Abstract: A twin fan turbomachine with two contrarotating fans, one upstream and one downstream, driven by two contrarotating coaxial shafts each of which connects one fan impeller to at least one low-pressure turbine impeller located downstream, the downstream fan impeller being connected by its downstream part to its drive shaft.

Abstract: A turbine engine arrangement is provided in which contra rotating shafts 104, 105 are respectively secured to a fan and a gearbox 106 which is also coupled to the shaft 104. In such circumstances the relative rotational speed ratio between the shafts 104, 105 can be determined with a first low pressure turbine 101 secured to the first shaft 104 arranged to rotate at a lower speed but provide high work whilst a second low pressure turbine 102 secured to the second shaft 105 rotates at a higher speed governed by the gearbox 106. By such an arrangement a smaller gearbox 106 may be used as less power is transferred through that gearbox 106 than with previous arrangements. By contra rotation of the turbines 101, 102 a lower flow deflection guide vane assembly 103 may be used and a further stator/guide vane assembly is not required between the turbines 101, 102.

Abstract: A turbojet has a low-pressure compressor and two fans at the front of an intermediate casing which fans are driven by two independent shafts. The compressor is disposed under the rear fan and has a downstream ring of rotor blades connecting the rear fan to the drive shaft of the rear fan, an upstream rotor blade ring that is connected to the drive shaft of the front fan and at least one ring of stator blades that is disposed between the rings of rotor blades and inside an outer stator that is supported by a second intermediate casing. Bearings are interposed respectively between the two intermediate casings and the two shafts.

Abstract: A method of assembling a gas turbine engine that includes rotatably coupling a first low-pressure turbine rotor to a high-pressure turbine, rotatably coupling a second low pressure turbine rotor to the first low-pressure turbine rotor, and rotatably coupling the second low-pressure turbine rotor to a turbine rear-frame such that a weight of the high-pressure turbine is transmitted to the turbine rear-frame.

Abstract: An aircraft propulsion system includes a gas turbine engine having a fan section, at least one row of FLADE fan blades disposed radially outwardly of and drivingly connected to the fan section, the row of FLADE fan blades radially extending across a FLADE duct circumscribing the fan section, an engine inlet including a fan inlet to the fan section and an annular FLADE inlet to the FLADE duct. A fixed geometry inlet duct is in direct flow communication with the engine inlet. The fan section may include only a single direction of rotation fan or alternatively axially spaced apart first and second counter-rotatable fans in which the FLADE fan blades are drivingly connected to one of the first and second counter-rotatable fans. The row of FLADE fan blades may be disposed between rows of variable first and second FLADE vanes.

Abstract: A method for assembling a turbine engine assembly is provided. The turbine engine includes a core engine. The method includes coupling a first rotor spool within the engine assembly wherein the first rotor spool includes a first shaft. The method further includes coupling a second rotor spool within the engine assembly wherein the second rotor spool includes a second shaft. The method also includes coupling a third rotor spool within the engine assembly wherein the third rotor spool includes a third shaft. At least one of the first, second, and third shafts include a first end, an opposing second end, and a tubular portion extending between the first and second ends. A reinforcing layer circumscribes a portion of the tubular portion wherein at least a portion of the reinforcing layer is a metallic matrix composite (MMC) material that includes reinforcing fibers. A turbine engine is also provided.

Abstract: A turbofan engine assembly includes a core gas turbine engine including a high-pressure compressor, a combustor, and a high-pressure turbine, a low-pressure turbine coupled to said core gas turbine engine, a counter-rotating booster compressor comprising a first rotor section configured to rotate in a first direction and a second rotor section configured to rotate in an opposite second direction, and a gearbox comprising an input and an output, said gearbox output coupled to at least one of said first and second rotor sections, said gearbox input coupled to said low-pressure turbine. A method of assembling the turbofan engine assembly described herein is also provided.

Abstract: A turbofan engine assembly includes a core gas turbine engine including a high-pressure compressor, a combustor, and a high-pressure turbine, a booster compressor coupled upstream from the core gas turbine engine, an intermediate-pressure turbine coupled to the booster compressor, the intermediate-pressure turbine disposed downstream from the core gas turbine engine, a counter-rotating fan assembly disposed upstream from the booster compressor, the counter-rotating fan assembly comprising a first fan configured to rotate in a first direction and a second fan configured to rotate in an opposite second direction, and an intermediate-pressure turbine disposed downstream from the core gas turbine engine, wherein the intermediate-pressure turbine drives the booster compressor and the second fan assembly. A method of assembling the above turbofan engine assembly is also described herein.

Abstract: A method for assembling a gas turbine engine includes coupling a low-pressure turbine to a core turbine engine, coupling a counter-rotating fan assembly including a forward fan assembly and an axially aft fan assembly to the low-pressure turbine such that the forward fan assembly rotates in a first direction and the aft fan assembly rotates in an opposite second direction, and coupling a booster compressor directly to the low-pressure turbine such that the booster compressor rotates in the first direction.

Abstract: A turbofan engine assembly includes a core gas turbine engine including a high-pressure compressor, a combustor, and a high-pressure turbine, a booster compressor coupled upstream from the core gas turbine engine, an intermediate-pressure turbine coupled to the booster compressor, the intermediate-pressure turbine disposed downstream from the core gas turbine engine, a counter-rotating fan assembly disposed upstream from the booster compressor, the counter-rotating fan assembly comprising a first fan configured to rotate in a first direction and a second fan configured to rotate in an opposite second direction, and a low-pressure turbine disposed downstream from the intermediate-pressure turbine, the low-pressure turbine configured to drive the counter-rotating fan assembly. A method of assembling the above turbofan engine assembly is also described herein.

Abstract: A method for assembling a gas turbine engine includes providing a first fan assembly including a plurality of rotor blades that are configured to rotate in a first rotational direction at a first rotational speed, rotatably coupling a second fan assembly axially aft of the first fan assembly, wherein the second fan assembly includes a plurality of rotor blades that are configured to rotate in a second rotational direction, and coupling a gearbox to the second fan assembly that is configured to rotate the second fan assembly at a second rotational speed that is different than the first rotational speed.

Abstract: The invention relates to aerodynamically matching the rear fan of a turbojet having two fans and a low-pressure compressor at the front of the intermediate casing, said fans being driven by two independent shafts. Said compressor is disposed between the blades of the two fans and includes at least one ring of rotor blades around the periphery of a wheel driven by the drive shaft for the front fan, and at least two grids of stator vanes disposed on either side of the ring of rotor blades and inside a grid carrier ring. A stationary outer grid and a stator of variable pitch connect the ring to the fan casing.

Abstract: A method for assembling a gas turbine engine that includes providing a first fan assembly configured to rotate in a first rotational direction, rotatably coupling a second fan assembly to the first fan assembly, wherein the second fan assembly is configured to rotate in a second rotational direction that is opposite the first rotational direction, coupling a first shaft to the first fan assembly and to a first turbine rotor that is configured to rotate in a first rotational direction, coupling a second shaft coupled to the second fan assembly and to a second turbine rotor that is configured to rotate in a second rotational direction that is opposite the first rotational direction, and coupling a lubrication system to the gas turbine engine such that a lubrication fluid is channeled through the first shaft to lubricate at least one of the first and second fan assemblies.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, and rotatably coupling a low-pressure turbine outer rotor to the inner rotor, wherein the outer rotor includes a second plurality of turbine blade rows that are configured to rotate in a second direction that is opposite the first rotational direction of the inner rotor and such that at least one of the second plurality of turbine blade rows is coupled axially forward of the first plurality of turbine blade rows.

Abstract: An exemplary embodiment of a FLADE counter-rotating fan aircraft gas turbine engine includes at least one row of FLADE fan blades disposed radially outwardly of and drivingly connected to one of axially spaced-apart first and second counter-rotatable fans. A core engine located downstream and axially aft of the first and second counter-rotatable fans is circumscribed by a fan bypass duct downstream and axially aft of the first and second counter-rotatable fans. The row of FLADE fan blades radially extend across a FLADE duct circumscribed about the first and second counter-rotatable fans and the fan bypass duct. A second low pressure turbine is drivingly connected to the first counter-rotatable fan and a first low pressure turbine is drivingly connected to the second counter-rotatable fan.

Abstract: The invention relates to a turbomachine (1) for an aircraft comprising a gas generator (4) designed such that hot gases are ejected from a combustion chamber (28) towards the upstream side of the turbomachine, and also comprising a plurality of hollow mixer struts (40) connected to an output (36) from the gas generator through which hot gases can pass, each mixer strut comprising an output mixer portion (40c) located inside the annular fan duct (8) so as to mix the secondary air and the hot gases ejected by these output mixer portions towards the downstream side of the turbomachine.

Abstract: A gas turbine engine comprising a combustion chamber section, a turbine section, and a compressor section. The turbine section surrounds the combustion chamber and the compressor section surrounds the turbine section.

Abstract: A method enables a gas turbine engine including a first rotor assembly and a second rotor assembly coupled in axial flow communication downstream from the first rotor assembly to be assembled. The method comprises coupling an upstream end of an extension duct to an outlet of the first rotor assembly, wherein the extension duct includes a plurality of panels coupled circumferentially, and coupling a downstream end of the extension duct to an inlet of the second rotor assembly using at least one fish mouth seal.

Abstract: An aircraft gas turbine engine includes a low pressure turbine having a low pressure turbine flowpath and low pressure inner and outer shaft turbines with counter rotatable low pressure inner and outer shaft turbine rotors, respectively. The low pressure inner and outer shaft turbine rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath which are drivingly connected to first and second fan blade rows by low pressure inner and outer shafts, respectively. At least one row of low pressure variable vanes is operably disposed across the low pressure turbine flowpath between the low pressure inner and outer shaft turbines. The low pressure first turbine blade rows of the low pressure inner shaft turbine may be in tandem with or interdigitated with the second turbine blade rows of the low pressure outer shaft turbines.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft, the counterrotatable booster compressor assembly including a first compressor blade row connected to the first drive shaft and a second compressor blade row interdigitated with the first compressor blade row and connected to the second drive shaft. A portion of each fan blade of the second fan blade row extends through a flowpath of the counterrotatable booster compressor so as to function as a compressor blade in the second compressor blade row.

Abstract: An aircraft gas turbine engine includes high and low pressure turbines having respective counter rotatable low pressure inner and outer rotors with low pressure inner and outer shafts in part rotatably disposed co-axially within a high pressure rotor and drivingly connected to first and second fan blade rows and first and second boosters respectively. A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the boosters axially located between the first and second fan blade rows. The engine has a high pressure compressor operable to produce an overall pressure ratio in a range of about 40-65 and a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds in a range of 1000 to 2500 feet per second.

Abstract: An aircraft gas turbine engine includes a low pressure turbine having a low pressure turbine flowpath and counter rotatable low pressure inner and outer shaft rotors having inner and outer shafts, respectively. The low pressure inner and outer shaft rotors include low pressure first and second turbine blade rows disposed across the turbine flowpath and drivingly connected to first and second fan blade rows by low pressure inner and outer shafts, respectively. At least one of the low pressure first and second turbine blade rows is interdigitated with an adjacent pair of one of the turbine blade rows. At least one row of non-rotatable low pressure vanes is disposed across the low pressure turbine flowpath between a non interdigitated adjacent pair of one of the turbine blade rows.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft. The counterrotatable booster compressor assembly includes a first compressor blade row connected to the first drive shaft, a plurality of fan shaft extensions connected to the second drive shaft for driving the second fan blade row, and at least one compressor blade integral with each fan shaft extension so as to form a second compressor blade row interdigitated with the first compressor blade row.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft. The counterrotatable booster compressor assembly includes a first compressor blade row connected to the first drive shaft, a plurality of fan shaft extensions connected to the second drive shaft for driving the second fan blade row, and at least one compressor blade integral with each fan shaft extension so as to form a second compressor blade row interdigitated with the first compressor blade row.

Abstract: A gas turbine engine comprising a combustion chamber section, a turbine section, and a compressor section. The turbine section surrounds the combustion chamber and the compressor section surrounds the turbine section.

Abstract: A counterrotatable booster compressor assembly for a gas turbine engine having a counterrotatable fan section with a first fan blade row connected to a first drive shaft and a second fan blade row axially spaced from the first fan blade row and connected to a second drive shaft, the counterrotatable booster compressor assembly including a first compressor blade row connected to the first drive shaft and a second compressor blade row interdigitated with the first compressor blade row and connected to the second drive shaft. A portion of each fan blade of the second fan blade row extends through a flowpath of the counterrotatable booster compressor so as to function as a compressor blade in the second compressor blade row.

Abstract: A rotatable turbine frame includes annular outer and inner bands having a plurality of circumferentially spaced apart struts extending therebetween. An annular drive shaft is fixedly joined to the inner band for transmitting output torque therefrom, and the struts are backwardly radially inclined relative to the direction of rotation of the frame so that gas flow between the struts tends to straighten the inclined struts to effect a compressive load component therein.

Abstract: A gas turbine engine rotor support system includes outer and inner rotors having respective outer and inner blades interdigitated in respective blade row stages. A stationary rear frame includes a rear support shaft. A rotatable aft frame is disposed forwardly of the rear frame includes an aft support shaft. And, a rotatable forward frame is disposed forwardly of the aft frame and includes a forward support shaft. The forward and aft frames support the outer and inner rotors, and a first bearing is disposed between the aft shaft and the rear shaft for supporting the aft shaft; a second bearing is disposed between the forward shaft and the aft shaft for supporting the forward shaft; and a third bearing is disposed between a core shaft of a core turbine and the forward shaft for supporting the core shaft thereon.

Abstract: A turbofan engine includes two oppositely rotating fan rotors which are arranged upstream of the core engine, and has a booster with two contra-rotating coaxial rotors. The radially exterior booster rotor is arranged between the rearward fan rotor and the low-pressure turbine in a torque-transmitting manner. The gearless construction permits an almost uniform loading of the low-pressure shafts.

Abstract: An internal combustion rotary engine has a housing which has a plurality of outer rotors and stators and bearings, the outer rotors having a plurality of curved channels on inner surfaces thereof and one of the stators having intake ports and a port for a spark plug; a shaft which is rotatably supported by the bearings and has a plurality of holes communicating with a center passage formed therein for flowing of a coolant; a central rotor being formed as a frustum and fixedly supported by the shaft and having a plurality of curved channels overlapping with the channels on the rotors at ends thereof, but in a reversed direction, and on a first conical surface at a middle portion thereof and a second conical surface with a larger slope than the first conical surface at a front portion thereof to define a combustion chamber.

Abstract: A gas turbine engine has outer and inner counterrotatable annular turbine rotors and pluralities of turbine blades disposed in alternating rows and mounted on and extending between the respective rotors. An interstage seal arrangement for the engine is composed of a sequence of alternating inner and outer seals. Each outer seal is disposed between an interior side of the outer rotor and outer ends of the blades in one plurality thereof being spaced radially inward at their outer ends from the outer rotor and being mounted at their inner ends to the exterior side of the inner rotor. Each inner seal is disposed between the exterior side of the inner rotor and inner ends of the blades in another plurality thereof being spaced radially outward at their inner ends from the inner rotor and being mounted at their outer ends to the interior side of the outer rotor.

Abstract: A turbopump system for pumping two separate fluids is disclosed having a r of concentric turbine rotors located within a turbine casing, each directly connected to a pump impeller. The centrifugal pumps, which may be located on opposite sides of the turbine casing are separately connected to their respective fluids such that the two fluids may be pumped simultaneously. A propellant fluid is directed through a generally annular space defined between the concentric turbine rotors to rotate them in opposite directions. Turbine blades attached to each of the turbine rotors extend into this annular space in order to drive the turbine rotors. The pitches of the turbine blades are such that the rotors are driven in opposite directions and one of the turbine rotors may be driven at a different speed than the other.

Abstract: A combined turborocket and ramjet propulsion unit comprises an air intake duct, and an air compressor with two counter-rotating rotor stages disposed downstream of the duct in an annular path formed between an outer casing and a central body which is maintained in position relative to the casing by structural arms. The rotor stages of the compressor are driven by a power turbine driven by combustion gases from a gas generator which burns liquid propellants. The power turbine includes two interleaved counter-rotating modules arraanged within the central body in a position longitudinally between the two compressor stages which they drive, and the gas generator is disposed axially within the central body to the rear of the turbine modules, the generator supplying the combustion gases to the turbine in an upstream direction through a central ejection chamber and an annular reversing chamber which redirects the gases rearwards through the turbine.

Abstract: A gas turbine engine power unit (10) comprising a gas turbine engine hot gas generator (10a), the exhaust efflux of which drives a turbine (20) having a power output shaft (22). The exhaust efflux from the turbine (20) is directed into a chamber (23) having two outlets (24,25) to atmosphere. The outlets (24,25) are provided with valves (26,28) which are operable to ensure that only one outlet (24,25) is open at a given time. The first outlet (24) contains a heat exchanger (15) whereas the second outlet (25) contains a further turbine (30). The further turbine (30) drives an air compressor (32) the output of which is directed to an air inlet (36) of the gas turbine engine (10a) in order to facilitate a boost in its output power.

Abstract: A gas turbine engine with counterrotating fans placed upstream from the engine so that gas circulates from upstream to downstream at the level of a LP compressor (1), then from downstream to upstream in a HP compressor (5), chamber an (6) and HP and LP turbines (7 and 8), then again downstream in first and second turbines (10, 11) before being expelled into the cold airflow coming from first and second fans (14, 15).

Abstract: A gas turbine engine intended for use as an aircraft engine which includes a gas generator which supplies interdigitated, contra-rotating radially inner and outer low-speed power turbines driving forward mounted contra-rotating front and rear fans, the gases flowing through the gas generator from the rear of the engine towards the front, and then flowing in a rearward direction through the power turbines. The front fan is driven by the radially inner power turbine via a central shaft extending towards the front of the engine and supporting a disc carrying the front fan, and the rear fan is driven by the radially outer power turbine via a forwardly extending drum which is rotatably mounted between the central shaft and a stationary structure of the engine and which carries the rear fan. This structure permits the pod mounting of a high by-pass ratio engine under wing of an aircraft.

Abstract: A new and improved gas turbine engine including a gas generator and a power turbine is disclosed. The power turbine includes a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom. The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving the first and second rotors in counterrotating directions.

Abstract: A turbojet bypass engine includes an upstream fan of a conventional type also a downstream fan arranged to rotate in the opposite direction to the upstream fan and driven by a free turbine interleaved with the low pressure turbine which drives the upstream fan. The two fans are supplied by separpate overlapping and inverleaved air flow paths which are formed largely within the nacelle structure of the engine, giving the nacelle an ovoide shape which allows such engines with a high bypass ratio to be mounted below the wings of an aircraft.

Abstract: A propulsion system for a reusable spacecraft is disclosed having turbojet, amjet and rocket modes of operation. Hydrogen or exhaust gases from a gas generator drives a gas turbine which powers an air compressor in the turbojet mode. An injection device injects hydrogen and exhaust from the gas driven turbine in the combustion chamber in the turbojet mode. In the ramjet mode, only hydrogen is injected into the combustion chamber. In the ramjet mode, hydrogen and oxygen are supplied to the rocket motor. An adjustable nozzle is provided to form a variable throat convergent-divergent nozzle in the turbojet and ramjet modes and to form a divergent nozzle in the rocket mode.

Abstract: A combination engine for a flying craft capable of flying at speeds within a wide range from subsonic speeds to hypersonic speeds, is equipped with a rocket engine section that is independent of an atmospheric air supply, and a turbo air jet engine section having a compressor which is driven by a turbine that is also independent of an external air supply. The combination may further include extra fuel supply nozzles for operation as a ram jet which uses the same flow channel as the turbo air jet engine section. A gas generator equipped with two valving mechanisms is arranged coaxially with the rocket engine section. The valving mechanisms are controllable to open the gas generator chamber to the rocket combustion chamber while closing off gas flow to the turbo air jet section or vice versa. Either section can work alone or in parallel and simultaneously with the other section.

Abstract: A combined air-hydrogen turbo-rocket engine is disclosed having a simplified construction in which the hydrogen driven turbine is formed integrally with the rotor wheel of the axial air compressor stages. The rotor stages are located downstream of a stator vane structure and are driven by gaseous hydrogen passing across the turbine blades. The hydrogen is subsequently injected into an air duct surrounding the axial air compressor and defining an airflow path having an air inlet. The hydrogen-air mixture is ignited and the burned gases are expanded through a converging-diverging exhaust nozzle.

Abstract: High by-pass fan jet engine includes counterrotating turbine blade sets for receiving hot combustion gases from a core engine portion and driving the fan blades through a planetary gear-type reduction gear assembly. Lightweight, highspeed, concentric, counterrotating shafts transmit power from both turbine blade sets to the fan via the reduction gear assembly. A pair of planetary gear assemblies with fixed fing gears and a rotatable common planetary gear carrier is used to drive a single set of fan blades via the carrier, and a pair of planetary gear assemblies with a fixed common planetary gear carrier and rotatable ring gears is used to drive two or more counterrotating fan blade sets via the ring gears. A low pressure "booster" compressor feeding the core engine portion can be directly driven from one of the two high speed shafts.

Abstract: A high bypass ratio turbofan engine having a fan section, a booster compressor disposed aft of the fan section relative to the flow of combustion gases through the engine, and a core section disposed aft of the booster compressor. A low pressure counterrotating turbine, disposed aft of the core section, is used for driving the fan section and the booster compressor. The counterrotating turbine includes at least one set of rotating turbine blades and at least one set of oppositely rotating counterrotating turbine blades. A twin spool shaft is provided for coupling the turbine blades to the booster compressor and for coupling the counterrotating turbine blades to the fan section. A reduction gear is disposed in the drive shaft for coupling the turbine blades to the fan section and for reducing the rotational speed of the turbine output power to match the rotational speed of the fan section thereby splitting the usable work of the turbine blades between the fan section and the booster.

Abstract: A coupling arrangement for supporting a power turbine in a gas turbine engine. The power turbine section comprises a stator structure, first and second rotors, and first and second bearings. The first and second rotors are coaxially disposed about a longitudinal axis of the stator. An annular gas flowpath is coaxially positioned between the first and second rotors. Annular arrays of turbine blades are coupled to the first and second rotors and extend into the flowpath so that a gas stream flowing through the flowpath reacts with the turbine blades causing the rotors to counterrotate. The first bearings are interposed between the first rotor and the stator structure to rotatably secure the first rotor to the stator structure. The second bearings are interposed between the second rotor and the first rotor to rotatably secure the second rotor to the first rotor.

Abstract: A propfan turbo-engine has two shrouded propfan rotors arranged upstream of a gas turbine which are driven by a contra-rotating working turbine. In addition to a favorable flow onto the blades, a low weight of the engine can be achieved.

Abstract: An unducted turbofan engine comprises first and second fan blades journaled for rotation from a folded position substantially parallel to the axis of rotation of the engine to a radially orientated propulsion condition. A releasable shroud normally retains the fan blades in the folded condition and releases the fan blades in a controlled sequence.

Abstract: A gas turbine engine having a core gas generator engine for generating combustion gases, with an additional power turbine, an additional fan section, and a booster compressor. The power turbine includes two counter rotatable turbine blade rows which are alternately interdigitized and serve to rotate counter rotating first and second drive shafts, respectively. The fan section also includes counter rotating spaced apart fan blade sections which are respectively connected to the first and second drive shafts. A booster compressor is axially positioned between the spaced apart fan sections. The booster compressor likewise includes first and second blade rows which are counter rotating alternately interdigitized and are likewise driven by the first and second drive shafts. The engine is supported from the two frames to permit the nacelle to be non-structural. The core engine is a modular unit so that it can be separable from the counter rotating turbine, fans and booster.

Abstract: A turboprop propulsion system where there is a core engine driving a power turbine that is connected through a planetary drive transmission to first and second counterrotating propellers. One of the propellers is connected to the ring gear, while the other propeller is connected to the planetary carrier, the sun gear being driven from the power turbine. The apparatus is provided with a substantially continuous gaseous flow path from the inlet to the exhaust nozzle of the engine, with the planetary drive transmission being spaced radially from the flow path. In some embodiments, the planetary transmission is positioned radially inwardly of the flow path, while in other embodiments, it is positioned radially outwardly of the flow path. The propellers can be in either the pusher configuration or the tractor configuration.

Abstract: The invention relates to intakes for turbopropeller gas turbine engines of the tractor type. A tractor turbopropeller gas turbine engine comprises a gas generator, an upstream propeller and a downstream propeller. The propellers are positioned upstream of the gas generator, and the propellers are driven via shaft means and gear means. An intake means supplies air to the gas generator, the intake means comprises an intake opening and an intake duct. The intake opening is positioned axially between the propellers, and a portion of the intake duct extends axially through the downstream propeller. This arrangement reduces the length of the shaft and the turbopropeller gas turbine engine, and reduces the weight accordingly. The intake opening is arranged to provide improved protection against ingestion of large foreign objects, such as birds, into the intake duct and gas generator.

Abstract: A gas turbine engine comprising a core gas generator engine for generating combustion gases, an additional power turbine, an additional fan section, and a booster compressor is disclosed. The power turbine includes first and second counterrotatable turbine blade rows effective for rotating first and second drive shafts, respectively. The fan section includes a first fan blade row connected to the first drive shaft and a second fan blade row connected to the second drive shaft. The booster compressor includes a first compressor blade row connected to the first drive shaft and a second compressor blade row connected to the second drive shaft.