Abstract: A turbine section for a gas turbine engine according to an example of the present disclosure includes, among other things, a first turbine rotor coupled to a first turbine shaft. The first turbine shaft is rotatable about a longitudinal axis. A second turbine rotor is coupled to a second turbine shaft. The second turbine shaft is rotatable about the longitudinal axis, and the second turbine rotor is axially aft of the first turbine rotor relative to the longitudinal axis. An aft bearing assembly rotatably supports the second turbine shaft. The second turbine rotor includes a disk assembly that carries at least one row of turbine blades. The disk assembly is mechanically attached to the second turbine shaft at an attachment point. The attachment point is axially aft of the aft bearing assembly such that an aft portion of the second turbine shaft is cantilevered from the aft bearing system with respect to the longitudinal axis. The disk assembly includes a metallic material.

Abstract: Aspects of the disclosure generally relate to a turbine engine and method of operating a turbine engine having an engine core including a compressor, combustor, and turbine in axial flow arrangement, whereby a working airflow passes through the engine core from the compressor to the turbine to define a flow direction through the engine core. The method includes generating a shockwave in the working airflow that propagates in the flow direction, and at least partially attenuating the shockwave.

Abstract: A turbojet engine including an intermediate casing including radial arms and a driveshaft for a gearbox for auxiliary machines of the turbojet engine, the driveshaft mounted in a radial arm, the radial arm including an intermediate bearing supporting the driveshaft, the intermediate bearing including plural rolling bearings supporting the driveshaft, the driveshaft including a first shaft element connected at one end to a mechanism providing mechanical transmission to an engine shaft of the turbojet engine, and a second shaft element connected at one end to a mechanism providing mechanical transmission to the box. The first shaft element and the second shaft element are linked by a connection including plural splines. A single plane transversely passes through the plural splines and the plural rolling bearings.

Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame.

Abstract: A differential gear system for a gas turbine engine includes an assembly of a ring gear, a sun gear, and a plurality of planet gears, and a planet carrier. The plurality of planet gears are enmeshed between the ring gear and the sun gear. The planet carrier acts as an input to rotatably drive the planet gears, the sun gear, and ring gear of the gas turbine engine.

Abstract: A tip turbine engine provides an axial compressor rotor that is counter-rotated relative to a fan. A planetary gearset couples rotation of a fan to an axial compressor rotor, such that the axial compressor rotor is driven by rotation of the fan in a rotational direction opposite that of the fan. By counter-rotating the axial compressor rotor, a final stage of compressor vanes between the final stage of compressor blades and inlets to the hollow fan blades of the fan are eliminated. As a result, the length of the axial compressor and the overall length of the tip turbine engine are decreased.

Abstract: An engine comprises a gas generator having an exhaust plenum and a propulsor comprising a propulsion fan coaxially mounted within an annular turbine. The annular turbine comprises a turbine duct and a plurality of turbine rotor blades rotationally coupled to the propulsion fan. A plurality of hollow struts extend axially and radially from the gas generator to the annular turbine. The hollow struts comprise flow ducts connecting the exhaust plenum to the turbine duct.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

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, and a fan drive gear system including a carrier for supporting a plurality of gears. A torque frame is attached to the carrier. A plurality of connectors extends between the carrier and torque frame for securing the torque frame to the carrier. A scupper captures lubricant during gear operation and directing lubricant into a space between at least one of the plurality of connectors and at least one of the torque frame and the carrier.

Abstract: A gas turbine apparatus may include one or more driving shafts; a compressor configured to receive power from the one or more driving shafts, suck air, and compress the sucked air at high pressure. The gas turbine may include a combustor configured to mix the compressed air with fuel, and combust the fuel mixture to generate combustion gas. A turbine is connected to the one or more driving shafts and is configured to rotate while the combustion gas passes through the turbine. The turbine may include one or more first turbine blades configured to rotate in a first direction and one or more second turbine blades configured to rotate in a second direction opposite to the first direction. The one or more first turbine blades and the one or more second turbine blades may be alternately arranged along an axial direction of the turbine.

Abstract: In a high pressure turbine used in combination with a contra-rotating low pressure turbine, in which the geometry of the inlet guide vane of the low pressure turbine induces a reflection of the incident shock wave from the high pressure turbine airfoil that propagates back upstream to impact the blade thereby denigrating both the performance of the high pressure turbine and its durability. The shock wave reflection is eliminated by fitting the downstream vanes with holes or indentations that allow jets of compressor bleed air to penetrate the boundary layer and induce a net velocity component across the main-flow direction in the near-wall region. This induced flow component acts to cancel the net flow component created by the moving shock from the upstream blade row thereby obviating the reflection that occurs naturally to cancel the flow.

Abstract: A turbine engine including two respectively upstream and downstream external impellers that are nonstreamlined, coaxial, and contrarotating is provided. The downstream impeller is retractable so as to reduce its diameter. The blades of the downstream impeller are mounted so as to pivot about a pivot, the axis of which forms a nonzero angle, notably perpendicular, with the axis of rotation of the impeller, the blades in the retracted position being tilted about the pivot.

Abstract: A fan-turbine rotor assembly includes one or more turbine ring rotors. Each turbine ring rotor is cast as a single integral annular ring. By forming the turbine as one or more rings, leakage between adjacent blade platforms is minimized which increases engine efficiency. Assembly of the turbine ring rotors to the diffuser ring includes axial installation and radial locking of each turbine ring rotor.

Abstract: A bypass aeronautical engine including a fan, an LP compressor, an HP compressor, a combustion chamber, an HP turbine, and at least one LP turbine stage is disclosed. The LP turbine stage includes a first movable bladed wheel carried by a first turbine shaft and a second movable bladed wheel turning in the opposite direction to the first wheel. The fan is driven by the second wheel via a second turbine shaft and by the first shaft via a module for reversing its direction of rotation. The reversing module includes two bevel pinions carried respectively by the first and second LP turbine shafts, and a reversing pinion.

Abstract: An exemplary gas turbine engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. The fan includes a first plurality of fan blades supported on a first fan support and a second plurality of fan blades supported on a second fan support. The first and second fan supports are rotatable about a fan axis independently of each other. A geared architecture includes at least one gear driven by the turbine section for rotating about a gear axis that is transverse to the fan axis. The gear drives the fan supports for rotating the respective pluralities of fan blades in opposite directions.

Abstract: Gas turbine engine systems involving tip fans are provided. In this regard, a representative gas turbine engine system includes: a multi-stage fan having a first rotatable set of blades and a second counter-rotatable set of blades, the first rotatable set of blades defining an inner fan and a tip fan; and an epicyclic differential gear assembly operative to receive a torque input and differentially apply the torque input to the first set of blades and the second set of blades.

Abstract: A compressor section includes a counter rotating low pressure compressor that includes outer and inner compressor blades interspersed with one another and are configured to rotate in an opposite direction than one another about an axis of rotation. A transmission couples at least one of the outer and inner compressor blades to a shaft. A turbine section includes a counter rotating low pressure turbine having an outer rotor that includes an outer set of turbine blades. An inner rotor has an inner set of turbine blades interspersed with the outer set of turbine blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system couples at least one of the outer and inner rotors to the shaft.

Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An inner rotor directly drives the shaft and includes an inner set of blades. An outer rotor has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system is engaged to the outer rotor and is positioned upstream of the inner set of blades.

Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame.

Abstract: An epicyclic gear train includes planetary gears rotatably mounted on spindles supported by an annular carrier and including axially spaced apart forward and aft sets of output teeth extending radially outwardly from a planetary gear hub and axially spaced apart forward and aft roller bearings disposed between planetary gears and spindles. The forward and aft roller bearings are axially aligned with or adjacent to spaced apart forward set of output teeth and input gear respectively. A ring gear meshes with forward set of output teeth and an external gear meshes with aft set of output teeth. An input gear fixedly attached to hub aft of aft set of output teeth and engaged with a sun gear. The output teeth, input gear, ring gear, external gear, and sun gear may all be helical. A turbofan gas turbine engine may include counter-rotatable first and second fan stages driven by a low pressure turbine through the gear train.

Abstract: An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a ratio of a number of stages in the low-pressure turbine to a number of stages in the high-pressure turbine is 2.

Abstract: Gas turbine engine systems involving tip fans are provided. In this regard, a representative gas turbine engine system includes: a multi-stage fan having a first rotatable set of blades and a second counter-rotatable set of blades, the second rotatable set of blades defining an inner fan and a tip fan and being located downstream of the first set of rotatable blades; and an epicyclic differential gear assembly operative to receive a torque input and differentially apply the torque input to the first set of blades and the second set of blades.

Abstract: A propulsion engine comprises a compressor, a combustor and a turbine arranged in flow series about a turbine axis. A shaft is coupled to the turbine along the turbine axis, and first and second propulsions fans are rotationally coupled to the turbine at opposite ends of the shaft. The first and second propulsion fans rotate substantially parallel to one another, and transversely with respect to the turbine axis.

Abstract: An aircraft gas turbine engine includes a counter-rotatable generator driven by a turbine includes a generator stator and counter-rotatable radially inner pole and outer magnet rotors. A gearbox may be used for counter-rotating the pole and magnet rotors. Counter-rotatable first and second sets of booster stages may be co-rotatable with corresponding ones of the pole and magnet rotors. The generator and the gearbox may be disposed within a booster cavity or a tail cone of the engine. The magnet rotor, the pole rotor, and the stator may be concentric. A particular embodiment of generator includes the magnet rotor encircling the pole rotor, the pole rotor encircling the stator, a rotor air gap between the magnet and pole rotors, and a transformer air gap between the pole rotor and the stator. The counter-rotatable generator may be driven directly by a counter-rotatable turbine.

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 two-stage turbofan system for use in a gas turbine engine comprises a first-stage fan shaft, a second-stage fan shaft, a stationary torque tube and a gear system. The second-stage fan shaft connects with a drive shaft in the gas turbine engine such that the second-stage fan shaft is driven at the speed of the drive shaft. The stationary torque tube is connected with a fan case in the gas turbine engine. The gear system is connected to the second-stage fan shaft and the torque tube. The first-stage fan shaft extends from the gear system such that the first-stage fan shaft is driven at a speed reduced from that of the drive shaft.

Abstract: A counter-rotatable low pressure turbine includes counter-rotatable outer and inner drum rotors. The outer drum rotor is connected to a sole shaft for transmitting torque and power out of the low pressure turbine. Low pressure outer drum turbine blade rows extend radially inwardly from an outer shell of the outer drum rotor. Low pressure inner drum turbine blade rows extend radially outwardly from the low pressure inner drum rotor. The outer drum turbine blade rows are interdigitated with the inner drum turbine blade rows. The drum rotors are geared together through an epicyclic gearbox for transmitting all the torque and power produced by the drum rotors to the shaft. The gearbox may be located aft of the drum rotors. A differential thrust bearing is disposed between the drum rotors. A single stage fan section of an engine is connected to the turbine by the shaft.

Abstract: A gas turbine aircraft engine with power variability is provided. The gas turbine aircraft engine comprises a compressor and a turbine mounted on a common shaft, and, a continuously variable transmission coupled to the shaft for transmitting power to a propulsion load.

Abstract: A representative gas turbine includes: a high pressure spool; a high pressure compressor; a high pressure turbine mechanically coupled to the high pressure spool; a lower pressure spool; a lower pressure turbine mechanically coupled to the lower pressure spool; a fan having a first set of fan blades and a second set of fan blades, the second set of fan blades being located downstream of the first set of fan blades and being operative to rotate at a rotational speed corresponding to a rotational speed of the lower pressure spool; and a gear assembly mechanically coupled to the lower pressure spool and engaging the first set of fan blades such that rotation of the lower pressure spool rotates the first set of fan blades at a lower rotational speed than the rotational speed of the second set of fan blades.

Abstract: A fan-turbine rotor assembly (24) includes one or more turbine ring rotors (32). Each turbine ring rotor is cast as a single integral annular ring. By forming the turbine as one or more rings, leakage between adjacent blade platforms is minimized which increases engine efficiency. Assembly of the turbine ring rotors to the diffuser ring (114) includes axial installation and radial locking of each turbine ring rotor.

Abstract: A gas turbine engine includes, in flow series, a generator section and a free power turbine. The generator section includes one or more generator turbine stages. One or more respective generator drive shafts extend axially forwardly from the generator turbine stages to one or more corresponding compressor stages. The free power turbine includes a first turbine stage and a contra-rotating second turbine stage. The gas turbine engine further includes a first drive shaft extending axially from the free power turbine to transmit rotational drive from the first turbine stage to a first propeller. The gas turbine engine further includes a second drive shaft extending from the free power turbine coaxially with the first drive shaft to transmit contra-rotational drive from the second turbine stage to a contra-rotating second propeller.

Abstract: A tip turbine engine (10) provides first and second turbines (32) rotatably driven by a combustor (30) generating a high-energy gas stream. The first turbine (32) is mounted at an outer periphery of a first fan (24a), such that the first fan is rotatably driven by the first turbine (32a). The second turbine (32b) is mounted at an outer periphery of a second fan (24b), and is rotatably driven by the high-energy gas stream. In one embodiment, the first turbine (32a) rotatably drives a plurality of stages of first compressor blades (54) in an axial compressor (22) in a first rotational direction, while the second turbine (32b) rotatably drives a plurality of stages of second compressor blades (52) in the axial compressor (22) in a second rotational direction opposite the first. By rotatably driving alternating stages of compressor blades in opposite directions, the efficiency of the axial compressor (22) is increased and/or the number of stages of compressor blades can be reduced.

Abstract: The present invention relates to a turbine engine comprising two respectively upstream and downstream external impellers (7, 9) that are nonstreamlined, coaxial and contrarotating. The engine is noteworthy in that the downstream impeller (7) is retractable so as to reduce its diameter. The blades (7a) of the downstream impeller are mounted so as to pivot about a pivot (71), the axis of which forms a nonzero angle, notably perpendicular, with the axis (2) of rotation of the impeller, the blades in the retracted position being tilted about the pivot (71).

Abstract: A twin-spool gas turbine engine including a high-pressure rotor and a low-pressure rotor, which rotors are mounted in bearings supported by an intermediate casing, at least one accessory gear box, and a drive device which drives radial and coaxial shafts for transmitting movement to the accessory gear box is disclosed. The drive device includes a high-pressure drive gear driving one of the radial shafts and connected to the high-pressure rotor and a low-pressure drive gear connected to the low-pressure rotor upstream of the high-pressure drive gear and driving the other of the radial shafts. The low-pressure drive gear is supported in the intermediate casing by an axial-positioning bearing.

Abstract: A gas turbine engine includes a gear system driven by a first and second counter rotating low pressure shaft and a fan driven by the gear system.

Abstract: A tip turbine engine assembly according to the present invention includes a load bearing engine support structure (12). The engine support structure (12) includes an engine support plane (P) that is substantially perpendicular to an engine centerline (A) and first rotationally fixed member (50) disposed about the engine centerline (A) and cantilevered from the engine support plane (P). A support member extends radially outward from the first rotationally fixed member (50) and structurally supports a second rotationally fixed member (58) that is coaxial with the first rotationally fixed member. A rotor is mounted on the first rotationally fixed member and rotates about the engine centerline (A).

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 counter-rotating blade stage in lieu of a stator stage may compensate for relatively low rotational speed of a gas turbine engine spool. A first spool may have at least two compressor blade stage and at least two turbine blade stage. A combustor is located between the at least two compressor blade stage and the at least two turbine blade stage along a core flowpath. The at least two counter-rotating compressor blade stage is interspersed with the first spool at least two compressor blade stage. A transmission couples the at least two additional compressor blade stage to the first spool for counter-rotation about the engine axis.

Abstract: A tip turbine engine (10) provides an axial compressor rotor (46) that is counter-rotated relative to a fan (24). A planetary gearset (90) couples rotation of a fan (46) to an axial compressor rotor (46), such that the axial compressor rotor (46) is driven by rotation of the fan in a rotational direction opposite that of the fan. By counter-rotating the axial compressor rotor, a final stage of compressor vanes (54) between the final stage of compressor blades (52) and inlets (66) to the hollow fan blades (28) of the fan are eliminated. As a result, the length of the axial compressor (22) and the overall length of the tip turbine engine (10) are decreased.

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 first fan assembly coupled to a low-pressure turbine, a second fan assembly disposed downstream from the first fan assembly, and a gearbox coupled between the second fan assembly and the low-pressure turbine. A method of assembling the above turbofan engine assembly is also described herein.

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 hydraulic stepper motor comprising three hydraulic ports each coupled to a piston that is actuatable between pressurised and depressurised configurations; and a contoured cam surface having at least one peak and trough and arranged to react against the pistons, the cam surface coupled to an output shaft; whereby actuation of one of the pistons moves the cam surface until a trough reacts against one or more pressurised pistons.

Abstract: With a two-shaft engine, the accessory drive gearbox (5) for the operation of a generator and other auxiliaries (6) is connected to the high-pressure shaft (1) via a high-pressure shaft gearbox (3), which is coupled to a low-pressure shaft gearbox (8) connecting to the low-pressure shaft (7), via an overrunning clutch (11), a conical-pulley variable transmission (10) settable on the basis of the shaft speed on both sides, and a first switchable clutch (9). Under certain operating conditions with low speed of the high-pressure shaft or in case of engine failure, the power of the low-pressure shaft, with the clutch (9) engaged, is transferred to the high-pressure shaft gearbox and hence to the accessory drive gearbox via the conical-pulley variable transmission, which compensates for shaft speed difference, and the overrunning clutch, in order to reliably further operate the generator and other accessories with low fuel consumption or by windmilling only.

Abstract: A fan for an aircraft engine, particularly a gas turbine aircraft engine, is disclosed. The fan has a fan rotor with fan blades, which extend radially outwardly from a hub to an outer flow path wall of a fan flow path. The fan has an auxiliary rotor that is positioned upstream of the fan rotor and has auxiliary blades that extend radially outwardly from a hub and end at a distance from the outer flow path wall of the fan flow path. The auxiliary rotor and the fan rotor are designed as separate rotors, such that the auxiliary rotor can be operated at a higher speed than the fan rotor.

Abstract: A gas turbine engine assembly is provided. The gas turbine engine assembly includes a core engine. The gas turbine engine assembly further includes a first rotor spool including a first fan assembly, an intermediate-pressure turbine, and a first shaft. The first fan assembly includes a single stage fan assembly coupled upstream from the core engine. The intermediate-pressure turbine includes a single stage turbine coupled downstream from the core engine. The first shaft extends between the first fan assembly and the intermediate-pressure turbine. The gas turbine engine assembly further includes a second rotor spool including a second fan assembly, a low-pressure turbine, and a second shaft. The second fan assembly includes a single stage fan assembly coupled upstream from the first fan assembly. The low-pressure turbine includes a single stage turbine coupled downstream from the intermediate-pressure turbine. The second shaft extends between the second fan assembly and the low-pressure turbine.

Abstract: A FLADE fan assembly includes radially inner and outer airfoils extending radially inwardly and outwardly respectively from an annular shroud circumferentially disposed about a centerline. Inner and outer chords extend between inner and outer leading and trailing edges of inner and outer airfoil cross-sections of the radially inner and outer airfoils respectively. Inner and outer stagger angles between the inner and outer chords respectively at the shroud and the centerline are different. The radially outer airfoils may outnumber the radially inner airfoils and particularly by a ratio in a range of 1.5:1 to about 4:1. Load paths or radii may extend radially through the inner and outer airfoils and through the rotating shroud between the inner and outer airfoils and may pass near or through the inner and outer leading edges and through the inner and outer trailing edges.

Abstract: A method for assembling a gas turbine engine includes coupling a gearbox to a low-pressure turbine, the gearbox includes a plurality of planetary gears intermeshed with the sun gear, each of the planetary gears includes a first gear portion having a first diameter and a second gear portion having a second diameter that is different than the first diameter.

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

Abstract: A gas turbine engine booster includes one or more rotatable booster stages having booster blades extending radially inwardly from a rotatable drum and one or more non-rotatable vane stages having booster vanes extending radially outwardly from a non-rotatable shell. The booster stages may be interdigitated with the vane stages. The booster may be co-rotatable or counter-rotatable with respect to a fan stage of an aircraft gas turbine engine. The booster may be driven by a single turbine or one of counter-rotatable turbines or though a gearbox by these turbines. The booster blades and the booster vanes extend across a core engine inlet duct having an entrance downstream of the first fan stage. A fan section with the booster may have counter-rotatable first and second fan stages with fan blades of the second fan stages connected to and mounted radially outwardly of the rotatable drum.