Abstract: Aspects of the invention relate to a system for attaching a fuel swirler assembly in a turbine engine combustor. According to embodiments of the invention, at least an axial upstream end of the fuel swirler assembly can be attached to a combustor support frame using one or more fasteners, such as bolts. The fuel swirler assembly can include a radial bracket provided on the swirler body to facilitate attachment. The axial downstream end of the fuel swirler assembly can be positioned within an opening defined in a swirler base plate or within an opening in a swirler support plate. Aspects of the invention can reduce the time and cost associated with installation and replacement of fuel swirlers in a combustor while avoiding concerns associated with known welded pin attachment systems.

Abstract: A mandrel is disclosed for fabricating a monolithic composite nacelle panel for use on an airplane. Also disclosed is the method of constructing the mandrel, the method of assembling the mandrel, and the method of disassembling the mandrel.

Abstract: Turboshaft engine including two subassemblies defining between them an annular chamber housing a seal. The two subassemblies are assembled under axial stress thereby defining an annular chamber housing the seal and an interposed part is inserted between the butting surfaces of the two parts of the annular chamber.

Abstract: A mounting system for a liner within a combination chamber of a gas turbine, where the liner is secured at one upstream end to a burner and at an opposite downstream end to a transition duct, the mounting system comprising: a first elastic head comprised of a first ring-shaped housing having first and second pluralities of radially spaced, axially extending teeth, with a first annular groove between said first and second pluralities of teeth, wherein said first elastic head is secured to said upstream end of said liner, with an end portion of said liner received within said first annular groove; and a second elastic head comprised of a second ring-shaped housing having third and fourth pluralities of radially-spaced, axially-extending teeth, with a second annular groove between said third and fourth pluralities of teeth, wherein said second elastic head is secured to said downstream end of said liner, with an opposite end portion of said liner received within said second annular groove.

Abstract: A method and apparatus for internally mounting a fuel manifold within a surrounding casing in a gas turbine engine which involves inserting the fuel manifold into the casing with support elements positioned in a retracted position, displacing the support elements into an extended position, and fastening said support elements to mount the manifold to an inner surface of said casing.

Abstract: A gas turbine fuel nozzle comprises a body and a sheath adapted to surround the body. A snap-on device is provided for releasably retaining the sheath on the body. The snap-on device is displaceable between a first position for allowing the sheath to be fitted over the body and a second position for retaining the sheath in place about the body.

Abstract: A device for fastening a fluid duct such as a fuel duct in an orifice of a turbojet casing, the device comprising a ring screwed into the orifice of the casing and having an annular rim coming to bear against a washer at the end of the duct, and a nut screwed onto the end of the duct so as to clamp the annular rim of the ring against the washer of the duct in order to fasten the duct to the casing.

Abstract: A fuel injector for a gas turbine engine is disclosed wherein the engine has an engine case that includes a reception bore for accommodating the fuel injector, and wherein the fuel injector includes a fuel inlet fitting having an annular mounting flange defining opposed upper and lower end surfaces and a cylindrical body portion which depends axially from the lower end surface of the mounting flange, the inlet fitting having integrally formed alignment structure located beneath the lower end surface of the mounting flange for guiding the fuel injector into an installed position within the reception bore of the engine case.

Abstract: The invention relates to an annular combustor (13) for a gas turbine (10), into which combustor (13) burners (14, 15) open on an inlet side, and which combustor (13) extends in the axial direction from the inlet side to an outlet side (33) and is lined on the insides with cooled liner segments (16, 17) for protection from the hot gases. In such a combustor, increased mechanical stability and flexibility in design and also simplification in manufacture and fitting are achieved by the liner segments (16, 17) being subdivided in the axial direction into a plurality of parts (16, 17) arranged one behind the other.

Abstract: A method of removably securing a reusable component to a replaceable component with a bonding agent is provided. The method includes, but is not limited to, preparing a first surface of the reusable component so that a bonding energy between the first surface and the bonding agent has a first level, preparing a second surface of the replaceable component so that a bonding energy between the second surface and the bonding agent has a second level, the second level being greater than the first level, and applying the bonding agent so that the bonding agent bonds the first and second surfaces to one another.

Abstract: An attachment device for coupling one or more removable turbine components to a carrier of a turbine engine. The attachment device may include a securement structure rotatable about an axis of rotation and including one or more arms for grasping an arm extending from a support structure. A fastener may extend from the support structure and be inserted into a fastener receiving recess in the securement structure for preventing the securement structure from rotating and becoming disengaged from the support structure. The attachment device may also include a fastener retaining device for preventing the fastener, which may be a threaded bolt, from becoming disengaged from the attachment device, traveling downstream in a turbine, and damaging a turbine assembly.

Abstract: A fastener shield for use in a fluid flow path within a gas turbine engine for reducing fluid drag and heating generated by fluid flow over a plurality of circumferentially spaced fasteners. The fastener shield has a radially-extending, downstream-facing mounting flange with a plurality of circumferentially spaced bolt holes positioned to receive respective engine mounting bolts therethrough and to attach the mounting flange to elements of the turbine engine. A curved, upstream-facing fastener shield cover is positioned in spaced-apart relation to the mounting flange for at least partially covering and separating an exposed, upstream-facing portion of the bolts from the fluid flow to thereby reduce drag and consequent heating of the bolts.

Abstract: A static structure for a miniature gas turbine engine includes a forward housing, a forward cover, a diffuser housing, a diffuser, a turbine nozzle, a combustor liner, a combustor housing and an exhaust pipe with relatively uncomplicated assembly interfaces. The static structure defines airflow and lubrication passages which communicate lubricant and airflow to each shaft bearing such that an adequate measured quantity of lubricant is supplied to the bearings in response to engine speed. The airflow through the airflow passage carries the lubrication supplied to the forward bearing toward the aft bearing to further lubricate the aft hot section bearing.

Abstract: Method and arrangement for providing a gas turbine (1) having a duct (11) for carrying gas from a gas turbine inlet (9) to a gas turbine outlet (10) and an outer housing (19, 20, 21) arranged radially outside a wall structure (12, 13, 14), which defines the radially outer limits of the gas duct (11). The gas turbine (1), between the inlet (9) and outlet (10), is constructed from a plurality of modules (6, 7, 8), each of which constitutes a part of the outer housing (19, 20, 21) and a part of the wall structure (12, 13, 14) of the gas duct. At least two adjacent parts of the wall structure (12, 13, 14) of the gas duct are arranged at a distance from one another.

Abstract: A method for assembling a gas turbine engine combustor facilitates reducing costs and time required for assembly. The combustor includes a spectacle plate, a plurality of swirlers, and a plurality of deflector plates. The method includes coupling an assembly fixture to at least one swirler, coupling the assembly fixture to the spectacle plate such that the swirler is maintained in alignment with respect to the spectacle plate during assembly of the combustor, and attaching the swirler to the spectacle plate.

Abstract: A method facilitates assembling a gas turbine engine. The method comprises coupling a combustor including a dome assembly and a combustor liner that extends downstream from the dome assembly to a combustor casing that is positioned radially outwardly from the combustor, coupling a ring support that includes a first radial flange, a second radial flange, and a plurality of beams that extend therebetween to the combustor casing, and coupling a primer nozzle including an injection tip to the combustor such that the primer nozzle extends axially through the dome assembly such that fuel may be discharged from the primer nozzle into the combustor during engine start-up operating conditions.

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: A combustion chamber (4) of a gas turbine (1), the combustion space (24) of which is bounded by an annular combustion chamber inner wall (28) and a combustion chamber outer wall (26), in which in order to generate a working medium (M) a supplied fuel is brought into reaction with supplied combustion air, and the combustion chamber wall (25) of which is provided on the inside with a lining formed from a plurality of heat shield elements (38), with the or each heat shield element (38) together with the combustion chamber wall (25) forming an inner space (40) to which a cooling medium (K) can be applied, is to be designed so as to provide a high level of system efficiency at the same time as having a comparatively simple structure and it should also be possible to disassemble the combustion chamber inner wall (28) in a time-saving manner.

Abstract: In a burner arrangement for the annular combustion chamber of a gas turbine, a plurality of burners are arranged above one another, in each case in pairs in the radial direction, on concentric rings. In a burner arrangement of this type, the mechanical integrity of the turbine casing is increased and operation of the combustion chamber is optimized by virtue of the two burners in a pair of burners being oriented out of the parallel position, in such a manner that their burner axes converge in the direction of flow.

Abstract: A gas turbo set comprises a compressor, combustion chamber and turbine, a common rotor shaft having arranged on it, within the compressor, a plurality of rows of compressor moving blades and, within the turbine, a plurality of rows of turbine moving blades, which alternate with guide vane rows of compressor guide vanes and turbine guide vanes fastened to a carrying structure surrounding the rotor shaft on the outside. Simplified mounting/demounting is achieved with the compressor guide vanes and/or the turbine guide vanes of a guide vane row held in each case in a plurality of separate annular segments together forming a complete ring with at least one guide vane row, and with the annular segments fastened releasably to the carrying structure such that the annular segments, together with the compressor or turbine guide vanes fastened to them, can be removed individually from the carrying structure, while maintaining the carrying structure.

Abstract: A multi-axial pivoting liner within the combustion system of a turbine engine allows the system to work with minimum thermal interference, especially during system operation at transient conditions, by allowing the liner to pivot and slide about its centerline and relative to the turbine scroll. The pivoting liner has the ability to control and minimize air leakage from part to part, for example, from the liner to the turbine scroll and liner to the surrounding structures, during various operating conditions. Additionally, the liner provides for easy assembly with no flow path steps. Finally, the pivoting liner tolerates thermal and mechanical stresses and minimizes thermal wear.

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 method facilitates assembling a gas turbine engine. The method comprises providing an engine frame including an integrally formed outer band, an inner band, and a plurality of circumferentially-spaced apart struts extending radially therebetween, and providing at least one fairing that is formed as an integral single piece casting and includes a first sidewall and a second sidewall connected at a leading edge and a trailing edge such that at least one cooling chamber is defined therebetween. The method also comprises coupling the at least one fairing around at least one strut such that the strut extends through the fairing at least one cooling chamber and such that during the coupling process the fairing is only transitioned axially around the strut rather being slid radially along the strut.

Abstract: A radial flow turbine including a rotor having a plurality of vanes defining an inlet, an outlet, and a flow path therebetween. A shroud is positioned to cover at least a portion of the flow path and a housing is positioned to at least partially support the rotor for rotation about a rotational axis. The housing at least partially defines a chamber for the receipt of a flow of products of combustion. A plurality of nozzle guide vane assemblies are positioned to provide fluid communication between the chamber and the inlet. Each of the plurality of nozzle guide vane assemblies is positioned adjacent another nozzle guide vane assembly to at least partially define one of a plurality of converging flow paths. Each nozzle guide vane assembly includes a guide vane positioned between the shroud and the housing and including an aperture therethrough. A bolt engages with the shroud and extends through the aperture to sandwich the guide vane between the shroud and the housing.

Abstract: A combustor baffle includes an annular splashplate having a center mounting tube therein for receiving a carbureted stream. A semi-ferrule is offset both laterally and transversely from the tube, and is open laterally outwardly therefrom. The semi-ferrule cooperates with an adjoining semi-ferrule for defining a port in which a pilot injector or igniter may be mounted.

Abstract: A method for assembling a gas turbine engine includes coupling an outer structure within the gas turbine engine, wherein the outer structure includes a socket extending from a radially inner surface of the outer structure, and coupling an inner structure to the outer structure by inserting a radial pin through the inner structure and into the socket such that the inner structure is aligned axially, circumferentially, and with respect to an engine centerline axis extending through the gas turbine engine.

Abstract: A screw includes a head and a shank, and includes a coolant passage which has an inflow opening and an outflow opening. The cooling pasage does not pass axially through the entire head and shank. By avoiding a coolant flow which passes axially through the entire screw, especially effective utilization of a coolant is achieved, in particular by coolant which also flows on the outer surface of the screw.

Abstract: The invention relates to an annular combustion chamber for a gas turbine wherein the annular combustion chamber extends in an axial direction, encloses a combustor and has on its inside facing the combustor a bearing structure on which a lining element secured to this lines the annular combustion chamber. The object is to disclose an annular combustion chamber with a lining element that meets the mechanical requirements while at the same time taking account of the system's maintenance-friendliness. The object is achieved in that the annular combustion chamber has a lining element wherein on the rear side facing away from the combustor of two edge areas on the lining element a plurality of interlocking means are located which have a hook width and wherein the lining element is secured to the corresponding bearing structure such that in order to release the lining element from the bearing structure the lining element is moved by the extent of the hook width of the interlocking means in the axial direction.

Abstract: A system for attaching liners to a carrier for forming inner surfaces of turbine combustor. The system may include one or more liners attached to a carrier using connectors engaged to the carrier on a cold side of the liner and may include one or more liners attached to the carrier using connectors engaged to a hot side of the liners. The carrier may include one or more access ports for accessing from the hot side connectors engaged to the carrier on the cold side. In at least one embodiment, the system may include attaching all liners, except for one liner, to a carrier using connectors coupled to the carrier on the cold side and may include attaching the one liner to the carrier using connectors coupled to the carrier on the hot side.

Abstract: An attachment and sealing system for securing a combustor cap assembly to a combustion chamber, while providing effective cooling to the combustor cap assembly aft end, is disclosed. The combustor cap assembly is secured within a combustion chamber by a plurality of pins such that the aft end region of the combustor cap assembly is cooled by a fluid medium that is injected through a plurality of first holes in the combustion chamber wall. A generally annular seal, which encompasses the combustor cap assembly, has a plurality of raised ridges, which are in sealing contact with the combustion chamber wall. Multiple embodiments are disclosed regarding the length of the generally annular seal and position and orientation of the plurality of first cooling holes in the combustion chamber wall.

Abstract: In a gas turbine power generator, a compressor rotor and a junction boss are assembled by fitting a convex faucet joint provided on the compressor rotor and a concave faucet joint provided on the junction boss, in a freely releasable manner, and the compressor rotor and a rotor shaft are assembled by fitting the convex faucet joint provided on the compressor rotor and a concave faucet joint provided on the rotor shaft, in a freely releasable manner, and those compressor rotor, the rotor shaft and a generator core are combined in series and fastened by using a tie bolt and a nut, which is connected to the junction boss of the turbine rotor.

Abstract: A two-stage gas turbine combustor is disclosed that incorporates an improved assembly and disassembly methodology that results in improved inspection and repair techniques. The combustor includes an upper liner, lower liner, cap assembly, and a venturi assembly that directs cooling air into the combustion chamber for use in the combustion process, wherein a venturi assembly is inserted within an upper liner and an upper liner is then inserted in a lower liner. The venturi assembly, upper liner, and lower liner are fixed together with a plurality of easily removable retaining pins. The improved assembly allows for easier inspection and repair of the upper liner, lower liner, and venturi assembly.

Abstract: A mounting system, for installing a starter-generator to an accessory housing, where the starter-generator has a mounting end with a connecting flange and a cooling air outlet radially inward of a peripheral starter-generator housing edge, and where the accessory housing has a matching connecting flange radially inward of a peripheral accessory housing edge. The mounting system includes a cooling air outlet collection shroud enveloping the peripheral starter-generator housing edge and the peripheral accessory housing edge, the shroud having a discharge outlet. A V-band clamp releasably secures the connecting flanges together and is enveloped within the shroud in a compact mounting system that collects and discharges the exhausted cooling air.

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: A gas turbine exhaust diffuser includes a substantially cylindrical outer casing provided with a forward casing flange; a substantially cylindrical liner spaced radially inwardly of, and secured to, the outer casing; and ceramic fiber insulation supported radially between the outer casing and the liner. A plurality of liner closeout segments are secured to the liner and include radial flange portions extending toward the forward casing flange. A plurality of spacer rings are supported between the forward casing flange and the liner closeout segments such that one of the spacer rings and the liner closeout segments present a flat face for direct attachment to an aft flange component of a gas turbine exhaust frame. This arrangement provides a thermally compliant load bearing interface between an air cooled gas turbine flange and an internally insulated diffuser casing flange in the high velocity section of the exhaust diffuser.

Abstract: Turbine insulation is disclosed, and comprises blankets, ribs and coupling members. The blankets define, with the turbine, an annular channel in receipt of the turbine vertical joint and a pair of longitudinal channels each in receipt of a respective turbine horizontal joint. A pair of ribs are positioned in each longitudinal channel, in volume-filling relation thereto and overlapping the blankets disposed, respectively, fore and aft of the vertical joint. A pair of ribs are positioned in volume-filling relation to the annular channel and overlapping the blankets disposed, respectively, on the upper and lower arcuate portions of the housing. A coupling member is positioned in each longitudinal channel at its intersection with the annular channel, in volume-filling relation to a void defined by the housing and the ribs and in lapped relation to the ribs.

Abstract: A method of making a gas turbine engine combustor using the steps of: providing a combustor; providing a circumferentially extending louver member for ducting a flow of compressed air through at least one inlet opening in a combustor wall from a source of compressed air outside the combustor; and mounting the louver member in a non-destructive manner for releasable connection and disconnection to an interior surface of the combustor wall and at least partially covering the at least one inlet opening, the louver having a number of outlet openings in flow communication with an inlet opening.

Abstract: The invention relates to a combustion-chamber arrangement for gas turbines, having an annular combustion chamber connected to at least one burner and leading into a turbine space, an annular-combustion-chamber wall which defines the annular combustion chamber being of double-shell construction, and an inner shell of the annular-combustion-chamber wall being composed of lining elements releasably arranged on an outer shell of the annular-combustion-chamber wall, and a gap space through which a cooling medium can flow being formed between the inner shell and the outer shell of the annular-combustion-chamber wall.

Abstract: The invention relates to an annular combustion chamber (4) for a gas turbine (1) wherein the annular combustion chamber (4) extends in an axial direction (A), encloses a combustor (7), and has on its inside facing the combustor (7) a bearing structure (26) on which a lining element (10) secured to this lines the annular combustion chamber (4). The object is to disclose an annular combustion chamber (4) with a lining element (10) that meets the mechanical requirements while at the same time taking account of the system's maintenance-friendliness.

Abstract: A fastener (31) is disclosed for a first article, for example a tile (29), to a second article for example an outer wall (27) of a wall structure (21, 22) of a combustor (15) of a gas turbine engine (10). The fastener (31) comprises an elongate main part, which may be in the form of a bolt (34). The main part can extend through the first and second articles. The main part defines a cooling pathway (46) therethrough to allow the passage of a cooling gas through the main part.

Abstract: A combustion turbine power plant includes a modular compressor assembly structured to enclose multiple configurations of vanes and blades, a modular combustor assembly structured to be disengagably coupleable to a fuel source, a modular first turbine assembly structured to enclose multiple configurations of vanes and blades, a modular diffuser, a modular second turbine unit, a work unit, and a modular heat exchange system. A work unit shaft is structured to have two ends that are disengagably coupleable to the modular second turbine unit and the work unit. A compressor turbine shaft is structured to have two ends that are disengagably coupleable to the modular compressor assembly and the modular compressor turbine assembly.

Abstract: The present invention relates to a combustion chamber head for a gas turbine with at least one combustion chamber wall 2,3, a metering panel 1, at least one heat shield 4, at least one sleeve 5 and a cowling 6, characterized in that the metering panel 1 forms one part with the combustion chamber outer wall 2 and the combustion chamber inner wall 3, in that the heat shield 4 is mounted onto the metering panel 1 from the downstream side of the metering panel 1, in that the sleeve 5 is mounted to the metering panel 1 from the upstream side, and in that the cowling 6 is separably attached to the metering panel 1.

Abstract: Interaction between mating gas turbine combustor components due to vibrations and assembly tolerances can cause excessive wear and premature replacement. Typically, wear reduction features, such as hardened materials, are used to minimize the amount of wear or direct such wear to a specific component. The present invention discloses an improved wear reduction device for the interface between a gas turbine combustor transition duct end frame and the turbine inlet. The improved wear reduction device incorporates “L-shaped” wear strip inserts, constructed of a hardened material, and fixed along the sidewalls of the end frame to protect it from wear by the mating turbine inlet seal. The improved wear strip insert provides a cost savings by not requiring costly manufacturing tooling to form the wear strip inserts, unlike previous wear reduction techniques for the same component interface.

Abstract: A combustor baffle includes an outer tube with external threads and a heat shield at opposite ends thereof. An inner tube is disposed inside the outer tube in a unitary assembly therewith. The outer tube is retained in a combustor dome by a retention nut, and the inner tube supports an air swirler with a brazed joint therewith. The brazed joint permits sacrifice of the baffle during disassembly for access to the threaded joint for final disassembly without damage to the dome or air swirler.

Abstract: Concentric installation of a pilot fuel assembly in an opening in a gas turbine combustor casing is achieved by providing a boss having at least two flat surfaces which are perpendicular to each other on the combustor casing surrounding the opening and a mounting flange having at least two flat surfaces which are perpendicular to each other on the pilot fuel assembly. The pilot fuel assembly is concentrically installed to the combustor casing by inserting the assembly into the combustor casing opening, and moving the pilot fuel assembly as far as it will go in a first direction substantially parallel to one of the flat boss surfaces. The distance between the other flat boss surface and one of the flat flange surfaces is then taken. Next, the pilot fuel assembly is moved in the direction opposite the first direction, at which point, the distance between the same two flat surfaces is again measured.

Abstract: The invention discloses an aeroengine nacelle afterbody of the type comprising a common nozzle for exhausting the mixed hot and cold gases originating respectively from the fan duct and from the aeroengine combustion chamber, the said nozzle having a longitudinal axis more or less coincident with the axis of the said engine. The common exhaust nozzle consists of two half-nozzles which are more or less symmetric with respect to a plane containing the said longitudinal axis, the said half-nozzles being, on the one hand, individually hinged via one of their longitudinal edges to a support structure hanging down from the wing structure or from the fuselage of the aircraft and, on the other hand, lockable together along their opposite longitudinal edges.

Abstract: The heat exchanger for exchanging heat between compressed air ejected from the compressor and exhaust gas ejected from the turbine rotor is connected to the exhaust gas outlet side on the back portion of the gas turbine. The heat exchanger has the core having a plurality of heat transfer plates for partitioning the first paths for compressed air and the second paths for exhaust gas inside the casing. On the side of the back portion of the core, the first inlets for compressed air are formed and on the side of the front portion of the core, the outlets for compressed air are formed. Between the core and the casing, the introduction path for introducing compressed air into the first inlets is formed. Exhaust gas flowing into the second paths from a position in front of the core is ejected from the back of the core. No pipe for connecting the gas turbine and heat exchanger is required and miniaturization of the gas turbine apparatus and a reduction in cost can be realized.

Abstract: A heat shield and support strut for a turbine engine afterburner. The heat shield can be replaced without moving or disconnecting the fuel tube. The heat shield includes a housing that may be telescoped over the fuel injection tube and is attachable to the strut without moving the housing transverse to a longitudinal axis of the fuel injection tube. The support strut includes a body having an outer end for attachment to an outer wall of the afterburner section, an inner end, and a passage extending through the body from its outer end to its inner end for receiving the fuel injection tube. The inner end has a cavity for receiving the head of the heat shield. The cavity is shaped for a close fit on the head in the cavity. The body has a flat mounting surface in the cavity for engagement by the flat mounting surface of the head.

Abstract: A cowl for directing flow of air entering an annular combustor assembly of a gas turbine engine and which is removable from the combustor without disassembling the combustor. The cowl is annular in shape and has a plurality of circumferentially spaced fastener holes for receiving fasteners which commonly connect and align the cowl, an annular dome, and a liner. At least one notch is in the cowl, each notch being sized sufficiently larger than a corresponding fastener to permit the cowl to be moved to or removed from an operative position while the corresponding fastener remains in position for connecting the dome and liner to one another.

Abstract: A combustion liner assembly for reducing nitric oxide (NOx) emissions is disclosed. This combustion liner utilizes an annular plenum outside the liner shell for the purpose of containing cooling air that has been heated and reintroducing it into the combustion process. The air plenum can be enlarged to increase the amount of cooling air contained within and still allow for installation of the liner assembly into the combustor by utilizing recesses that locally reduce the diameter of the air plenum to allow it to pass by liner mounting pegs within the combustor assembly.