Abstract: Aspects of the disclosure are directed to a gas turbine engine, having a fan that rotates about a central longitudinal axis. The gas turbine engine may comprise a nacelle having an outer wall that radially circumscribes the fan. The gas turbine engine may further comprise a fan drive gear system that drives the fan, and a rotating shaft that is coupled to and drives the fan drive gear system. The gas turbine engine may further comprise an air vent line in fluid communication with a bearing compartment to remove air from the bearing compartment and provide the removed air to a deoiler that removes oil droplets from the removed air and provides deoiler filtered air. A vent output line receives the deoiler filtered air and discharges the deoiler filtered air along a radially interior surface of the nacelle outer wall.

Abstract: A gas turbine engine comprising: an inner core nacelle; an outer fan nacelle; a bypass duct between the inner core nacelle and the outer fan nacelle; at least one bifurcation that extends between the inner core nacelle and the outer fan nacelle; and a cooling system, wherein the cooling system comprises at least one pipe for conveying a fluid to be cooled, the at least one pipe forming part of a fluid system of the engine, wherein the at least one pipe passes through the at least one bifurcation, and wherein at least a portion of one or more of the pipes is arranged to bring the fluid to be cooled into a heat exchanging relationship with a fluid, e.g. air, flowing in the bypass duct.

Abstract: The gas turbine engine can have a bearing cavity, an ejector having an air/oil path fluidly connected to the bearing cavity, and a nozzle fluidly coupled with the air/oil path, the nozzle connected to a compressed air source.

Abstract: A gas turbine engine including a lubricant system defining a lubricant circuit through which a lubricant flows in fluid communication with a bearing assembly of the engine. The lubricant system selectively bypasses thermal communication of the lubricant and a heat sink based at least on a temperature of the lubricant within the lubricant circuit.

Abstract: An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine. The system includes an oil tank located within a core of the engine, an oil filling reservoir located within a bifurcation that extends from the core of the engine, an oil tank filling pipe that leads from the oil filling reservoir to a tank filling port on the oil tank, and an oil pump. The system is configured so that the oil pump pumps oil from the oil filling reservoir through the oil tank filling pipe and the tank filling port to fill the oil tank. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: A gas turbine engine includes a fan, a fan shaft coupled with the fan and arranged along an engine central axis, and a frame supporting the fan shaft. The frame defines a lateral frame stiffness (LFS). An epicyclic gear system is coupled to the fan shaft, and a non-rotatable flexible coupling and a rotatable flexible coupling support the epicyclic gear system. The non-rotatable flexible coupling and the rotatable flexible coupling each have a stiffness of a common stiffness type under a common type of motion with respect to the engine central axis. The stiffness is defined with respect to the LFS. The stiffness of the rotatable flexible coupling is greater than the stiffness of the non-rotatable flexible coupling.

Abstract: A method of sizing a heat exchanger for a geared architecture gas turbine engine includes sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75.

Abstract: Disclosed is a mid-turbine frame of a gas turbine engine, including: a vane extending along an axis; a first fitting; a second fitting spaced apart from the first fitting along the axis; a tube assembly located within the vane wherein a cavity is defined between the vane and the tube assembly, the tube assembly including: an outer tube coupled to the first fitting and extending therefrom to the second fitting; an inner tube located within the outer tube and extending from the first fitting to the second fitting; and an outer passage defined between the outer tube and the inner tube, and an inner passage defined within the inner tube, wherein the cavity, the outer passage and the inner passage are fluidly isolated from one another, and the cavity is configured to direct a cooling flow therethrough, and the outer tube is configured to direct a buffer flow therethrough.

Abstract: A gas turbine engine has a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction, and includes a lubricant pump to supply an air/oil mixture to an inlet of a deaerator. The deaerator includes a separator for separating oil and air, delivering separated air to an air outlet, and delivering separated oil back into an oil tank. The separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank. Air within the oil tank moves outwardly through an air exit into the deaerator. A method of designing a gas turbine engine is also disclosed.

Abstract: The present invention relates to a local thermal energy consumer assembly and a local thermal energy generator assembly to be connected to a thermal energy circuit comprising a hot and a cold conduit. The local thermal energy consumer assembly is selectively connected, via a pump or a valve to the hot conduit. The local thermal energy generator assembly is selectively connected, via a pump or a valve to the cold conduit. The use of either the valve or the pump is controlled by determining a local pressure difference between heat transfer liquid of the hot and the cold conduits.

Abstract: The present invention relates to a district thermal energy distribution system comprising a thermal energy circuit comprising a hot and a cold conduit for allowing flow of heat transfer liquid therethrough, a thermal energy consumer heat exchanger and a thermal energy generator heat exchanger. The thermal energy consumer heat exchanger is selectively connected to the hot conduit via a thermal energy consumer valve or a thermal energy consumer pump. The thermal energy generator heat exchanger is selectively connected to the cold conduit via a thermal energy generator valve or a thermal energy generator pump.

Abstract: A turbomachine comprises an inter-flow compartment extending radially between first and second intermediate walls, gas bypass ducts passing through the second intermediate wall and communicating with evacuation outlets opening up into a secondary gas flow; an oil reservoir arranged in the inter-flow compartment, and a surface air-oil heat exchanger communicating with the oil reservoir for fluid circulation. Circumferentially in relation to the inter-flow compartment, the surface air-oil heat exchanger extends at least partially between the evacuation outlets which pass through the first intermediate wall.

Abstract: A skid-mounted depressurizing system includes a main process module, a mechanical control system, a thermodynamic balance system and an intelligent control system. The main process module includes multiple main process pipelines, each of the multiple main process pipelines is provided with a pressure reducing valve set. During depressurization, process medium enters at least one of the multiple main process pipelines, the mechanical control system outputs a torque according to the control instruction of the intelligent control system to control each valve of the multiple main process pipelines to perform an action switch and an opening degree adjustment, the thermodynamic balance system is configured to provide the medium of different properties to the multiple main process pipelines according to the control instruction of the intelligent control system to control temperature, pressure, cleanliness degree and/or sealing degree of the main process module.

Abstract: A heat exhaust system for an on-aircraft electric generator, includes: a fuel pump which is driven by electric power generated by an electric generator installed in an aircraft, and which measures and supplies fuel to an engine of the aircraft; a first heat exchanger which releases heat generated by the electric generator into the fuel supplied to the engine; a fuel flow rate determiner which determines a flow rate of the fuel supplied to the engine by the fuel pump; an electric generator output determiner which determines an output from the electric generator; and a heat release controller which controls heat release at the first heat exchanger by using at least one of the flow rate of the fuel determined by the fuel flow rate determiner and the output determined by the electric generator output determiner as a control parameter.

Abstract: An outlet tube arrangement for an aircraft nacelle may comprise a vent tube defining a vent tube outlet, the vent tube outlet configured to be flush with an outer surface of the aircraft nacelle, and an inner tube disposed within the vent tube, the inner tube configured to increase momentum of fluid flowing through the inner tube to urge the fluid to separate from the inner tube without wetting the aircraft nacelle upon exiting from the vent tube, wherein the inner tube comprises an inner tube inlet comprising an inlet cross-sectional area and an inner tube outlet comprising an outlet cross-sectional area, the inlet cross-sectional area being less than a cross-sectional area of the vent tube at the inner tube inlet, the inlet cross-sectional area being greater than the outlet cross-sectional area, and the inner tube outlet being flush with the vent tube outlet.

Abstract: Oil tanks for gas turbine engines and associated assembly methods are disclosed. In one exemplary embodiment, an oil tank is configured to be installed to occupy a radially-inner space defined by an annular radial air inlet duct of a reverse flow gas turbine engine. The oil tank may comprise a first tank portion and a second tank portion assembled together to cooperatively define an interior volume of the tank. An optional intermediate spacer may be disposed between the first tank portion and the second tank portion in order to form an oil tank of a larger size.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a shaft and first and second bearing structures that support the shaft. Each of the first bearing structure and the second bearing structure includes a bearing compartment that contains a lubricant and a seal that contains the lubricant within the bearing compartments. A buffer system is configured to pressurize the seals to prevent the lubricant from escaping the bearing compartments. The buffer system includes a first circuit configured to supply a first buffer supply air to the first bearing structure, a second circuit configured to supply a second buffer supply air to the second bearing structure, and a controller configured to select between at least two bleed air supplies to communicate the first buffer supply air and the second buffer supply air.

Abstract: A gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications.

Abstract: The present disclosure is directed to a method of turbine section thermal management for a gas turbine engine. The engine includes a first turbine bearing defining an outer air bearing disposed radially adjacent to a low speed turbine rotor hub of a low speed turbine rotor and an inner bearing disposed radially adjacent to a high pressure (HP) shaft coupled to a high speed turbine rotor, wherein a first manifold is in fluid communication from a pressure plenum of a combustion section to the first turbine bearing, and wherein a second manifold is in fluid communication from the first turbine bearing to a pressure regulating valve and an outer diameter secondary flowpath of the turbine section, and wherein a third manifold is in fluid communication from the pressure plenum of the combustion section to the pressure regulating valve.

Abstract: A system for controlling a plurality of hydraulic effectors operably connected to an engine to control engine parameters. The system also includes a plurality of sensors operably connected to measure a state or parameter of each effector, a pump configured to supply fluid to the plurality of effectors, and a controller operably connected to the plurality of sensors, the plurality of effectors, and the pump. The controller executes a method for an adaptive model-based control for controlling each effector, The method includes receiving a request indicative of a desired state for each effector, receiving a weighting associated each request, obtaining information about a current state of each effector, and updating an adaptive model based control (MBC) based upon the information. The method also includes generating a control command for an effector based upon the adaptive MBC and commanding the effector based upon the control command.

Abstract: A bore basket assembly for a gas turbine engine. The assembly includes an outer cylindrical member. The assembly also includes an inner cylindrical member at least partially surrounded by the outer cylindrical member to define an annulus therebetween, one of the cylindrical members operatively coupled to an aft rotor stage, the other of the cylindrical members operatively coupled to a forward rotor stage, the cylindrical members moveable in an axially telescoping manner relative to each other.

Abstract: Oil slinger systems include a seal runner comprising an annular radial member having a radius (R) and an outer axially extending member having an axial length (L) such that a ratio (L/R) is between 0.8 and 1.4, the annular radial member disposed at a first angle with respect to the outer axially extending member, a heat shield in mechanical communication with the seal runner, and a volume bounded by an outer face of the annular radial member and an inner face of the heat shield. Methods of radial convective cooling include pumping a cooling liquid through the oil slinger system and convectively cooling the oil slinger.

Abstract: A gas turbine engine includes a buffer system that communicates a buffer supply air to a portion of the gas turbine engine. The buffer system includes a first bleed air supply having a first pressure, a second bleed air supply having a second pressure that is greater than the first pressure, and a valve that selects between the first bleed air supply and the second bleed air supply to communicate the buffer supply air to the portion of the gas turbine engine.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor comprising a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor comprising a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor.

Abstract: A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool. The LP spool and the HP spool are independently rotatable about an axis. The LP pressure spool has an LP compressor and an LP turbine. The HP spool has an HP turbine and an HP compressor. The LP compressor is axially positioned between the HP compressor and an accessory gear box (AGB). The AGB is drivingly connected to the HP spool through the center of the LP compressor.

Abstract: Disclosed is a lubricant supply system, for a bearing damper in a bearing housing, the bearing housing in an engine bearing compartment of a gas turbine engine, the bearing compartment rotatably supporting an engine component, including a first interface, a second interface, a bearing supply conduit fluidly coupled to the second interface and extending from the first interface to the second interface, the bearing supply conduit supplying lubricant to the bearing housing, a damper supply conduit located within the bearing supply conduit, extending between the interfaces, the damper supply conduit supplying lubricant to the bearing damper, and fluid in the bearing supply conduit is capable of insulting fluid in the damper supply conduit 108 from heat transferred through the interfaces.

Abstract: A component for a gas turbine engine is disclosed. The component includes a platform and a rail having an outer radial surface and an inner radial surface, with the inner radial surface connected to the platform. The rail includes a plurality of apertures spaced circumferentially between the outer radial surface and the inner radial surface, the apertures disposed to form a truss-like structure.

Abstract: A guide vane of a gas turbine engine, in particular of an aircraft engine, which has a pressure-side wall, a suction-side wall, a guide vane root, a guide vane tip, a guide vane leading edge area that is impinged by a cooling air flow of a cooling system, a guide vane trailing edge area that is facing away from the guide vane leading edge area, and at least one channel for conducting a fluid to be cooled arranged in an internal space of the guide vane. At that, during operation of the gas turbine engine, a first part of the cooling air flow flows around a pressure-side wall, and a second part of the cooling air flow flows around the suction-side wall, and a third part of the cooling air flow flows through the internal space including the channel. What is further suggested is a gas turbine engine with at least one such static guide vane.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor including a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor including a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; and a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit extended within the high speed turbine rotor.

Abstract: A turbomachine includes a compressor configured to compress air received at an intake portion to form a compressed airflow that exits into an outlet portion. A combustor is operably connected with the compressor, and receives the compressed airflow. A turbine is operably connected with the combustor, and receives the combustion gas flow. The turbine has a plurality of wheels and a plurality of buckets. The turbine receives compressor bleed off air to cool the wheels and buckets. A cooling system is operatively connected to the turbine. The cooling system includes a plurality of heat pipes located axially upstream of at least one of the wheels. The heat pipes are operatively connected to a bearing cooler system. The heat pipes and the bearing cooler system are configured to transfer heat from the compressor bleed off air to one or more heat exchangers.

Abstract: A gas turbine engine includes a spool which rotates in use, a static structure and an air bearing, the air bearing being provided at an interface between the spool and the static structure.

Abstract: A gas turbine engine comprises a fan drive turbine, a fan rotor, and a gear reduction driven by the fan drive turbine to, in turn, drive the gear architecture. A main oil supply system supplies oil to components within the gear reduction, and an auxiliary oil supply system. The auxiliary oil system operates to ensure that the gear reduction will be adequately supplied with lubricant for at least 30 seconds at power should the main oil supply system fail.

Abstract: An annular duct including a modular annular heat exchanger for a gas turbine engine is provided, where the modular annular heat exchanger includes a plurality of radial modules in circumferentially adjacent arrangement. Each radial module includes a cooled fluid inlet plenum segment, a plurality of blades, and a cooled fluid outlet plenum segment. The plurality of blades is configured in circumferentially adjacent arrangement and defines an angular space that is conformal between each circumferentially adjacent blade. The cooled fluid inlet plenum segment, the plurality of blades, and the cooled fluid outlet plenum segment are in serial axial flow arrangement and define an internal cooled fluid flowpath and an external cooling fluid flowpath parallel to the internal cooled fluid flowpath. Each radial module further includes an inner annular ring segment and an outer annular ring segment. The inner annular ring segment and the outer annular ring segment define a plurality of blade retainers.

Abstract: A method is disclosed for cooling a gas turbine having a turbine, wherein a rotor, which rotates about a machine axis, carries a plurality of rotating blades, which are mounted by blade roots and extend with their airfoils into a hot gas path of the gas turbine. The rotor is concentrically surrounded by a turbine vane carrier carrying a plurality of stationary vanes, whereby the rotating blades and the stationary vanes are arranged in alternating rows in axial direction. An extended lifetime with external cooling is achieved by providing first and second cooling systems for the turbine.

Abstract: An oil circuit of an aircraft engine, includes a tank, a lubricant pump and a recirculation pump. The lubricant pump supplies oil via a supply line from the tank to a load of the engine and to an accessory gearbox. Oil is resupplied via a recirculation conduit to the tank by the recirculation pump. The recirculation conduit is connected to the accessory gearbox via a further recirculation conduit. A valve is provided that can be switched between an open switching state and a closed state. In the open switching state of the valve, oil is supplied to the tank via the recirculation conduit. In the closed state of the valve device, an oil feed to the tank via the recirculation conduit is blocked, and oil from the recirculation conduit is supplied to the accessory gearbox via the further recirculation conduit.

Abstract: An aerospace plane having an elongate body supporting a pair of wings each having at least two angled, trailing edge portions. The pair of wings are adapted to extend away from the elongated body in opposing directions. A landing gear assembly is operatively associated with the elongated body to be moveable from a retracted position where the landing gear assembly is substantially locatable within the elongated body and an extended position where the landing gear assembly extends at least partially away from the elongated body. At least one engine adapted to generate thrust. At least one stabilizer adapted to assist with movement of the aerospace plane during flight. The at least one engine is located at least partially within an intake housing adapted to direct air into said at least one engine.

Abstract: A seal runner may have an inner annular member, an outer annular member, and a radial wall member joined to the inner annular member according to an angle and extending from the inner annular member to the outer annular member. The seal runner may be positioned axially along a shaft and a bearing may be positioned axially along the shaft such that it is radially between the inner annular member and the outer annular member. In this regard, the seal runner may resemble a “C-shape.” A seal may interface with a seal land radially outboard of the outer annular member. Thus, the seal may be positioned axially near to the bearing so that the axial length of a shaft and accompanying bearing compartment hardware may be minimized.

Abstract: A device for protecting against oil leaks towards the rotors of a turbomachine turbine includes an upstream cavity in which an oil/air mixture can circulate, the upstream cavity being delimited downstream by a journal of the turbine rotor and on the outside by a sealing flange of the journal extending axially in an upstream direction, a downstream cavity opening towards the discs of the turbine and communicating with the upstream cavity via a plurality of ventilation holes provided in the journal, and an upstream annular deflector fixed to the sealing flange of the journal and extending radially inwards towards the ventilation holes The upstream deflector has oil passage ports provided at the end of same secured to the sealing flange of the journal.

Abstract: A gas turbine engine includes a bearing compartment with a bottom disposed opposite a top. An outlet is formed in the bottom of the bearing compartment. A seal is positioned between the bottom of the bearing compartment and the top of the bearing compartment. A scavenge line is connected to the outlet. The scavenge line forms a fluid trap that extends in a first direction from the outlet and the bottom of the bearing compartment and then extends in a second direction to a position disposed vertically between the bottom of the bearing compartment and the seal.

Abstract: A rotor, in particular a drum of a low-pressure compressor of a turbojet aero engine, is disclosed. The rotor includes an outer annular wall delimiting a primary annular flow of the turbine engine, sealing devices with two rubbing strips or annular ribs formed on the wall. The rubbing strips cooperate by abrasion with inner shrouds. In addition, the annular wall includes rows of intake orifices for leakages which are arranged between each pair of rubbing strips in order to aspirate the recirculation leakages there. A plenum for leakages is formed inside the rotor by means of a composite partitioning, then evacuates the parasitic flow downstream of the turbine engine via the central shaft.

Abstract: A system includes a gas delivery disk coupled to a second shaft, wherein the second shaft and the gas delivery disk are disposed about an axis. The gas delivery disk includes an inner axial opening configured to facilitate a first axial flow through a first passage within the second shaft, a duct configured to supply a bearing flow in a radial direction toward the axis, and a bearing face disposed within the first passage and radially interior to the inner axial opening. The bearing face is configured to receive the bearing flow, and the bearing face is configured to form a gas bearing between the bearing face and a first shaft disposed about the axis.

Abstract: A disclosed lubrication system for a gas turbine engine includes a primary passage defining a flow path for lubricant to a gas turbine engine and a bypass passage defining a flow path for lubricant around the gas turbine engine. The lubrication system further includes a primary lubrication pump including a reduced total flow capacity of lubricant with a reduced bypass lubricant flow relative to the total overall flow capacity.

Abstract: A single, unitary fuel/oil manifold for a gas turbine engine is provided which comprises one or more interfaces for mounting various fuel system and lubricating system components directly to the fuel/oil manifold. The fuel/oil manifold also defines fluid passages for transferring fuel or lubricant from one component to another component. Packing numerous fuel system and lubricating system components within the fuel/oil manifold reduces cost and weight and simplifies maintenance.

Abstract: A gas turbine system (1A) includes a gas turbine unit (2) and a cooling fluid generator (5). The gas turbine unit (2) includes a first compressor (21) and a first expansion turbine (23) coupled to each other by a first shaft (22), a combustor (26), and a fuel tank (30). A fuel held in the fuel tank (30) circulates through a fuel circulation passage (31). A working fluid that has a pressure increased by the first compressor (1) is extracted from the gas turbine unit (2). The cooling fluid generator (5) includes a cooler (55) for cooling, with the fuel flowing through the fuel circulation passage (31), the working fluid that has been extracted from the gas turbine unit (2), and a second expansion turbine (53) for expanding the working fluid that has flowed out of the cooler (55).

Abstract: In one aspect, the present disclosure is directed to a bearing compartment sealing system including a bearing for supporting a shaft. A housing encloses the bearing and defines a compartment for holding lubricant therein. At least two seals are located between the shaft and the housing. The at least two seals, the housing, and the shaft collectively enclose the compartment. Only one of the at least two seals is a hydrodynamic seal.

Abstract: A seal support and duct assembly of a gas turbine engine includes a seal support housing disposed about a central axis, a duct housing attached to the seal support housing, the duct housing and the seal support housing defining the seal support and duct assembly, an outer annular cavity, an inlet that supplies the outer annular cavity with cooling buffer air, an inner annular cavity disposed radially inward of the outer annular cavity, a first plurality of outlets that provide the cooling buffer air from the outer annular cavity to the inner annular cavity, and a second plurality of outlets that provide the cooling buffer air from the inner annular cavity to an area surrounding a bearing compartment.

Abstract: An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.

Abstract: A lubrication system for an engine including a scavenge pump configured to provide lubricant to the engine during a normal condition, an auxiliary pump configured to provide lubricant to the engine during an abnormal condition, a pressure sensor disposed at an outlet of the auxiliary pump for detecting a pressure created by the auxiliary pump, and a restriction device disposed in a conduit fluidly connecting the scavenge pump to the outlet of the auxiliary pump and configured to increase a first pressure produced by the auxiliary pump at the pressure sensor, wherein the first pressure is greater than a second pressure created by the scavenge pump at the pressure sensor.

Abstract: A field of turbine engines, and more particularly to an engine including a compressor; a combustion chamber; a first turbine connected to the compressor by a first rotary shaft; an actuator device for actuating the first rotary shaft in order to keep the first turbine and the compressor in rotation while the combustion chamber is extinguished; and a lubrication circuit for lubricating the engine. The circuit passes through at least one heat source suitable for heating the lubricant in the lubrication circuit while the first turbine and the compressor are rotating with the combustion chamber extinguished.