Abstract: A turbine engine heat exchanger has an array of heat exchanger plates mounted to an inner case wall for providing heat transfer from a bleed flowpath to a bypass flowpath. Each plate has: first and second faces along the bypass flowpath; a proximal edge mounted to the inner case wall; an inlet along the proximal edge; an outlet along the proximal edge; and a branch segment of the bleed flowpath passing from the inlet to the outlet.

Abstract: A gas turbine engine includes a compressor, a combustor, and a turbine. The compressor compresses gases entering the gas turbine engine. The combustor receives the compressed gases from the compressor and mixes fuel with the compressed gases. The turbine receives the hot, high pressure combustion products created by the combustor by igniting the fuel mixed with the compressed gases. The turbine extracts mechanical work from the hot, high pressure combustion products to drive the compressor and a fan, shaft, or propeller.

Abstract: There is provided a load transfer interface in an aircraft engine for transferring a load from a bearing housing to an engine casing. The load transfer interface comprises a first component operatively coupled to and receiving the load from the bearing housing. The first component has a first annular body with a spigot extending axially from the first annular body. The interface comprises a second component operatively coupled to the first component and to the engine casing. The second component has a second annular body with a spigot-receiving cavity disposed therein. The spigot-receiving cavity is shaped and positioned to receive the spigot of the first component. The second component receives the load from the first component and transfers the load to the engine casing.

Abstract: Method for manufacturing a CMC, i.e. ceramic matrix composite material, part provided with at least one cutout, as well as to such a CMC part provided with at least one cutout, the method comprising the following steps: providing (E1) a fibrous reinforcement (10), forming (E2?) a cavity in a portion of the fibrous reinforcement (10), injecting (E3) a slip comprising at least a ceramic powder and a solvent, the slip being injected so as to impregnate the fibrous reinforcement (10?) and to fill the cavity of the fibrous reinforcement (10?), drying (E4) the obtained assembly, carrying out a densification (E6) by infiltration of a liquid densification material and solidification of said densification material, machining (E7) at least one cutout in the obtained blank (30) within the volume corresponding to the cavity of the fibrous reinforcement (10).

Abstract: Systems and methods of operating systems are provided. For example, a system comprises a fuel cooling loop including a cold fuel flowpath having a fuel flowing therethrough, a fuel cooler heat exchanger for cooling the fuel in fluid communication with the cold fuel flowpath, and a cold fuel tank disposed along the cold fuel flowpath for accumulating at least a portion of the cooled fuel. The system further comprises a fuel heating loop including a hot fuel flowpath for a flow of the fuel, a fuel heater heat exchanger for heating the fuel in fluid communication with the hot fuel flowpath, and a hot fuel tank disposed along the hot fuel flowpath for accumulating at least a portion of the heated fuel. The fuel cooling loop is coupled to the fuel heating loop such that the fuel circulates through both the fuel cooling loop and the fuel heating loop.

Abstract: A manufacturing method is provided during which a preform component for a turbine engine is provided. A cooling aperture is formed in the preform component. The cooling aperture includes a centerline, an inlet and an outlet. The cooling aperture extends longitudinally along the centerline through a wall of the preform component from the inlet to the outlet. The forming of the cooling aperture includes forming a first portion of the cooling aperture using a machining tool implement with a first toolpath that is angularly offset from the centerline by a first angle between thirty-five degrees and ninety degrees.

Abstract: A gas turbine engine includes a compressor section, a combustor section, and a turbine section operably coupled to the compressor section. A primary flow path is defined through the compressor section, the combustor section, and the turbine section. An engine case surrounds the compressor section, the combustor section, and the turbine section. The gas turbine engine also includes a means for providing an active variable cooling flow through a bypass duct external to the engine case to a secondary flow cavity of the turbine section.

Abstract: A system for modulating airflow into a bore defined by a rotor of a gas turbine engine defining an axial direction, a circumferential direction, and a radial direction is provided. The system includes a movable member positioned forward of a first stage of rotor blades of the rotor. The movable member is movable between at least a first position and a second position to modulate airflow into the bore via a plurality of opening in fluid communication with the bore.

Abstract: An assembly adapted for use with a gas turbine engine includes a static component and a metering band. The static component is fixed relative to an axis. The metering band is arranged to extend circumferentially at least partway about the axis and is coupled with the static component. The metering band defines at least a portion of a cooling passageway for air to flow through.

Abstract: A heat exchanger array includes a first row of heat exchangers, a second row of heat exchangers, and side curtains. The first row heat exchangers are spaced apart to define first gaps. The second row heat exchangers are spaced apart to define second gaps and are positioned downstream of and staggered from the first row heat exchangers such that the second row heat exchangers are aligned with the first gaps and the first row heat exchangers are aligned with the second gaps. Each side curtain is in close proximity to a first row heat exchanger and a second row heat exchanger. The side curtains define a neck region upstream of and aligned with each first row heat exchanger and each second row heat exchanger. Each neck region has a neck area that is less than a frontal area of the heat exchanger with which it is aligned.

Abstract: A gas turbine engine includes a platform that has a gas path side, a non-gas path side, a first mate face, and a second mate face. The second mate face has a beveled edge sloping towards the first mate face. The gas turbine engine also includes a coverplate that includes a first bend, a flat portion substantially parallel to the first mate face and a first wing substantially parallel to the second mate face.

Abstract: A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed.

Abstract: Aspects of the disclosure regard an exhaust nozzle for a gas turbine engine that includes an outer nozzle wall, a flow channel which is limited radially outwards by the nozzle wall, a centerbody arranged in the flow channel, and at least two struts connecting the centerbody to the nozzle wall. One of the struts is connected to the nozzle wall by a first connection, the first connection constraining movement of the strut relative to the nozzle wall in the radial direction and in the circumferential direction. The at least one other strut is connected to the nozzle wall by a second connection, the second connection constraining movement of the strut relative to the nozzle wall in the circumferential direction but allowing movement of the strut relative to the nozzle wall in the radial direction.

Abstract: The invention relates to a pressurized-air supply unit for an air-jet cooling device cooling an outer casing of a turbomachine turbine. This unit is notable in that it is monobloc, in that it comprises: —a body which has an interior wall provided with air-ejection perforations, an exterior wall and outlet ducts which are configured to be coupled to cooling lines of the cooling device, —an elbowed air-conveying pipe coupled by its outlet opening to said exterior wall of the body, and in that said unit comprises at least one air-distribution partition, arranged in the outlet opening and connecting the internal face of the portion of the air-conveying pipe that is situated facing the body to the internal face of the opposite portion of the air-conveying pipe.

Abstract: Cooling device (21) extending circumferentially around a turbomachine casing, such as a turbine casing, comprising a support (24) extending axially and designed to be secured to the casing, at least one cooling tube extending circumferentially, at least one securing member (25), comprising a radially inner portion (36) at least partially surrounding the tube, and a radially outer portion (37) secured to the support (24), the radially outer portion (37) of the securing member (25) being secured to the support (24) via the intermediary of a connection member (30) comprising a central portion (31) having a circumferential end portion (32) and a second circumferential end portion (32) which are circumferentially opposite, the first end portion (32) and the second end portion (32) each extending in an opposite axial direction, the first and second end portions (32, 32) each being secured to the support (24), the radially outer portion (37) of the securing member (25) being secured to the central portion (31) of th

Abstract: A combined power generation system improves the generation efficiency of a pressure difference power generation facility by using at least one of air for cooling a turbine of a gas turbine power generation facility and waste heat of flue gas generated by the gas turbine power generation facility. Working fluid to be used in a supercritical fluid power generation facility is cooled by using cold energy of liquefied natural gas. The system includes an air discharge channel via which compressed air is discharged; a fuel gas heater for heating the natural gas to be introduced into the pressure difference power generation facility by performing a heat exchange between the discharged air and the natural gas being heated; and a cooling air inflow channel for guiding the cooled air passed through the fuel gas heater to a turbine of the gas turbine power generation facility.

Abstract: A diffuser and deswirl flow structure includes a plurality of tube structures with an outer wall that is hollow and elongate and that extends between a first portion and a second portion. The plurality of tube structures is disposed in an annular arrangement about the longitudinal axis. The flow structure also includes a plurality of flow passages extending through the tube structures. The plurality of flow passages extend from the first portion to the second portion, respectively. The plurality of flow passages respectfully include a diffuser portion, which is proximate the first portion and configured to diffuse a fluid flow from a compressor wheel. The plurality of flow passages respectfully include a deswirl portion, which is proximate the second portion and configured to deswirl the fluid flow from the diffuser portion. The outer wall defines the diffuser portion and the deswirl portion. The outer wall is self-supporting.

Abstract: An exemplary gas turbine power plant includes a gas turbine with a compressor having a compressor inlet. A combustion chamber follows the compressor and a turbine follows the combustion chamber. A cross section of the compressor inlet includes an inner sector and an outer sector in relation to the axis of rotation of the compressor. A plurality of feed ducts introduces oxygen-reduced gas into the inner sector of the compressor inlet. The plurality of feed ducts is arranged in the compressor inlet so as to be distributed in a circumferential direction on a circle concentrically with respect to the axis of the gas turbine.

Abstract: A turbine blade for a gas turbine engine having a plurality of cooling holes defined therein, the plurality of cooling holes are located in an airfoil of the turbine blade according to the coordinates of Table 1.

Abstract: A gas turbine engine exhaust case assembly comprises: an exhaust case having an axis and defining an annular gas path; a bearing housing coaxially supported within the exhaust case; an exhaust cone coaxial and rearward of the exhaust case; a heat shield joined to the exhaust cone, the heat shield being disposed radially between the exhaust cone and the bearing housing; and a mounting bracket extending from the exhaust case and joining the exhaust case and the exhaust cone together.

Abstract: A turbine airfoil includes a body that has inner and outer platforms and an airfoil section that extends between the inner and outer platforms. There are film cooling holes that define external breakout points from the body. The external breakout points are located in accordance with Cartesian coordinates of at least points 222 through 256 set forth in Table 1 herein.

Abstract: An impeller tube-type nozzle for a gas turbine, with an inlet section, a retraction section and an outlet section. The inlet section is a section of annular channel, the retraction section comprises multiple gas flow channels separated by multiple blades, each gas flow channel is encircled by an outer peripheral wall face, an inner peripheral wall face, a suction face of one of two adjacent blades and a pressure face of the other of the two adjacent blades, and inlets of the gas flow channels have a fan-shaped cross section. For each gas flow channel, along the direction of gas flow from the inlet of the gas flow channel to the outlet the fan-shaped cross section gradually smoothly transitions into a circular cross section.

Abstract: Nozzle segments and methods of cooling airfoils of nozzle segments are provided. For example, a turbine nozzle segment includes an inner band defining an inner band cavity and/or an outer band defining an outer band cavity. The inner band may define an inner band aperture extending from the inner band cavity through the inner band, and the outer band may define an outer band aperture extending from the outer band cavity through the outer band. Inner and/or outer band cooling passages may extend through a trailing edge portion of a CMC airfoil of the nozzle segment. An inlet of any inner band cooling passage is defined adjacent an inner band aperture, and an inlet of any outer band cooling passage is defined adjacent an outer band aperture. The cooling passage inlets are aligned with the adjacent inner or outer band apertures to provide cooling fluid from the respective cavity.

Abstract: A turbocharge includes: a compressor wheel; a turbine wheel configured to rotate with the compressor wheel; a turbine housing disposed so as to cover the turbine wheel; a bearing supporting a rotational shaft of the turbine wheel rotatably; and a bearing housing accommodating the bearing. One of the turbine housing or the bearing housing includes a fin portion protruding toward the other one of the turbine housing or the bearing housing so as to extend along an axial direction of the rotational shaft, and, between the turbine housing and the bearing housing, a cavity is formed on each side of the fin portion with respect to a radial direction of the rotational shaft.

Abstract: An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber. A tangential onboard injector (TOBI) may be disposed in the first airflow path. A radial onboard injector (ROBI) may be disposed in the second airflow path.

Abstract: A guide vane for an aircraft turbomachine, the aerodynamic part of the vane being defined by an extrados body and an intrados body, the part comprising an internal lubricant cooling passage equipped with flow disturbing studs, including a first series of studs made in a single piece with the extrados body, and a second series of studs made in a single piece with the intrados body, the studs of the second series defining a second inter-stud space between them through which studs of the first series pass while the studs of the second series pass through the first inter-stud space. Furthermore, the end of the studs is located at a distance from the intrados body, and the end of the studs is located at a distance from the extrados body.

Abstract: An intercooled cooling system for a gas turbine engine includes a heat exchanger in fluid communication with a cooling airflow source directed through the heat exchanger and an auxiliary compressor fluidly coupled to the heat exchanger via a discharge duct to compress the cooling airflow exiting the heat exchanger. A compressor discharge pathway directs a first portion of the cooling airflow from the auxiliary compressor to a first cooling location of the gas turbine engine, and a bypass pathway is fluidly coupled to the discharge duct between the heat exchanger and the auxiliary compressor to direct a second portion of the cooling airflow to a second cooling location of the gas turbine without passing through the auxiliary compressor.

Abstract: In order to compensate for the increase in a leakage rate of ventilation air originating in a circuit via a labyrinth seal that gradually wears at the bottom of a ventilation chamber, and the reduction of a purge rate via an opening upstream from the chamber, a second injection opening is added to a main injection opening. The second injection opening communicates with the chamber via a second labyrinth seal that wears at the same time as the first, and thus gradually increases the air flow rate passing through the injector in order to compensate for the additional leakage rate through the first injector and maintain the ventilation at its initial level.

Abstract: Combustors and panels for use in combustor sections of gas turbine engines, the panels having a panel body having a first wall with a first sidewall and a second sidewall extending each extending from the first wall in the same direction at edges of the first wall and a plurality of cavity walls arranged between the first sidewall and the second sidewall and extending from the panel in the direction of the first and second sidewalls, wherein a plurality of cooling cavities are defined by the cavity walls, the first wall, and at least one of the first sidewall and the second sidewall, wherein each cooling cavity extends from an inlet hole to an outlet hole, wherein the outlet hole is formed in the first wall.

Abstract: Cooling systems for high pressure compressor systems are provided. The cooling systems may comprise tangential on board injectors (“TOBIs”). The TOBIs may comprise one or more fluid channels configured to conduct cooling fluid flow to components of the compressor, including, for example, disk-hub portions of the compressor. In this regard, the TOBI may be configured to exhaust cooling air in a manner such that the exhausted air has a similar linear velocity of the disk-hub portion. The cooling air may also be exhausted in a manner that is substantially parallel to the disk-hub portion.

Abstract: A guide vane for a twin-spool aircraft turbomachine has an aerodynamic part that includes an internal lubricant cooling passage extending along a principal lubricant flow direction. The aerodynamic part is made in a single piece and also includes heat transfer fins arranged in the passage connecting the intrados and extrados walls and extending approximately parallel to the direction, these fins being distributed in successive rows along the principal direction and made such that for two rows of staggered directly consecutive fins, a first row includes fins forming a positive acute angle A1 with a dummy reference plane, while a second row includes fins forming a negative acute angle A2 with this plane.

Abstract: According to an aspect, a system includes a starter air valve in fluid communication with an air turbine starter to drive motoring of a gas turbine engine responsive to a compressed air flow from a compressed air source. The system also includes a variable-position electromechanical device operable to adjust positioning of the starter air valve and a discrete-position electromechanical device operable to adjust positioning of the starter air valve and limit a motoring speed of the gas turbine engine below a resonance speed of the gas turbine engine responsive to a pulse width modulation control based on a failure of the variable-position electromechanical device.

Abstract: Combustor panels including panel bodies with first and second sides, a pin array extending from the first side, wherein each pin extends a first height, has a pin diameter, and is separated from adjacent pins by a pin array separation distance. A structural protrusion extends from the first side. No pins of the pin array are located within a flashing distance that is equal to a protrusion separation distance plus half of the pin diameter, wherein a location of the pin is measured from a center point of the pin to a closest point on the exterior surface of the structural protrusion. At least one pin array extension is integrally formed with the structural protrusion, the pin array extension extending along the first side to a position that replaces a pin of the pin array that would be within the flashing distance.

Abstract: A high pressure compressor has a downstream most end. A housing surrounds the compressor section and a turbine section. A low pressure turbine has a downstream most end. A first tap selectively taps high pressure cooling air from a location downstream of the downstream most end in the high pressure compressor and passes the high pressure cooling air through a heat exchanger. A second tap taps compressed air from a location upstream of the downstream most end in the high pressure compressor, and passes air over the heat exchanger, cooling the high pressure cooling air. A chamber is defined between the core engine housing and a nacelle airflow wall, and the second tap air flows through the chamber. The second tap air moves from the chamber into a core engine flow at a location downstream of the downstream most end of the low pressure turbine.

Abstract: Components for gas turbine engines are provided. The components include an airfoil extending between a first end and a second end and from a leading edge to a trailing edge. The airfoil has a pressure side and a suction side. The airfoil has a non-leading edge stagnation line that transitions from a location along a leading edge surface of the airfoil at the first end to a location on the airfoil on the pressure side of the airfoil at the second end. A leading edge feed cavity is located in the leading edge of the airfoil and extending between the first end and the second end of the airfoil. A stagnation impingement cavity is formed extending along a surface of the airfoil that aligns with the non-leading edge stagnation line.

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor is connected to be driven with at least one rotor in the main compressor section. A source of pressurized air is selectively sent to the cooling compressor to drive a rotor of the cooling compressor to rotate, and to in turn drive the at least one rotor of the main compressor section at start-up of the gas turbine engine. An intercooling system is also disclosed.

Abstract: To remove water accumulated in a cooling air system while preventing performance loss of a gas turbine, the gas turbine is provided with a cooling air system that connects an intermediate stage or the exit of a compressor to a turbine to supply compressed air bled from the compressor to the turbine, a TCA cooler, i.e., a heat exchanger, that cools the compressed air on the route of the cooling air system, and a drain water discharge valve, i.e., a first drain water discharge valve and a second drain water discharge valve disposed downstream of the compressed air of the TCA cooler, wherein at least for a predetermined period of time after a rated speed of the gas turbine at start up has been reached, the drain water discharge valve is set to an open state, and thereafter, the drain water discharge valve is set to a closed state.

Abstract: A flow transfer apparatus for transferring cooling flow from a primary gas flowpath to a turbine rotor. The apparatus includes a first supply plenum communicating with the primary gas flowpath and first inducers, the first inducers configured to accelerate a first fluid flow from the first supply plenum and discharge it toward the rotor with a tangential velocity; a second supply plenum communicating with the primary gas flowpath and second inducers, the second inducers configured to accelerate a second fluid flow from the second supply plenum towards the rotor with a tangential velocity; and a cooling modulation valve operable to selectively permit or block the second fluid flow from the primary gas flowpath to the second supply plenum. The valve includes a flow control structure disposed in the primary gas flowpath and an actuation structure extending to a location radially outside of a casing defining the primary gas flowpath.

Abstract: A turbine engine assembly is provided. The assembly includes a booster compressor that discharges a flow of first compressed air at a first temperature, and a high-pressure compressor (HPC). The HPC receives a first portion of the flow of first compressed air and discharges the first portion at a second temperature higher than the first temperature such that a flow of second compressed air is formed. The assembly also includes a heat exchanger (HEX) that receives a second portion of the flow of first compressed air from the booster compressor and a first portion of the flow of second compressed air from the HPC such that heat is transferred therebetween. The HEX discharges the first portion of the flow of second compressed air at a third temperature lower than the second temperature and higher than the first temperature such that a flow of cooled bleed air is formed.

Abstract: A gas turbine engine comprises a main compressor section having a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses ng air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A gas turbomachine includes a casing assembly surrounding a portion of the gas turbomachine and a counter-flow cooling system arranged within the casing. The counter-flow cooling system is configured and disposed to guide cooling fluid through the casing assembly in a first axial direction and return cooling fluid through the casing assembly in a second axial direction that is opposite the first axial direction.

Abstract: The chamber (4) of the turbine engine (1) comprises a continuous detonation wave engine (6) provided with an annular detonation chamber (7) and associated means (8, 9) that can be used to generate a continuous production of hot gases from a detonation mixture of fuel and air. The continuous detonation wave engine (6) is arranged such as to form, from a flow of incoming air (E), a first flow (F1) which enters the detonation chamber (7) and which is used by the engine (6) and a second flow (F2) which bypasses the chamber. The turbine engine (1) also includes auxiliary means (10) for mixing the hot gases (F3) leaving the detonation chamber (7) with the second flow of air (F2) before directing same towards the turbine (5). A plurality of detonation chambers (7) are arranged concentrically to one another relative to the axis of the turbine engine.

Abstract: A micro jet gas film generation apparatus aims to eject a gas to a work object which is desired for cooling or insulating from heat. The micro jet gas film generation apparatus has a spout formed at a diameter of 5-100 ?m to generate a gas film on the work object. As the diameter of the spout of the micro jet gas film generation apparatus is small, the micro jet gas film generated from the spout cannot produce a large eddy due to the lack of sufficient energy, hence can maintain a thin film after a long distance ejection to improve cooling and heat insulation performance. Moreover, due to the small diameter of the spout, it also consumes less gas and can reduce the amount of gas required.

Abstract: A buffer air cooler system for gas turbine engines disposed in a bypass duct of the engine, includes a housing for containing the buffer air cooler therein and an inlet portion attached to the housing. In one embodiment, the inlet portion has a double-skin configuration in at least one region of a top, bottom and sides of the inlet portion.

Abstract: In one aspect, a turbine engine is provided. The engine includes a fan, a core engine case configured to receive a core airflow, and an inner fan case. An inner air bypass is defined between the core engine case and the inner fan case, and the engine further includes an outer fan case, an outer air bypass defined between the inner fan case and the outer fan case, a first heat exchanger arranged in the core engine case, the first heat exchanger providing heat exchange between the liquid and the core airflow to cool the core airflow, and a second heat exchanger arranged in the outer air bypass. The second heat exchanger is fluidly coupled to the first heat exchanger to receive the cooled core airflow. The second heat exchanger provides heat exchange between cooled core airflow and the second portion of airflow to further cool the cooled core airflow.

Abstract: An engine exhaust system attachment is provided that includes a housing a fan, a filter, and a conduit. The housing is configured to enclose an exhaust turbine, an exhaust manifold, and at least a portion of an exhaust pipe. The exhaust turbine, the exhaust manifold, and the exhaust pipe are connected to receive exhaust gas from an engine and are mounted to an engine frame of a device. The fan is configured for mounting to the device to move air. The filter is configured for mounting to the fan to receive the moved air and to provide filtered air. The conduit is configured for connecting the filter to the housing to provide the filtered air to the housing.

Abstract: A gas turbine engine has, in flow series, a compressor section, a combustor, and a turbine section. The gas turbine engine further has a system (i) for cooling the turbine section and (ii) for providing tip clearance control between turbine blades of the turbine section and a plurality of circumferentially distributed segments which form an annular shroud surrounding the outer tips of the turbine blades. The system includes a turbine section cooling sub-system which diverts a first cooling air flow received from the compressor section to a heat exchanger and then to the turbine section to cool components thereof. The first cooling air flow by-passes the combustor and is cooled in the heat exchanger. The turbine section cooling subsystem has a first valve arrangement which regulates the first cooling air flow. The system further includes a tip clearance control sub-system which supplies a second cooling air flow to an engine case to which the segments are mounted.

Abstract: A cooling system is provided for a transition (420) of a gas turbine engine (410). The cooling system includes a cowling (460) configured to receive an air flow (111) from an outlet of a compressor section of the gas turbine engine (410). The cowling (460) is positioned adjacent to a region of the transition (420) to cool the transition region upon circulation of the air flow within the cowling (460). The cooling system further includes a manifold (121) to directly couple the air flow (111) from the compressor section outlet to an inlet (462) of the cowling (460). The cowling (460) is configured to circulate the air flow (111) within an interior space (426) of the cowling (460) that extends radially outward from an inner diameter (423) of the cowling to an outer diameter (424) of the cowling at an outer surface.

Abstract: An impingement plate for a turbine vane with an integrated cooling flow metering device is disclosed. The impingement plate—which covers the outer end of the vane and allows some cooling air to pass through to the vane's top surface—is re-designed to incorporate an orifice plate for metering the amount of cooling air flow which enters a cooling passage in the vane. The multi-hole orifice pattern in the metering device is designed to optimize the downstream airflow pattern, thus improving heat transfer from the vane to the cooling air. The reduced cooling air flow through the vane results in increased turbine engine efficiency, and the re-designed impingement plate can be used with the existing vane design.

Abstract: A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.