Abstract: A gas turbine combustion liner is disclosed having an alternate interface region between it and a transition duct where the cooling effectiveness along the aft end of the combustion liner is improved, resulting in extended component life, while utilizing a simpler combustion liner geometry. The region of the combustion liner proximate its second end comprises a plurality of spring seals that seal against a transition duct while admitting a cooling fluid to pass into a passage, formed between the combustion liner and spring seals, that feeds a plurality of cooling holes located in the combustion liner proximate the liner second end. Depending on the cooling requirements, the cooling holes can be angled both axially and circumferentially to maximize the cooling effectiveness.

Abstract: A turbine fuel ring assembly includes a fuel distribution ring, at least one fuel supply tube attached to the fuel distribution ring and at least one attachment leg connected to the fuel distribution ring. The fuel ring has a hollow interior and a plurality of apertures for expelling a fluid. The attachment leg is configured to allow flexibility due to thermal expansion induced under certain load conditions such as during engine start-up or shut-down. Further, the configuration of the attachment legs provides improved stress distribution characteristics. The fuel supply tube includes a rectangular passage and a round passage that are disposed substantially transverse to each other and in fluid communication with each other and with the hollow interior of the fuel distribution ring. The rectangular passage and the round passage have substantially identical cross-sectional areas. The fuel supply tube is configured to avoid structural interferences with neighboring components.

Abstract: A combustor assembly for a gas turbine engine comprises a casing arrangement defining an aperture and a seal element mounted on the casing arrangement through which a component of the combustor can extend. The seal element comprises a relatively flexible force absorbing portion and a relatively rigid sealing portion. The force absorbing portion is provided between the sealing portion and the casing arrangement. The force absorbing portion is resiliently deformable to absorb force during movement of the casing arrangement. Such resilient deformation restricts transmission of the forces to the sealing portion. The seal element may be configured to carry a seal member.

Abstract: Aspects of the invention relate to a system for attaching a ceramic liner to a non-ceramic combustor head-end component while accommodating different rates of thermal expansion of these components. One end of the liner can be received within a slot formed by the combustor head-end component. The liner can be held in place by a plurality of pins with each pin passing through a pair of aligned openings in the liner and the head-end component. The end of the liner can be spaced from each of the walls of the slot so as to form a series of gaps in fluid communication. The gaps and the pins can accommodate differential thermal expansion between the liner and combustor head-end component. If desired, at least some of the gaps can be tightly controlled so as to regulate air leakage around the end of the liner.

Abstract: A system and method is provided for lightoff detection in a turbine engine that provides improved accuracy and consistency. The system and method use engine temperature and engine rotation speed sensor data to determine a first time interval during which the lightoff is estimated to have occurred. A peak gradient in the temperature sensor data within the first time interval and a peak gradient in the engine rotation sensor data within the first time interval are then determined. The temperature peak gradient time and the engine rotation peak gradient time are then used as boundaries to determine a narrowed time interval within the first time interval. The minimum gradient in the engine rotation sensor data in the narrowed time interval is determined, with the minimum gradient time in the engine rotation sensor data comprising an accurate estimate of lightoff time for the turbine engine.

Abstract: The invention relates to a device for passive control of the thermal expansion of the extension casing of a turbo-jet engine, said extension casing surrounding the casing of the high pressure compressor of the turbo-jet engine, and including a flange for attachment to an upstream flange of the casing of the combustion chamber. It is characterised in that at least one circumferential cavity is provided between both said flanges wherein circulates a gas flux tapped at the inlet of the combustion chamber. One uses thus a natural circulation generated by the differential pressure. Thanks to the device of the invention, the flange is passively controlled, and the stresses resulting from a differential expansion between the skin of the casing and its attachment flange are reduced.

Abstract: A sealing device for use between a gas turbine combustor transition duct aft frame and a turbine inlet having improved durability, reduced wear on the mating turbine vane, and reduced manufacturing costs, is disclosed. The sealing device has a circumferential length, an axial width, and a radial height and contains a plurality of channels extending axially along the seal inner surface for passing a controlled amount of cooling air to a turbine inlet. The sealing device is formed of abradable honeycomb having a plurality of honeycomb cells with the honeycomb cells oriented to ensure maximum control against cooling air leakage while also providing maximum flexibility during assembly. The sealing device is captured between the transition duct aft frame, bulkhead, and turbine vane platform, thereby allowing easy replacement of the seal without requiring major disassembly of the transition duct aft frame section.

Abstract: A combustor includes an outer wall and an inner liner joined to an inner shell in turn mounted to an inner casing. The casing includes a first rabbet at an end flange in which is mounted a corresponding flange of the inner shell. The inner shell also includes a second rabbet which receives an end flange of the inner liner. The inner shell is trapped in the first rabbet by an inner retainer. And, the inner liner is trapped in the surrounding second rabbet for aft-mounting the liner and shell to the inner casing.

Abstract: A simplified floating collar is provided comprising an collar with opposed flanges. The arrangement offers reduced cost, and simplicity, and therefore facilitates manufacturing.

Abstract: A simplified floating collar mounting arrangement for receiving a fuel nozzle swirler body of a gas turbine engine combustor is provided. The assembly comprises a floating collar mounted between a spaced-apart mounting flange and cap, and slidably trapped therebetween such that relative radial movement is permitted. The arrangement offers reduced part count and simplicity, and therefore improves reliability.

Abstract: The endcover assembly for a combustor includes an endcover having an aperture, an insert disposed in the stepped bore opening through an internal face of the endcover, a fuel cartridge extending through the insert and aperture and a seal cover on the external face of the endcover. Seal cover/endcover annular seals are disposed in registering axial cavities to seal against fluid leakage externally of the endcover assembly. Generally C-shaped and W-shaped seals are disposed in the cavities between the stepped shoulders of the insert and aperture to seal against leakage through the internal face of the endcover. The seals maintain sealing engagement during assembly and thermal distortion during turbine usage.

Abstract: An interface region between a combustion liner and a transition duct of a gas turbine combustor is disclosed having improved cooling such that component life is increased and metal temperatures are lowered. An aft end of a combustion liner is telescopically received within the transition duct such that a combustion liner seal is in contact with an inner wall of the transition duct inlet ring. Increasing the dedicated cooling air supply to the combustion liner aft end, coupled with a modified combustion liner aft end geometry, significantly reduces turbulence and flow re-circulation, thereby resulting in lower metal temperatures and increased component life. Multiple embodiments of the interface region are disclosed depending on the amount of cooling required.

Abstract: An outer combustor wall (36) includes an outer support shell (40) surrounding an outer combustion liner (42). The shell (40) includes a first radial aft flange (46), and the liner (42) includes a second radial aft flange (48) axially adjoining the first flange. An annular retainer (50) and seal (62) are joined to the first flange for slidingly trapping the second flange therebetween for thermally floating the liner on the shell.

Abstract: A gas turbine engine comprises a turbine scroll inside a combustor housing, discouragers with 90-degree bending angles, a radial nozzle contacting a forward bayonet on the forward side of the turbine scroll at a bayonet engagement point and a B-width measured between the discouragers. Retaining ring maintains the size of the B-width. The turbine scroll may have eight surfaces sealing at four locations and a provision to maintain constant “B-width” for the thin sheet metal scroll within the combustion system. The design allows the scroll to operate at high temperature while maintaining lowest possible thermal and mechanical stresses. It can be easily assembled with excellent capability to control gas leakage and minimal component interface wearing or fretting. The gas turbine engine is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially high performance, high cycle flight vehicles.

Abstract: The invention relates to a combustion chamber, in particular of a gas turbine, comprising an external wall structure that surrounds an internal area and an internal wall, which is formed by the surface of a housing that is located in the internal area. Said combustion chamber can be cooled essentially by convection by an air stream that flows between the external wall structure and the internal wall, whereby the air stream is conducted in a closed cooling air canal. The housing can be mounted in the internal area of the combustion chamber by means of a suspension device, which comprises a large number of fixing elements that are arranged on the periphery of the housing and are braced against the wall structure.

Abstract: A combustor dome assembly for a gas turbine engine having a longitudinal centerline axis extending therethrough, including: an annular dome plate having an inner portion, an outer portion, a forward surface and a plurality of circumferentially spaced openings formed therein, wherein a radial section is defined between adjacent openings; an annular outer cowl connected to the outer portion of the dome plate; an annular inner cowl connected to the inner portion of the dome plate; and, a swirler located between the forward surface of the dome plate and the inner and outer cowls in substantial alignment with each of the openings in the dome plate. Each swirler further includes a forward portion and an aft portion and is retained by at least one tab member located upstream thereof so as to be movable in a radial and axial direction.

Abstract: A gas turbine engine with: a compressor section; a turbine section; a combustor, disposed between the compressor and turbine sections, having at least one combustor mounting assembly connecting the combustor to the engine. Each combustor mounting assembly has: a longitudinal axis; and an articulating joint having a first and second portion constrained from relative translation transverse to the longitudinal axis, and where said first and second portion have a multiple rotational degrees of freedom relative to each other about axes transverse to the longitudinal axis.

Abstract: The invention relates to an annular combustion chamber for a turbomachine of axis X, in which the chamber end wall is fixed to inner and outer side walls by means of a plurality of radially-inner attachment fittings and by means of a plurality of radially-outer attachment fittings disposed upstream from said chamber end wall by bolting a zone of each of said fittings to said side walls. At least one of said pluralities of attachment fittings is integrated in a ring secured to the chamber end wall and comprises a plurality of pairs of tabs which extend upstream beyond the bolting zones, the two tabs in each pair being disposed circumferentially on either side of a bolting zone and being interconnected at their upstream end to the upstream end of a fixing plate which extends between said two tabs and which presents a bolting zone for being fixed to the adjacent side wall.

Abstract: In a gas turbine (1) with an annular combustion chamber (4), the combustion area (24) of which is bounded by an annular combustion chamber outer wall (26) on the one hand and by an annular combustion chamber inner wall (28) located therein on the other hand, it should be possible to dismantle the combustion chamber inner wall (28) comparatively quickly and easily. For this purpose according to the invention the combustion chamber inner wall (28) is formed by a plurality of wall elements attached to a support structure, wherein the support structure is formed by a plurality of sub-components abutting each other at a horizontal parting joint and the abutting sub-components (30) of the combustion chamber inner wall (28) are connected to each other at their horizontal parting joint by means of a plurality of screw connections (32) oriented at an angle to the inner wall surface.

Abstract: An assembly for a gas turbine engine, the assembly comprising a combustor and a vane assembly downstream from the combustor. The vane assembly includes at least one airfoil radially extending between an inner and outer platform defining an annular gas path therebetween. At least the outer platform forms a first sliding joint connection with an adjacent outer combustor wall such that radial sealing between the outer platform and the outer combustor wall is provided at engine operating temperature, while permitting relative axial displacement therebetween.

Abstract: The present invention provides a fixing or sealing ring for maximizing cooling at the end of a combustion chamber wall. For this purpose, the ring is constituted by a sleeve which is fixed around the end of a wall of the combustion chamber by means of a plurality of orifices for receiving fasteners or by means of any other system for connecting the ring to the wall. The sleeve includes at least one recess in its face facing the wall of the combustion chamber, thereby reducing the area of the sleeve that presses against the wall, and co-operating with said wall to form an open cavity in which a cooling air stream can flow.

Abstract: A combustor dome assembly for a gas turbine engine having a longitudinal centerline axis extending therethrough includes: an annular dome plate having an inner portion, an outer portion, a forward surface, and a plurality of circumferentially spaced openings formed therein, wherein a radial section defined between each of the openings includes a cooling trough formed therein; and, a free floating swirler located upstream of the dome plate in substantial alignment with each of the openings in the dome plate, the swirler including a forward portion and an aft portion. The aft portion of each swirler is configured so as to permit air flow to the cooling trough during operation of the gas turbine engine. In particular, the swirler aft portion includes a substantially annular flange extending adjacent to the forward surface of the dome plate, wherein opposing portions of a circumference for the flange adjacent the radial sections of the dome plate are substantially linear.

Abstract: An assembly for mounting an igniter in a gas turbine engine combustor between an outer casing and an outer liner, wherein a longitudinal centerline axis extends through the gas turbine engine. The igniter mounting assembly includes a first spring member encircling a portion of said igniter and being positioned between a surface adjacent the outer casing and an outer surface of the outer liner, a first ring member connected to a first end of the first spring member adjacent the outer surface of the outer liner, and a second ring member connected to a second end of the first spring member adjacent the surface adjacent the outer casing. Accordingly, the igniter is able to maintain substantial alignment with respect to an opening in the outer liner while moving radially and/or axially with respect to the outer liner as the outer casing experiences thermal growth greater than the outer liner.

Abstract: A structure for attaching together or sealing a space between a first component and a second component that have different rates or amounts of dimensional change upon being exposed to temperatures other than ambient temperature. The structure comprises a first attachment structure associated with the first component that slidably engages a second attachment structure associated with the second component, thereby allowing for an independent floating movement of the second component relative to the first component. The structure can comprise split rings, laminar rings, or multiple split rings.

Abstract: A wear cover for use on a gas turbine transition duct support bracket is disclosed. The wear cover, which is manufactured from a single piece of sheet metal, does not require complex manufacturing processes or tools, thereby reducing component cost. The wear cover comprises first and second leg portions substantially parallel to each other and each generally perpendicular to a third leg portion. Each of the leg portions serve to protect a surface of a support bracket finger from excessive wear associated with frictional contact with mounting blocks of a transition duct. The wear cover is designed to slip onto a support bracket finger through a first opening and is fixed to the finger by a means such as welding.

Abstract: A mounting assembly for a forward end of a liner in a combustor of a gas turbine engine including a dome and a cowl, wherein a longitudinal centerline axis extends through the gas turbine engine. The mounting assembly includes a pin member extending through each one of a plurality of circumferentially spaced openings formed in the forward end of the liner, an aft portion of the cowl, and a portion of the dome, with each pin member including a head portion at one end thereof. A nut is adjustably connected to an end of each pin member opposite the head portion. A bushing is located on each pin member at a position intermediate the head portion and the nut, wherein the openings in the liner forward end are sized to fit around the bushings.

Abstract: An assembly for providing a seal at an aft end of a combustor liner for a gas turbine engine including a longitudinal centerline axis extending therethrough. The sealing assembly includes a substantially annular first sealing member positioned between an aft portion of a support member and the liner aft end so as to seat on a designated surface portion of the liner aft end and a substantially annular second sealing member positioned between the support member aft portion and a turbine nozzle located downstream of the liner aft end so as to seat on a designated surface portion of the support member aft portion. Accordingly, the first sealing member is maintained in its seated position as the support member aft portion moves radially with respect to the liner aft end and the second sealing member is maintained in its seated position as the support member aft portion moves axially with respect to the turbine nozzle.

Abstract: Aspects of the present invention relate to hollow elongated turbine engine components, such as transition ducts, having a cooling system. The component can have first and second ends, and can be defined by an outer peripheral wall. The cooling system extends longitudinally within the wall between the first and second ends. While configured to cool the entire component, the cooling system is particularly configured to reduce thermal gradients that can occur in regions proximate to the first and second ends of the component caused by higher loads in those regions. The cooling system can include at least one pair of cooling channels. Each channel pair includes a first channel and a second channel. The first and second channels can be arranged such that a coolant supplied to each channel can exchange heat with itself in at least the region substantially proximate to one of the ends.

Abstract: In order to provide a gas turbine tail tube seal which can prevent an outer shroud and an inner shroud in a first row stationary blade from being damaged by heat and from being worn, and to provide a gas turbine having the above-mentioned gas turbine tail tube seal, in the gas turbine tail tube seal, the width of a combustion gas flow channel in the gas turbine tail tube seal is smaller than the width of a combustion gas flow channel made between an outer surface of the inner shroud and an inner surface of the outer shroud, and the combustion gas flow channel is provided with inclined surfaces having a gradually enlarging cross section toward the first row stationary blade in at least a downstream end section.

Abstract: A sealing structure for sealing a gap between a combustor liner having an outer surface and a surrounding structure surrounding the combustor liner, includes: an annular sealing member held on the surrounding structure so as to be in contact with the outer surface of the combustor liner. The sealing member includes a plurality of sealing segments arranged in an annular configuration. The sealing structure seals the gap around the combustor liner of the combustor and absorbs, the difference in thermal expansion between the combustor liner and the surrounding structure.

Abstract: A gas turbine engine (40) having an active combustion control system (54) including a sensor element (44) positioned proximate the combustion chamber (46) as part of a sensor assembly (84) including a spring-action bellows (74). The sensor assembly is installed through an opening (69) in the casing (48) of the engine to make contact with the combustor wall (62). The bellows is compressed from its resting position by tightening mounting bolts (70) against a mounting flange (66) of the sensor assembly. The spring action of the bellows ensures contact between the sensing assembly and the combustor wall in spite of differential thermal growth between the casing and the combustor. The sensor element may be replaced by simply removing the mounting bolts without further disassembly of the engine.

Abstract: An improved sealing device for use between a transition duct aft frame and turbine inlet is disclosed. The improved sealing device includes a plurality of corrugated metal seals that are secured to the transition duct aft frame along the are-like sections of the aft frame. The corrugated seals contact both the transition duct aft frame and turbine inlet to provide a more effective seal and to control the amount of compressor discharge air introduced to cool the turbine vane platforms. In the preferred embodiment, the corrugated seals are coated to reduce wear, fretting, and galling and are used in conjunction with a transition duct thermally free aft frame and pivot bearing mounting system. An alternate embodiment of the seal is also disclosed that utilizes an alternate cooling hole pattern for providing additional cooling air to the turbine vane platform.

Abstract: A flame-passage device (10) for non-annular gas turbine combustion chambers comprises a tubular body (11), which is provided with a plurality of cooling holes (12), for cooling on the swirl-cooling type. The tubular body (11) is inserted in a flanged pipe (27), which connects the cases (31, 32) of two successive combustion chambers (17, 18), and has a first end (13) with a cylindrical shape, and a second end (14) with an oval shape, wherein the second, oval end (14) is provided with three rings (23, 24, 25), for anchorage of the corresponding combustion chamber (17) to the case (31).

Abstract: In a turbomachine comprising a metal material shell containing in a gas flow direction F: a fuel injection assembly; a composite material combustion chamber; and a metal material sectorized nozzle forming the inlet stage with fixed blades of a high pressure turbine, provision is made for the combustion chamber to be held by a sectorized flexible sleeve of metal material having one end fixed to the combustion chamber by first fixing means and a flange-forming opposite end fixed to the shell by second fixing means. The first fixing means also serve to connect the combustion chamber to the sectorized nozzle.

Abstract: Thermal stress relief in a nozzle head of a gas turbine fuel nozzle is provided by forming a slit through each selected air passages of the nozzle head. The strategic location of stress-relief slits contributes to extend the fatigue life of the nozzle with minimal cost and impact on the nozzle aerodynamics.

Abstract: An internal combustion engine includes an insulated combustion chamber having a fuel mixture inlet and a spark plug nearby the inlet. A series of baffles is configured within the combustion chamber to absorb a shockwave caused by ignition of fuel mixture by the spark plug. A turbine receives reduced-pressure combustion gases from an exhaust-side of the baffles and there is a power takeoff at the turbine.

Abstract: A heat engine unit has a combustion chamber, the combustion chamber being in communication with a previous component of the heat engine unit, e.g., a compressor, a heat exchanger or a recuperator, at a first end, and in communication with subsequent stages of the heat engine unit at a second end. The combustion chamber comprises at least one part, and is allowed to expand when exposed to heat by being movably engaged at the first end. The combustion chamber is supported at the second end by way of a contact surface formed between the second end of the combustion chamber and supporting means to allow tilting of the combustion chamber.

Abstract: An improved sealing device for use between a transition duct aft frame and turbine inlet is disclosed. The improved sealing device includes a plurality of corrugated metal seals that are secured to the transition duct aft frame along the arc-like sections of the aft frame. The corrugated seals contact both the transition duct aft frame and turbine inlet to provide a more effective seal and to control the amount of compressor discharge air introduced to cool the turbine vane platforms. In the preferred embodiment, the corrugated seals are coated to reduce wear, fretting, and galling and are used in conjunction with a transition duct thermally free aft frame and pivot bearing mounting system. An alternate embodiment of the seal is also disclosed that utilizes an alternate cooling hole pattern for providing additional cooling air to the turbine vane platform. A further alternate embodiment of the seal incorporates a group of wear reduction coatings from which to select.

Abstract: A crossfire tube assembly is provided having an increased resistance to wear and oxidation. This increased resistance is provided by the application of coatings to the mating surfaces between the crossfire tube and crossfire tube clip. In the preferred embodiment, a physical vapor deposition (PVD) Titanium Aluminum Nitride coating is applied to the outer walls of the crossfire tubes which is used in combination with an air plasma spray (APS) Aluminum Bronze coating applied to the crossfire tube clip open end that mates with the crossfire tubes. The coating combination serves to protect the components from frictional wear and extend their life. In an alternate embodiment, the crossfire tubes also contain a plurality of air purge holes that provide a layer of cooling air along the crossfire tube inner wall.

Abstract: A support assembly for a gas turbine engine combustor including an inner liner and an inner casing spaced therefrom, wherein a longitudinal centerline axis extends through the gas turbine engine. The support assembly includes an annular inner support cone located adjacent an aft end of said inner liner, an annular nozzle support connected to the inner support cone, and a plurality of support members connected at a first end to a forward end of the inner liner and connected at a second end to the inner support cone.

Abstract: A fuel nozzle for a gas turbine engine includes a fuel tube enclosed by a housing. The fuel tube and housing are connected by a movable joint including a tubular seat formed on the housing and a boss formed on the fuel tube so as to be disposed within the tubular seat. The boss includes a contact surface that is in sliding engagement with the tubular seat and is rotatable about an axis that is perpendicular to the seat's longitudinal axis. The fuel nozzle further includes a fixed joint that connects the housing to the same section of the fuel tube as the movable joint.

Abstract: A multiplex injector system comprising an injector head, a first fuel path located in the injector head, and a first set of injector tips located in the injector head and in fluid communication with the first fuel path. The first set of injector tips includes at least one first injector tip. The multiplex injector further includes a second fuel path located in the injector head and a second set of injector tips located in the injector head and in fluid communication with the second fuel path. The second set of injector tips includes at least one second injector tip. A flow of fuel in each of the first and second fuel paths can be selectively controlled to control the flow of fuel through the first and second sets of injector tips.

Abstract: In a tail tube seal structure of gas turbine, a U-shaped groove is provided at one side of a tail tube seal where a flange of a tail tube outlet is fitted, and a pi-shaped groove is provided at other side of the tail tube seal where a gas pass side flange end is fitted, thereby composing the seal of the connection area. Inclined cooling holes are drilled in the tail tube seal in addition to the cooling holes existing conventionally. The cooling air flows in from the inclined holes and cools the gas pass side of the groove due to the film effect. Therefore, the difference in thermal expansion between the groove and flange end is decreased, the wear of this area is decreased, and the reliability of the seal is enhanced.

Abstract: A flame-passage device (10) for non-annular gas turbine combustion chambers comprises a tubular body (11), which is provided with a plurality of cooling holes (12), for cooling on the swirl-cooling type. The tubular body (11) is inserted in a flanged pipe (27), which connects the cases (31, 32) of two successive combustion chambers (17, 18), and has a first end (13) with a cylindrical shape, and a second end (14) with an oval shape, wherein the second, oval end (14) is provided with three rings (23, 24, 25), for anchorage of the corresponding combustion chamber (17) to the case (31).

Abstract: In a turbomachine comprising an annular shell of metal material containing in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material; and an annular nozzle of metal material forming the inlet stage with fixed blades of a high pressure turbine, provision is made for the combustion chamber to be held in position inside the annular metal shell by a plurality of flexible metal tongues each comprising three branches connected together in a star configuration, the ends of two of these three branches being fixed securely to a downstream end of the combustion chamber via respective first and second fixing means, and the end of the third branch being fixed securely to the annular shell via third fixing means.

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means on a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said metal fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells.

Abstract: A crossfire tube assembly with telescoping inner and outer crossfire tubes with an enhanced cooling mechanism for connecting adjacent combustors in a gas turbine is disclosed. The enhanced cooling configuration includes a plurality of channels formed in the telescoping region of the inner and outer crossfire tubes of the assembly to improve heat transfer and reduce local operating temperatures such that component life is extended.

Abstract: A turbine includes a sealing element with a receiving area for sealing the guide blade vanes which are adjacent to each other in the peripheral direction of the turbine. The foot plates of the guide blade vanes extend into the receiving area. The edge area of the foot plates does not have to be reinforced compared to a conventional seal, which enables the entire foot plate to be cooled homogeneously. A closed cooling system can therefore be used for cooling, especially with steam.

Abstract: A turbine installation, especially a gas turbine installation, includes foot plates of the guide blades of adjacent turbine stages being interconnected with a clip-type sealing element on their rear sides facing away from the gas area. This provides a simple seal between adjacent foot plates which is effective regardless of the thermal expansion of the foot plates. The clip-type sealing element is also suitable for sealing the tiles of a combustor of the turbine installation together.

Abstract: A turbomachine comprises an annular shell of metal material containing in a gas flow direction F a fuel injection assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine, said nozzle being supported by the annular shell and being fixed thereto by first releasable fixing means, and provision being made for the combustion chamber to be mounted in floating manner inside the annular shell and held in position solely by the nozzle to which it is fixed in resilient manner by second releasable fixing means.