Abstract: The invention relates to a system (1) for lubricating an aeronautical engine (5) and a reduction gearbox (4) associated with the engine (5), the system (1) comprising an oil reservoir (2) feeding at least one first supply pump (3) supplying a first circuit (6) of the gearbox (4) opening into at least one chamber (4a) of the gearbox (4) and, in parallel, a second circuit (7) of the engine (5) opening into chambers (5a) of the engine (5). The second circuit (7) comprises a jet pump (9) of variable cross section supplied at least by the first supply pump (3), bypassing the first circuit (6), a second driven supply pump (10) being integrated into the second circuit (7) downstream of the jet pump (9), a portion of a flow (Qp) in the first circuit (6) being drawn off by the jet pump (9) to supply the second circuit (7).

Abstract: Provided is a gas turbine engine including a main passage through which compressed air extracted from a compressor is guided to a bearing; and an oil supply device that supplies oil to the main passage. The oil supply device includes a pressure reducer in the main passage; an oil tank a divider that is movable and divides an internal space of the container into an oil chamber storing the oil and an air chamber; an extruder including a cylinder and a piston that divides an internal space of the cylinder into a first chamber and a second chamber; a switching mechanism that switches between a first state where the first chamber communicates with a low-pressure space whose pressure is lower than pressure in the air chamber and a second state where the first chamber communicates with a high-pressure space whose pressure is higher than pressure in the downstream passage.

Abstract: A lubricant system for supplying lubrication to a component in a turbine engine includes a lubricant reservoir, a supply line fluidly coupling the lubricant reservoir to the component in the turbine engine, a scavenge line fluidly coupling the component to the lubricant reservoir, and a bypass line fluidly coupling the supply line to the scavenge line and bypassing the component.

Abstract: A gas turbine engine has a low pressure turbine driving a fan rotor through a gear reduction gearbox. The low pressure turbine has an input shaft driving a sun gear. The sun gear drives intermediate gears. There is an output shaft driven by one of a carrier and a ring gear in the gear reduction to in turn drive a fan drive shaft to drive the fan. A bearing compartment housing encloses the gear reduction gearbox. A fan shaft bearing supports the fan drive shaft. An oil supply system delivers oil to at least one of the sun gear, the intermediate gears and the ring gear. A scavenge system returns oil from the bearing compartment housing to a sump. An ejector pump communicates with a source of pressurized fluid to assist in driving oil from within the bearing compartment housing into a scavenge tube.

Abstract: A jet for lubricating a piece of a turbomachine for an aircraft such as an airplane, including a circulation duct for flowing a pressurised fluid, includes a first duct chamber, a second duct chamber, a first nozzle for passing between the first duct chamber and the second duct chamber, the first nozzle having a fixed minimum passage cross-section, and a second nozzle for passing from the second chamber to the outlet port formed by the second nozzle, the second nozzle including a fixed minimum passage cross-section. The ratio of the cross-section of the first nozzle to the cross-section of the second nozzle is between 0.16 and 3.61.

Abstract: A gas turbine engine includes a fan assembly, a compressor assembly, a combustion chamber, a turbine assembly, a bypass duct conveying rearward a bypass airstream driven by the fan assembly when the gas turbine engine is in use, a fairing extending across at least a portion of the bypass duct downstream of the fan assembly, and a heat exchanger having an inlet fluidly connected to the compressor assembly and an outlet fluidly connected to a pneumatic actuator of the gas turbine engine. The fairing has a leading edge and a trailing edge. The heat exchanger is disposed adjacent the trailing edge of the fairing.

Abstract: An oil pumping system for use with a gas turbine engine includes an electric machine, an oil pump assembly, and a variable frequency drive controller. The variable frequency drive is connected with the electric machine and the oil pump assembly. The variable frequency drive controller is programmed to control a torque and speed of the pump motor independent of the gas turbine engine speed so that a flow of oil moved by the oil pump assembly is controlled to cool or lubricate components of the gas turbine engine independent of the gas turbine engine speed.

Abstract: A lubrication system for an aircraft engine includes an engine lubricant tank including at least a supply port hydraulically connectable to the aircraft engine, a lubricant makeup port, and an overfill port, an auxiliary lubricant tank, a lubricant makeup conduit hydraulically connecting the auxiliary lubricant tank to the lubricant makeup port. The lubricant makeup conduit includes a pump operable to move lubricant from the auxiliary lubricant tank to the lubricant makeup port, and an overfill conduit hydraulically connecting the overfill port to the auxiliary lubricant tank. A method of operating a lubrication system of an aircraft engine of an aircraft is also disclosed.

Abstract: The invention concerns an epicycloidal gear train (10) comprising a central pinion (26), an outer crown (28) and satellite pinions (32) in engagement with the central pinion (26) and the outer crown (28) and each mounted freely rotatable on a satellite carrier (36), the gear train (10) comprising an annular cup (56) being integral with the satellite carrier (36) and open radially inward. According to the invention, the cup (56) is divided into a circumferential succession of first basins (60) of the first oil circuit and second basins (62) of the second oil circuit, the first basins (60) being fluidly separated from the second cells (62) and characterised in that the cup (56) comprises two annular walls (56a, 56b) axially facing, the annular wall (56b) of which is furthest from a median transverse plane (74) of the gear train (10) has openings (76) opening into the second basins (62).

Abstract: A gas turbine engine with a lubrication system includes a ball bearing assembly and a rotor circumscribing a rotational axis and journaled within the ball bearing assembly. The gas turbine engine also includes a lubrication system located radially outward from a rotational axis and radially outward and adjacent to the ball bearing assembly, which includes a lubrication channel having an inlet and an outlet and a dispersion cone adjacent to the outlet of the lubrication channel.

Abstract: Dynamic sealing device for an aircraft turbomachine, comprising two rotors (3, 8) which are configured to rotate about the one same axis (A) and different speeds, the said rotors extending at least in part around one another and dynamic sealing means (14) being mounted between the said rotors, characterized in that: —the said sealing means comprise annular rings (38) extending about the said axis, these rings being mounted with the ability to slide on a first (3) of the said rotors, in a radial direction with respect to the said axis, and being able to bear radially via their external periphery against a second (8) of the said rotors which at least partially surrounds the said first rotor, and in that it further comprises: —oil discharge means comprising an annular cavity (44) formed in the said second rotor, the said annular cavity opening radially towards the inside in the vicinity of the said sealing means and being in fluidic communication with oil passage ports (45) formed in the bottom of the cavity and

Abstract: The invention relates to an arrangement, in a pod of an aircraft propulsion assembly, for drawing in air and trapping foreign bodies. Said arrangement includes a main air inlet duct (11) separating into, on one hand, a channel (13) for leading air to a compressor and, on the other hand, a bypass channel (12) capable of trapping foreign bodies (5) that enter said main duct (11). Said arrangement comprises a heat exchanger (6) that extends along a section of the bypass channel (12). Said heat exchanger (6) carries out surface heat exchange along said section and is coupled with an external oil system in order to cool the oil thereof by heat exchange with the air (4) flowing in the bypass channel (12). Said bypass channel (12) has an air outlet (12a) acting as a means for discharging the foreign bodies (5).

Abstract: System for an aircraft, the system comprising: a turbine engine (2) with a hydraulic lubrication circuit (30); and a fuel cell (28) with a hydraulic circuit (40) for setting and maintaining the operating temperature of the fuel cell (28); wherein the hydraulic lubrication circuit (30) and the hydraulic circuit (40) form a single and common oil circuit and the oil circuit comprises a pump (50) with a heating element integrated in the pump for heating the oil. The pump may be an electric pump with an electric motor, wherein the electric motor comprises a coil fed with DC current forming a heating element for heating the oil. The pump may comprise a body and the heating element may be an electric resistor embedded into the body of the pump and in direct contact with the oil.

Abstract: The present disclosure concerns an oil distribution assembly for a system having rotating components requiring a supply of oil, such as bearing components of a gas turbine engine. Example embodiments include an oil distribution assembly for a gas turbine engine, comprising: an oil distributor mounted for rotation about a rotation axis of the assembly and comprising an internal volume having a plurality of channels extending along an inner radial surface of the oil distributor; and an oil injector arranged to direct a supply of oil to the plurality of channels in a direction having a component in a first direction along the rotation axis; wherein each of the plurality of channels is angled relative to the rotation axis such that a radial distance between each channel and the rotation axis increases along the rotation axis.

Abstract: A ventilation system includes a cavity, a fluid motive force device, and a motive fluid supply system. The cavity includes different ventilation level requirements for a plurality of modes of operation. The fluid motive force device includes a suction port, an outlet port, and a motive fluid inlet port. The suction port is coupled in flow communication with the cavity to be vented. A flow supply to the motive fluid inlet port determines a ventilation flow through the suction port. The motive fluid supply system is coupled in flow communication with the motive fluid inlet port. An operation of the motive fluid supply system determines a flow of motive fluid from the motive fluid supply system to the motive fluid inlet port. The flow of motive fluid to the motive fluid inlet port generates a ventilation flow through the suction port approximately matching a current ventilation demand of said cavity.

Abstract: An oil tank (100) for a gas turbine engine is provided. The oil tank (100) comprises an oil inlet (102), an oil outlet (104), and a body including a coiled portion (106) interposed between the oil inlet (102) and the oil outlet (104). Oil is received by the coiled portion (106). The coiled portion (106) acts to at least partially de-aerate oil received from the oil inlet (104).

Abstract: The heat exchanger system includes a first heat exchanger assembly including a plurality of airfoil members circumferentially spaced in a flow stream of an annular duct. Each airfoil member including a radially inner end and a radially outer end and a first internal flowpath configured to channel a flow of cooled fluid therethrough. The heat exchanger assembly includes a second heat exchanger assembly including a plurality of fin members extending proximate the flow stream and a second internal flowpath configured to channel a flow of cooled fluid therethrough. The heat exchanger assembly includes a header system including a conduit path configured to couple the first heat exchanger assembly and the second heat exchanger assembly in flow communication. The header system includes an inlet connection configured to receive a flow of hot fluid from thermal loads and an outlet connection configured to direct cooled fluid to thermal loads.

Abstract: A gas turbine engine includes a fan, a compressor section, and a turbine section configured to drive the compressor section and the fan. A buffer system is configured to communicate 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 an ejector that selectively augments the first bleed air supply to prepare the buffer supply air for communication to the portion of the gas turbine engine. A method and a buffer system are also disclosed.

Abstract: A system for managing thermal transfer in at least one of an aircraft or a gas turbine engine includes a first engine system utilizing an oil for heat transfer. The oil of the first system has a temperature limit of at least about 500° F. The system additionally includes a fuel system having a deoxygenation unit for deoxygenating fuel in the fuel system, as well as a fuel-oil heat exchanger located downstream of the deoxygenation unit. The fuel-oil heat exchanger is in thermal communication with the oil in the first engine system and the fuel in the fuel system for transferring heat from the oil in the first engine system to the fuel in the fuel system.

Abstract: A turbine engine includes a fan. The fan includes a first shaft driven in rotation via a reduction gear, the first shaft being guided in rotation with respect to a support of a fixed structure via an upstream bearing and a downstream bearing, the downstream bearing guiding in rotation the first shaft with respect to a downstream branch of the support, the first shaft and the support together defining an enclosure. The reduction gear includes a ring gear fixed to the structure via a shell. The turbine engine also includes a lubricator composed of at least one pipe for conveying a liquid lubricant, wherein the pipe is located in the enclosure, the pipe passing through an orifice made in the shell and an opening made in the downstream branch.

Abstract: A turbine engine system includes a gutter and a gear train with an axial centerline. The gutter is disposed radially outside of the axial centerline. The gutter includes an inner surface and a channel that receives fluid directed out of the gear train. The inner surface at least partially defines a bore in which the gear train is arranged. The channel extends radially into the gutter from the inner surface, and circumferentially to a channel outlet. The bore has a cross-sectional bore area, and the channel has a cross-sectional channel area that is substantially equal to or less than about two percent of the bore area.

Abstract: A strut includes a strut passage extending through the strut along a radial length of the strut. A tube is disposed within the strut passage and is spaced from an interior surface of the strut passage. A grommet is disposed about the tube and is in communication with the interior surface of the strut passage. The grommet defines a compressible zone including a hollow space extending radially through the grommet. The compressible zone is disposed between the tube and the interior surface of the strut passage.

Abstract: A gas turbine engine includes a core engine defined about an axis, a gear system driven by the core engine, and a pylon variable area flow system. A fan is driven by the gear system. The variable area flow system operates to effect the bypass flow.

Abstract: A gas turbine engine includes a fan drive turbine driving a fan rotor through a main gear reduction. A primary lubricant system supplies lubricant to the main gear reduction. An auxiliary oil pump supplies oil to the main gear reduction. An auxiliary pump epicyclic gear train drives the auxiliary pump when the fan rotor is rotating in either direction. The main gear reduction is separate from the auxiliary pump epicyclic gear train.

Abstract: A turbofan including a low-pressure shaft; a high-pressure shaft; a fan shaft; a reducing mechanism that couples the low-pressure shaft to the fan shaft; and a maximum of six chambers that accommodate the bearings of the low-pressure shaft, the bearings of the high-pressure shaft and the bearings of the fan as well as the reducing mechanism.

Abstract: A lubrication system for a drive train of a wind turbine including a gearbox is provided. The gearbox has a first stage with a first internal oil level and a second stage with a second internal oil level. The second internal oil level is at a geodetic higher level than the first internal oil level. The gearbox further includes a supply passage for supplying oil from the first stage to the second stage of the gearbox.

Abstract: An aircraft twin-spool turbomachine comprises an inter flow compartment (50) housing a fuel tank (54), the exterior of the inter flow compartment being delimited by a cover (52) delimiting the interior of a bypass flow (14b) of the turbomachine, the cover comprising a closable access (64) to a remote lubricant tank filling orifice (74). The turbomachine also comprises: an overflow evacuation orifice (82) located on the tank lower than the lubricant inlet orifice (68); a lubricant overflow evacuation conduit (84) routed downwards from the overflow evacuation orifice (82); a controlled evacuation valve (88), associated with the overflow evacuation orifice (82); and a control device (90, 94) for the evacuation valve (88), the control device being accessible through the access (64).

Abstract: Device of the reduction gear or differential type for a turbine engine of an aircraft. A central sun gear and a row of planet gears, each of the planet gears are rotationally guided by at least one bearing extending around a tubular support. The tubular support includes an inner cavity for the reception of oil and through-orifices for the passage of oil of the tubular support to the at least one bearing. A member for circulating the oil is mounted in the cavity and includes at least one oil duct intended to be connected to an oil supply. The oil is intended to flow from the at least one duct into the inner cavity to form an oil film in contact with said tubular support, and to flow through the orifices.

Abstract: Apparatus for a gas turbine engine, the apparatus comprising: a core engine casing having a longitudinal axis and including: an inner wall defining at least part of a core airflow path through the gas turbine engine; an outer wall defining an external surface of the core engine casing, a first cavity being defined between the inner wall and the outer wall of the core engine casing; a plurality of guide vanes extending radially from the outer wall of the core engine casing; a torque box defined within the first cavity of the core engine casing and at least partially overlapping axially with the plurality of guide vanes, the torque box defining a second cavity; and an accessory gear box positioned within the second cavity of the torque box.

Abstract: A lubrication circuit comprising an oil tank (4) integrated into a first lubricated chamber (5) comprises an evacuation conduit (20) on a recovery conduit (10) from lubricated chambers of the circuit leading to the reservoir (4), opening up into an outlet (21) when an excessive pressure in the tank, that is a symptom of progressive filling, is reached. Application to aircraft engine lubrication circuits, particularly to guard against fuel leaks in the lubricant as a result of a defective heat exchanger.

Abstract: A feed system for feeding lubricating oil to members of a turbine engine including a reduction gearbox, the feed system including a nonpositive-displacement pump device for connecting upstream to an oil tank and driven in rotation at a speed that is not correlated with an operating speed of the turbine engine; a separator node connected to the outlet of the nonpositive-displacement pump device; a first delivery branch for lubricating the RGB connected to the nonpositive-displacement pump device via the separator node; a second delivery branch for lubricating other members connected to the nonpositive-displacement pump device via the separator node, the second delivery branch including a positive-displacement pump; and at least one fluid metering device having a metering slot fed by the nonpositive-displacement pump device via the separator node for the purpose of feeding the RGB.

Abstract: A fan drive gear system for a turbofan engine is disclosed and includes a gear assembly and a front center body. The front center body is an annular case that supports the gear assembly. The front center body includes a passage portion that defines a portion of a core flow path, a forward flange configured for attachment to a first case structure forward of the front center body, and gutter portion disposed on a radially inner side of the front center body for collecting lubricant exhausted from the geared assembly.

Abstract: An aircraft turbine engine includes a fan and a reduction gear including a plurality of rotary elements and driving the fan. The turbine engine further includes a gearbox and a mechanical power takeoff box driving the gearbox. The turbine engine includes a first toothed wheel and a second toothed wheel forming part of the mechanical power takeoff box and meshing with the first toothed wheel.

Abstract: The invention relates to a turbine engine, comprising: a device for bleeding a compressor, particularly a high-pressure compressor, comprising at least one valve for bleeding said compressor, an outlet of said valve being connected to a bleed system designed to discharge a bleed airflow; and at least one device for cooling at least one unit, comprising a heat exchanger, an air outlet of which is connected to an ejector of the jet pump type, which comprises a feed-through duct for a secondary airflow coming from said air outlet, as well as a nozzle for spraying a primary airflow inside said feed-through duct, wherein the nozzle is connected to said bleed system.

Abstract: A ventilation fan includes a shaft, a rotor, a motor housing, a bearing housing, and an air bearing. The shaft has a shaft body that extends between a first shaft end and a second shaft end. The shaft body defines a first port and a bore. The rotor is disposed about the shaft. The motor housing is disposed about the shaft and is axially spaced apart from the rotor. The bearing housing is disposed about the shaft. The air bearing is disposed proximate the second shaft end and is disposed between the bearing arm and the second shaft end.

Abstract: A gas turbine engine valve assembly having an oil shut-off valve and a compensating valve fluidly connected thereto. The compensating valve, in a first position, restricts access to a pressure-reducing oil flow path leading to an oil reservoir, and in a second position, provides access to the pressure-reducing oil flow path. The pressure-reducing oil flow path is defined by fluid flow between housing apertures, a laminar hole and an orifice. The valve assembly is operable between a first configuration, wherein the oil shut-off valve is in a fluid transfer position and a piston of the compensating valve is in a closed position to direct oil toward the engine component, and a second configuration, wherein the oil shut-off valve is in the shut-off position and the piston is in a open position to direct oil toward the oil reservoir via the pressure-reducing oil flow path.

Abstract: A lubrication device of a mechanical system, the lubrication device having a main tank, at least one recovery tank, a lubrication circuit, a flow generator supplying the lubrication circuit with a lubrication liquid located in the main tank and a delivery system for each recovery tank. Each delivery system has an ejector arranged in a recovery tank and first and second conduits linking the ejector firstly to the lubrication circuit and secondly to the main tank in order to deliver the lubrication liquid located in the recovery tank to the main tank using the ejector.

Abstract: A bypass turbojet engine including a low pressure shaft supported by at least two low pressure bearings, a high pressure shaft supported by at least two high pressure bearings, a fan shaft supported by at least two fan bearings, a reduction system coupling the low pressure shaft, with the fan shaft, enclosures housing the low pressure bearings, the high pressure bearings, the fan bearings and the reduction system, and a lubrication assembly including a closed oil circuit configured to supply oil to the enclosures in order to cool the bearings and the reduction system and at most five recovery pumps configured to recover oil from the enclosures.

Abstract: The present invention relates to a bearing chamber housing for bearing a shaft of a turbomachine, having an additively built-up housing wall and a cover of the oil chamber, wherein, in each case referred to an axis of rotation of the shaft, the housing wall bounds radially outwardly an oil chamber of the bearing chamber housing and the cover of the oil chamber disposed on an axial end of the housing wall axially bounds a region of the oil chamber; the region bounds radially inwardly an inner shell of the bearing chamber housing.

Abstract: The invention relates to a lubrication system for turbomachine. The hydromechanical cutoff device is configured to close when a rotation speed of a turbomachine shaft reduces and becomes lower than a first threshold, and wherein the hydromechanical cutoff device is configured to open when the rotation speed of the shaft increases and becomes higher than a second threshold higher than the first threshold.

Abstract: A gas turbine engine includes a gearbox including a sun gear, annulus gear, plurality of planet gears and a planet gear carrier. Each planet gear is rotatably mounted in the planet gear carrier by a planet gear bearing. A lubrication system is arranged to supply lubricant to the planet gear bearings and at least one of the sun gear, the planet gears and the annulus gears. The planet gear carrier includes an annular extension. The lubrication system includes an annular member arranged around annular extension to define an annular chamber which is sealed at its ends to the planet gear carrier and which contains a reserve of lubricant. The annular member has an inlet to supply lubricant into the annular chamber and the planet carrier has a first passage to supply lubricant to the sun, planet or annulus gear and a second passage to supply lubricant to planet gear bearings.

Abstract: An oil supply system includes an oil supply assembly in a turbine engine. The assembly includes scavenge cavity and an oil supply tube located in and protected by a strut in the turbine engine to supply oil to a bearing chamber.

Abstract: A fluid film damper may comprise a sleeve, an annular support housing surrounding the sleeve, an annular volume between the annular support housing and the sleeve, a first supply conduit in fluid communication with a first inlet to the annular volume, a second supply conduit in fluid communication with a second inlet to the annular volume, wherein the first inlet is disposed opposite the annular volume from the second inlet, and a common oil supply conduit in fluid communication with the annular volume via a check valve, wherein the common oil supply conduit supplies the fluid to the annular volume via the first supply conduit and the second supply conduit.

Abstract: A turbomachine includes: a rotational shaft; a pair of thrust collars disposed around the rotational shaft; a thrust bearing disposed around the rotational shaft at an axial directional position between the pair of thrust collars; and an oil-drain channel disposed around the rotational shaft, for draining lubricant oil after lubricating a sliding portion between the thrust bearing and the thrust collars. The oil-drain channel includes: an oil-drain port portion for discharging the lubricant oil inside the oil-drain channel downward; and an oil-guide channel portion disposed above the oil-drain port portion and configured to guide the lubricant oil from the sliding portion in a circumferential direction of the rotational shaft to the oil-drain port portion.

Abstract: A system is provided that includes a support structure and a nozzle. The support structure includes a cavity surface and an aperture. The cavity surface at least partially forms a boundary of a cavity. The aperture extends partially into the support structure from the cavity surface. The nozzle includes a mount, a neck and a head. The mount is seated within the aperture. The neck is connected to the mount and extends axially along a centerline away from the surface to the head. The head is configured to inject fluid out of the nozzle and into the cavity.

Abstract: A sump assembly for use in a gas turbine engine includes a housing and a shaft assembly. The housing is arranged about a central axis of the sump assembly to define a cavity configured to house oil and high-pressure gasses. The shaft assembly is mounted to rotate about the central axis and to direct the high-pressure gasses into the housing.

Abstract: A gas turbine engine includes a fan section and a core turbine engine operable with the fan section. The gas turbine engine also includes a sump pressurization assembly including a valve, a first duct in airflow communication with the valve and a first pressurized air source, and a second duct in airflow communication with the valve and a second pressurized air source. Further, the sump pressurization assembly includes a supply duct selectively in airflow communication with the first duct and the second duct through the valve. The supply duct is located at least partially inward of a core air flowpath of the core turbine engine for pressurizing a sump of the gas turbine engine.

Abstract: An engine turbocharger (10) includes a radially oriented oil deflector (50) that incorporates an oil guiding feature (70) to improve control of a continuous turbulent mass of oil being impelled against the deflector by an oil flinger (35). The oil deflector (50) has a downwardly extending tongue (60) to assist in directing oil flow from a thrust bearing chamber. The feature (70), formed on the tongue (60), enhances otherwise gravitationally directed oil flow into an engine oil sump. The feature is vertically oriented, and redirects oil thrown axially and rotationally against the deflector, providing improved piston ring sealing through more efficient evacuation of oil from the thrust bearing chamber. The feature may be formed by stamping, as either an integral disruptive surface rib, or as an otherwise radially oriented flange protruding axially from the surface of the tongue. Such flange may be integral with, or otherwise attached to, the oil deflector tongue, as for example via welding.

Abstract: An aero gas turbine engine with a fan stage and a bearing rotatably supporting the fan stage, where the bearing is cooled by passing cooling air through the bearing, and the spent bearing cooling air is re-injected into the air flow path at a location where the effect is minimalized. The location is just downstream from a leading edge of a stator vane and along a pressure side surface of the vane. A fine mist of oil can be discharged into the bearing cooling air to also lubricate the bearing, where both the oil and the spent cooling air is reintroduced into the air flow path.

Abstract: A gas turbine provided with an air bleeder tube (1) that, during startup, bleeds a portion of the compressed air of a compressor from the compressor and discharged the bled air into a cylindrical exhaust duct (20), wherein the air bleeder tube (1) is disposed at a portion that does not obstruct the flow of the main flow of combustion gas.