Abstract: A system for controlling discharge doors of a turbomachine includes: a ring and a device connecting the ring to each discharge door. The ring is made up of bars connected end-to-end by rotary links ensuring that a bar pivoting on itself causes the other bars to pivot on themselves, with each discharge door being connected to a corresponding bar in order to open or close when this bar pivots.

Abstract: A supersonic aircraft is disclosed herein. The supersonic aircraft comprises one or more fuselages, one or more wings, at least one wing-mounted engine, at least one inlet mounted on the upper surface of or above the wing, an upper wing surface propulsion system fairing, and a lower wing surface propulsion system fairing or lower wing surface modification to account for the presence of the engine.

Abstract: An after-fan system for an engine may comprise an after-fan turbine an electrical generator operationally coupled to the after-fan turbine, and an electric motor electrically coupled to the electrical generator. The electrical generator may be configured to generate an electrical current in response to rotation of the after-fan turbine. The electric motor may be configured to generate torque.

Abstract: A propulsion unit for an aircraft includes a gas generator; an air inlet; and a sealing system inserted axially between the air inlet and the gas generator. The sealing system has at least: a compression ring with a first axial end configured for mechanical connection to one of the gas generator and the air inlet. The compression ring also has a second opposite axial end configured to be axially separated from the other of the gas generator and the air inlet by an annular space. An annular seal is mounted in the space and configured to be axially compressed by reduction of the axial dimension of said space.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A pylon system for coupling an engine to a vehicle is provided. The pylon system includes a vehicle pylon to be coupled to the vehicle. The vehicle pylon includes a seal along a portion of the vehicle pylon. The pylon system includes an engine pylon including an inboard longeron and an outboard longeron. The inboard longeron is coupled to the outboard longeron at a first end of the engine pylon and is spaced apart from the outboard longeron at a second end of the engine pylon. The engine pylon is to be coupled to the engine, and the engine pylon is slidably coupled to the seal such that the engine pylon is movable relative to the vehicle pylon between at least a first position and a second position.

Abstract: A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes 3-5 rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-6.5. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.3-1.6.

Abstract: A gas turbine engine comprises, in axial flow sequence, a fan assembly, a compressor module, a turbine module, and an exhaust module. The gas turbine engine comprises first and second engine mount planes. The first and second engine mount planes are configured to support gas turbine engine in machine body. The gas turbine engine has axial length L, with axial length extending from inlet face of engine to exhaust face of engine. The first engine mount plane is positioned axially along gas turbine engine at first engine mount plane position, with first engine mount plane position within a range of between 0.24*L to 0.32*L from the inlet face of the engine. The second engine mount plane is positioned axially along gas turbine engine at second engine mount plane position, with second engine mount plane position within a range of between 0.73*L to 0.79*L from the inlet face of the engine.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor section including a propulsor, a gear reduction, a first spool including a first compressor, a first turbine and a first shaft, and a second spool including a second compressor, a second turbine and a second shaft. The gear reduction includes a carrier and a plurality of gears. The plurality of gears include a sun gear, a ring gear, and a plurality of intermediate gears that engage the sun gear and the ring gear. At least two of the plurality of gears are double helical gears in meshed engagement.

Abstract: A gas turbine engine includes a bypass duct extending about a longitudinal centerline of the gas turbine engine. The bypass duct includes at least one bypass duct wall defining at least a portion of a bypass flow path through the bypass duct. The at least one bypass duct wall includes a scoop extending into the bypass flow path. The gas turbine engine further includes a bleed conduit including an inlet connected to the bypass duct within the scoop of the at least one bypass duct wall and at least one louver mounted to the bleed conduit within the inlet. The at least one louver extends between a leading edge and a trailing edge opposite the leading edge. The leading edge is located within the bypass flow path and the trailing edge is located within the bleed conduit.

Abstract: Embodiments of the invention are shown in the figures, where a gas turbine engine for an aircraft includes: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan including a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan about a rotational axis at a lower rotational speed than the core shaft, wherein the gearbox is an epicyclic gearbox with a plurality of planet gears arranged in first and second sets of planet gears, the planet gears of the first set being positioned displaced relative to the planet gears of the second set in a direction parallel to the rotational axis.

Abstract: A confluence structure of an aircraft bypass turbine engine which includes a confluence plate with a downstream end supported by a portion that is movable in the direction of the axis by a control mechanism which can optionally be adjusted in flight. A mobile portion of a sleeve delimiting the secondary stream on the outside, and an inner projection of the outer casing can also slide axially in certain embodiments. This provides a wide range of options for modifying the gas dilution and operating conditions of the engine.

Abstract: An after-fan system for an engine may comprise an after-fan turbine an electrical generator operationally coupled to the after-fan turbine, and an electric motor electrically coupled to the electrical generator. The electrical generator may be configured to generate an electrical current in response to rotation of the after-fan turbine. The electric motor may be configured to generate torque.

Abstract: A gas turbine engine is provided that includes an engine core and a bypass duct. The engine core includes a compressor section, a combustor section, a turbine section and a core flowpath. The compressor section includes a radial flow compressor rotor. The core flowpath extends through the compressor section, the combustor section and the turbine section from a core inlet to a core exhaust. The bypass duct includes a bypass flowpath that extends outside of the engine core from a bypass inlet to a bypass exhaust. The bypass inlet is disposed along the compressor section and fluidly coupled with core flowpath.

Abstract: A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals R 1 R 3 = ( E ? F ? P ) ? ( 1 - RqR Sec . - Fan 2 ) ( 1 - RqR Prim . - Fan 2 ) ? ( T P T S ) ? ( E ? B ? A ) .

Abstract: The present disclosure provides an ejector nozzle and an ejector including the same. The ejector nozzle includes a first tube having a first flow path into which a fluid is introduced, and a second tube provided outside the first tube and having an inner diameter larger than an inner diameter of the first tube, the second tube defining a second flow path between the first tube and the second tube, in which the first tube further includes a communication port that penetrates the first tube to allow the first flow path to communicate with the second flow path and is openably and closably provided, and in which when the communication port is opened, a part of the fluid flowing in the first flow path is allowed to flow along the second flow path.

Abstract: Device and methods for guiding bleed air in a turbofan gas turbine engine are disclosed. The devices provided include louvers and baffles that guide bleed air toward a bypass duct of the turbofan engine. The louvers and baffles have a geometric configuration that promotes desirable flow conditions and reduced energy loss.

Abstract: A confluence structure of an aircraft bypass turbine engine which includes a confluence plate with a downstream end supported by a portion that is movable in the direction of the axis by a control mechanism which can optionally be adjusted in flight. A mobile portion of a sleeve delimiting the secondary stream on the outside, and an inner projection of the outer casing can also slide axially in certain embodiments. This provides a wide range of options for modifying the gas dilution and operating conditions of the engine.

Abstract: A disclosed gas turbine engine includes a gas generator section for generating a gas stream flow and a propulsor section for generating propulsive thrust as a mass flow rate of air through a bypass flow path. The propulsor section includes a fan driven by a power turbine through a speed reduction device at a second rotational speed lower than a first rotational speed of the power turbine. An Engine Unit Thrust Parameter (“EUTP”) defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is between 0.05 and 0.13 during operation of the gas turbine engine.

Abstract: A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The fan defines a fan diameter and the turbine section includes a fan drive turbine with a diameter less than 0.50 the size of the fan diameter. A geared architecture is driven by the turbine section for rotating the fan about the axis.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine includes a propulsor section, a compressor section including a first compressor, a front center body case structure that defines a first core flow path for core airflow into the first compressor, an intermediate case structure, a geared architecture, a propulsor shaft, a first bearing structure mounted to the front center body case structure, a bearing compartment passage structure in communication with the first bearing structure through the front center body case structure, and a turbine section.

Abstract: A gas turbine engine comprising a planetary gear train, and a core engine casing. The gear train has a ratio of greater than approximately 3.0, with an input to the gear train being operatively connected to the compressor section, and an output from the gear train being operatively connected to the fan. The core engine casing encloses the compressor section and the turbine section. The fan has a diameter F, and the core engine casing has a diameter C. The core engine casing diameter C varies along an axial length of the core engine casing, and a ratio (C/F) of the core engine casing diameter C to the fan diameter F is within the range 0.2<(C/F)<0.4, along an axial length of the core engine casing.

Abstract: A method for gathering flight load data for a gas turbine engine includes providing a plurality of blocker doors rotatably mounted to a translating sleeve and a plurality of blocker door islands, each blocker door island disposed between circumferentially adjacent blocker doors of the plurality of blocker doors and mounted to at least one mounting point of the translating sleeve, wherein a blocker door island of the plurality of blocker door islands is a data acquisition unit; and connecting the data acquisition unit to at least one sensor in communication with the translating sleeve.

Abstract: A double flow turbojet including: a low pressure compressor; a series of casings downstream of this low pressure compressor to delimit a primary flow path for circulating a primary stream, and including an upstream edge delimiting an inlet opening; a high pressure compressor in the primary flow path; a shroud surrounding the series of casings to delimit a flow path for circulating an intermediate stream, and having an upstream edge delimiting a circular inlet opening situated upstream of the high pressure compressor; a secondary flow path casing surrounding the shroud to delimit a secondary flow path for circulating a secondary stream; an exhaust casing including radial arms collecting the air coming from the intermediate flow path.

Abstract: A valve assembly for a turbine engine includes a valve arranged in an air flow path. The valve has a shaft. The valve is movable between an open position and a closed position. A hydraulic actuator is coupled to the shaft. The hydraulic actuator is configured to utilize fuel as a working fluid. The hydraulic actuator is configured to move the valve between the open position and the closed position.

Abstract: A gas turbine engine turbine has a high pressure turbine configured to rotate with a high pressure compressor as a high pressure spool in a first direction about a central axis and a low pressure turbine configured to rotate with a low pressure compressor as a low pressure spool in the first direction about the central axis. A power density is greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/cubic inches. A fan is connected to the low pressure spool via a speed changing mechanism and rotates in the first direction.

Abstract: A bleed system of a gas turbine engine includes a bleed duct having a duct inlet located at a flowpath of a gas turbine engine, and a bleed outlet located outside of the flowpath, and extending circumferentially around a central longitudinal axis. A plurality of bleed doors are located at the bleed outlet and are arrayed along a circumferential length on the bleed duct. Each bleed door includes a first circumferential end, and a second circumferential end. The bleed doors are arrayed such that when the bleed doors are in a closed position the first circumferential end is located at the second circumferential end of an adjacent bleed door of the bleed doors. Each bleed door includes a pivot, such that each bleed door rotates about the pivot from the closed position covering the duct outlet to an opened position allowing a bleed airflow to pass through the duct outlet.

Abstract: An engine including a motor and a nacelle and a duct between the nacelle and the motor. The nacelle includes a fixed structure, a mobile assembly that is mobile between an advanced position and a retracted position to define a window between the duct and the outside, and a plurality of blades that are mobile in rotation between a stowed position and a deployed position, each one extending on either side of its axis of rotation with a first arm and a second arm. In the stowed position, the first arm is outside the duct and the second arm is inside the nacelle, and where, in the deployed position, the first arm is across the duct and the second arm projects out of the nacelle. With such blades, the flow of air is optimally directed towards the front without it being necessary to provide cascades.

Abstract: A boost compressor assembly may comprise an outer annular structure and a plurality of blades. Each blade in the plurality of blades may be moveably coupled to the outer annular structure. The plurality of blades may be configured to deploy in response to the boost compressor assembly rotating. The plurality of blades may be configured to retract when the boost compressor assembly stops rotating.

Abstract: A turbofan gas turbine engine includes a front center body case structure. A geared architecture is at least partially supported by the front center body case structure. A bearing structure rotationally support a fan shaft. A fan rotor bearing support structure extends radially inward from the front center body and at least partially supports the bearing structure. The fan rotor bearing support structure includes a passage that communicates buffer supply air from a hollow strut to the bearing structure.

Abstract: A heat exchanger for cooling a component is coupled with a cowl at least partially surrounding the component. The cowl defines a cowl plenum and a peripheral gap for receiving the heat exchanger. The heat exchanger includes a housing defining a heat exchange plenum for receiving a cool fluid stream and a plurality of heat exchange tubes passing through the heat exchange plenum for receiving a hot fluid stream. A discharge manifold defines a discharge plenum that provides fluid communication between the heat exchange plenum and the cowl plenum through a fluid outlet. In addition, an impingement baffle at least partially defines the discharge manifold and defines a plurality of cooling holes for impinging cooling air on the component proximate the heat exchanger.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

Abstract: A gas turbine engine coupled to an aircraft includes an engine core arranged axially along an axis, a bypass duct arranged circumferentially around the engine core to define a bypass channel, and a heat exchanger system. The bypass channel is arranged to conduct bypass air around the engine core to provide thrust for the gas turbine engine. The heat exchanger system is configured to provide cooling for the engine core.

Abstract: A method for controlling flow through an exhaust nozzle includes: providing a centerbody including a maximum diameter section; providing an inner shroud surrounding the centerbody, including at least a middle section of decreased diameter and terminating at an aft edge; providing an outer shroud. wherein the centerbody and the inner shroud collectively define a throat, and the outer shroud and the centerbody collectively define an exit; selectively translating the inner shroud and outer shroud to vary the throat; and selectively translating the outer shroud to vary the ratio of the exit to the throat; wherein, when the inner shroud is in a forward position, its aft edge is forward of the maximum diameter section of the centerbody, such that the throat of the nozzle is formed between the aft edge of the inner shroud and the centerbody.

Abstract: A turbofan engine having: an engine core having a centre axis and including in flow series a compressor, a combustor and a turbine; and a bypass duct surrounding the engine core, the bypass duct has a bypass duct exit area at its downstream end. The engine further includes an exhaust nozzle assembly including: coaxially arranged inner mixer and outer exhaust nozzles, the exhaust nozzle being axially downstream of said mixer nozzle; a core flow duct defined by the mixer nozzle, the core flow duct having a core exit area; and an exhaust duct defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust throat area.

Abstract: A turbofan engine includes a fan rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a fan drive gear system including a carrier for supporting a plurality of gears, and a scupper capturing lubricant during gear operation and directing lubricant into the carrier. A fan drive gear system is also disclosed.

Abstract: Disclosed herein is a method for controlling a surge margin of a gas turbine and an extraction device for a gas turbine. The method for controlling the surge margin of the gas turbine and the extraction device for the gas turbine may support stable operation of a compressor unit in the gas turbine, thereby improving the efficiency of the gas turbine and minimizing vibration and noise of the gas turbine.

Abstract: A double flow turbojet includes a fan including a disk centered on an axis of the fan which is provided with fan blades on its periphery, the blades having a leading edge, and an air inlet sleeve extending upstream of the fan and configured to delimit a gas flow designed to enter into the fan the air inlet sleeve having a collecting surface, the turbojet having an aspect ratio S 2 S col included in the interval [ 1.0 ; 1.0 + 0.4 ? ( L D ) ] , where L/D is the form factor of the air inlet sleeve.

Abstract: A gas turbine engine includes a fan and a compressor section that is fluidly connected to the fan. The compressor includes a high pressure compressor and a low pressure compressor. A combustor is fluidly connected to the compressor section. A turbine section is fluidly connected to the combustor. The turbine section includes a high pressure turbine coupled to the high pressure compressor via a first shaft. A low pressure turbine is coupled to the low pressure compressor via a second shaft. A geared architecture interconnects between the second shaft and the fan. The gas turbine engine is a high bypass geared aircraft engine having a bypass ratio of greater than six (6). The low pressure turbine has a pressure ratio that is greater than 5, and the geared architecture includes a gear reduction ratio of greater than 2.5:1.

Abstract: A propulsion and electric power generation system includes a gas turbine propulsion engine, an electrical generator, an aircraft power distribution system, a plurality of auxiliary fans, and a controller. The gas turbine propulsion engine includes at least a low-pressure turbine coupled to a fan via a low-pressure spool, and the low-pressure turbine is configured to generate mechanical power. The electrical generator is directly connected to the low-pressure spool and generates a total amount of electrical power (Pe). The aircraft power distribution system receives a first fraction (Pa) of the total amount of electrical power. The auxiliary fans receive a second fraction (Pf) of the total amount of electrical power. The controller is configured to control a ratio of Pf to Pa (Pf/Pa) such that the ratio spans a range from less than 0.6 to at least 0.9.

Abstract: According to various aspects, an unmanned aerial vehicle may include one or more electric drive motors. Each drive motor of the one or more electric drive motors may include a first drive rotor structure; a second drive rotor structure, the first drive rotor structure being coaxially aligned with the second drive rotor structure, the first drive rotor structure including a first magnetic arrangement and the second drive rotor structure including a second magnetic arrangement.

Abstract: A method of operating a twin engine helicopter power plant, the power plant comprising: two turboshaft engines each having an engine shaft with a turbine at a distal end and a one-way clutch at a proximal end; a gear box having an input driven by the one way clutch of each engine and an output driving a helicopter rotor; a bypass clutch disposed between the proximal end of each engine shaft and the input of the gear box; and power plant management system controls for activating the bypass clutch; the method comprising: detecting when a rotary speed of an associated engine shaft is less than a rotary speed of the gear box input; activating the bypass clutch to drive the associated engine shaft using the rotation of the gear box input; and starting an associated engine by injecting fuel when the bypass clutch is activated.

Abstract: A method of extracting work from a convertible gas turbine engine having a core flowpath and a bypass flowpath. The method comprises operating the convertible gas turbine engine at a first volumetric flow rate through the core flowpath and a second volumetric flow rate through the bypass flowpath to produce a first work output of the convertible gas turbine engine; extracting the first work output via an unshrouded fan and a shaft at a first fan to shaft extraction ratio; altering the second volumetric flowrate through the bypass flowpath while maintaining the first work output; and extracting the first work output via an unshrouded fan and a shaft at a second fan to shaft extraction ratio.

Abstract: The invention relates to a hub (2) of an intermediate casing (1) for a bypass turbomachine comprising: —a bleed stream duct (18), —a bleed valve, comprising a mobile door at the inlet orifice to the bleed stream duct (18), —a set of bleed vanes (22) which are mounted with the ability to rotate about a pivot (26) in the bleed stream duct (18) between an open configuration in which a flow of air coming from the inlet orifice (4) passes between the bleed vanes (22) and a closed configuration, the pivot (26) for each bleed vane (22) being closer to its leading edge (BA) than to its trailing edge (BF).

Abstract: An exhaust system or nozzle for use in a gas turbine engine is disclosed herein. The exhaust system is adapted to adjust various streams of pressurized air produced by the gas turbine engine to control operation of the gas turbine.

Abstract: A gas turbine engine comprises a fan on an engine central axis. Plural gas generators are downstream of the fan, each along a respective central axis, mutually offset, and offset from the engine central axis. A fan drive turbine is on the engine central axis, downstream of the dual gas generators, and driven by output from the dual gas generators, to drive the fan.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: A gas turbine engine has a fan, first and second compressor stages, first and second turbine stages. The first turbine stage drives the second compressor stage as a high spool. The second turbine stage drives the first compressor stage as part of a low spool. A gear train drives the fan with the low spool, such that the fan and first compressor stage rotate in the same direction. The high spool operates at higher pressures than the low spool. A lubrication system is also disclosed.

Abstract: An exhaust system for a variable cycle aircraft engine. The exhaust system comprises a core exhaust for bypass air and hot gases of combustion. The core exhaust includes a convergent-divergent nozzle. The convergent-divergent nozzle is formed from a plurality of flaps and seals. The exhaust system comprises a third air duct for a third stream of air. The third stream of air is selectively exhausted from the third duct through a secondary nozzle or divergent slots in the convergent-divergent nozzle, or both depending upon the flight mode. A diverter valve is positioned in the third stream duct to selectively control the flow of third stream air through the secondary nozzle, the divergent slots and combinations thereof.