Abstract: A gas turbine engine reverse-flow sheet metal combustor for an aero gas turbine engine has a louverless single-piece liner design which includes a plurality of effusion patches, containing multiple rows of effusion cooling holes, bounded by cooled effusionless bands to aid in repairability of the combustor. The cooling of the bands improves the durability of the combustor and provides a method of repair for the combustor.

Abstract: An annular turbomachine combustion chamber comprising an inner wall, an outer wall, and a chamber end wall disposed between said inner and outer walls in the upstream region of said chamber, the chamber end wall presenting a central portion and an outer assembly flange, and the outer wall presenting an upstream assembly flange, the chamber end wall and the outer wall being assembled together by their assembly flanges, in which the free edge of said outer assembly flange is curved inwards. In addition, the outer face of the outer assembly flange presents a shoulder.

Abstract: A combustor includes a first wall, a second wall, an injector grommet, a combustor longitudinal axis, and a connection assembly indirectly connecting the first wall to the second wall. A portion of the injector grommet is disposed within a groove of the connection assembly.

Abstract: A gas turbine combustion chamber includes an annular combustion chamber (3) confined by an outer (1) and an inner combustion chamber wall (2) and is provided with several fuel injection nozzles (4). Air deflectors (6) are arranged in a front-side area of the annular combustion chamber (3) over the entire combustion chamber annulus (5) concentrically around a combustion chamber axis (12) and fuel injection nozzles (4) are distributed essentially uniformly over the combustion chamber annulus (5).

Abstract: A combustion chamber which includes a chamber endwall which takes the form of a substantially frustoconical component including a succession of adjacent planar portions is disclosed. The invention can be applied to a combustion chamber of a turbine engine.

Abstract: A combustor is provided having one or more apertures adapted to asymmetrically introduce air adjacent the fuel nozzles to reduce smoke resulting from unburned hydrocarbons in a gas turbine engine combustion system by impeding escape of unburned hydrocarbons.

Abstract: A method for assembling a gas turbine combustor system is provided. The method includes providing a combustion liner including a center axis, an outer wall, a first end, and a second end. The outer wall is orientated substantially parallel to the center axis. The method also includes coupling a transition piece to the liner second end. The transition piece includes an outer wall. The method further includes coupling a plurality of lean-direct injectors along at least one of the liner outer wall and the transition piece outer wall such that the injectors are spaced axially apart along the wall.

Abstract: A combustor having liners with a plurality of angled effusion holes defined therethrough at a first angle with respect to a surface of the liners and at a second angle with respect to a corresponding radial plane. A density of the effusion holes defined in a primary section receiving the fuel nozzles is at least equal to a density of the effusion holes defined in a secondary downstream section.

Abstract: A method of assembling a combustor for use in a turbine engine is disclosed. The method includes providing a dome assembly ring including a plurality of assembly ring openings, positioning a plurality of elongated rings on the dome assembly ring, providing a cowl assembly including an inner cowl portion and an outer cowl portion that are fabricated from sheet metal, and coupling the inner and outer cowl portions to the dome assembly ring such that each of the plurality of elongated rings is coupled to the dome assembly ring.

Abstract: A one-piece combustion chamber having an end wall subdivided into sectors. The combustion chamber comprises a chamber end wall interconnecting the inner and outer walls, and it includes windows closed by removable walls, each carrying a plurality of injector systems.

Abstract: An annular combustion chamber comprising axial walls interconnected by a chamber end wall having a coefficient of thermal expansion that is different from that of the axial walls, the chamber end wall being provided with a plurality of inner and outer fastener tabs secured by fastener systems to the end portions of the axial walls. Each fastener system comprises a bolt, a nut tightened onto one of the ends of the bolt, and a slideway bushing disposed around the bolt between the nut and the end portion of the corresponding axial wall, a determined amount of radial clearance being provided between the nut and the end portion of the axial wall so as to enable the chamber end wall to expand radially freely in operation relative to the axial walls.

Abstract: The combustion chamber includes a chamber endwall perforated by at least one passage hole. An injection system is mounted so that it can slide diametrically with respect to the passage hole. A deflector is mounted on the chamber endwall. The deflector includes a fitting ring which includes first tenons. A sleeve which is coaxial to the fitting ring includes second tenons. The first and the second tenons allow the passage of the first tenons between the second tenons and then the engagement of the first tenons behind the second tenons. The injection system is slidably mounted in the sleeve. Antirotation means are provided to prevent a rotation of the injection system with respect to the sleeve.

Abstract: In a gas turbine plant and a gas turbine power generator plant, as a device of reducing the heat transfer from the gas turbine to the compressor, for instance, an air flow path for circulating part of compressed air from the compressor toward the gas turbine is formed between the gas turbine and the compressor. With the air flow path, which is a heat transfer reducing device, the heat conducted from the gas turbine to the compressor can be reduced by cooling or interception, and accordingly the temperature on the compressor side does not rise high, making it possible to prevent the performance of the compressor from deteriorating.

Abstract: A gas turbine engine comprises a rotary compressor (6) which is subject to forward thrust forces and is mounted on an axial shaft (50) which has a forward facing surface (67) located in a thrust bearing (52). A bleed of air under pressure (70, 74, 78) is provided from the compressor to the thrust bearing to form an air cushion (66, 68) opposing the forward thrust of the compressor transmitted by the shaft. The air cushion pressure automatically increases with engine speed. Other air cushion bearings (54, 56) are also provided along the shaft.

Abstract: A gas turbine engine comprises a compressor section, a combustion section disposed downstream from the compressor section, and a turbine section disposed downstream from the combustion section. The turbine section includes a turbine disk defining a plurality of turbine disk slots for accommodating turbine blades. The plurality of turbine disk slots each include an inlet having a rounded periphery at a bottom portion thereof.

Abstract: A turbine engine annular combustion chamber, including longitudinal walls connected by a transverse chamber bottom and a single-piece cowling, the chamber bottom and the cowling each including an inner flange and an outer flange each with a plurality of holes made in it for the passage of fixing systems for fixing the cowling on the chamber bottom. Each fixing system has associated with it at least one tangential slot formed on the corresponding flange of the chamber bottom and/or on the corresponding flange of the cowling, each slot being formed in the immediate vicinity of the corresponding fixing system.

Abstract: A diffuser for decelerating a compressed gas, is provided, which includes a ring divided at circumferential positions around the ring into several arcuate sections, the complete ring is assembled from these sections. The ring defines one or more passages, the cross-section of which increases in the direction of flow of a fluid through the diffuser. The interface between adjacent sections is configured so that relative movement between the sections is prevented. The interface is configured to have a series of interlocking serrations along the axial length of the diffuser. In an embodiment, there are two arcuate sections involving two interfaces, and the peaks and troughs of the serrations of each of the two sections are arranged at an angle to a longitudinal axis of the diffuser. A small relative movement between the sections is achieved when the two angles at the two interfaces have opposite signs.

Abstract: The component elements of a jet engine combustion chamber are assembled together in a manner that is designed to facilitate maintenance while also improving aerodynamic performance. In an embodiment, the outer cap and the inner cap have tongues and the chamber end wail has corresponding tongues, and the tongues of the caps are assembled to the tongues of the chamber end wall.

Abstract: A combustor having a combustor wall with a plurality of angled effusion holes defined therethrough. The tangential component of the hole direction of the effusion holes corresponds to a same rotational direction about the central axis of the combustor. The effusion holes directional arrangement is angled from a radial plane and the combustor liner surfaces in order to promote swirl at the combustor exit.

Abstract: A gas turbine engine combustor liner having a plurality of holes defined therein for directing air into the combustion chamber. The plurality of holes provide a greater cooling air flow in regions intermediate each diffuser pipe than in other areas of the combustor liner.

Abstract: The engine comprises a low pressure stage (1) and a high pressure stage (2) with two independent rotors (3,4) and associated fixed fluid preparation members (9,10,13,14), combining “compressor” means (15,16) and “expander” means (17,18) and at least one combustion chamber (19,20) in a central volume defined axially by the rotors (3,4) and laterally by three substantially cylindrical, coaxial walls (22,23,24), defining in pairs an outer duct (25) connecting the outlet of the LP compressor (15) to the inlet of the HP compressor (16) and an inner duct (26) connecting the outlet of the HP compressor (16) to the inlet of a first combustion chamber (19), whose outlet supplies the HP expander (18) with gases produced by combustion guided to the inlet of the LP expander (17), whose outlet is in communication with the outside via at least one exhaust opening (41) for the combustion gases.

Abstract: A turbomachine assembly comprising: an annular compression section; a casing formed by an outer annular shell (204) and an inner annular shell (206) secured inside the outer shell by means of a plurality of radial support arms (208); an annular combustion section housed inside the casing; and an annular turbine section.

Abstract: A low-emission method for producing power using a gas turbine includes premixing a plurality of fuel and air mixtures, injecting the fuel and air mixtures into a combustion chamber using a plurality of fuel nozzles, and adjusting a ratio of fuel and air injected by at least one of the nozzles to control a fuel/air concentration distribution within the combustion chamber.

Abstract: A combustor includes two or more sets of effusion cooling holes that extend through the inner and outer liners. Each set of effusion cooling holes includes one or more initial rows of effusion cooling holes, one or more final rows of effusion cooling holes disposed downstream of the one or more initial rows, and a plurality of interposed rows of effusion cooling holes disposed between the initial and final rows. Each effusion cooling hole is disposed at a tangential angle relative to an axial line. The tangential angle of the effusion cooling holes in each set of effusion cooling holes gradually transitions from a substantially transverse tangential angle in each initial row to a substantially axial tangential angle in each final row.

Abstract: A gas turbine engine fuel nozzle includes an axis of symmetry extending therethrough, the nozzle body including a first passage extending coaxially therethrough, a second passage, and a third passage, the second passage circumscribing the first passage, the third passage formed radially outward of the second passage, and a nozzle tip coupled to the nozzle body, the nozzle tip including at least one primary discharge opening in flow communication with the first passage, at least one secondary discharge opening in flow communication with the second passage, and at least one tertiary discharge opening in flow communication with the third passage.

Abstract: A transition piece seal assembly for sealing an interface between at least one transition piece extending between a turbine combustor and a first stage turbine nozzle. The seal assembly includes an aft frame having on a first axial side thereof at least one axially projecting can shaped receptacle for axially receiving an aft end of a transition piece and a generally planar mounting surface on a second axial side thereof for being disposed in opposed facing relation to the first stage nozzle; and at least one resilient seal element disposed on an inner peripheral surface of the can shape receptacle so as to be disposed between the transition piece aft end and the can shaped receptacle.

Abstract: A method of operating a gas turbine engine is provided. The method includes discharging pilot fuel into a combustion chamber from a nozzle through at least one pilot fuel outlet defined in a tip of the nozzle, discharging steam from the nozzle through a plurality of steam outlets that are spaced circumferentially about the plurality of pilot fuel outlets, and discharging primary fuel from the nozzle through a plurality of primary fuel outlets that are circumferentially aligned with the plurality of steam outlets. To facilitate mixing the primary fuel with the steam, the primary fuel is discharged from the nozzle tip at an oblique angle with respect to a centerline extending through the nozzle tip.

Abstract: A heat shield for a gas turbine engine combustor includes an apparatus for providing a spiral flow to improve at least the cooling of the heat shield.

Abstract: A gas turbine engine combustor liner having a plurality of holes defined therein for directing air into the combustion chamber. The plurality of holes provide a greater cooling air flow in fuel nozzle regions than in other areas of the combustor liner.

Abstract: A method and system for operating at partial load a gas turbine system having a compressor, a combustor, and a turbine. The method and system may include the steps of lowering a flow of fuel to the combustor, extracting air from the compressor so as to lower a flow of air to the compressor, and returning the extracted air to the turbine or a component of the gas turbine system other than the combustor. Extracting air from the compressor raises a combustion temperature within the combustor. Raising the combustion temperature maintains a combustion exhaust below a predetermined level, maintains stable combustion, and extends turbine turndown values.

Abstract: A cover plate for use with a rotor disk in a gas turbine engine includes an enclosed chamber associated with a web on the rotor disk. The enclosed chamber ensures that adequate cooling air is delivered by rotation of the cover plate.

Abstract: A brush seal assembly is formed of a bristle array receiving at least one housing member. Alignment structures are formed on the housing and bristle assembly to ensure that the bristle assembly will be properly positioned in a working environment.

Abstract: This invention provides a seal assembly for sealing a gap between first and second relatively movable members. The first member is provided with a sealing surface, and the second member is provided with a sacrificial abradable layer that is positioned opposite the sealing surface of the first member. When the seal assembly is in operation the sealing surface of the first member moves relative to, and abrades, the abradable layer of the second member thereby maintaining a seal between the members. The present invention is characterised in that, the second member and the abradable layer comprise the same composite material. The second member and the abradable layer may be co-cured together during manufacture.

Abstract: An annular combustion chamber for a turbomachine comprises inner and outer annular walls united by a transverse wall, the inner and outer walls are extended at their downstream ends by inner and outer fastener flanges for fastening respectively to inner and outer shells of a turbomachine casing in order to hold the combustion chamber in position, the inner flange being provided with a plurality of holes for feeding cooling air to a high pressure turbine of the turbomachine, the air feed holes through the inner flange being distributed circumferentially over at least two rows disposed in a staggered configuration.

Abstract: A blade outer air seal system includes a body that extends between two circumferential sides, a leading edge and a trailing edge, and a radially inner side and a radially outer side. An attachment section associated with the body and includes at least one engagement surface that is transverse to the radially outer side.

Abstract: A system for controlling the operating lines of a gas turbine engine comprising a compression system which compresses an incoming flow of a gas, a bleed system which extracts gas from a bleed location in the compression system prior to the gas entering a throttling orifice, a recirculation system which reintroduces at least a portion of the gas extracted by the bleed system into a feed location in the compression system or upstream from it, a control system which controls the recirculated gas such that the operating lines of the compression system are changed such that the stall margin of the compression system is increased.

Abstract: A reverse flow combustor for a gas turbine engine having an outer combustor liner and an inner combustor liner defining an annular combustion chamber, and a compound-angle frustoconical portion in the outer liner having a first and second conical slopes towards an engine centreline.

Abstract: Impingement hole patterns are provided in a combustor dome panel to provide efficient cooling of a combustor dome heat shield. The impingement holes are substantially aligned or located above partial height protuberances provided on the back face of the heat shield.

Abstract: A method of assembling a dome assembly for use in a turbine engine combustor is disclosed. The method includes providing a dome assembly ring, an inner liner portion, and an outer liner portion. The method also includes coupling the inner liner portion and the outer liner portion to the dome assembly ring, positioning a plurality of rings on the dome assembly ring, and coupling at least one of an inner cowl and an outer cowl to the dome assembly ring such that each of the plurality of rings and at least one of the inner and outer cowls are removable without uncoupling the dome assembly ring from either the inner or outer liner portions.

Abstract: A dual mode combustor of a gas turbine engine contains at least one dual mode combustor device having a combustion chamber, a fuel air mixing element, a high frequency solenoid valve and a fuel injector. During a first mode of operation the dual mode combustor device operates in a steady, constant pressure deflagration mode, receiving its fuel from the fuel injector. In a second mode of operation the dual mode combustor device operates in a pulse detonation mode, receiving its fuel from the high frequency solenoid valve.

Abstract: A tip turbine engine (10) according to the present invention includes an engine support structure (12) for cantilevering a load bearing shaft from an engine support plane (P). An engine support structure (12) defines the engine support plane (P), which is perpendicular to an engine centerline (A). The shaft (60) is rotationally fixed, coaxial with the engine centerline (A), and structurally supported by the engine support structure (12) such that a load borne by the shaft (60) is transferred along the engine support plane.

Abstract: An annular turbine engine combustion chamber configured so that in an axial half-section, values of first acute angles formed between a line that is effectively the median of the half-section located between an external axial wall and an internal axial wall, and the principal directions, in this half-section, of the holes in an external portion of a chamber base, decrease as a function of the distance between the holes and this line that is effectively the median. Further, values of second acute angles formed between the line that is effectively the median and the principal directions, in this half-section, of the holes in an internal portion of the chamber base, decrease as a function of the distance between the holes and this line that is effectively the median.

Abstract: A combustion chamber 10 for a gas turbine that operates in a highly diluted mode of combustion has a substantially toroidal plenum 30. At least one combustion fluid injector is located at the periphery of the plenum. The shape of the plenum and the disposition of the combustion fuel injectors are such that a vortex flow 2 is established in the combustion chamber in use. Gasses are entrained in the vortex which provides high enough levels of flue gas re-circulation to allow for the onset of highly diluted combustion.

Abstract: A combustor for a gas turbine engine includes a sheet metal combustor wall having a plurality of cooling apertures therein immediately upstream of a corner between two intersecting combustor wall portions.

Abstract: Annular combustion chamber of a turbomachine, including an inner wall, an outer wall (28) and a chamber bottom (30) arranged between said walls in the upstream region of said chamber, in which the chamber bottom (30) is divided into several adjacent sectors (130), each sector being attached to said walls (28), these sectors (130) presenting lateral edges such that the lateral edges (130a, 130b) of two adjacent sectors overlap. More precisely, each sector (130) includes a lip (60) extending along one of its lateral edges (130a), this lip projecting relative to one of the faces of this sector and covering the lateral edge (130b) of the adjacent sector.

Abstract: Annular combustion chamber (24) of a turbomachine, including an inner wall, an outer wall (28), a chamber bottom (30) disposed between said walls in the upstream region of said chamber, and two attachment flanges (27, 29) disposed downstream of the chamber bottom and respectively enabling said walls to be attached to other parts of the turbomachine, each wall being divided into several adjacent sectors (128) and each sector being attached to the chamber bottom and to one of the attachment flanges. Advantageously, said adjacent sectors (128) overlap at their lateral edges and there exists a degree of radial play between two adjacent sectors. In addition, the lateral edges (128a, 128b) of the sectors (128) are inclined circumferentially relative to the principal axis of the combustion chamber.

Abstract: The invention relates to an annular wall (10) intended to connect transversely longitudinal walls of an annular combustion chamber of a turbine engine. The wall (10) is essentially flat, inclined in relation to a longitudinal axis of the turbine engine, and comprises a plurality of deflectors (16), each formed by an essentially rectangular flat sheet. The deflectors are mounted on the wall and each comprise an aperture for the installation of a fuel injection system, a plurality of multi-perforation holes (18) formed in relation to the deflectors (16) around their aperture so as to allow a passage of air intended for the cooling of the deflectors, and means (20) to force the flow of air for cooling the deflectors to flow radially around the fuel injection systems.

Abstract: A fabricated cowl for a double annular combustor of a gas turbine engine having a longitudinal axis extending therethrough, including an outer annular portion, an inner annular portion, and a middle annular portion. A plurality of circumferentially spaced radial slots are formed in at least one of the annular portions. The outer annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of outer radial members connecting the outer annular portion and the middle annular portion are provided between adjacent windows therein. Similarly, the inner annular portion includes a plurality of circumferentially spaced windows of desired size formed therein, wherein a plurality of inner radial members connecting the inner annular portion and the middle annular portion are provided between adjacent windows therein.

Abstract: The invention relates to a turbomachine annular combustion chamber (1) comprising an outer axial wall (2), an inner axial wall (4), and a chamber bottom end (8) connecting the said walls, the chamber bottom end being provided with several passages (34,36,72) enabling initiation of a cooling air film (D2) along the hot inner surface (30) of the outer axial wall and initiation of a cooling air film (D1) along the hot inner surface (32) of the inner axial wall, the outer axial wall and the inner axial wall being multi-perforated roughly over their full length in order to enable reinforcement of the cooling air films. According to the invention, the outer axial wall and the inner axial wall are provided, in an upstream part, with a first zone (54,40) of perforations (38) formed such that cooling air is introduced inside the combustion chamber in reverse flow.

Abstract: A combustion chamber 10 for a gas turbine operates in a highly diluted mode of combustion has a substantially toroidal plenum 30. At least one combustion fluid injector is located at the periphery of the plenum. The shape of the plenum and the disposition of the combustion fuel injectors are such that a vortex flow 2 is established in the combustion chamber in use. Gasses are entrained in the vortex which provides high enough levels of the gas re-circulation to allow for the onset of highly diluted combustion.