Abstract: A combustor structure of an embodiment includes: a plurality of combustors; a rear liner which guides a combustion gas from the plurality of combustors to a downstream side; and a scroll which guides the combustion gas from the rear liner in an axial direction and a circumferential direction of a turbine rotor. Each of the combustors includes: a combustor liner; a fuel supply part which supplies fuel into the combustor liner; and an oxidant supply part which supplies a swirling flow of an oxidant from the circumference of the fuel supply part into the combustor liner. The plurality of combustors include: the combustor in which a swirling direction of the swirling flow of the oxidant is clockwise; and the combustor in which a swirling direction of the swirling flow of the oxidant is counterclockwise.

Abstract: A trapped vortex combustor for use in a gas turbine engine includes an outer vortex chamber wall and a dome attached to, or formed integrally with, the outer vortex chamber wall. The dome, the outer vortex chamber wall, or both define at least in part an outer trapped vortex chamber and a channel. The channel extends along the circumferential direction at a forward end of the outer vortex chamber wall, the channel configured to receive an airflow through or around the outer vortex chamber wall, the dome, or both and provide such airflow as a continuous annular airflow to the inner surface of the outer vortex chamber wall. The dome further defines a fuel nozzle opening, with all openings in the dome outward of the fuel nozzle opening along the radial direction, excepting any effusion cooling holes having a diameter less than about 0.035 inches, being in airflow communication with the channel.

Abstract: A liner cooling structure of a duct assembly reduces pressure loss generated in the compressed air flow for cooling the liner. The duct assembly includes a liner, a transition piece, and a flow sleeve, and the transition piece and the flow sleeve form a transition piece channel through which a main stream of compressed air is introduced to the duct assembly. The liner cooling structure includes a first flow passage through which the main stream of compressed air passes in a first direction; and a second flow passage formed as a plurality of inlet holes in the flow sleeve to communicate with the first flow passage and configured to pass an auxiliary stream of compressed air in a second direction from outside the flow sleeve to inside the flow sleeve, the auxiliary stream joining the main stream such that the second direction forms an acute angle with the first direction.

Abstract: A method of operating a combustion burner to heat a furnace. Fuel and combustion air are supplied into a combustion zone and ignited. Additional combustion air is supplied into the atmosphere outside of the combustion zone. The amount of additional combustion air supplied outside of the combustion zone is decreased as a temperature of the atmosphere inside the furnace increases such that the content of nitrogen oxides (NOx), as corrected for 3% O2 (cNOx (3% O2)), in the gases generated by combustion of the fuel and the combustion air and emitted from the furnace maintained below a predetermined value.

Abstract: A combustor for a turbine engine includes an inner liner panel with a multiple of inner dilution passages. The multiple of dilution passages includes a repeating pattern of a first major inner air passage, a minor inner air passage, and a second major inner air passage. A combustor for a turbine engine includes an outer liner panel with a multiple of outer dilution passages. The multiple of outer dilution passages includes a repeating pattern of a first major outer air passage, a first minor outer air passage, a second major outer air passage, and a second minor outer air passage.

Abstract: A combustion system and a method of combustion for a gas turbine engine includes a combustor liner defining a combustion chamber. A plurality of fuel nozzle sets extend into, and supply fuel flow to, the combustion chamber. A pilot fuel nozzle injects a first fuel spray in a tangential direction relative to the combustion liner and toward the upstream end of the combustion chamber. A main fuel nozzle injects a second fuel spray toward the exit end of the combustion chamber. At ignition conditions, a majority of the fuel flow is injected through the pilot fuel nozzles, and at high power conditions a majority of the fuel flow is injected through the main fuel nozzles. At high power conditions, a fuel rich mixture is supplied to the combustion chamber, and a row of quench jets are configured to supply air to the combustion chamber, providing rich-quench-lean combustion.

Abstract: A spur gear transmission with at least two intermeshing spur gears the teeth of which are in meshing engagement with one another, and each of which are rotatable via an axis of rotation. The spur gear transmission has an enveloping wall which surrounds the two spur gears in circumferential direction and in the direction of the axes of rotation. The enveloping wall has an inner contour which is adapted to the outer diameter of the spur gears in such a manner that two annular gaps which merge into each other are formed between the enveloping wall and the spur gears. One annular gap respectively is arranged at least substantially concentrically with respect to a respective axis of rotation. The enveloping wall is formed by a two-part housing which is closed along a parting line and which has an upper part and a lower part.

Abstract: An injector includes: an upstream support plate into which a fuel gas is to be introduced and which has a shape of a tapered cylinder with a diameter which enlarges; a downstream support plate that defines a plenum at an inner side along with the upstream support plate; and premixing tubes, each of which is supported by the upstream support plate and the downstream support plate and is configured to introduce air. The premixing tubes are disposed in circular rows at equal intervals in a circumferential direction, and portions of the premixing tubes which are located in the plenum include fuel introducing holes.

Abstract: An annular wall for an annular turbine engine combustion chamber, including one annular row of dilution orifices including larger area orifices and smaller area orifices, and a multi-perforation formed from micro-perforations distributed as an upstream row, a downstream row, and at least one intermediate row interrupted by the orifices, a geometric ratio being defined as the quotient obtained by dividing the maximum spacing L between any two points on the edge of the orifice measured along a direction parallel to an axis of the wall by the maximum spacing l between any two points on the edge of this orifice measured along a direction perpendicular to the axis, the geometric ratio of the larger area orifices being greater than or equal to 1, and the geometric ratio of the smaller area orifices being greater than the geometric ratio of the larger area orifices.

Abstract: The invention refers to a reheat burner that includes a flow channel for a hot gas flow with a lance arranged along said flow channel, protruding into the flow channel for injecting a fuel over an injection plane perpendicular to a channel longitudinal axis, wherein the channel and lance define a vortex generation zone upstream of the injection plane and a mixing zone downstream of the injection plane in the hot gas flow direction. The mixing zone provides at least one axially region having different cross sectional areas along its longitudinal axis with continuously changing shape, or having non circular cross section areas which change location along its longitudinal axis by continuously rotation around the longitudinal axis.

Abstract: A combustor with a liner where each of the walls has a respective circumferential row of dilution holes defined therethrough adjacent a junction between the primary zone and the dilution zone. In the primary zone, the inner surface of each of the walls is covered by at least one heat shield attached thereto and spaced apart therefrom to allow air circulation between the inner surface and the at least one heat shield, the walls each having a plurality of cooling holes defined therethrough having a smaller diameter than that of the dilution holes. In the dilution zone, the inner surface of each of the walls is free of heat shields, and the walls each have a plurality of effusion cooling holes defined therethrough and having a smaller diameter than that of the dilution holes.

Abstract: A combustor for a gas turbine includes a fuel nozzle having a central swirler that circumferentially surrounds a downstream end of the fuel nozzle. A primary combustion zone is defined within the central swirler. The combustor further includes an outer swirler that circumferentially surrounds at least a portion of the central swirler and a venturi that is disposed downstream from the primary combustion zone. The venturi includes an inner surface. The central swirler imparts angular swirl to a compressed working fluid so as to rapidly mix and react the fuel rich primary zone products with the working fluid. The outer swirler imparts angular swirl to a compressed working fluid so as to provide a cooling boundary layer along the inner surface of the venturi.

Abstract: An air and fuel injection system for the back of a turbomachine's annular combustion chamber, including a central injector and a peripheral annular fuel injector including at least one fuel injection aperture made in an annular wall and emerging in a peripheral annular channel separated from a central channel by the annular wall and having an annular space of admission of air. The system also includes multiple air ejection apertures made in the annular wall downstream from the fuel ejection aperture or apertures in reference to the flow of the air stream, for an additional injection of air into the channel.

Abstract: A method for operating a hydrogen-fueled gas turbine is provided wherein a supply of fuel is passed to a gas turbine combustor, and a supply of nitrogen and sufficient air to provide at least sufficient compressed air to the gas turbine for fuel combustion is passed to a compressor. A sufficient portion of the compressor discharge flow is passed to a combustor for fuel rich combustion of the fuel flow to the combustor and the fuel is combusted to produce hot combustion gases that are, in turn, passed to a turbine.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes a forward bulkhead, an inner radial combustor wall, and an outer radial combustor wall. The bulkhead includes a plurality of circumferentially disposed injector apertures. The inner radial combustor wall includes a plurality of inner quench aperture sets. Each inner quench aperture set includes a first inner quench aperture and a second inner quench aperture separated from each other by an inner interset distance. Each inner quench aperture set is separated from an adjacent inner quench aperture set by an inner intraset distance. The inner intraset distance is different than the inner interset distance. The outer radial combustor wall includes a plurality of circumferentially disposed outer quench apertures. The outer radial combustor wall is disposed radially outside of the inner radial combustor wall, thereby defining an annular combustion region therebetween.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes an annular inner liner; an annular outer liner circumscribing the annular inner liner; and a combustor dome having a first edge coupled to the annular inner liner and a second edge coupled to the annular outer liner, the combustor dome forming a combustion chamber with the annular inner liner and the annular outer liner. The combustion chamber accommodates fluid flow through the annular inner and annular outer liners. The combustion chamber converges in the direction of the air flow to reduce a diameter of the combustion chamber. The combustor dome is configured to bifurcate the air flow at the combustor dome into a first stream directed to the annular inner liner and a second stream directed to the annular outer liner.

Abstract: A double annular radially staged combustor for a gas turbine engine having a reduced internal flow transitioning area between an outboard primary combustion zone and al inboard secondary combustion zone. The primary combustion zone is an annular shell form with a radially inboard annular exit. An outboard ring with thru holes and an inboard annular channel yields a quick quench zone. The secondary combustion zone, adjacent to and downstream of the quick quench zone is of an annular shell form, generally radially inboard of the primary combustion zone. In a preferred embodiment the secondary combustion zone exiting hot gases flows into a dilution zone prior to exiting the combustor.

Abstract: A combustor assembly includes a convergent segment followed by a divergent segment to advantageously improve combustion. The combustor assembly includes a first segment beginning at a forward end that transitions to a second segment past a transition segment in a direction along a combustor axis toward an aft end. The reduction in cross-sectional area within the first segment provides desirable fuel and air mixing properties. The convergent first segment in combination with the divergent second segment decreases residence time of fuel-air mixture within the combustor chamber that decreases production of undesirable emissions from the combustor assembly.

Abstract: Fuel and air are injected in a first poloidal flow in a first poloidal direction within a first annular zone of an annular combustor. A first combustion gas from the at least partial combustion of the fuel and air is discharged into an annular transition zone of the annular combustor and transformed to a second combustion gas therein within an at least partial second poloidal flow followed by an at least partial third poloidal flow in the annular transition zone, wherein the direction of the second poloidal flow is opposite to that of the first and third poloidal flows. The second combustion gas is discharged into a second annular zone of the annular combustor, and then transformed to a third combustion gas therein before being discharged therefrom, responsive to which a back pressure is generated in the annular combustor.

Abstract: An annular combustor for a gas turbine has a combustion chamber having an interior volume that, in longitudinal section, includes a forward volume, an intermediate volume and an aft volume. The forward volume represents from about 30% to about 40% of the combustor interior volume, the intermediate volume represents from about 10% to about 20% of the combustor interior volume, and the aft volume represents from about 40% to about 60% of the combustor interior volume.

Abstract: A system and method of reducing gas turbine nitric oxide emissions includes a first combustion stage configured to burn air vitiated with diluents to generate first combustion stage products. A second combustion stage is configured to burn the first combustion stage products in combination with enriched oxygen to generate second combustion stage products having a lower level of nitric oxide emissions than that achievable through combustion with vitiated air alone or through combustion staging alone.

Abstract: An oxyfiring system and method for capturing carbon dioxide in a combustion process is disclosed. The oxyfiring system comprises (a) an oxidation reactor for oxidizing a reduced metal oxide; (b) a decomposition reactor wherein a decomposition fuel is combusted and oxidized metal oxide sorbents are reduced with oxygen being released and a flue gas with an oxygen enriched carbon dioxide stream is produced; (c) a fuel combustion reactor for combusting a primary fuel and the oxygen enriched carbon dioxide stream into a primary flue gas; and (d) separation apparatus for separating a portion of the primary flue gas so that a carbon dioxide enriched stream can be prepared. The method comprises providing a primary fuel and an oxygen enriched carbon dioxide stream to a fuel combustion reactor. The primary fuel and oxygen enriched carbon dioxide stream are combusted into a primary flue gas stream which is split into a first flue gas portion and a second flue gas portion.

Abstract: A combustor for a gas turbine engine is provided. The combustor includes an inner liner; an outer liner circumscribing the inner liner; and a combustor dome having a first edge coupled to the inner liner and a second edge coupled to the outer liner. The combustor dome forms a combustion chamber with the inner liner and the outer liner. The combustion chamber receives air flow through the inner and outer liners, and the combustor dome is configured to bifurcate the air flow at the combustor dome into a first stream directed to the inner liner and a second stream directed to the outer liner.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

Abstract: A combustor may include an outer liner having a first group of air admission holes and defining a plurality of outer liner regions. The combustor may further include an inner liner having a second group of air admission holes and defining a plurality of inner liner regions. The first group of air admission holes within a respective outer liner region may include a first plunged air admission hole approximately axially aligned with the respective fuel injector, a second plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a first adjacent outer liner region, the first air admission hole being downstream of the second air admission hole, and a third plunged air admission hole approximately on the outer boundary line between the respective outer liner region and a second adjacent outer liner region.

Abstract: A combustor assembly includes a convergent segment followed by a divergent segment to advantageously improve combustion. The combustor assembly includes a first segment beginning at a forward end that transitions to a second segment past a transition segment in a direction along a combustor axis toward an aft end. The reduction in cross-sectional area within the first segment provides desirable fuel and air mixing properties. The convergent first segment in combination with the divergent second segment decreases residence time of fuel-air mixture within the combustor chamber that decreases production of undesirable emissions from the combustor assembly.

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

Abstract: A gas turbine combustion system is provided comprising a combustion chamber (16) having a central axis (44), a primary combustion stage (28) located at a front end (32) of the combustion chamber (16) for injecting fuel, air, or mixtures thereof substantially along the central axis (44), a plurality of secondary combustion stages (30A-D) spaced apart in flow series along a length of the combustion chamber (16), wherein each of the plurality of secondary combustion stages (30A-D) comprises a plurality of circumferentially-spaced secondary injectors (48) for injecting fuel, air, or mixtures thereof, toward the central axis (44), and wherein an internal diameter of the combustion chamber (16) decreases from at least a first one of the plurality of secondary combustion stages (30AD) to at least a second one of the plurality of secondary combustion stages (30A-D).

Abstract: A combustion apparatus and duct structure for a gas turbine engine are provided. The combustion apparatus comprises a combustion system to receive fuel and air, ignite at least a portion of the fuel and air and output a stream of first combustion products and any remaining fuel and air. The combustion apparatus further comprises structure positioned adjacent to the combustion system for receiving and accelerating the first combustion products and any remaining fuel and air from the combustion system. The duct structure receives the first combustion products and any remaining fuel and air from the combustion apparatus, allows any remaining fuel and air to combust to generate second combustion products, accelerates the first and second combustion products and outputs the first and second combustion products to a first row blade assembly.

Abstract: A combustion apparatus and duct structure for a gas turbine engine are provided. The combustion apparatus comprises a combustion system to receive fuel and air, ignite at least a portion of the fuel and air and output a stream of first combustion products and any remaining fuel and air. The combustion apparatus further comprises structure positioned adjacent to the combustion system for receiving and accelerating the first combustion products and any remaining fuel and air from the combustion system. The duct structure receives the first combustion products and any remaining fuel and air from the combustion apparatus, allows any remaining fuel and air to combust to generate second combustion products, accelerates the first and second combustion products and outputs the first and second combustion products to a first row blade assembly.

Abstract: A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.

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

Abstract: Gate valve (10) for a system controlling cooling in a turbomachine, comprising a gate (20) mounted so that it can pivot about an axis (16) between a position in which it shuts off an air passage orifice (72) and a position in which it opens the orifice (72), and means (12, 14, 18) for rotating the gate (20) about the axis (16), these means comprising two superposed coaxial rotary members (14, 18) collaborating with one another via cam surfaces (40, 46) designed such that rotating the first rotary member (14) from the shut-off position causes a translational movement of the second rotary member (18) and of the gate (20) along the axis of rotation (16) followed by a rotation of this member (18) and of the gate (20) about this axis. Application of the invention to the cooling of turbojet jet pipe nozzle flaps.

Abstract: A combustor includes an outer liner having a first row and a second row of circumferentially distributed air admission holes penetrating therethrough. The combustor further includes an inner liner circumscribed by the outer liner and having a first row and a second row of circumferentially distributed air admission holes penetrating therethrough. The combustor further includes a plurality of fuel injectors extending into the combustion chamber and configured to deliver an air-fuel mixture to the combustion chamber. Each of the plurality of fuel injectors are associated with two air admission holes in the first row of the outer liner, two air admission holes in the first row of the inner liner, four air admission holes in the second row of the outer liner, and two air admission holes of the second row of the inner liner.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: An annular combustor has an annular inner shell, an annular outer shell and a dome end wall connecting the inner shell with the outer shell. A dam is positioned between the dome end wall and the exit end and extends radially from at least one of the inner shell and the outer shell to provide a reduced flow area passageway within the combustion chamber.

Abstract: A combustor for a gas turbine engine includes inner and outer liners with a row 112, 114 of dilution air holes penetrating through each liner. The row 114 of holes in the outer liner comprise at least a set of large size, major outer holes 116 and may also include a set of smaller size minor outer holes 118 circumferentially intermediate neighboring pairs of the major outer holes. The row 112 of holes in the inner liner include dilution air holes circumferentially offset from the major outer holes and may also include a set of minor holes 122 circumferentially intermediate major inner holes 120. The major and minor holes admit respective major and minor jets of dilution air into the combustor. The unique distribution of major and minor holes and the corresponding major and minor dilution air jets helps to minimize NOx emissions and regulates the spatial temperature profile of the exhaust gases discharged from the combustor.

Abstract: A bypass air injection scheme for a combustor of a gas turbine. Combustor includes a body with an inner liner and a casing enclosing the body with a passageway defined therebetween. A predetermined amount of the compressor discharge air passing through the passageway is extracted through a manifold. A conduit feeds the extracted air into an injection manifold having a plurality of injection tubes for injecting the extracted air into the combustor bypassing the reactor. The injection tubes and the injection manifold are disposed in a substantially common axial plane.

Abstract: The present invention deals with a turboreactor comprising at least a precombustion chamber, a combustion chamber, means for directing combustion gases from the precombustion chamber into the combustion chamber, an air compressor, and means for dividing the air flow from the air compressor while guiding it in the combustion chamber. The turboreactor comprises advantageously a contrarotating stator. The invention relates also to a process in which the blades are heated and pushed in a thermoplastic layer of the body of the rotor and/or stator.

Abstract: A combustor for a gas turbine engine operates with high combustion efficiency and low carbon monoxide, nitrous oxide, and smoke emissions during low, intermediate, and high engine power operations is described. The combustor includes a mixer assembly including a pilot mixer and a main mixer. The pilot mixer includes a pilot fuel injector, at least one swirler, and an air splitter. The main mixer extends circumferentially around the pilot mixer and includes a plurality of fuel injection ports and a conical air swirler upstream from the fuel injection ports. During idle engine power operation, the pilot mixer is aerodynamically isolated from the main mixer, and only air is supplied to the main mixer. During increased power operations, fuel is also supplied to the main mixer, and the main mixer conical swirler facilitates radial and circumferential fuel-air mixing to provide a substantially uniform fuel and air distribution for combustion.

Abstract: A structure of a 50 pound thrust level gas turbine jet engine combustor is a small type 50 pound gas turbine jet engine combustor. The combustor includes a combustor outer liner, a combustor inner liner and a hub. Pressure swirl fuel atomizer and thin film cooling devices are further used. In the combustor, a primary jet flow is matched with a jet tube so as to build a stable recirculation in the primary zone. 12 SIMPLEX fuel atomizers are installed in the inner liner for jetting fuel radially. The fuel must be jetted properly, if it is too close the front plate, the droplet will impact the front plate, then the downstream will possible destroy the recirculation structure. Therefore, if the fuel is jetted in proper positions, then the retaining time of fuel will be reduced greatly and the burning in the primary zone can be enhances. By analyzing and testing, it is appreciated that high temperature will induce at the middle section near the wall of the inner liner.

Abstract: A gas turbine engine for generating electricity having low NOx emissions that includes a turbine system linearly and axially connected by a power shaft to a compressor section and having a combustion section mounted vertically relative to the turbine section and the compressor section. The combustion system includes a plurality of individual combustors mounted in a circular or annular array around the upper cap or dome of the combustion system, each of said combustors being a dual mode, two-stage, emitting low levels of NOx. Each combustor exhausts its combustion gases into a common central plenum chamber that is vertically oriented relative to the turbine engine centerline. The plenum provides the hot gases to the turbine blades through an annular chamber 360 degrees around the shaft. The gas turbine engine vertical combustion system provides for a highly efficient, low nitric oxide emissions while allowing for uniform mixing of the combusting gases powering the turbine system.

Abstract: A fuel injection system for a gas turbine engine combustor, wherein the combustor includes a dome inlet module having a plurality of flow passages formed therein and at least one cavity formed in a liner downstream of said dome inlet module. The fuel injection system includes a fuel supply and a plurality of fuel injector bars positioned circumferentially around and interfacing with the inlet dome module. The fuel injector bars are in flow communication with the fuel supply, with each of the fuel injector bars further including a body portion having an upstream end, a downstream end, and a pair of sides. At least one injector is formed in the downstream end of the body portion and in flow communication with the fuel supply, whereby fuel is provided to the cavity through the fuel injector bars in accordance with a Rich-Quench-Lean (RQL) process.

Abstract: A combustor includes outer and inner liners joined at one end to an annular dome. A swirl cup is mounted to the dome and includes a flare cone extending therethrough. The flare cone has an obtuse flare angle for discharging fuel in a wide spray from an outlet of the cone. A flared beat shield surrounds the cone, and extends in part aft from the cone outlet, and is offset in part forwardly from the cone outlet to define a radially outwardly facing step.

Abstract: A combustor liner includes a row of primary dilution holes disposed aft of a forward end thereof, and a row of secondary dilution holes disposed between the primary holes and an aft end of the liner. The secondary holes vary in size circumferentially to a maximum size offset circumferentially from the primary holes.

Abstract: A method of combusting fuel in the combustor of a gas turbine engine is disclosed which includes injecting fuel and a first predetermined amount of airflow into the combustion chamber to form a fuel-rich, highly mixed, uniform distribution fuel-air spray pattern flowing downstream in the combustion chamber and introducing a second predetermined amount of airflow into the fuel air spray pattern from combustor air inlets positioned at a first predetermined distance downstream from the combustor dome with the first predetermined distance being greater than 0.75 times the dome height and said second predetermined amount of airflow being sufficient to cause rapid mixing and quenching of the rich fuel-air mixture to a lean fuel-air mixture.

Abstract: A hybrid combustor for a gas turbine engine includes a plurality of circularly arrayed ceramic can combustors whose outlets communicate with the inlet of an annular, metal combustor. The combustion process is continuous through the plurality of can combustors and into the single annular combustor. Preferably only fuel-rich combustion occurs within each of the can combustors, and fuel-lean combustion continues within the single annular combustor.

Abstract: In a combustion chamber, the hot gases are prepared sequentially via two stages (20, 40). Arranged at the end of the first stage (20) in the direction of flow is a cross-sectional constriction (30) via which the hot gases (21) prepared in the first stage (20) are passed over into the second stage (40). The Mach number at the outlet (31) of this cross-sectional constriction (30) corresponds to the area ratio of outlet area (A2) over inlet area (A1). This results in a low-reflection configuration in which low-frequency vibrations are absorbed to a significant extent. And acoustic energy reflected from the turbine is substantially reduced.

Abstract: A combustor cooperating with a compressor in driving a gas turbine includes a cylindrical outer combustor casing. A combustion liner, having an upstream rich section, a quench section and a downstream lean section, is disposed within the outer combustor casing defining a combustion chamber having at least a core quench region and an outer quench region. A first plurality of quench holes are disposed within the liner at the quench section having a first diameter to provide cooling jet penetration to the core region of the quench section of the combustion chamber. A second plurality of quench holes are disposed within the liner at the quench section having a second diameter to provide cooling jet penetration to the outer region of the quench section of the combustion chamber.

Abstract: A combustor for generating hot gas in order to rotate a turbine is able to burn fuel with a high vanadium content without emitting appreciable vanadium pentoxide to the turbine. The combustor includes a primary combustion zone for burning fuel in a rich fuel/air environment a quenching zone for cooling the hot gas and unburned fuel a secondary combustion zone for burning previously-unburned fuel. The fuel burned in the secondary combustion zone raises the temperature in the secondary combustion zone above the temperature of the quenching zone. The result of burning the fuel in a rich environment in the primary combustion zone is that most of the vanadium present in the fuel forms vanadium trioxide or vanadium tetroxide rather than vanadium pentoxide. Cooling the hot gas quickly in the quenching zone removes most of the remaining vanadium pentoxide by transforming it to vanadium trioxide or vanadium tetroxide.