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: A control method of a gas turbine which is provided with a multi can type combustor consisting of a plurality of unit combustors, each of which has a first stage combustion section for diffusion combustion and a second stage combustion section for premix combustion. Combustion states in the first and second stage combustion sections are monitored to take in quantities representative of the combustion conditions as control factors. The combustion in the second stage combustion section is assumed to be abnormal when the taken in quantities exceed an allowable limit and the combustion in the first and second stage combustion sections is changed to independent combustion only in the first stage combustion section when the number of unit combustors, each of which is assumed to be abnormal, exceeds a predetermined number of the unit combustors.

Abstract: A double annular combustor for a gas turbine engine is provided with annular arrays of main and pilot fuel injection modules. The main fuel injection modules are of the premix type so as to vaporise fuel. However, the pilot fuel injection modules are configured so as to function as both premix and airspray fuel injectors. During starting and low power conditions, the pilot fuel injectors are operational alone in their airspray mode. However during high power conditions, both the main and pilot fuel injection modules function as premix injectors. The arrangement reduces noxious emissions.

Abstract: In a gas-turbine annular combustion chamber (4) which is arranged downstream of a compressor (1) and is equipped on its front plate with at least one row of premix burners (5) arranged in an annular form, in each case a combustion-air duct (15) designed as a diffuser leads directly downstream of the compressor outlet from the guide vanes (9) of the last compressor row to each burner (5), at the downstream end of which combustion-air duct (15) at least one longitudinal-vortex generator (16) is located, at least one fuel injection means (17) being provided in or downstream of the longitudinal-vortex generator (16). A mixing duct (19) which ends in the combustion chamber (4) and has a constant height (H) and a length (L) which corresponds approximately to twice the value of the hydraulic duct diameter (D) is arranged downstream of the fuel injection means (17). The overall size of the gas turbine in the region of the combustion chamber (4) can thereby be substantially reduced.

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 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 multi-swirl combustor plate is disclosed. The multi-swirl combustor plate contains a plurality of swirlers arranged around an interior section of the combustor plate, and a plurality of non-swirled holes arranged around an exterior section of the plate that had been cast as a solid section. The swirlers are preferably arranged in concentric circles, and the non-swirled holes are also preferably created in concentric circles disposed outside of the swirlers. In accordance with the present invention, each of the swirlers also preferably comprises an outer wall having a thickness greater than that of the multi-swirl plate and is an extension of the multi-swirl plate disposed in the downstream direction. Similarly, the swirler hub also has a thickness greater than that of the plate is an extension disposed in the downstream direction.

Abstract: A gas turbine combustor having a pre-mixing zone and a combustion zone separated by a barrier plate. A plurality of concentrically arranged annular passages, into which lean mixtures of fuel and air are introduced, are disposed in the pre-mixing zone. Each of the annular passages has a discharge end in which a plurality of swirlers are circumferentially arranged that allow fuel/air mixtures flowing through the annular passages to enter the combustion zone, the swirlers acting as flame holders. A toroidal fuel manifold upstream of each of the annular passages allows the fuel/air ratio in each annular passage to be individually controlled. The firing temperature of the combustor is increased by sequentially supplying fuel to each annular passage in turn, with the innermost passages being supplied with fuel first. A pilot fuel assembly is centrally disposed with respect to the annular passages and has swirler vanes and an ignitor.

Abstract: In a method of regulating the rotational speed of a gas turbine during load disconnection, the fuel line (8) of which to the combustion chamber (10) can be controlled via regulating valves (5) and via burner control valves (6), and the combustion chamber of which is operated with premixing burners (11), some of the premixing burners are used as idling premixing burners (11a) and the rest are operated only as main burners. During load disconnection, the main burners are shut off by means of the burner control valve (6), and the idling premixing burners are supplied with fuel via a regulated bypass line (3).

Abstract: A combustor, for a gas turbine engine, employing regenerative combustor cooling. The combustion gas flow direction extends generally longitudinally aft of the combustor fuel nozzle. A coolant flowpath between the combustor casing and the combustor liner has: 1) a longitudinally aft inlet in fluid communication with a source of compressor-derived cooling air, of lower temperature and higher pressure than diffused air from the combustor diffuser; and 2) a longitudinally forward outlet in fluid communication with the combustor fuel nozzle for "spent" cooling air to be used for combustion.

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: The present invention has an object to provide a gas turbine combustor which is able to effect stable combustion in a wide range of fuel flow rate. A burner 1 is provided with fuel nozzles 31 and 32. When a fuel flow rate is small, diffusion flame is formed with fuel supplied from the nozzle 31 with a ring-shaped flame stabilizer 11. Next, fuel is supplied from the nozzle 32 to mix with air, reach to the flame stabilizer 11 and be held by the diffusion flame already formed, whereby stable premixed flames are formed in the flame stabilizer 11 from a range of low fuel air ratio. Further, when flame is propagated from the burner 1 to the burner 2, a fuel air ratio at the outer periphery side of the burner 1 is locally raised by the fuel supplied from the nozzle 31, whereby the combustion stability can be raised in a wide range of fuel flow rate and propagation of flame to adjacent burners becomes easy.

Abstract: A multiple annular combustor for a gas turbine engine is disclosed having a plurality of concentrically disposed annular combustion zones about a centerline axis. The multiple annular combustor includes a first annular combustion zone, a second annular combustion zone, and a centerbody disposed between the first and second annular combustion zones. The centerbody has a first wall extending axially from an upstream end to a downstream end and a second wall radially spaced from the first wall extending axially from an upstream end to a downstream end, where the downstream ends of the first and second walls are connected by an aft wall and a pair of side walls to form a passage therebetween. The centerbody is geometrically configured to cause a thermoelastic neutral axis to be positioned radially with respect to the first and second walls thereof so that thermal stress imposed on the first and second walls is minimized.

Abstract: An improved dual-head combustion chamber is disclosed having a generally annular configuration extending about a central axis with a low power head, operating during low power engine conditions and a radially displaced high power head operative under high power engine operating conditions. The low power head has N number of fuel/air injector assemblies arranged in an annular array and spaced apart in a circumferential direction about the central axis. The fuel/air injector assemblies of the low power head have an air permeability of P1. The high power head also is arranged in a generally annular array with N number of first fuel/air injector assemblies and N number of second fuel/air injector assemblies with each of the second fuel/air injector assemblies aligned with a fuel/air injector assembly of the low power head along a radius line extending from the central axis.

Abstract: A method of supplying fuel to a combustion chamber of a turbojet engine having a low power head with a plurality of low power fuel injectors and a high power head having a plurality of high power fuel injectors is disclosed in which fuel is supplied to the plurality of low power fuel injectors during low power operation of the engine, fuel is supplied to a first fuel circuit in the plurality of high power fuel injectors during high power operation of the turbojet engine and fuel is also supplied to a second fuel circuit in the plurality of high power fuel injectors during low power operation of the turbojet engine, the second fuel circuit being separate from the first fuel circuit. This method of supplying fuel to the dual head combustion chamber optimizes operation of the combustion chamber in all modes of engine operation.

Abstract: A dome assembly for a multiple annular combustor is disclosed as including an annular dome plate having at least two radial domes, wherein each of the radial domes include a plurality of circumferentially spaced openings therein. A heat shield is positioned within each of the openings, with a threaded forward end located upstream of the dome plate and an aft end located downstream of the dome plate. A retainer nut with threads formed on an annular surface thereof is matingly engagable with the forward end of each heat shield. When the retainer nut is tightened onto the heat shield forward end, the heat shield is mechanically attached to the dome plate. Preferably, at least one centerbody extends axially downstream from either the radially outer or inner side of the aft end. The dome assembly includes a ferrule within each of the dome plate openings for receiving an air/fuel mixer therein, the ferrule having an annular sealing flange extending radially therefrom.

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 low NOx generating combustor in which a first lean mixture of fuel and air is pre-heated by transferring heat from hot gas discharging from the combustor. The preheated first fuel/air mixture is then catalyzed in a catalytic reactor and then combusted so as to produce a hot gas having a temperature in excess of the ignition temperature of the fuel. Second and third lean mixtures of fuel and air are then sequentially introduced into the hot gas, thereby raising their temperatures above the ignition temperature and causing homogeneous combustion of the second and third fuel air mixtures. This homogeneous combustion is enhanced by the presence of the free radicals created during the catalyzation of the first fuel/air mixture. In addition, the catalytic reactor acts as a pilot that imparts stability to the combustion of the lean second and third fuel/air mixtures.

Abstract: A dome assembly for a single annular combustor of a gas turbine engine is disclosed as having a first dome wall in flow communication with compressed air supplied to the combustor, the first dome wall including a central opening therein and at least one cooling passage therethrough. A baffle is spaced downstream of and connected to the first dome wall at radially outward and inward ends, the baffle also including a central opening therein. A second dome wall defining the central opening in the first dome wall is provided which extends upstream of the first dome wall. A venturi is located within the central opening of the first dome wall, with the venturi including a flange extending radially outward from the central opening, wherein the second dome wall is connected to the flange at an upstream end.

Abstract: A combustor includes a pilot burner provided at a center portion thereof, and a premixer surrounding the pilot burner. The premixer has a plurality of premix chambers separated from one another in a peripheral direction. Each of the premix chambers includes a fuel injection nozzle provided at an inlet portion thereof, and a blade-like structural member disposed downstream of the fuel injection nozzle. Fuel and air, supplied from an inlet of the premixer, jointly form a premixture, and a Lanchester vortex is formed rearwardly of the blade-like structural member. The Lanchester vortex forms circulating flows at the outlet of the premixer, and also forms a main stream outside these circulating flows.

Abstract: A fuel control system for a gas turbine engine having a staged combustor includes apparatus and methods to operate the combustor during pilot operation, staged operation, and transition between the two modes of operation. Various construction details are developed which provide a fuel control having a switch with built in hysteresis for determining the transition points between operating modes. In one embodiment, the combustor is switched from pilot to staged operation at a thrust level greater than the thrust level at which the combustor is switched from staged to pilot operation. During transitions between operating modes, the changes in fuel flows are limited by a plurality of rate limits to prevent combustor fuel spikes and blow-outs from occurring. During engine starts, the fuel lines to the main combustor are pre-filled to provide responsiveness during the transition from pilot to staged operation.

Abstract: A separator-diffuser for a gas turbine engine is disclosed wherein the diffuser includes axial walls which function as a guide vane stage for the oxidizer compressor. By incorporating the last, or downstream, guide vane stage into the diffuser structure, the overall axial length and bulk of the gas turbine engine may be significantly reduced. The separator-diffuser has a wall defining a generally annular oxidizer flow passage extending around the central axis of the gas turbine engine, the passage having an oxidizer inlet located downstream of the last rotor stage of the oxidizer compressor so as to enable the oxidizer to pass from the oxidizer compressor into the oxidizer flow passage. A generally annular separator member is located in the oxidizer flow passage downstream of the oxidizer inlet so as to divide the oxidizer flow passage into at least two separate oxidizer flow paths, each flow path having a separate oxidizer outlet.

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: The present invention relates to a system for cooling a high power fuel injector of a dual fuel injector wherein the system comprises a first fuel supply circuit having a first conduit connecting a fuel feed supply and the high power fuel injector, the first conduit having a terminal end adjacent to a distal end of the high power fuel injector, and a second conduit connecting the terminal end of the first conduit to the low power fuel injector such that all of the fuel supplied to the low power injector first passes through the high power fuel injector. The system also has a second fuel supply circuit, separate from the first fuel supply circuit, which comprises a third conduit connecting the fuel feed supply and the fuel injection orifices of the high power fuel injector so as to supply fuel to the fuel injection orifices.

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 gas flow separator for an annular combustion chamber of a gas turbine engine is disclosed in which the radially spaced apart first and second gas flow separator walls which extend axially into the combustion chamber from the combustion chamber upstream end wall each define a plurality of axially extending channels. Each of the channels extends generally axially in an upstream direction from the downstream edge of each of the separator walls such that their axial lengths are less than the axial lengths of the respective separator walls. Both the depth and the width of each of the channels increases from its upstream end point to the downstream edges of the separator walls. The channels may be circumferentially aligned with each other, or the channels of one separator wall may be circumferentially offset from the channels of the other separator wall.

Abstract: Truncated pie shaped bulkhead liner sections 60 are each divided into two liner segments 62. The division occurs adjacent fuel nozzle opening 20. Upstream extending lips 71, 75 and 77 abut the bulkhead 14. Cooling air passes through cooling flow openings in the bulkhead with all the flow continuing toward the shell 38, 40 edges of the liner segments.

Abstract: An improved gas turbine combustor of the type including primary and secondary combustion chambers with a venturi including a throat portion located therebetween; a plurality of primary fuel injection nozzles secured to a combustor cap in an annular array upstream of the primary combustion chamber; and a centerbody including a secondary fuel nozzle, said centerbody extending from said combustor cap to said secondary combustion chamber; the improvement comprising a plurality of tertiary fuel injection nozzles arranged in a circular array about a longitudinal axis of the combustor, at a location downstream of said venturi throat portion, for injecting fuel into the secondary combustion chamber.

Abstract: A combustor for a gas turbine having a centrally located fuel nozzle and inner, middle and outer concentric cylindrical liners, the inner liner enclosing a primary combustion zone. The combustor has an air inlet that forms two passages, each of which has a circumferential array of individually rotatable swirl vanes. The swirl vanes are mounted on axially oriented primary fuel pegs that extend through the air inlet passages. The middle and outer liners form an outer annular passage in which radially oriented secondary fuel pegs are disposed. The middle and inner liners form an inner annular passage that is supplied with cooling air. A perforated circumferentially extending baffle is locating in the inner annular passage and directs the cooling air to flow over the inner liner.

Abstract: A film starter structure for a combustor of a gas turbine engine which includes a plurality of circumferentially spaced, axially extending ribs formed on a radially inner surface of a forward section of an outer combustor liner adjacent a combustor dome. An annular ring overlays the ribs for defining a plurality of air passages. A support extends from the combustor dome and supports the outer liner about the dome. Compressor discharge air is introduced into the air passages and exits the air passages along the inner surface of the outer liner for establishing a cooling film barrier on the outer combustor liner surface. A spring seal between the combustor dome and the inner ring seats the dome within the ring and establishes a seal for preventing leakage air therebetween and allowing independent radial expansion of the liner and dome by compressing the spring seal.

Abstract: A gas turbine engine combustion chamber has primary and secondary combustion zones. An annular secondary fuel and air mixing duct surrounds the primary combustion zone. The annular secondary fuel and air mixing duct is defined at its radially outer extremity by an annular wall. The annular wall is at least partially formed by an annular manifold. The annular manifold has a number of radially inwardly extending fuel injectors which inject fuel into the secondary fuel and air mixing duct. The annular fuel manifold is mechanically isolated from the remaining portion of the annular wall by an annular gap. The annular fuel manifold is supported from the combustor casing by a fuel supply pipe which is secured to the combustor casing and the annular fuel manifold.

Abstract: A fuel control system for a gas turbine engine having a staged combustor includes apparatus and methods to operate the combustor during pilot operation, staged operation, and transition between the two modes of operation. Various construction details are developed which provide a fuel control having a switch with built in hysteresis for determining the transition points between operating modes. In one embodiment, the combustor is switched from pilot to staged operation at a thrust level greater than the thrust level at which the combustor is switched from staged to pilot operation. During transitions between operating modes, the changes in fuel flows are limited by a plurality of rate limits to prevent combustor fuel spikes and blow-outs from occurring. During engine starts, the fuel lines to the main combustor are pre-filled to provide responsiveness during the transition from pilot to staged operation.

Abstract: A gas turbine engine (10) includes a multi-stage combustion chamber (19) having a pilot stage (22) and a main stage (24), wherein during the low power setting only the pilot stage (22) is operating. During the non-operation of the main stage (24) a problem of coking occurs in a fuel supply system in general, and in a check valve (30) distributing fuel to the main stage, in particular. The check valve (30) is cooled by pilot fuel flowing through an auxiliary chamber (52) and is insulated by an air jacket (60) that is formed between the outer housing and the inner housing of the check valve.

Abstract: A double annular combustor having concentrically disposed inner and outer annular combustors is provided with inner and outer domes. A centerbody is disposed between the inner and outer domes and is constructed of a plurality of substantially rectangular segments positioned such that circumferential gaps between adjacent centerbody segments align alternatively with and between carburetors positioned in the inner and outer domes.

Abstract: A separator for an annular gas turbine engine combustion chamber is disclosed having a plurality of separator segments arranged in a generally annular configuration to form an annular separator extending from an upstream end wall of a combustion chamber into the combustion chamber so as to separate first and second combustion zones within the combustion chamber. Each of the separator segments has a hollow interior defined by a front wall, an outer wall, an inner wall which has at least a portion tapering or converging towards the outer wall in a direction extending away from the front wall, and first and second sides. The sides define an array of holes communicating with the hollow interior, the positioning of the array of holes on a first sides being different from that of the array of holes on the second sides such that the axes of the holes are out of alignment with each other.

Abstract: Apparatus for retro-fitting low NOx combustors in a conventional gas turbine using the existing openings in the cylinder casing and without any significant alteration of the turbine apparatus. Each one of the low NOx combustors is mounted by a nozzle block to the cylinder casing over an existing opening sized for a conventional combustor, the nozzle block having a plurality of fuel nozzles for spraying fuel into said combustor. The nozzle block has a pilot nozzle aperture and a plurality of annular fuel passages such that when the nozzle block is mechanically connected to the combustor the nozzles are in flow communication with the fuel passages. The nozzle block is sized such that one end can be flushly mounted to the cylinder casing over the existing opening.

Abstract: A gas turbine combustor suppresses the generation of NOx by mixing a fuel and air homogeneously. A pilot nozzle is provided at the center of the gas turbine combustor. A plurality of main nozzles surround the pilot nozzle. A diverging cone projects from the vicinity of the injection port of the pilot nozzle such that the flame from the main nozzles is sustained by the pilot nozzle and NOx is suppressed. The main nozzles can be provided upstream of the injection port of the pilot nozzle in which case an annular premixing nozzle having a throttled exit end is disposed downstream of the main nozzles. The main nozzle may also be in the form of a fuel nozzle having a plurality of tubes, a gaseous fuel being injected through one of the tubes, and liquid fuel being injected into the annular premixing nozzle through the other of the tubes so that the liquid fuel is atomized to mix the fuel and air homogeneously.

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 fuel supply system for a gas turbine engine having a multi-stage combustor provides responsiveness by blocking the drainage of fuel within the fuel lines during pilot only operation. In a particular embodiment, the fuel supply system includes a fuel source, a flow splitter valve, a main flow divider valve interconnected with a plurality of main fuel injectors, a pilot flow divider valve interconnected with a plurality of fuel injectors, and check valves disposed in the fuel lines connecting the main flow divider and the main fuel injectors. The check valves block the flow of fuel to the plurality of main fuel injectors during pilot only operation. A method of operating the fuel supply system includes built in hysteresis. The method includes the steps of switching from pilot to staged at a first thrust level TL.sub.1, and switching from staged to pilot at a second thrust level TL.sub.2 <TL.sub.1.

Abstract: A combustor for a gas turbine having a centrally located fuel nozzle and inner, middle and outer concentric cylindrical liners, the inner liner enclosing a primary combustion zone. The combustor has an air inlet that forms two passages, each of which has a circumferential array of individually rotatable swirl vanes. The swirl vanes are mounted on axially oriented primary fuel pegs that extend through the air inlet passages. The middle and outer liners form an outer annular passage in which radially oriented secondary fuel pegs are disposed. The middle and inner liners form an inner annular passage that is supplied with cooling air. A perforated circumferentially extending baffle is locating in the inner annular passage and directs the cooling air to flow over the inner liner.

Abstract: A double annular combustor having concentrically disposed inner and outer annular combustors is provided with inner and outer domes. A centerbody is disposed between the inner and outer domes and is constructed of a plurality of annular segments. Each segment includes an upper wall, a lower wall, an upstream wall and a downstream end. The upper and lower walls include flanges extending therefrom to form cavities, within which are cooling holes. Pins are also provided which extend between the upper and lower walls of the centerbody to augment the cooling and structural connection thereof.

Abstract: The radial jets of a fuel nozzle for the combustor of a gas turbine engine are non-axisymmetrically disposed around the tip of the fuel nozzle to distribute the fuel into the swirler unevenly around the circumference to reduce pattern factor.

Abstract: A gas turbine combustor having a pre-mixing zone and a combustion zone separated by a barrier plate. A plurality of concentrically arranged annular passages, into which lean mixtures of fuel and air are introduced, are disposed in the pre-mixing zone. Each of the annular passages has a discharge end in which a plurality of swirlers are circumferentially arranged that allow fuel/air mixtures flowing through the annular passages to enter the combustion zone, the swirlers acting as flame holders. A toroidal fuel manifold upstream of each of the annular passages allows the fuel/air ratio in each annular passage to be individually controlled. The firing temperature of the combustor is increased by sequentially supplying fuel to each annular passage in turn, with the innermost passages being supplied with fuel first. A pilot fuel assembly is centrally disposed with respect to the annular passages and has swirler vanes and an ignitor.

Abstract: In a method for operating a silo combustion chamber (A) of a gas turbine, this combustion chamber is equipped with a number of premixing burners (B), these burners being arranged so that they are subdivided in groups within the combustion chamber (A). At least one group is equipped with controllable premixing burners. Very high demands are made on the quality of the combustion during the operation of such a combustion chamber (A), particularly in the full-load range, in order to meet the stringent NO.sub.x emission regulations. For this purpose, the last group to be switched on is controlled at full load in accordance with the ambient conditions. A load control system also acts at part load. By this means, it is possible to prevent a fluctuation of the NO.sub.x emissions due to the varying ambient conditions.

Abstract: A combustor cap assembly for a gas turbine includes an outer sleeve (12); an impingement plate (14) fixed to a forward end of the outer sleeve. The impingement plate is provided with a plurality of primary fuel nozzle openings (18) and a single secondary fuel nozzle opening (20). A plurality of open-ended nozzle cups (28) are fixed to the impingement plate (14) in alignment with respective primary fuel nozzle openings (18); and a nozzle collar (32) is fixed to a rearward end of each nozzle cup (28).

Abstract: A gas turbine combustor is capable of reducing NOx over a whole load range of a gas turbine. The combustor comprises two burner systems. It is possible to independently control a flow rate of combustion air supplied to a burner of the respective systems and a flow rate of bypass air jetted from through a bypass air port. The air supplied to the burners is mixed with fuel in a mixing chamber and then ejected into a combustion chamber so as to form a flame. The amount of bypass air is controlled in accordance with an operation load, a humidity of air and a heating value of fuel. It is therefore possible to prevent any blow off and reduce NOx over the whole load range of the gas turbine.

Abstract: In accordance with the present invention, a diffuser for a gas turbine engine having a compressor providing compressed airflow thereto is disclosed having an outer wall and an inner wall, with a plurality of circumferential splitter vanes disposed therebetween to define a plurality of channels for diffusing airflow therethrough. The splitter vanes have an aft end and a leading edge, with the aft end being thicker and tapering in size to the leading edge. The splitter vanes also have a relatively thick bluff base region in order to better spread the airflow radially.

Abstract: A generally annular combustion chamber for a gas turbine engine is disclosed in which a plurality of generally cylindrical walls extend forwardly from an upstream end wall of the combustion chamber such that each cylindrical wall defines a cavity which is in communication with the interior of the combustion chamber. A first fuel injection head is located in each of the cylindrical walls so as to inject fuel into the cavity which is mixed with air and passes into the combustion chamber. The first fuel injection heads are located at a first axial position with respect to a longitudinal axis passing through the combustion chamber. A plurality of second fuel injection heads are located adjacent to the upstream end wall of the combustion chamber so as to spray fuel directly into the combustion chamber. The second fuel injection heads are located axially downstream of the axial positions of the first fuel injection heads.

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: 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.