Abstract: A fuel nozzle for a gas turbine engine combustor includes a fuel nozzle assembly including an inflow tube disposed along and a nozzle axis, the inflow tube defining an inner air passage and a liquid swirler concentrically disposed about the inflow tube, the liquid swirler including a liquid swirler inner wall having a liquid swirler inner wall end portion angled radially outward from the nozzle axis, a liquid swirler outer wall having a liquid swirler outer wall end portion angled radially outward from the nozzle axis, and an annular liquid passage defined therebetween. The fuel nozzle assembly further includes a radial air swirler (RAS) concentrically disposed about the liquid swirler outer wall, the RAS including an RAS inner wall having an RAS inner wall end portion angled radially inward toward the nozzle axis, an RAS outer wall having an end cap at a downstream-most position, and an annular gas passage defined therebetween.

Abstract: A combustion section of a gas turbine includes an annular combustion chamber, a burner arrangement upstream of the combustion chamber, an annular compressor diffusor, and an annular combustion casing. The combustion casing has an outer section wall, an upstream section wall, and an inner section wall. The inner section wall extends along the rotor axis and is arranged between the burner arrangement and the compressor diffusor. A shielding is attached to the inner section wall at the side facing the rotor axis with a free space in-between at least. An outer diffusor wall of the compressor diffusor extends in axial direction up to the upstream half of the shielding.

Abstract: A fuel nozzle for a gas turbine engine combustor includes a fuel nozzle assembly including an inflow tube disposed along and a nozzle axis, the inflow tube defining an inner air passage, and a liquid swirler concentrically disposed about the inflow tube, the liquid swirler including a liquid swirler inner wall, liquid swirler outer wall having a liquid swirler outer wall end portion angled radially inward toward the nozzle axis, and an annular liquid passage defined therebetween. The fuel nozzle assembly further includes a radial air swirler (RAS) concentrically disposed about the liquid swirler outer wall, the RAS including an RAS inner wall having an RAS inner wall end portion angled radially inward toward the nozzle axis such that it is parallel to the liquid swirler outer wall end portion, an RAS outer wall having an end cap at a downstream-most position, and an annular gas passage defined therebetween. The end cap includes a radiused inner surface and an outer surface.

Abstract: An electrical assembly for an aeronautical turbomachine, including an electric machine configured to be disposed in a turbomachine and comprising a stator and a rotor comprising magnets, the assembly including a short-circuit detecting means, a hot air injecting means configured to draw hot air off the turbomachine at a temperature greater than the temperature of demagnetization of the magnets of the rotor, and to inject the drawn hot air onto the magnets of said rotor when the short-circuit detecting means detects the presence of a short-circuit in the electric machine, and a cool air injecting means, configured to draw cool air off the turbomachine and to inject it into an inner chamber of the turbomachine, the temperature of the cool air drawn by the cool air injecting means being less than the temperature of the hot air drawn by the hot air injecting means.

Abstract: A heat management system includes a fuel tank storing a fuel; a first heat exchanger thermally coupled to the fuel tank; a hydraulic pump for circulating a hydraulic fluid; a hydraulic circuit including first and second hydraulic lines fluidly coupled to the first heat exchanger and the hydraulic pump, such that the first heat exchanger brings the hydraulic fluid and the fuel into a heat exchange relationship; an oil circuit; and a second heat exchanger thermally coupled to the oil circuit and at least one of the first and second hydraulic lines, such that the second heat exchanger brings the hydraulic fluid and the oil into a heat exchange relationship, thereby allowing heat transfer between the fuel and the oil via the hydraulic fluid.

Abstract: An assembly is provided for a turbine engine. This assembly includes a rotating structure rotatable about an axis, a stationary structure and a bearing. The stationary structure forms a bearing compartment, a buffer cavity, a low pressure cavity and a high pressure cavity with the rotating structure. The buffer cavity surrounds the bearing compartment. The low pressure cavity surrounds the buffer cavity. The high pressure cavity surrounds the low pressure cavity. The bearing is within the bearing compartment. The bearing rotatably couples the rotating structure to the stationary structure.

Abstract: An integrated oil system manifold is provided, which includes a combination of a reverse rotation protection element, a first pump check valve, a second pump check valve, and a pressure relief valve integrated with a housing. Furthermore, a combination of a temperature control element, a filter, a heat exchanger, an integral main pump assembly, and an integral auxiliary pump assembly may also be included in the integrated manifold. An associated method is also provided.

Abstract: A rotating detonation combustor includes a combustion chamber configured for a rotating detonation process to produce a flow of combustion gas and an air plenum configured to contain a volume of air. The rotating detonation combustor also includes a flow passage coupled in flow communication between the combustion chamber and the air plenum and configured to channel an airflow from the air plenum. The rotating detonation combustor also includes a fuel inlet coupled in flow communication with the flow passage and configured to channel a fuel flow into the flow passage. The flow passage includes a plurality of fuel mixing mechanisms configured to mix the airflow and the fuel flow within the combustion chamber.

Abstract: A method of operating a gas turbine engine including an engine core and turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a main gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft, the main gearbox including gears and journal bearings; a recirculating lubrication system arranged to supply oil to lubricate the gears and journal bearings, the lubrication system including a first oil tank arranged to supply oil to the gears and journal bearings and a second oil tank arranged to supply oil to the journal bearings only. The method includes, at cruise conditions: transferring 200-600 kJ/m3 of heat to the fuel from the oil through the heat exchange system so as to control the oil temperature.

Abstract: An electrical power generation system including: a micro-turbine alternator, including: a decomposition chamber; a turbine including blades driven by combustion gases from the decomposition chamber; a blower operably connected to the decomposition chamber to provide a blown airflow thereto; one or more shafts connecting the turbine to the blower such that rotation of the turbine drives rotation of the blower; an electric generator disposed along the one or more shafts such that electrical power is generated via rotation of the one or more shafts; and a housing enclosing the electric generator within a generator cavity formed therein. The blower is configured to blow air through the generator cavity through or around the electric generator. The air is configured to exit the generator cavity to enter the turbine.

Abstract: A method is provided for operating a gas turbine engine. The method includes: determining data indicative of an operation of a cooling system of the gas turbine engine during a shutdown of the gas turbine engine, following the shutdown of the gas turbine engine, or both; and modifying a startup schedule of the gas turbine engine in response to the determined data indicative of the operation of the cooling system of the gas turbine engine.

Abstract: A bypass turbine engine includes a fixed casing, a first shaft (low-pressure shaft), a second shaft (high-pressure shaft), at least one accessory to be driven by a motor powered with electrical energy, a first intermediate shaft tapping mechanical power off the low-pressure shaft, a second intermediate shaft tapping mechanical power off the high-pressure shaft, and an electrical energy generator assembly coupled to the first and second intermediate shafts so as to receive mechanical power from the first and second intermediate shafts. The generator assembly converts the mechanical power received from the first and second intermediate shafts into electrical energy to power the motor or motors, which comes simultaneously from the mechanical power tapped off the low-pressure shaft and the mechanical power tapped off the high-pressure shaft. The generator assembly is housed in an arm in the lower part of the turbine engine and extending vertically into a bypass flow duct.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A gas turbine engine according to a first embodiment, having a case; a core within the case; an exhaust nozzle that is fluidly coupled to the case; a duct having a duct outlet that is fluidly coupled to the exhaust nozzle and a duct body extending from the duct outlet to a duct inlet; an air/oil cooler having an airflow outlet disposed at the duct inlet and a cooler body extending from the airflow outlet to an airflow inlet; and an airflow control baffle disposed at one of: the duct inlet; the duct outlet; and the airflow inlet of the cooler, wherein the airflow control baffle is configured to open when oil temperature in the engine is above a first threshold and close when the oil temperature is below a second threshold that is lower than the first threshold.

Abstract: A fuel system is disclosed for a gas turbine engine, which includes an augmentor pump having an inlet communicating with a fuel supply source and a discharge communicating with an augmentation stage of the engine, wherein the augmentor pump is connected to an accessory drive gearbox mounted to the engine, and a high speed clutch for selectively engaging and disengaging the augmentor pump and the accessory drive gearbox.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Abstract: A combination of a gas turbine engine and power electronics, includes an engine core and oil circuit to cool and lubricate bearings of the engine core, and a fuel circuit for supplying fuel to the combustor. The fuel circuit includes a low pressure pump for pressurising the fuel to a low pressure, and a high pressure pump to receive the low pressure fuel and increase the pressure to a high pressure for supply to a fuel metering system and the combustor. The engine includes a fuel-oil heat exchanger having a fuel side on the fuel circuit between an outlet of the low pressure pump and an inlet of the high pressure pump, and an oil side on the oil circuit to transfer heat from the oil circuit to the fuel circuit. The power electronics transfers heat to a cooling flow formed by a portion of the low pressure fuel.

Abstract: A gas turbine engine including a fuel delivery system arranged to provide fuel; a combustor arranged to combust at least a proportion of the fuel; a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough; and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. The method includes using at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger to heat the fuel so as to adjust the fuel viscosity to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: A heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

Abstract: A heat exchange system of a turbomachine, includes a heat exchanger including a support wall, a plurality of fins each extending in a radial direction from a radially outer surface of the support wall, and a cover covering the fins, wherein the cover is connected, upstream in the direction of flow of the air flow, to a first profiled wall, and downstream to a second profiled wall, the first profiled wall being arranged upstream from the fins and configured to guide and slow down the flow of air entering the heat exchanger through the fins, and the second profiled wall being arranged downstream from the fins and configured so as to accelerate the flow of air leaving the heat exchanger, wherein the cover has an at least partially curvilinear aerodynamic profile and an outer peripheral surface having surface continuity with radially outer surfaces of the first and second walls.

Abstract: A turbine engine that includes an engine core, a core casing surrounding the engine core, at least one cowl surrounding the core casing, a power converter located between the at least one cowl and the core casing, an electrical generator disposed within the engine core, at least one electrical strut having an outer wall defining an interior and extending radially between the engine core and the core casing; and at least one electrical conduit located within the interior and electrically coupling the generator to the power converter.

Abstract: In accordance with at least one aspect of this disclosure, a fluid system of an aircraft includes a primary fluid conduit that conveys a primary fluid, and a leak detection system disposed around at least a portion of the primary fluid conduit and forming one or more detection volumes. The leak detection system determines whether there is a primary fluid leak into the one or more detection volumes by sensing a pressure change in the one or more detection volumes.

Abstract: A propulsion assembly with a primary duct and a secondary duct, an outer fairing delimiting an outer compartment with an inlet opening, an inner fairing delimiting an inner compartment with an air exhaust aperture, a suction system having an inlet, an outlet and mobile elements that draw the air in via the inlet and deliver the air via the outlet, a drive that drives the mobile elements, a transfer pipe connected to the outlet, a distribution pipe disposed in the inner compartment, connected to the transfer pipe and having nozzles, a supply pipe that opens into the outer compartment and is connected to the inlet, and a regulator that regulates the flow rate of air in the suction system. The outside air is then driven by the suction system and passes through the two compartments in a forced manner.

Abstract: An auxiliary tank for an aircraft turbine engine has a curved general shape, and includes a cylindrical or frustoconical radially outer wall and three cylindrical or frustoconical radially inner walls arranged facing the radially outer wall, the three radially inner walls having a middle wall and two side walls arranged to either side of the middle wall. The inner volume of the tank forms a U-shape, the branches of which are formed by lateral volume portions between the side walls and the radially outer wall, and the base of which is formed by a middle volume portion between the middle wall and the radially outer wall.

Abstract: A heat engine comprising a compressor providing a flow of compressed air from a core flowpath of the heat engine; a cooled cooling air (CCA) heat exchanger system to which the flow of compressed air is provided from the compressor; a coolant supply system providing a flow of coolant to the CCA heat exchanger in thermal communication with the flow of compressed air at the CCA heat exchanger, in which the coolant supply system and CCA heat exchanger together define a CCA circuit through which the compressed air flows in thermal communication with the coolant; and a hot section disposed downstream of the compressor section along the core flowpath through which combustion gases flow, in which the hot section defines a secondary flowpath through which the flow of compressed air from the CCA heat exchanger is provided.

Abstract: Methods, apparatus, systems and articles of manufacture are disclosed to illustrate a clearance design process and strategy with CCA-ACC optimization for exhaust gas temperature (EGT) and performance improvement. In some examples, an apparatus includes a case surrounding at least part of a turbine engine, the at least part of the turbine engine including a turbine or a compressor. The apparatus further includes a first source to obtain external air; a second source to obtain cooled cooling air; a heat exchanger to control temperature of cooled cooling air; and a case cooler to provide active clearance control air to the case to control deflection of the case, wherein the active clearance control air is a combination of the external air and the cooled cooling air, the case cooler coupled to the heat exchanger using a first valve, the first valve triggered by a first control signal.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inner structure that extends axially along and circumferentially about an axial centerline. The nacelle inner structure includes an internal compartment and a cowl. The internal compartment is configured to house a core of a gas turbine engine. The cowl is configured to form an outer radial periphery of the internal compartment. An aft end portion of the cowl is also configured to form an outer radial periphery of a compartment exhaust to the internal compartment. The aft end portion of the cowl includes a plurality of axial fingers arranged circumferentially about the axial centerline in an array.

Abstract: A heat exchange system for a turbomachine includes a heat exchanger that has a support wall extending along a longitudinal direction L and a plurality of fins each extending along a radial direction from a radially external surface of the support wall. The heat exchanger further includes a first profiled wall arranged upstream from the fins and configured to guide and slow down the flow of air entering the heat exchanger through the fins. A second profiled wall is arranged downstream from the fins and configured to accelerate the flow of air leaving the heat exchanger. Each first and second profiled wall is attached to the support wall via support elements extending radially from the radially external surface.

Abstract: A starting/generating system for an aircraft turbine engine, the starting/generating system comprising at least one brushless drive motor/generator, at least one control module and at least one power module, the power module being configured to supply/receive electric power from the brushless drive motor/generator, the control module being connected to the brushless drive motor/generator by a control cable in order to control its operation, in which system the power module is configured to be mounted in the housing of the non-pressurized zone so as to be located adjacent to the brushless drive motor/generator and the control module is configured to be mounted in a pressurized zone of the aircraft, the control module being connected to the power module by a two-way communication cable in order to control its operation.

Abstract: A heat exchanger includes a plurality of fins arranged as a network and delimiting corridors, and an envelope having an internal wall and an external wall, the internal and external walls delimiting between them a channel for a flow of a first fluid in a main direction, the network of fins being arranged in the channel and connected to the internal and external walls, at least one passage for a flow of a second fluid being embedded in at least one of the internal and external walls, the channel being, in the main direction, divergent and then convergent.

Abstract: A propulsion unit for an aircraft includes a nacelle, a turbojet engine, an annular flow path for circulating a secondary air flow, and a precooler device communicating with a motor enclosure and including a scoop opening into the annular flow path. The propulsion unit includes a compressed air supply circuit arranged in the propulsion unit for injecting a flow of compressed air into the scoop of the precooler device. A method for ventilating a motor enclosure of a propulsion unit includes injecting compressed air into a scoop of the precooler device when the turbojet engine is stopped.

Abstract: Aircraft and aircraft blower systems are described. The blower systems include a shaft, a motor having a stator and a rotor, the rotor being operably coupled to the shaft, a fan operably coupled to the shaft and configured to be driven by rotation of the shaft, one or more bearings arranged along the shaft, and a high pressure cooling source configured to supply high pressure air to the one or more bearings.

Abstract: Fuel injectors having a fuel-air heat exchange system and methods thereof for use in a gas turbine engine. The fuel-air heat exchanger allows heat transfer between a flow of cooling air used to cool components of the engine and a flow of fuel used to drive the engine to transfer heat to the flow of fuel and cool the cooling air.

Abstract: A gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, and a front mount and a rear mount, the front and rear mounts being configured to connect the gas turbine engine to the aircraft, wherein the front mount is coupled to a casing of the engine core and the front mount is located at substantially the same axial position as a centre of gravity (CG) of the gas turbine engine or forward of the centre of gravity of the gas turbine engine.

Abstract: A heat transfer system includes a heat exchanger located at least partially within a coolant flowpath. The heat exchanger defines at least in part a first flowpath and a second flowpath, the first flowpath configured to be in fluid communication with the coolant flowpath, and the second flowpath configured to receive a flow of a motive fluid. The heat transfer system further includes a throttling device that is in fluid communication with the second flowpath of the heat exchanger. The heat exchanger receives at least a portion of the flow of the motive fluid from the heat exchanger. The throttling device is also in fluid communication with the coolant flowpath at a location upstream of the heat exchanger for providing the flow of motive fluid to the coolant flowpath at the location upstream of the heat exchanger.

Abstract: The present invention discloses a soundproof cabin of a turbine engine. The soundproof cabin is sleeved on the turbine engine. The soundproof cabin includes a cabin body, an intake noise reduction unit and a ventilation noise reduction unit, wherein the intake noise reduction unit and the ventilation noise reduction unit are disposed on the cabin body, the surrounding of which is filled with soundproof materials, the intake noise reduction unit is used to reduce the induction noise of the turbine engine, the ventilation noise reduction unit is used to reduce the noise of the ventilation system of the turbine engine.

Abstract: A hot fairing structure for a pre-diffuser includes a ring-strut-ring structure that comprises a multiple of hollow struts; and a multiple of diffusion passage ducts attached to the ring-strut-ring structure. A pre-diffuser for a gas turbine engine includes an exit guide vane ring having a multiple of exit guide vanes defined around an engine longitudinal axis; a ring-strut-ring structure adjacent to the exit guide vane ring to form a multiple of diffusion passages defined around the engine longitudinal axis, an inlet to each of the multiple of diffusion passages smaller than an exit from each of the multiple diffusion passage through the ring-strut-ring structure; a diffusion passage duct attached to the ring-strut-ring structure at the exit from each of the multiple diffusion passage.

Abstract: A heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

Abstract: A hot fairing structure for a pre-diffuser includes a ring-strut-ring structure that comprises a multiple of hollow struts and a multiple of inlets to a respective diffusion passage, one of the multiple of inlets formed between each one of the multiple of hollow struts located between two diffusion passages.

Abstract: A dispatchable storage combined cycle power plant comprises a topping cycle that combusts fuel to generate electricity and produce hot exhaust gases, a steam power system, a heat source other than the topping cycle, and a thermal energy storage system. Heat from the heat source, from the thermal energy storage system, or from the heat source and the thermal energy storage system is used to generate steam in the steam power system. Heat from the topping cycle may be used in series with or in parallel with the thermal energy storage system and/or the heat source to generate the steam, and additionally to super heat the steam.

Abstract: A gas turbine engine including: a tail cone; a low pressure compressor; a low pressure turbine; a low speed spool interconnecting the low pressure compressor and the low pressure turbine; and an electric generator located within the tail cone, the electric generator being operably connected to the low speed spool, wherein the electric generator includes a coolant cavity in thermal communication with one or more components of the electric generator; a structural support housing at least partially enclosing the electric generator, the structural support housing including a forward wall located on a forward end of the structural support housing, wherein the forward wall includes a first opening; a first coolant conveying tube extending through the first opening to fluidly connect to the coolant cavity; and a first electrical connector tube extending through the first opening within the first coolant conveying tube to electrically connect to the electric generator.

Abstract: An air bleed device for an aircraft engine, including a frame and a flap able to rotate about an axis in relation to the frame, the device further including a return system configured to bias the flap in a determined position about the axis and including a torsion spring, a first end of which is connected to the flap and a second end of which is connected to the frame. The second end is connected to the frame by adjusting the preload of the spring by screwing when the flap is in the determined position.

Abstract: A method is provided for maintaining a gas turbine engine installed on an aircraft, the gas turbine engine including an electric machine mounted at least partially inward of a core air flowpath of the gas turbine engine. The method includes removing a rotor mount connecting a rotor of the electric machine to a rotary component of the gas turbine engine; removing a stator mount connecting a stator of the electric machine to a stationary component of the gas turbine engine; and removing in situ the electric machine.

Abstract: An electrical power unit cooling system including at least one cooling fin configured to be thermally connected to a bus housing of the electrical power unit, the bus housing being configured to be attached to a surface of an aircraft fan housing such that the at least one cooling fin protrudes through the fan housing in order to be cooled by air flow within the fan housing.

Abstract: An engine assembly for an aircraft, including an internal combustion engine having a liquid coolant system in fluid communication with a heat exchanger, an exhaust duct in fluid communication with air passages of the heat exchanger, a fan in fluid communication with the exhaust duct for driving a cooling air flow through the air passages of the heat exchanger and into the exhaust duct, and an intermediate duct in fluid communication with an exhaust of the engine and having an outlet positioned within the exhaust duct downstream of the fan and upstream of the outlet of the exhaust duct. The outlet of the intermediate duct is spaced inwardly from a peripheral wall of the exhaust duct. The engine assembly may be configured as an auxiliary power unit. A method of discharging air and exhaust gases in an auxiliary power unit having an internal combustion engine is also discussed.

Abstract: A system is provided. The system includes a first environmental control sub-system, operating in a first mode, that receives a first medium at a first flow amount and a first pressure. The system also includes a second environmental control sub-system, operating in a second mode, that receives a second medium at a second flow amount and a second pressure. The first flow amount is greater than the second flow amount, and the second pressure is greater than the first pressure.

Abstract: A method and apparatus using a cooler to selectively cool P3 air directed into an impeller rear cavity of a gas turbine engine during engine high power levels, may include connection of the cooler to a low pressure compressor bleed-off valve apparatus that is modulated to be closed during engine high power levels to create a pressure differential over the bleed-off valve apparatus, to drive a stream of low pressure compressor air as a cooling work fluid under such pressure differential to selectively flow through the cooler.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a tie-rod and a threaded retainer. The tie-rod includes a tie-rod threaded portion and a tie-rod unthreaded portion. The threaded retainer includes a retainer threaded portion and a retainer unthreaded portion. The retainer threaded portion is mated with the tie-rod threaded portion. The retainer unthreaded portion is radially engaged with the tie-rod unthreaded portion.

Abstract: A thermal management system for a gas turbine engine and/or an aircraft is provided including a thermal transport bus having a heat exchange fluid flowing therethrough. The thermal management system also includes a plurality of heat source exchangers and at least one heat sink exchanger. The plurality of heat source exchangers and the at least one heat sink exchanger are in thermal communication with the heat exchange fluid in the thermal transport bus. The plurality of heat source exchangers are arranged along the thermal transport bus and configured to transfer heat from one or more accessory systems to the heat exchange fluid, and the at least one heat sink exchanger is located downstream of the plurality of heat source exchangers and configured to remove heat from the heat exchange fluid.

Abstract: A ducted fan gas turbine engine has a propulsive fan, a fan case surrounding the fan, a core engine, and a plurality of support structures which transmit loads from the core engine to the fan case. Each support structure has two support struts which extend from the core engine to a joint radially outwardly of the fan case. The support struts are spaced apart at the core engine but converge to meet at the joint. Each support structure further has two structural elements which extend from the joint to respective fixing positions on the fan case at opposite sides of the joint.