Abstract: A turbomachine engine according to aspects of the present disclosure is provided. The engine includes a fan assembly including a plurality of fan blades, and a core engine surrounded by an outer casing. The core engine includes a power output component operably connected to the fan assembly, a first input power source and a second input power source. The first input power source is counter-rotatable relative to the second input power source. The core engine includes a gear assembly operably connected to the power output component and configured to receive power from the first input power source and the second input power source.

Abstract: A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings.

Abstract: A differential gearbox assembly for a turbine engine having a fan shaft and a drive shaft. The differential gearbox assembly includes an epicyclic gear assembly coupling the fan shaft to the drive shaft. The epicyclic gear assembly includes a sun gear, a planet gear constrained by a planet carrier, and a ring gear. The sun gear is coupled to the drive shaft and the planet carrier is coupled to the fan shaft. The sun gear, the planet gear, and the ring gear all rotate about the drive shaft. The differential gearbox assembly includes an electric machine assembly that includes an input coupled to the epicyclic gear assembly. The electric machine assembly provides mechanical power to the fan shaft through the epicyclic gear assembly.

Abstract: A system and method for inspecting fan blade clearances in a fan stator module are presented. The method includes rotating a fan rotor assembly that includes an attached bracket situated between adjacent fan blades connected to the fan rotor assembly. The bracket includes a laser module and an inclinometer configured to measure a circumferential position of the laser module. The method further includes projecting a laser beam from the laser module to the fan case to determine a distance measurement between the laser module and the fan case at multiple circumferential points around the fan case. The distance measurement and circumferential position data is processed by a computing device to determine a clearance distance between a fan blade and the fan case.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil body extending between leading and trailing edges and extending from a root section, and the airfoil body defining pressure and suction sides. The airfoil body defines a recessed region extending inwardly from at least one of the pressure and suction sides, and the airfoil body includes one or more ribs that define a plurality of pockets within a perimeter of the recessed region. A plurality of cover skins is welded to the airfoil body along the one or more ribs to enclose respective ones of the plurality of pockets. The plurality of cover skins are formed from a common cover having a perimeter that is dimensioned to mate with the perimeter of the recess.

Abstract: A gas turbine engine comprises an engine core and an electric fan array. The engine core includes a compressor, a combustor, and a turbine. The compressor compresses and delivers air to the combustor. The combustor mixes fuel with the compressed air received from the compressor and ignites the fuel. The hot, high-pressure products of the combustion reaction in the combustor are directed into the turbine to cause the turbine to rotate about an axis and drive the compressor. The electric fan array may include at least two non-concentric fans having parallel shafts.

Abstract: An assembly for mounting an auxiliary component to an engine case of a gas turbine includes a support bracket, the support bracket having a first end configured for attachment to a first flange of the engine case, a second end configured for attachment to a second flange of the engine case, and an intermediate portion located intermediate the first end and the second end; a bearing member disposed within the intermediate portion; and a mechanical fuse disposed within the bearing member.

Abstract: Turbines for an aircraft include a supply device of compressed air configured to supply compressed air to the aircraft, a cooling system of the compressed air supplied to the aircraft, having a scoop configured to collect cooling air in a flow duct of a secondary flow, and a management system configured to be supplied with the cooling air collected from the flow duct and configured to control radial clearances between a turbine casing and turbine rotor vanes tips. The management system is supplied with cooling air by the scoop of the cooling system.

Abstract: A propulsion system includes a drive shaft, a fan, a fan shaft, and a reduction device coupling the drive and fan shafts. The reduction device includes a first reduction stage and a second reduction state, and include a sun gear, centered on an axis of rotation of the drive and fan shafts and driven in rotation by the drive shaft, a ring gear, coaxial with the sun gear and that drives the fan shaft in rotation about the axis, and planet gears distributed circumferentially about the axis between the sun and ring gears, each planet gear including a first portion which is meshed with the sun gear and a second portion which is meshed with the ring gear, a diameter of the first portion being different from a diameter of the second portion, and an oil transfer bearing positioned between the fan and the reduction device.

Abstract: A turbomachine having an air flow measurement system. The system comprises a plurality of acoustic sensors upstream of a compressor of the turbomachine. The acoustic sensors are configured to transmit acoustic waveforms through an intake airflow to further acoustic sensors. The acoustic sensors are located in a single plane and mounted at angular spacing about a central axis of an intake.

Abstract: A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft, and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.

Abstract: A cooling device (11) for a turbine of a turbomachine extending along an axis includes at least one radially inner metal sheet (14) and one radially outer metal sheet (15) that are joined to one another and delimit, between them, cooling air circulation channels (17) extending circumferentially from a connection region (16). Each channel (17) includes at least one air inlet and air ejection orifices (19), that are designed to be oriented toward a region to be cooled. The cooling device also includes at least one cooling duct (21) intended for the circulation of cooling air, the duct (21) located radially outside said metal sheets (14, 15) and close to or in contact with the metal sheets so as to cool said metal sheets using the cooling air circulating in the duct (21), the cooling duct (21) extending axially and arranged toward the circumferential end regions of the channels (17).

Abstract: A structural subassembly which has a bearing which comprises a statically arranged outer ring and a rotatably arranged inner ring, wherein the inner ring is connected for conjoint rotation to a component that is rotatable about a longitudinal axis or said inner ring forms part of such a component, and wherein the longitudinal axis defines an axial direction of the bearing. The structural subassembly furthermore comprises a housing flange of a support structure, to which flange the statically arranged outer ring is connected. Provision is made for the outer ring to be of two-part design, wherein each part of the outer ring has a connecting element which is connected to the housing flange, wherein the housing flange is arranged between the two connecting elements in the axial direction.

Abstract: A variable exhaust nozzle for use with a gas turbine engine includes an outer shroud and an inner plug that can move relative to the outer shroud. The relative movement of the inner plug and the outer shroud changes the shape of the variable exhaust nozzle from one that converges in area to one that converges and then diverges in area.

Abstract: The present disclosure is directed to a rotating detonation combustion system for a propulsion system including a plurality of combustors in adjacent arrangement along the circumferential direction. Each combustor defines a combustor centerline extended through each combustor, and each combustor comprises an outer wall defining a combustion chamber and a combustion inlet. Each combustion chamber is defined by an annular gap and a combustion chamber length together defining a volume of each combustion chamber. Each combustor defines a plurality of nozzle assemblies each disposed at the combustion inlet in adjacent arrangement around each combustor centerline. Each nozzle assembly defines a nozzle wall extended along a lengthwise direction, a nozzle inlet, a nozzle outlet, and a throat therebetween, and each nozzle assembly defines a converging-diverging nozzle. A first array of combustors defines a first volume and a second array of combustors defines a second volume different from the first volume.

Abstract: Aircraft propulsion systems include a fan shaft connected to a fan, the fan shaft defining a centerline axis of the aircraft propulsion system, one or more offset cores arranged at an angle to the centerline axis, the one or more offset cores each comprising a hydrogen burning combustor, a centerline cavity defined along the centerline axis, and a heat exchanger arranged within the centerline cavity. In operation, a portion of air is directed from the fan into the centerline cavity to provide a first working fluid to the heat exchanger within the centerline cavity.

Abstract: A case for a gas turbine engine includes a cast case section cast case section configured to be welded between a forward case section and an aft case section.

Abstract: An aircraft turbine engine includes a gas generator and a fan arranged upstream of the gas generator and configured to generate a main gas flow, of which one portion flows in a stream from the gas generator to form a primary flow, and of which another portion flows in a stream around the gas generator to form a secondary flow. The gas generator includes a low pressure body having a rotor driving the fan by means of a crown of a mechanical planetary reduction gear. The turbine engine further includes an electrical machine mounted coaxially around the reduction gear and having a rotor rotated by the crown of the reduction gear, and a stator extending around the rotor of the electrical machine.

Abstract: The gas turbine engine for an aircraft includes at least a low pressure spool with a low pressure turbine shaft operatively connected to at least one turbine, the low pressure turbine shaft rotatable about an engine axis, and a low pressure compressor operatively connected to a low pressure compressor shaft that is independently rotatable relative to the low pressure turbine shaft. A differential gearbox has an input operatively connected to the low pressure turbine shaft, a first output and a second output, the first output of the differential gearbox operatively connected to the low pressure compressor shaft and the second output of the differential gearbox operatively connected to an output shaft of the gas turbine engine. The differential gearbox permits the output shaft, the low pressure compressor shaft and the low pressure turbine shaft to rotate at different speeds.

Abstract: A blade for a turbomachine includes a blade root portion for securing the blade to a rotatable annular outer drum rotor. The blade root portion includes one or more radial retention features for radially retaining each of the blade root portions within the rotatable annular outer drum rotor. Further, at least one of the radial retention feature(s) includes at least one sealing member integrated therewith.

Abstract: A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A blade splitter (shroud) on the fan partially spans the fan blade thus separating the core and bypass streams downstream while leaving a gap upstream for communication between the flows. The communication gap expands the operational range of the fan over fans without the communication gap.

Abstract: A fan module with variable pitch blades for a longitudinal axis propulsion assembly. The module includes a rotor through which a rotor shaft runs along the longitudinal axis X and carries the blades of the fan. A system for changing the pitch of the blades includes a mechanism connected to the blades of the fan and a controller acting on the mechanism. The controller has a fixed body and a movable body translatable along the longitudinal axis X relative to the fixed body, and a load transfer bearing arranged between the mechanism and the controller. The fan module includes a power shaft driving the rotor shaft via an epicyclic train speed reducer, the speed reducer including a sun gear connected to the power shaft and a satellite carrier including a fixed annular ferrule secured to a fixed casing. The fixed body of the controller is mounted on the annular ferrule.

Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a low speed spool and a high speed spool; a rotor assembly coupled to the low speed spool; an electric machine rotatable with the low speed spool for extracting power from the low speed spool, for adding power to the low speed spool, or both; and an inter-spool clutch positioned between the low speed spool and the high speed spool for selectively coupling the low speed spool to the high speed spool.

Abstract: An assembly is provided for a turbine engine. This assembly includes a primary duct, a bleed duct, a plurality of secondary ducts, a heat exchanger and a flow regulator. The bleed duct extends from a bleed duct inlet to a bleed duct outlet. The bleed duct inlet is fluidly coupled with the primary duct. The secondary ducts are arranged in parallel between the bleed duct outlet and the primary duct. The secondary ducts include a first duct and a second duct. The heat exchanger is configured with the second duct. The flow regulator is configured to direct at least a majority of fluid flowing through the bleed duct outlet to: (A) the first duct during a first mode of operation; and (B) the second duct during a second mode of operation.

Abstract: Systems, methods, lift fans, and aircraft involving active flow control of a ducted fan or fan-in-wing configuration are described.

Abstract: A system includes a gas turbine engine having a low speed spool and a high speed spool. The system also includes a spool coupling system configured to mechanically link the low speed spool and the high speed spool. A controller is operable to determine a mode of operation of the gas turbine engine, monitor for a spool coupling activation condition associated with the mode of operation, and activate the spool coupling system based on the controller detecting the spool coupling activation condition. Engagement and power transfer between the low speed spool and the high speed spool occurs based on activation of the spool coupling system and reaching an engagement condition of the spool coupling system.

Abstract: A rotor blade for a gas turbine engine includes an airfoil section and a root section extending along a longitudinal direction between an upstream surface and a downstream surface. The root section further extends along a radial direction between an inner surface positioned at an inner end of the root section and an outer end coupled to the airfoil section. Moreover, the root section extends along a circumferential direction between a first side surface and a second side surface. Furthermore, the root section defines a longitudinal centerline extending along the longitudinal direction and positioned equidistant from the inner surface and the outer end in the radial direction. The root section includes a first portion formed from a composite material and a second portion formed from a metallic material, with the longitudinal centerline extending through the second portion of the root section.

Abstract: A turbine engine heat exchanger has an array of heat exchanger plates mounted to an inner case wall for providing heat transfer from a bleed flowpath to a bypass flowpath. Each plate has: first and second faces along the bypass flowpath; a proximal edge mounted to the inner case wall; an inlet along the proximal edge; an outlet along the proximal edge; and a branch segment of the bleed flowpath passing from the inlet to the outlet.

Abstract: A gas turbine engine is disclosed. In various embodiments, the gas turbine engine includes a high pressure turbine having a first stage and a second stage, the first stage connected to a high pressure turbine first stage spool and the second stage connected to a high pressure turbine second stage spool; a high pressure compressor connected to a high pressure compressor spool; and a differential system having a first stage input gear connected to the high pressure turbine first stage spool, a second stage input gear connected to the high pressure turbine second stage spool and an output gear connected to the high pressure compressor spool.

Abstract: A lubrication system may comprise a shaft including a shaft gear, a pump configured to supply a flow of lubricant to a component benefiting from lubrication and an input gear train coupled to the shaft gear and configured to drive the gear pump with a unidirectional rotation in response to a forward rotation and an reverse rotation of the shaft gear. The input gear grain may comprise a primary gear coupled to a common shaft. The input gear train may further comprise a forward shaft and a forward gear about the common shaft, wherein the forward gear is coupled to the forward shaft.

Abstract: A turbomachine includes a rotatable annular outer drum rotor connected to a first plurality of blades. The rotatable annular outer drum rotor is constructed of, at least, a first drum segment and a second drum segment. The turbomachine further includes a retaining ring arranged and secured between the first and second drum segments of the rotatable annular outer drum rotor for radially retaining each of the first plurality of blades via their respective blade root portions within the rotatable annular outer drum rotor.

Abstract: Fairing installations disclosed herein may include a damper for mitigating vibration of a cantilevered fairing disposed in a bypass duct of a gas turbine engine. The bypass duct may include a first shroud radially spaced apart from a second shroud to define a bypass passage between the first and second shrouds. The fairing may be disposed in the bypass passage and cantilevered from the first shroud. The fairing may have a secured end secured to the first shroud and a free end proximate the second shroud. The damper may be engaged with the free end of the fairing to damp movement of the free end of the fairing.

Abstract: A turbomachine engine according to aspects of the present disclosure is provided. The engine includes a fan assembly including a plurality of fan blades, and a core engine surrounded by an outer casing. The core engine includes a power output component operably connected to the fan assembly, a first input power source and a second input power source. The first input power source is counter-rotatable relative to the second input power source. The core engine includes a gear assembly operably connected to the power output component and configured to receive power from the first input power source and the second input power source.

Abstract: A gas turbine engine includes a combustor section arranged between a compressor section and a turbine section. The compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged upstream of the combustor section. A fan section has an array of twenty-six or fewer fan blades. The low pressure compressor is downstream from the fan section. An airfoil is arranged in the low pressure compressor and includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between an axial stacking offset and span position that includes a curve with a negative slope from 90% span to 100% span. The negative slope leans forward relative to an engine axis.

Abstract: A mechanical reduction gear of a turbine engine, in particular for an aircraft, includes a sun gear having external toothing, a ring gear which extends around the sun gear and has internal toothing, planet gears which are in mesh with the sun gear and the ring gear and which have toothings of different diameters, a planet carrier which supports bearings for guiding the planet gears in rotation, and a lubrication circuit of the reduction gear. The planet carrier has a hydrodynamic abutment on which the sun gear is configured to bear axially.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a tip speed of from 1 to 3 times that of the outer fan.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The inner fan is configured to have, in operation, a rate of rotation of from 3 to 8 times that of the outer fan.

Abstract: Intercooling systems and methods for an aircraft engine are provided. An intercooling system includes: a first inlet configured to receive a first air flow of ambient air into the aircraft engine; a second inlet separate from the first inlet and configured to receive a second air flow of ambient air into the aircraft engine separately from the first air flow of ambient air; and a heat exchanger configured to facilitate heat transfer between at least a portion of the first air flow compressed by a compressor section of the aircraft engine and the second air flow.

Abstract: A turbofan gas turbine engine includes a fan section having a fan, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture including an epicyclic gear train, a turbine section including a low pressure turbine and a high pressure turbine, the fan driven by the low pressure turbine through the geared architecture, and a power density between 4.84 lbf/in3 and 5.5 lbf/in3.

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: An active laminar flow control arrangement may comprise a variable speed constant frequency (VSCF) converter comprising a housing, a compressor, an electric motor operably coupled to the compressor, and a laminar flow control duct. An airflow is received by the housing from the laminar flow control duct in response to the electric motor driving the compressor for cooling the VSCF converter.

Abstract: The invention relates to a module of an aircraft bypass engine including an annular flow path of a secondary flow of gas wherein is mounted an annular row of stator blades configured to straighten the gas flow, an arm extending in the flow path downstream from said annular row of blades and being connected physically to one of the stator blades so as to form a unitary aerodynamic assembly, wherein at least one of the two blades adjacent to the blade of the unitary aerodynamic assembly is of a variable pitch type and capable of being rotationally moved about a substantially radial axis with respect to a longitudinal axis of the engine.

Abstract: A propulsion unit for an aircraft including a nacelle with a D-shaped structure housing a thrust-reversing device with movable vanes, the nacelle containing two D-shaped half-structures each including an external reverser half-cowl. The propulsion unit contains an assembly box-type structure, attached in downstream cantilevered fashion to the turbojet engine fan casing, the box-type structure including two guide rails guiding the deflection vanes, a locking device between the assembly box-type structure and the half-beams in the six o'clock position of the D-shaped half-structures, the assembly box-type structure being arranged in the propulsion unit such that the guide rails guiding the vanes of the assembly box-type structure are situated in the continuation of the guide rails in the six o'clock position guiding the vanes attached to the fan casing, to provide continuity between the rails.

Abstract: A frangible strut for a gas turbine engine includes a first end to couple to a first structure of the gas turbine engine, and a second end opposite the first end. The second end is to couple to a second structure of the gas turbine engine. The second end includes an outer portion coaxial with and spaced radially apart from an inner portion. The inner portion is to receive a fastening device to couple the second end to the second structure. The outer portion is coupled to a body of the frangible strut. The inner portion is interconnected to the outer portion by a plurality of frangible members that are spaced apart about a perimeter of the inner portion, and each of the frangible members are configured to release the inner portion from the outer portion. The frangible strut includes the body interconnecting the first end and the second end.

Abstract: The invention relates to a jet engine with a fixed housing in which a primary flow is formed in which incoming air is burned in at least one combustion chamber, in said housing a secondary flow being formed in which incoming air is accelerated by a fan and, said secondary flow being expelled at the outlet cone of the housing together with the exhaust gas from the combustion chamber, said fan being mounted on a main shaft rotatably about an axis and having a plurality of substantially radially-extending fan blades.

Abstract: A turbojet engine (2), such as an aircraft turbojet engine, comprising: an epicyclic gear train (36); a turbine rotating a transmission shaft (34) constrained to rotate with the inner planetary gear (60) of the epicyclic gear train (36); a fan (18) rigidly connected to the ring gear (66) of the epicyclic gear train (36); and an electric machine (70) comprising a rotor (72) and a stator (74), the rotor (72) being rigidly connected to the planet carrier (68) of the epicyclic gear train (36). Alternatively, the fan (18) is rigidly connected to the planet carrier and the electric machine (70) is rigidly connected to the ring gear. The invention also relates to methods for using said turbojet engine, in particular for controlling the reduction ratio between the transmission shaft and the turbine, in order to recover kinetic energy or for taxiing.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. The outer fan has a diameter of from 2.5 to 3.5 times a diameter of the inner fan.

Abstract: Conventional commercial engine exhaust systems are defined with axi-symmetric surfaces (e.g., conical or nearly conical surfaces), which create an annular exhaust for the fan (bypass) nozzle of roughly constant duct-height around the circumference. In one example configuration, the fan sleeve has been sheared upward (towards the wing or pylon) causing a larger area and duct height near the pylon relative to the portion away from the pylon. For a given thrust generated by the turbofan engine housed in the nacelle, the shear toward the pylon mount realigns the thrust in the direction of flight which may, in some examples, reduce noise experienced downstream of the turbofan engine and decreases fuel consumed in the engine core.

Abstract: An axial flow turbomachine (102) for producing thrust to propel an aircraft is shown. The turbomachine has an inner duct (202) and an outer duct (204), both of which are annular and concentric with one another. An inner fan (206) is located in the inner duct, and is configured to produce a primary pressurised flow (P). An outer fan (207) is located in an outer duct, and is configured to produce a secondary pressurised flow (S). The outer fan has a hollow hub (208) through which the inner duct passes. A swept area of the outer fan is from 2 to 20 times greater than a swept area of the inner fan.

Abstract: An anterior part of a nacelle of an aircraft propulsion unit having an air inlet lip at the front end, an internal structure and an external panel extending the air inlet lip, and an annular rigidifying frame having an external peripheral edge connected to the external panel is disclosed. The anterior part of the nacelle includes a shock-absorbing element connected rigidly on the one hand to an internal peripheral edge of the rigidifying frame and on the other hand to the internal structure.