combustion chamber and the downstream end of the sleeve.

Abstract: A gas turbine fuel nozzle and method of fuel flow provide resistance to coking of the fuel by means of a multiple passage heat transfer cooling circuit. Fuel streams to primary and secondary sprays of pilot and main nozzle tips are arranged to transfer heat between the pilot primary fuel stream and each of the main secondary fuel stream and the pilot secondary fuel stream. This protects the fuel in the streams from coking during both low flow, lower engine heat conditions and high flow, high engine heat conditions. This nozzle and flow can protect engines with both single tip and dual tip nozzles.

Abstract: A fuel valve comprising main fuel valve connected to a burner staging valve such that fuel flow from the main fuel valve is divided between pilot and main domes in a dual annular combustor. The burner staging valve includes a movable spool located within a fixed sleeve such that the spool may be moved both laterally and rotated within predetermined limits. The movable spool and sleeve include pilot, main and inlet windows arranged such that the inlet windows receive fuel from the main fuel valve and the pilot and main windows split the fuel flow between the pilot and main domes of the combustor.

Abstract: A combustion system with low pollutant emission, of the pre-mixing type for gas turbines, in which the combustion air passes from the interspace to the pre-mixing chamber via apertures in this latter, their degree of opening being varied, in accordance with the fuel quantity used, by corresponding apertures in a rotatable drum, and in which a series of small holes fed with additional fuel is provided in the interspace. Further flame stabilization expedients and differential cooling of the combustion chamber are also provided.

Abstract: Past systems have attempted to provide a combustion system to reduce NOx pollution, however, such methods have failed to attain adequate reductions in NOx pollution. The present system or structure for for reducing NOx includes a pilot fuel system and a premixed supply system to reduce the NOx to a workable level. The pilot system uses a device for supplying combustible fuel to the fuel injection nozzle generally along a combustion axis during all or a portion of the operating conditions of the engine. Low NOx is maintained by using a device for supplying combustible fuel into each of a plurality of spaces formed between a plurality of swirler vanes. Low NOx is further maintained by having the cooling fluid used to cool the fuel injection nozzle mixed with the fuel and air from the spaces prior to entering the combustion chamber.

Abstract: The combustion dynamics and efficiency of gas turbine having an annular combustor 26 provided with fuel injection nozzles 50 that inject fuel generally tangentially is improved by providing the walls 32, 34, 39 of the combustion 26 with cooling air film injectors 70, 86; 72, 88; 74, 90 at substantially equally angularly spaced locations about each such wall and which are oriented to generally tangentially inject a film-like air stream on the associated wall 32, 34, 39.