Abstract: A three stage lean burn combustion chamber (28) comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (76,78,80,92). The primary fuel and air mixing ducts (78,78) comprise first and second radial flow swirlers (60,62). The primary fuel injectors (64,66) inject fuel into the first and second swirlers (60,62). The primary fuel injectors (44,66) and the first and second swirlers (60,62) are arranged such the fuel to air ratio of the fuel and air flowing from the passages (61) of the first swirler (60) into the primary combustion zone (36) is different to the fuel to air ratio of the fuel and air flowing from the passages (65) of the second swirler (62) into the primary combustion zone (36).

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: A method and apparatus are provided for rejecting waste heat (12) from a system (10) including a combustion engine (14), an oxidizer supply (16), and a fuel supply (18). The method includes the steps of providing a fuel flow (32) from the fuel supply (18), providing an oxidizer flow (30) from the oxidizer supply (16), splitting one of the fuel flow and the oxidizer flow into cooling flow (34) and a combustion flow (36), rejecting system waste heat (12) to the cooling flow (34), combusting the combustion flow (36) with the other of the fuel flow (32) and the oxidizer flow (30) to supply combustion gas for the combustion engine (14), and directing the cooling flow (34) into the combustion gas to reduce the temperature of the combustion gas.

Abstract: Lean premixed combustion of a hydrocarbon fuel and air is combined with lean direct injection of hydrocarbon fuel and air 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. After this low temperature reaction has been completed, additional fuel and air are injected into the products of combustion downstream of the main reaction zone in order to raise the temperature of the mixture to the level required to operate an advanced, high efficiency, heavy duty industrial gas turbine at high load.

Abstract: A catalytic combustion chamber is provided with at least two catalytic combustion zones arranged in flow series. In a first mode of operation fuel is supplied from first fuel injectors, positioned upstream of the first catalytic combustion zone, into the catalytic combustion chamber and is burnt in the first catalytic combustion zone in order to preheat the subsequent catalytic combustion zones. In the second mode of operation the supply of fuel to the first fuel injectors is reduced and fuel is supplied from second fuel injectors positioned between the first catalytic combustion zone and the second catalytic combustion zone into the space between the first catalytic combustion zone and the second catalytic combustion zone. This prevents the first catalytic zone becoming overheated, and reduces the possibility of the second and third catalytic combustion zones becoming overheated and allows the optimum catalyst to be selected for the first catalytic combustion zone.

Abstract: A multi-stage topping combustor unit having rich burn, and lean burn zones followed by a quench/dilution zone. Combustible fuel gas is supplied only to the rich burn zone, and an oxygen-containing air-flue gas mixture is supplied to all three zones in a fixed flow proportioning that establishes a desired temperature profile through the combustor unit that minimizes combustor NOx and CO emissions. The combustor unit casing is cooled by an augmented air and oxygen stream discharged into the rich burn zone only, or into both the rich and lean burn zones. The cooling air flow is regulated by a control device/valve that allows the cooling air flow, and hence the augmenting oxygen entering the rich burn zone or the rich and lean burn zones, to be reduced as the fuel gas flow/firing temperature is reduced.

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 gaseous fuel (B) is fed to a a premixing burner which has a main line (10) for the feed of oxygenous gas (G) into at least one mixing zone and/or a combustion zone (VMZ, EMZ, VZ). In the process, at least a first portion of the fuel is fed to the main line (10), in which it is mixed with oxygenous primary gas (PG) inside at least one premixing zone (VMZ) and forms a premixture. The concentration of the fuel in the premixture (VM), under the respective operating conditions of the burner, is selected in such a way that the premixture lies outside the ignition limit. The premixture (VM) is fed to a final mixing zone (EMZ) and is mixed therein with the residual fuel (RB) likewise supplied and/or with oxygenous secondary gas (SG). The resulting final mixture (EM) is fed to a following combustion zone (VZ).

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 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: A low-emission combustion chamber for gas turbine engines comprises an outer casing with an upstream end wall with a pilot fuel injector, a first flow swirler, an igniting members for initiating a stable diffusion frame in a pilot zone, at least one second coaxial swirler, main fuel injectors, secondary air inlets, and a main combustion zone. For obtaining a still further reduced emissions of primarily nitrogen oxides, the pilot zone is confined radially outwardly by a surrounding wall which constitutes the radially inner confinement of an axial outlet portion of a radial vaporization channel within the second swirler and a third radial flow swirler is adapted to supply the secondary air in a rotary motion opposite to that of the main flow of fuel and air.

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: 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: A combustor for a gas turbine having primary and secondary combustion zones. The combustor has a centrally disposed dual fuel nozzle that can supply a fuel rich mixture of either liquid or gaseous fuel to the primary combustion zone. The combustor also 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 dual fuel spray bars for supplying a lean mixture of either gaseous or liquid fuel to the secondary combustion zone via a second annular pre-mixing passage. The dual fuel spray bars are aerodynamically shaped and have passages for distributing gas and liquid fuel to a number of fuel discharge ports. The gas fuel discharge ports are formed in two rows on either side of the spray bar. The liquid fuel discharge ports are formed by a row of spray nozzles arranged along the downstream edge of the spray bar.

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 method and apparatus is disclosed for controlling staged combustion systems of the type used for combusting two reactants wherein at least one reactant is supplied as two flows, the first of which is combusted with the flow of the other reactant in a primary stage and the other of which is combusted with the exhaust of the primary stage in a secondary stage. In the present invention, the respective flows are monitored and controlled to provide a desired equivalence ratio for both the first stage and for the burner overall which is adaptable in response to variable input conditions. By controlling the flows in this way, a calculated energy output or adiabatic flame temperature can be produced. The present invention can also control the reactant flows to each stage in response to measured parameters such as flame temperature or emission levels. In this way, the flows can be varied in order to drive the measured parameters to a desired level.

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: A gas turbine combustor apparatus utilized in conjunction with a biomass fueled pressurized gasifier, the fuel gas and primary air injected into a primary combustion chamber at independently controlled rates in order to provide a rich burn of the fuel gas at a constant temperature which inhibits formation of nitrogen oxides, the combustion product of the primary combustion chamber flowing downstream through a chamber combustion nozzle and then to a secondary combustion chamber, a secondary air injected into the secondary combustion chamber at an independently controlled rate in order to provide a lean burn of the combustion product of the primary combustion chamber, also inhibiting the formation of nitrogen oxides. The combustion product of this lean burn being diluted by injection of a tertiary air independently controlled, the diluted combustion product expelled to a land based stationary turbine. The combustor designed to efficiently burn low BTU biomass fuel gas.

Abstract: A combustion chamber 42 for an axial flow rotary machine 10 is disclosed. Various construction details are developed for controlling the production of emissions such as carbon monoxide, unburnt hydrocarbons and oxides of nitrogen. In one embodiment, the combustion chamber has a pilot zone 108 which extends axially and a main zone 106 axially adjacent to the pilot zone which has a divergent outer wall having a divergent end region 178 through which locally swirling gases are injected into swirling gases from the pilot zone 108.

Abstract: A combustion chamber apparatus 1 for a gas turbine has a pilot stage and a main stage for low-emission lean-burn operation. The pilot stage is separate from the main stage and is formed by a flame tube 2 with an upstream first fuel injector 3a and with a secondary combustion zone 7a. The main stage includes a vaporization chamber 4 and a combustion chamber 5, and a second fuel injector 3b is provided at an upstream end of the vaporization chamber 4. The vaporization chamber 4 produces a homogenous fuel-air mixture which will not burn until it is in the combustion chamber 5. In lean-burn operation of the apparatus, the separate arrangement of the two chambers 4, 5 significantly reduces the formation of nitrogen oxides. For use in aircraft gas turbine engines, the apparatus has an annular shape with several main and pilot stages.

Abstract: A combustor for a gas turbine having primary and secondary combustion zones. The combustor has a centrally disposed dual fuel nozzle that can supply a fuel rich mixture of either liquid and gaseous fuel to the primary combustion zone. The combustor also has primary gas fuel spray bars for supplying a lean mixture of gaseous fuel to the primary combustion zone via a first annular pre-mixing passage and secondary gas fuel spray bars for supplying a lean mixture of gaseous fuel to the secondary combustion zone via a second annular pre-mixing passage. In addition, the combustor also has a plurality of liquid fuel spray nozzles for introducing a lean mixture of liquid fuel into the secondary combustion zone via the second annular pre-mixing passage. The liquid fuel spray nozzles are disposed in fan shaped channels that are arranged in a circumferential array and that are connected to the second annular pre-mixing passage.

Abstract: A gas turbine combustor of the pre-mixed combustion system in which the pre-mixed fuel and the air are combusted. The gas turbine combustor comprises main cylindrical nozzles provided in the end wall on the upstream side of a cylindrical combustion chamber, auxiliary nozzles formed to surround the main nozzles, a main pre-mixed gas supply for supplying a pre-mixed gas to the main nozzles, and an auxiliary pre-mixed gas supply for supplying a pre-mixed gas having an fuel/air ratio smaller than that of the main pre-mixed gas to the auxiliary nozzles, and wherein it is allowed to stably burn a lean pre-mixed gas having an fuel/air ratio of greater than 1 from a low-load condition through and up to a high-load condition of the gas turbine.

Abstract: A gas turbine engine combustion chamber is disclosed having multiple sets of fuel injectors. A set of first fuel injectors are arranged in two radially inner and outer generally circular first arrays such that the fuel injectors of one array are located circumferentially between the fuel injectors of the other array. The combustion chamber also includes a set of second fuel injectors which are also arranged in two generally circular, radially inner and outer arrays such that, in each array, the second fuel injectors are located circumferentially between the first fuel injectors. In a circumferential direction, the positions of the first and second fuel injectors vary between the radially outward array and the radially inward array. One set of fuel injectors may operate under idle and low power operating conditions, while the other set of fuel injectors inject fuel into the combustion chamber during full power take off or cruise operations.

Abstract: Gas turbine engine combustion chamber has staged combustion to reduce nitrous oxides and includes a first radial flow swirler and a second radial flow swirler located axially of an annular mixing zone with each swirler having vanes for rotating the incoming air in substantially opposite directions relative to each other; first and second fuel injectors are provided with a first fuel injectors located in one of the passages of each of the first and second swirlers and with the second fuel injectors located upstream of the passages of the first and second swirlers.

Abstract: A combustion system for a gas turbine engine provides a reaction limiting means for preventing the combustion process within the combustion chamber of the engine from proceeding towards completion by limiting the reaction rate or duration of the reaction process sufficiently such that a clean combustion gas is produced having substantially less carbon particles than for a combustor effective for essentially complete combustion. The invention also provides an efficient method of removing carbon deposits which do form by controlling the combustor shutdown technique in such a way as to cause the carbon to be oxidized. The preferred embodiment provides an aircraft emergency power unit having a high pressure air storage tank, an aviation fuel storage tank, and a combustor with a reaction limiting means which combusts pressurized air and aviation fuel in a fuel rich ratio to produce a motive combustion gas in a reaction limiting means.

Abstract: A low pollution combustion chamber for a turbojet engine is disclosed in which first and second combustion chambers are oriented in a counter-flow direction and both combustion chambers communicate with a separate and distinct exhaust chamber. The first combustion chamber has a fuel injector, as well as a primary oxidizer intake, and the wall bounding the first combustion chamber defines a plurality of dilution oxidizer intake orifices. The wall also defines a first exhaust orifice located approximately midway between the upstream and downstream ends of the first combustion chamber. A second combustion chamber has a second fuel injector, as well as a second primary oxidizer intake, but the wall defining the second combustion chamber does not define any dilution oxidizer intake orifices. A second exhaust orifice also communicates with the exhaust chamber.

Abstract: This disclosure describes a single shell, double dome combustor design configuration with the intermediate dividing wall between the two annuli removed, thereby reducing weight and cost, and improving durability. An annular daisy type mixer is positioned between the domes to improve the mixing of air and combustor gases prior to the exiting of the gases from the combustor to the turbine vanes. The daisy mixer positioned downstream of the domes to provides effective, low pressure loss mixing at a a controlled axial rate, by creating effective cross-flow barriers to avoid uncontrolled fuel/air ratio transfer between the inner and outer annuli, and thereby allows a concentric and controlled axial air admission in conjunction with the individual swirlers.

Abstract: A combustion chamber suitable for use in a gas turbine engine includes a first enclosure with a fuel injector for low-power operation and a primary-oxidizer intake, a second enclosure with a fuel injector for full-power operation and a primary-oxidizer intake, and a third enclosure from which to evacuate burnt gases and which communicates with the first and second enclosures. The wall of the first enclosure includes intake orifices for a dilution oxidizer, but the wall bounding the second enclosure lacks any orifices other than the primary-oxidizer intake orifices.

Abstract: A centerbody cup construction for use in a dry low NOx gas turbine combustor includes a pair of inner and outer generally cylindrical members, the inner member having a diverging cup portion attached to its forward end extending toward the outer cylindrical member, and a coolant manifold arranged between the inner and outer cylindrical members radially adjacent the centerbody cup portion.

Abstract: A burner comprises at least one swirl member for mixing a fuel with an air previously to burning of the fuel, and at least one eddy generating device which is arranged in a flow of a mixture of the fuel and air to generate an eddy flow in the flow so that the eddy flow maintains a shape of a flame of the burned fuel. The at least one eddy generating device is arranged apart from a downstream end of the swirl member by a fixed sufficient distance so that the shape of the flame maintained by the eddy flow is prevented from moving toward the downstream end of the swirler member.

Abstract: A method of operating a double dome combustor includes channeling compressed airflow through an inner dome for generating inner combustion gases having an inner reference velocity greater than an outer reference velocity of outer combustion gases generated from compressed airflow channeled through an outer dome, and diffusing the outer and inner combustion gases in outer and inner combustion zones. One double dome combustor effective for practicing the method in accordance with the present invention includes outer and inner combustor liners and domes, with the domes having outer and inner carburetors disposed therein, respectively. The inner carburetors are sized for generating inner combustion gases in the inner combustion zone having an inner reference velocity greater than an outer reference velocity of the outer combustion gases generated in the outer combustion zone.

Abstract: A combustor includes spaced outer and inner liners defining therebetween a combustion zone, and having a plurality of circumferentially spaced tubular domes disposed between forward ends thereof. Each of the domes includes an annular inlet for receiving pilot airflow, a swirler having a tubular ferrule and primary swirl vanes for receiving the pilot airflow, and an outlet disposed in flow communication with the combustion zone. A pilot fuel injector is disposed in each of the domes for injecting pilot fuel into the pilot airflow for generating a pilot fuel/air mixture dischargeable from the dome outlet. A plurality of main fuel spraybars inject main fuel into compressed airflow channeled between the domes and the outer and inner liners for generating a main fuel/air mixture dischargeable concentrically around the pilot fuel/air mixture.

Abstract: A method and apparatus are provided for reducing NOx emissions in dual stage, dual mode gas turbine combustors. The gas turbine combustor includes first and second combustion chambers separated by a reduced diameter throat section. The throat section is formed by converging and diverging wall sections constructed of porous material. A liner is provided in surrounding relationship to the throat region to form a plenum by which predetermined amounts of air are metered into the plenum and permitted to pass through the porous throat wall sections.

Abstract: A gas turbine combustor and a method of operating the gas turbine combustor, with the gas turbine combustor comprising a combustion sleeve, a combustion sub-chamber disposed at an upstream side of the combustion sleeve and having an air and fuel supply system, and a combustion main chamber disposed on a downstream side of the sub-chamber and having an air and fuel supply system. The combustion sub-chamber and combustion main chamber are formed in such a manner that the start-up of the gas turbine is effected by the hot combustion gas generated in the sub-chamber. The combustion sub-chamber is provided with an inner sleeve protruding from a side wall of the sub-chamber such that a tip of the inner sleeve is located in the vicinity of a downstream end of the sub-chamber, and a fuel nozzle is provided in a tip of the inner sleeve for increasing the fuel-to-air ratio during an initiating period of an increase in the gas turbine rotation speed.

Abstract: A combustion assembly includes a combustor having inner and outer liners, and pilot stage and main stage combustion means disposed between the liners. A turbine nozzle is joined to downstream ends of the combustor inner and outer liners and the main stage combustion means is close-coupled to the turbine nozzle for obtaining short combustion residence time of main stage combustion gases for reducing NO.sub.x emissions. In a preferred and exemplary embodiment of the invention, the combustion assembly includes first and second pluralities of circumferentially spaced fuel injectors and air swirlers disposed radially outwardly of a plurality of circumferentially spaced hollow flameholders having fuel discharge holes. Pilot stage combustion is effected downstream of the first and second fuel injectors and swirlers, and main stage combustion is effected downstream of the flameholders.

Abstract: A combustor for driving a gas turbine includes a first burning stage and a second burning stage. The combustor further comprises a primary combustion chamber in which air from the associated compressor and fuel are burnt, a pre-mixing chamber in which air from the associated compressor and fuel are pre-mixed, a main combustion chamber in which fuel/air mixture is burnt and forwarded to the gas turbine, an air passage provided in a wall of the main combustion chamber, through which air from the compressor flows to cool the wall thereof, and an intercommunicating air passage for intercommunicating the air passage to the pre-mixing chamber, whereby cooling air for the main combustion chamber wall is used to supplement the air in the pre-mixing chamber.

Abstract: A gas turbine combustor comprising a head combustion chamber and a rear combustion chamber connected to a downstream side of the head combustion chamber, a first stage burner for premixing first stage fuel and air and supplying the resultant fuel and air premixture into the head combustion chamber to effect first stage premix combution, a second stage burner for premixing second stage fuel and air and the resultant fuel and air premixture into the rear combustion chamber to effect premix combustion in addition to the first stage premix combustion, and a device for regulating flow rates of combustion air to be premixed with first and second stage fuel. The combustor is further provided with a pilot burner in the head combustion chamber to form pilot flame and stabilize the first stage premix combustion.

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 dual mode, dual stage low NOx combustor, includes a primary combustion chamber and a secondary combustion chamber separated by a throat region of reduced diameter. Fuel is input into the primary combustors by an annular array of diffusion type nozzles whereas in accordance with the invention fuel is input into the secondary combustor by a central combination premix and diffusion nozzle.

Abstract: Improved performance in a hot gas generator is achieved by disposing a dual fuel injector 42 having spaced fuel discharge ports 48, 60 on the longitudinal axis 28 of a vessel 18 having a narrow inlet 22, an opposite narrow outlet 14 and an intermediate, enlarged chamber 24 which serves as a combustion chamber.

Abstract: A coordinated valve splitter assembly includes a linear trim primary valve in parallel with an equal-percentage-trim valve to split fuel gas between primary nozzles and a secondary nozzle of a gas turbine combustor. In one embodiment, coordination of the actuation of the two valves is accomplished through rigid connection of the valve shafts. Shaped-cam and electronic actuation of the two valves is also disclosed.

Abstract: A dilution air dispensing apparatus for a double dome combustor with dilution air dispensing holes that are staggered on the opposite walls of the centerbody. The wakes on the downstream side of the jets provide avenues of access for the deflected combustion gases to continue their travel across the combustor and into the regions downstream of the opposite domes. The gases then mix with the remaining undeflected gases from the domes and with the spent dilution air before reaching a combustor exit plane.

Abstract: A gas turbine combustor for reducing a production of NOx. The combustor includes a head combustion chamber and a rear combustion chamber which is larger in diameter than the head combustion chamber. The head combustion chamber is provided with an axially extending hollow frustoconical tubular member to form an annular combustion space therein, air holes for axially jetting air into the annular combustion chamber, air holes formed on a peripheral wall for injecting air and a plurality of fuel nozzles projected into the annular combustion space for injecting fuel into vortex formed by the air jet and the injected air flow whereby the flame is stabilized and lean combustion can be effected. The rear combustion chamber has a fuel and air supply means on the upstream side which includes air inlets formed by whirling vanes and fuel nozzles disposed in the air inlets so that fuel and air are mixed well.

Abstract: In the combustion space of a combustion chamber of a gas turbine operated with liquid fuel, at least one after-burner (4) is employed in each case in combination with one or more primary burners (2, 2a). The after-burner (4) and at least its fuel spray cone (15), which acts directly into the central combustion chamber (6), are screened by an unswirled enveloping airstream (14) against the hot gases (13) from the combustion in the primary burners (2, 2a). The after-burner (4) itself is not automatically operating, i.e. the ignition of its mixture (14/15) takes place further downstream, preferably at the beginning of the mixing chamber (7), as a result of which a turbulence-free flow with uniform pressure and temperature profile is provided for acting on the turbine (9).

Abstract: In a gas turbine apparatus of a premixed gas combustion type, a plurality of fuel supply systems are provided for supplying fuels to a combustion chamber of the gas turbine apparatus after the fuels are mixed with a high-pressure air, and the supply of fuels from the plurality of fuel supply systems is controlled according to programmed patterns which are predetermined for the fuel supply systems. According to a variation of the turbine load, the patterns of fuel supplies are executed in a transferred manner at a number of transfer points along a variation of the turbine load, with equal amounts of fuel reduced and increased stepwisely in the precedingly and subsequently executed patterns at each of the transfer points.

Abstract: A combustion chamber is defined by a wall having a closed end into which liquid fuel and oxidant are introduced and an open end through which exhaust gases flow. A baffle at the open end causes exhaust gas to flow adjacent the wall. Injector devices direct liquid fuel radially inwardly of the chamber between the closed and open ends. Unburnt fuel attached to the wall cools the walls and passes to the open end before being entrained by the exhaust gas flow through the baffle.

Abstract: A method of starting a gas turbine plant, which plant has at least one combustor including a primary combustion chamber into which primary fuel nozzles open and a secondary combustion chamber into which secondary fuel nozzles open, a compressor for supplying the combustor with compressed combustion air, and a gas turbine driven by the combustion gas generated in the combustor and adapted to drive a load such as an electric power generator. When the gas turbine is being accelerated to a rated speed or while the load is still below a normal load range, a fuel is supplied only to the primary fuel nozzle, whereas, in other loaded operation range, the fuel is supplied to both the primary and secondary fuel nozzles. Before the fuel supply to the secondary fuel nozzles is commenced, a part of the compressed air is bled through at least one air bleed pipe leading from the compressor and having an air bleed valve.

Abstract: A premixture of compressed air and gaseous fuel is supplied and conbusted into a head combustion chamber of a combustor for gas turbine and main combustion of lean, low flame temperature is carried out in a rear combustion chamber following the heat combustion chamber to prevent the formation of high flame temperature which tends to promote NOx generation.