Abstract: An annular combustor includes an annular outer shell that includes a first flange that defines an inner diameter IDOS and an annular inner shell that includes a second flange that defines an outer diameter ODIS. An annular hood includes a radially outer hood flange and a radially inner hood flange. A bulkhead includes a radially outer bulkhead flange that defines an outer diameter ODB and a radially inner bulkhead flange that defines an inner diameter IDB. The first flange is secured at a radially outer joint between the radially outer hood flange and the radially outer bulkhead flange. The second flange is secured at a radially inner joint between the radially inner hood flange and the radially inner bulkhead flange. The IDOS and the ODB define a ratio R1 of IDOS/ODB that is 0.998622-1.001129, and the IDB and the ODIS define a ratio R2 of IDB/ODIS that is 0.998812-1.001388.

Abstract: There is provided an integrated fuel nozzle and ignition assembly for a gas turbine engine comprising a body having a fuel nozzle portion and an igniter portion. The fuel nozzle portion defines a fuel passage extending therethrough between a fuel inlet and a fuel outlet for directing a fuel flow into a combustion chamber. The igniter portion projects laterally from the fuel nozzle portion on a side thereof and comprises an igniter receiving cavity positioned adjacently and laterally on a side of the fuel passage. The assembly further comprises an igniter secured in the igniter receiving cavity for igniting the fuel flow discharged by the fuel passage.

Abstract: Disclosed herein are screw shaft turbine compressors having (i) a compressor section, (ii) a turbine section, (iii) a combustion section coupling to the compressor section and the turbine section, and (iv) a grooved shaft. The grooved shaft can include one or more grooves for providing fuel from the compressor section to the combustion section and for allowing exhaust to leave the combustion section and exit the turbine section. A method for generating different speed to torque ratios on the shaft and a system for generating torque on the shaft are further disclosed.

Abstract: A gas turbine engine has an annular reverse-flow combustor with a combustor inner liner enclosing a combustion chamber. The inner liner having a dome portion at an upstream end of the combustor and a downstream combustor exit defined between a small exit duct portion and a large exit duct portion. At least one of the dome portion, the small exit duct portion and the large exit duct portion is made of a separately formed hemi-toroidal shell composed of a ceramic matrix composite.

Abstract: A combustion chamber includes at least one annular wall. The annular wall includes a plurality of segments and each segment has a first portion, a second portion and a third portion. The first portion of each segment extends circumferentially, the third portion of each segment extends circumferentially and the second portion of each segment extends radially to connect the first portion and third portion of the segment. The segments are arranged such that the first portion of each segment overlaps the third portion of an adjacent segment and the first portion of each segment is bolted to the third portion of the adjacent segment. The segments are cast and define a double wall combustion chamber and enable easier construction of the combustion chamber and replacement of damaged segments.

Abstract: A combustor liner with an input end and an output end includes an annular inner wall and an annular outer wall. At least one of the inner wall and outer wall is three-dimensionally contoured. The inner wall and the outer wall form a combustion chamber with the contours creating alternating expanding and constricting regions inside the chamber causing combustion gases to flow in the circumferential and axial directions.

Abstract: An annular shroud (42), having an internal face intended to cover the back end wall (33) of an annular combustion chamber (12) of a turbomachine (14) fitted with a centrifugal compressor, together with an external face opposite said internal face, and where said shroud includes multiple apertures (54) intended to allow fuel injectors (38, 40) supported by said back end wall (33) to pass through, together with multiple bosses (56) which extend as projections on said external face radially towards the interior respectively from the respective radially internal edges (58) of said apertures (54), such that each of said bosses (56) delimits an extension (60) of the corresponding aperture (54) open radially towards the exterior so as to form an air intake scoop.

Abstract: The present invention relates to a segment component in high-temperature casting material for an annular combustion chamber of an aircraft engine, characterized by a combustion-chamber wall which in operation shields a fuel flame extending along a burner axis from the environment, with the combustion-chamber wall having a bulge which points in a direction facing away from the burner axis. The invention furthermore relates to an annular combustion chamber, an aircraft engine with an annular combustion chamber as well as a method for the manufacture of an annular combustion chamber.

Abstract: A precision counter-swirl combustor that includes an annular combustor having a forward end, an aft end opposite the forward end, and an interior. The aft end being proximal to a gas turbine. The combustor further includes a fuel inlet and swirler operatively connected to the forward end and at least one air inlet. The air inlet is equipped with a chute that extends into the interior of said combustor. The combustor is secured to a fixed structure proximate the forward end of the combustor.

Abstract: A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed.

Abstract: A combustor system is provided for a gas turbine engine. The combustor system includes an annular combustion chamber, and outer and inner casings housing the combustion chamber. The outer casing is outwardly radially spaced from the combustion chamber and the inner casing being inwardly radially spaced from the combustion chamber. The combustor system further includes an attachment arrangement at the rearward end of the combustion chamber for attaching the combustion chamber to the engine. The attachment arrangement axially restrains the rearward end of the combustion chamber relative to the casings. The combustor system further includes a first stop surface formed at the forward end of the combustion chamber adjacent a selected one of the outer and inner casings, and a corresponding first baulking surface axially forward of the first stop surface formed by the selected casing.

Abstract: A fuel nozzle for use with a combustion chamber defined within a gas turbine engine and a method for assembling the same are provided. The fuel nozzle includes a swirler assembly and a burner tube that is coupled to a support flange in a detachable manner such that an inner surface of the burner tube circumscribes an outer surface of the swirler assembly during fuel nozzle operation.

Abstract: A system comprises a turbine combustor cap. The turbine combustor cap includes a plurality of segments. Each segment of the plurality of segments has edges abutting at least two fuel nozzle receptacles. Moreover, each segment of the plurality of segments does not completely surrounding any one fuel nozzle receptacle.

Abstract: A turbo machine is provided that includes a turbine nozzle having a wall. The wall provides spaced apart first and second surfaces. A combustor includes a liner having an annular lip that engages one of the first and second surfaces. A floating ring seal is supported by the liner in a slip-fit relationship, for example, using a retainer. The floating ring seal engages the other of the first and second surfaces. The floating ring seal is slidably moveable relative to the liner in a floating direction in response to movement of the liner in a radial direction.

Abstract: A method of managing transient events regularly seen during gas turbine operation that may cause undesirable operation and hardware damage. During certain transient operations, a lag may be seen between reference exhaust temperature and actual turbine exhaust temperature. This lag can result in an under-fired condition within the combustion system of variable magnitude and duration. Either fuel split schedules or a control algorithm can be positioned during these transients to prevent combustion dynamics or loss of flame. Combustion dynamics are known to cause damage that may require hardware replacement. Once the transient has completed, normal control operation is resumed.

Abstract: A stationary crankshaft, cylindrical rotor, low friction, adjustable timing, pistonless, rotary internal combustion engine. A cylindrical rotor rotates freely about a stator, also cylindrical in shape. A plurality of cavities are circumferentially disposed along the external circumference of the stator and the internal circumference of the rotor that, when rotationally aligned, form combustion chambers. Rotation of the rotor is induced by electrical-spark induced combustion of a fuel/air mixture in the combustion chambers. The combustive exhaust is vented into an exhaust manifold for transport to an exhaust disposal system, such as a catalytic converter. Cooling, fuel pressurization, and electrical generation can be internal to the engine or supplied externally. Engine speed, torque, and other operational requirements can be accommodated by coupling multiple rotational units together and engine vibration can be virtually eliminated by offsetting the combustion chambers on coupled units.

Abstract: A rotational system for a miniature gas turbine engine includes a rotor shaft rotationally mounted to a forward bearing and an aft bearing. The rotor shaft is located axially with a single threaded rotor nut which provides an axial preload maintained by a fixed compressor wheel and spacers. A preload spring provides dynamic compensation for tolerance variation and undesirable axial movement during rotation of the shaft. A permanent magnet generator is mounted to the rotor shaft behind the forward bearing to generate electrical power for the engine and other accessories. The rotor shaft is inertia welded to a turbine wheel to form a cavity which provides a thermally conductive path away from the aft bearing. A seal ring mounted within a groove within the rotor shaft provides a pneumatic seal to further protect the aft bearing from thermal transfer between the turbine section and the compressor section.

Abstract: A combustor end cap assembly having an improved cooling configuration is disclosed. Embodiments of the present invention are directed towards an apparatus and method for cooling an effusion plate of the combustor end cap assembly. The combustor end cap assembly also incorporates an impingement plate having a plurality of cooling holes with the impingement plate positioned a predetermined distance from the effusion plate. The cooling fluid passes through the impingement plate and is directed towards and onto the effusion plate for cooling of the effusion plate.

Abstract: A combustor section is provided and includes one or more annular quaternary fuel manifolds mounted within an annular passage defined between a casing and a cap assembly of a combustor through which air and/or a fuel/air mixture flows upstream from a fuel nozzle support, the manifold including a body to accommodate quaternary fuel therein, the body defining injection holes through which the quaternary fuel is injected into a section of the passage at a location upstream from the fuel nozzle support.

Abstract: A thermal machine, especially a gas turbine, includes an annular combustor which is outwardly delimited by an outer shell and an inner shell (33) and through which a hot gas axially flows. The outer shell and inner shell (33) are each provided with a concentric cooling shroud (31) which is attached at a distance on their outer side, forming a cooling passage (32) through which cooling passage (32) cooling air flows in a direction which is opposite to the hot gas flow. The cooling of the combustor is improved by at least one of the cooling shrouds (31), on the side on which the cooling air enters the cooling passage (32), having an outwardly curved, rounded inlet edge (37) for improving the inflow conditions.

Abstract: A gas-turbine burner having a plurality of main swirl generators, each having an inlet flow opening formed by the main swirl generator edge, is provided. In order to achieve a uniform flow of combustion air through the main swirl generator, the gas-turbine burner has an inlet-flow guide means with a flow guide surface which runs from one of the inlet-flow openings to an adjacent inlet-flow opening, to which the main swirl generator edges which form the inlet-flow openings are connected, and the flow guide surface widens from there radially upwards. The main swirl generators are central-symmetrically arranged around a pilot burner and the flow guide surface runs radially outside the main swirl generators.

Abstract: A combustor includes a first flow path and a second flow path. The first flow path places a diffuser in indirect fluid communication with a combustor cap assembly by way of an intervening lower combustor annular passageway. The second flow path places the diffuser in direct fluid communication with the combustor cap assembly.

Abstract: An air starter for turbomachine, including a forward casing, an aft casing, an annular exhaust flow path opening up between an aft end of the forward casing and a forward end of the aft casing, and a cylindrical outlet mesh of the exhaust flow path is disclosed. The forward and aft ends of the outlet mesh include devices for axially retaining the forward casing and the aft casing respectively to the mesh. At least one of the devices for axially retaining one of the casings to the mesh enables a relative rotation of the mesh and this casing.

Abstract: A thermal machine including a wall defining a hot gas duct for transferring a hot gas stream and a cooling jacket disposed at a distance from the wall on an outside of the hot gas duct so as to define a cooling duct with an inlet and an outlet. The cooling duct is configured to conduct a cooling medium along an external face of the wall from the inlet to an outlet in a direction counter to a flow of hot gas in the hot gas duct. An impingement cooling plate is disposed at the inlet of the cooling duct and includes cooling baffle holes configured such that cooling medium entering the cooling duct through the cooling baffle holes flows in a direction perpendicular to the wall. The impingement cooling plate is positioned such that an inflow-side edge sealingly abuts the wall of the hot gas duct so as to reduce a transverse flow of the cooling medium through the cooling duct.

Abstract: A turbomachine includes a compressor, a combustor operatively connected to the compressor, an end cover mounted to the combustor, and an injection nozzle assembly operatively connected to the combustor. The injection nozzle assembly includes a first end portion that extends to a second end portion, and a plurality of tube elements provided at the second end portion. Each of the plurality of tube elements defining a fluid passage includes a body having a first end section that extends to a second end section. The second end section projects beyond the second end portion of the injection nozzle assembly.

Abstract: A system includes an air flow conditioner configured to mount in an air chamber separated from a combustion chamber of a turbine combustor. The air flow conditioner comprises a perforated annular wall configured to direct an air flow in both an axial direction and a radial direction relative to an axis of the turbine combustor. In addition, the air flow conditioner is configured to uniformly supply the air flow into air inlets of one or more fuel nozzles.

Abstract: A turbomachine including an annular combustion chamber fitted with fuel injectors and nozzle vanes arranged at the outlet from the chamber, the number of nozzle vanes being an integer multiple of the number of fuel injectors, and the head of each injector being situated angularly half-way between the leading edges of two consecutive nozzle vanes, these leading edges being in alignment with primary air holes and/or with dilution air holes.

Abstract: A combustor assembly for a turbine engine includes a cap assembly which gradually increases the volume of compressed air provided to the combustor assembly, and which slows the compressed air in a controlled manner before delivering the compressed air into a combustor area. The cap assembly includes inner and outer sleeves which are mounted in a concentric fashion. The annular space formed between the inner and outer sleeves gradually increases from an aft end to a forward end. A stepped portion is formed on the inner sleeve of the cap assembly, and this stepped portion is joined to a corresponding combustor liner.

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 for a gas turbine engine includes an annular combustor shell having an inner liner and an outer liner respectively with inner and outer flanges at least partly overlapping to form a dome end portion of the shell, at least the outer flange including intersecting upstream and downstream wall portions defining a corner therebetween, the upstream wall portion having a plurality of cooling apertures defined therethrough immediately upstream of the corner, and the cooling apertures being oriented to direct a cooling air flow from outside the combustor shell therethrough and adjacent an inner surface of the downstream wall portion.

Abstract: A turbomachine includes a compressor, a turbine, and a combustor operatively connected to the turbine. The turbomachine further includes a cap member mounted to the combustor. The cap member includes a first surface and a second surface. A combustion chamber is defined within the combustor. An injection nozzle is supported at the second surface of the cap member. The injection nozzle includes a first end that extends through an inner flow path to a second end. The first end is configured to receive an amount of a first fluid and the second end is configured to receive an amount of a second fluid. A mixture of the first and second fluids is discharged from the second end of the injection nozzle.

Abstract: An annular gas turbine engine combustion chamber including an outer wall and an inner wall connected by a wall forming the chamber bottom is disclosed. The walls delimit sources of combustion with axes inclined relative to the axis of the chamber. The chamber-bottom wall, of frustoconical shape, is pierced with orifices for the fuel injection systems, the planes of the orifices being perpendicular to the axes of the sources of combustion. The combustion chamber also includes heat-protection baffles centered on each of the orifices with a shoulder by which they rest against a flat surface portion along the periphery of the orifices. The chamber-bottom wall is conformed in a succession of adjacent flat facets having a common edge, with one facet per orifice, and the shoulder of the deflectors pressing against the plane of the facets.

Abstract: A fuel injection nozzle for a gas turbine engine has a central fuel ejection nozzle and a plurality of airflow passages within the spray tip that include a first and second group of circumferentially spaced apart fuel-spray forming airflow passages disposed on opposite sides of a transverse axis and oriented towards each other such as to produce opposed fuel spray shaping air jets which generate a shaped final fuel spray.

Abstract: An annular combustion chamber has an inner wall and an outer wall of ceramic matrix composite material, together with a chamber end wall that is connected to the inner and outer walls. Elastically-deformable link elements connect the inner wall and the outer chamber walls to inner and outer metal casings. Each of the inner and outer chamber walls is subdivided circumferentially into adjacent sectors along longitudinal edges, each sector extending continuously from the end wall to the opposite end of the chamber and being folded outwards from the chamber at each of its longitudinal edges to form a portion having a U-shaped cross-section terminated by a folded-back margin that is spaced apart from the outer face of the corresponding chamber wall. The link elements are connected to the inner and outer chamber walls by being fastened to the sector margins.

Abstract: A combustor may include an outer liner having a first row and a second row of circumferentially distributed air admission holes. The second row of the outer liner may be downstream of the first row of the outer liner, and the air admission holes of the first row of the outer liner may be larger than the air admission holes of the second row of the outer liner. An inner liner is circumscribed by the outer liner and has a third and fourth row of circumferentially distributed air admission holes. The air admission holes of the third row of the inner liner may be larger than the air admission holes of the fourth row of the inner liner. The inner and outer liners may form a combustion chamber, and at least a portion of the air admission holes of the first, second, third, or fourth rows may be plunged.

Abstract: An assembled annular combustion chamber comprises an annular inner wall and an annular outer wall made of ceramic matrix composite material together with a chamber end wall connected to the inner and outer walls and provided with orifices for receiving injectors. Elastically-deformable link parts connect the inner wall and the outer wall of the chamber to inner and outer casings that are made of metal. The assembly formed by the inner wall, the outer wall, and the combustion chamber end wall is subdivided circumferentially into adjacent chamber sectors, each sector being made as a single piece of ceramic composite material and comprising an inner wall sector, an outer wall sector, and a chamber end wall sector. The link parts connect the inner metal casing and the outer metal casing respectively to each inner wall sector of the combustion chamber and to each outer wall sector of the chamber. The chamber end wall sectors are in contact with a one-piece ring to which they are connected.

Abstract: A gas turbine system includes a combustion chamber (2), a turbine (3), a radially inward and/or radially outward axial gap (4; 16, 17) between the combustion chamber (2) and the turbine (3), at which gap inner and/or outer combustion chamber walls (7, 8) and inner and/or outer turbine walls (11, 12) end, and a cooling gas supply (5), which via the gap (4; 16, 17) introduces a cooling gas into the turbine gas path (13) and/or into the combustion chamber gas path (9). An end section (21, 22, 23, 24) of the respective turbine wall (11, 12) and/or of the respective chamber wall (7, 8) adjoining the gap (4; 16, 17) is radially inwardly and/or radially outwardly, alternatingly formed.

Abstract: A turbomachine including an annular combustion chamber, a centrifugal compressor, an annular diffuser enabling the flow of gas leaving the compressor to be diffused and directing this flow of gas towards the combustion chamber, and an outer casing surrounding the outside of the combustion chamber is disclosed. The diffuser includes an upstream portion oriented radially and presenting diffusion passages connected to the outlet from the compressor, an intermediate portion that is elbow-shaped, and a downstream portion having a series of circularly spaced-apart flow-straightening vanes. The flow stream in the downstream portion is delimited externally by the outer casing. The outer casing presents a shoulder against which the outer downstream edge of the intermediate portion comes into abutment.

Abstract: A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have an internal passage extending between an inlet to an outlet. An axis of the transition duct body may be generally linear such that gases expelled from the transition duct body flow in a proper direction into the downstream turbine blades.

Abstract: A combustor includes a chamber mixing and burning fuel and air and an air hole plate disposed on a wall surface of the chamber. The, air hole plate includes a plurality of rows disposed concentrically of a plurality of air holes jetting coaxial jets of fuel and air into the chamber. A first fuel nozzle and a second fuel nozzle are disposed near a fuel hole for jetting fuel into an air hole row on an inner peripheral side. The first fuel nozzle is structured to suppress turbulence of a surrounding air flow and the second fuel nozzle is structured to promote turbulence of a surrounding air flow. In accordance with the aspect of the present invention, good flame stability can be maintained while further reducing NOx in a combustor using coaxial jets.

Abstract: A fuel nozzle is provided. The fuel nozzle includes a heat shield, a fuel tube and a plurality of support members. The support members are radially interposed between the heat shield and the fuel tube. The support members are preferably cylindrical tubes, thus creating voids or pockets between the heat shield and the fuel tube. The cylindrical tubes are connected to one another at a first end are free at an opposed end. As such, the tubes can move or slide relative to one another. Further, the tubes preferably only contact one another with at most line contacts. The fuel nozzle may also include a tip portion that includes a tip heat shield that extends radially outward from the primary heat shield. The tip heat shield defines a cavity that connects with the central cavity of the heat shield. The fuel tube extends through the tip heat shield.

Abstract: A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine has an internal passage from an inlet to an outlet. The outlet may include canted sides that reduce formation of damaging vibration in downstream turbine blades caused by downstream wake between adjacent transition ducts and by pressure differentials between adjacent transition ducts that include turning sections.

Abstract: A transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have an internal passage extending between an inlet to an outlet. An axis of the transition duct body may be generally linear such that gases expelled from the transition duct body flow in a proper direction into the downstream turbine blades. The linear transition duct may include an outlet with exhaust mouths that are configured such that sides of the transition duct are coplanar with adjacent transition ducts, thereby eliminating destructive turbulence between adjacent, linear transition ducts.

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: The present application provides a liner for a combustor. The combustor liner may include a mouth, a passage for a flow of fuel, a passage for a flow of air, a first zone for mixing the flow of fuel and the flow of air, and the first zone having a constant diameter. The combustor liner may have at least one transition zone with first and second side walls angled substantially in the same direction, and an off center exit. This configuration will provide a more stable flow when exiting the combustor liner.

Abstract: The present invention is directed to a lean direct injection (LDI) combustion system for a gas turbine using a shell and tube heat exchanger concept to construct a shell and tube lean direct injector (“LDI”) for the combustion system. One side of the LDI injector, either the shell side or the tube side, carries an oxidizer, such as air, to the combustor, while the other side of the LDI injector carries fuel to the combustor. Straight or angled holes drilled in an end plate of the combustor allow the fuel to enter the combustor and mix with air being injected into the combustor.

Abstract: A deflector takes the form of a plate provided with a hole. The plate is a portion of a conical surface of revolution about a cone axis with a concave face and a convex face. The plate includes a contour with four sides. Two of these the sides are concentric circular arcs centered on the cone axis. The other two sides are segments of generatrices of the cone which connect the first sides. The invention is applicable to a deflector of a chamber endwall of a combustion chamber of a turbine engine.

Abstract: A turbomachine including an annular combustion chamber is disclosed. The combustion chamber includes a radially inner wall and a radially outer wall exhibiting symmetry of revolution. An annular chamber end wall equipped with a fuel injection device connects upstream ends of the inner and outer walls. A suspension device suspends the upstream end of the combustion chamber from an outer casing.

Abstract: A gas turbine engine combustor has forward bulkhead extending between inboard and outboard walls and cooperating therewith to define a combustor interior volume or combustion chamber. At least one of the walls has an exterior shell and an interior shell including a number of panels. Each panel has interior and exterior surfaces and a perimeter having leading and trailing edges and first and second lateral edges. A number of cooling passageways have inlets on the panel exterior surface and outlets on the panel interior surface. The shell has a plurality of holes for directing air to a space between the shell and heat shield and adapted for preferentially directing said air toward leading edge portions of first stage vanes of a turbine section.

Abstract: A combustion device used in gas turbine engines includes an annular combustor that contains the combustion process of air and fuel and then guides the hot gas products to a first stage turbine subsection of a gas turbine engine. The annular combustor has an inner/outer shell having corrugated surfaces that extend radially outward and inward across an entire hot gas stream inside the annular combustor. The corrugations twist about the engine centerline in a longitudinal direction of travel of the engine. The resulting flow path accelerates and turns the hot gas stream to conditions suitable for introduction into the first stage turbine blades, which eliminate the need for first stage turbine vanes. The annular combustor is configured with a system of fuel and air inlet passages and nozzles that results in a staged combustion of premixed fuel and air.