Abstract: Disclosed is a method of monitoring relative blade pitch angle alignment of a set of at least two rotor blades in a rotary device. The rotary device may be a wind turbine generator. Also disclosed are operational schemes based on the observed relative blade pitch angle alignments.

Abstract: A high rise building escape drone is shown and described. The high rise building escape drone includes of a frame. The frame has a spine secured to an upper housing. The upper housing secures a motor and a CPU operably connected to the motor. The motor is rotatably coupled to at least one propeller. A seat is secured to the spine below the upper housing. A control panel is secured to the seat. The control panel is operably coupled to the CPU and is capable to control the drone. A plurality of feet are secured to a bottom of the spine such that the feet support the drone.

Abstract: A control system and methods for controlling the position of one or more engine-mounting links of an engine-mounting linkage system are provided. In one aspect, an engine-mounting linkage system includes one or more engine-mounting links that each have an adjustable inclination angle. An inclination angle of a link may be adjusted by an actuator of the control system. One or more controllers of the control system can control the actuator and thus the inclination angle of the link by determining a control command based at least in part on an output received from one or more sensors of the control system. The controllers can then cause the actuator to change the inclination angle of the link based at least in part on the determined control command.

Abstract: A gas turbine engine including: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order; a fan defining a fan axis and comprising a plurality of fan blades rotatable about the fan axis; and a pitch change mechanism operable with the plurality of fan blades, the pitch change mechanism including a plurality of linkages, the plurality of linkages including a first linkage coupled to a first fan blade of the plurality of fan blades and a second linkage coupled to a second fan blade of the plurality of fan blades; and a non-uniform blade actuator system operable with one or more of the plurality of linkages to control a pitch of the first fan blade relative to a pitch of the second fan blade.

Abstract: A drone or an aircraft for personal air mobility includes at least one horizontal planar structure having four ducted propeller rotors with vertical axis, which are substantially coplanar with each other. Each of the rotors has a rotating ring rotatably mounted within a circular opening. The rotating ring is configured as to define an annular wall for ducting of air flow produced by the rotor. Each of the rotors has one or more blades which extend towards the central axis of the rotor, and have tips terminating at a distance from the axis of the propeller, such that each rotor is an annular propeller. The rotating ring of each rotor is controlled by an actuator.

Abstract: Provided is an unmanned aircraft which is capable of stably taking off. In one example, the unmanned aircraft has an aircraft body which includes an information acquisition device, a plurality of rotary wings, and a protective member disposed around the rotary wings. Each of the rotary wings has a rotation axis which is tilted with respect to a vertical direction by a given angle, such that the rotary wings generate a forward thrust force. The aircraft body has a lower edge which includes a middle lower edge and a forward lower edge, wherein the protective member has a lower edge including the forward lower edge. The forward lower edge is located above the middle lower edge, and the forward lower edge has a rear end located backward of a rear end of the rotary wing.

Abstract: A multi-rotor system for providing air thrust is disclosed comprising at least one multi-rotor assembly. The multi-rotor assembly comprising at least two rotors rotatable about a common axis wherein the outer radius of a first rotor is substantially similar to the inner radius of the second rotor. An airborne vehicle is also disclosed that is adapted to perform vertical takeoff and landing (VTOL). The airborne vehicle comprising at least two multi-rotor system disposed substantially symmetrically around the center of gravity of the vehicle.

Abstract: A gas turbine engine includes a fan assembly having a plurality of fan blades; and a turbomachine. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbomachine further including a first input power source; a second input power source configured to counter-rotate relative to the first input power source; a power output component operably connected to the fan assembly; and a gear assembly located forward of the combustion section of the turbomachine, the gear assembly configured to receive power from the first input power source and the second input power source and provide power to the power output component, the gear assembly comprising a helical gear.

Abstract: A variable pitch fan assembly includes a plurality of rotating trunnion assemblies, a plurality of counterweight assemblies, a first unison ring gear engaged with the trunnion assemblies, and a second unison ring member that restricts out of synch movement of the counterweights relative to each other.

Abstract: A transmission having a housing (18), at least a first input (20) rotatable about a first input axis (IA1), and at least a first output (26) rotatable about a first output axis (OA1). The transmission further includes a first outer side gear (14) connected to and drivable by the first input, a first planet gear set comprising a first inner planet gear (11) and a first outer planet gear (15) with the inner and outer planet gears being rigidly connected to, and in axial alignment with, one another. The first planet gear set is rotatable (22, 24) via the first outer side gear and is also rotatable (23) about the first output axis. The transmission further includes a second outer side gear (16) in mesh with the first outer planet gear, and a second inner side gear (12) connected to the first output and being rotatable about the first output axis via the first inner planet gear.

Abstract: A method of operating an aircraft engine coupled to a wing of an aircraft including: setting a pitch of a plurality of rotor blades of a rotor assembly of the aircraft engine at non-uniform pitch angles along a circumferential direction of the aircraft engine such that the plurality of rotor blades define a first pitch at a first position and a second pitch at a second position, wherein the second position is 180 degrees offset from the first position, and wherein the first pitch is different from the second pitch.

Abstract: An electric propulsion system for a vertical takeoff-and-landing aircraft having an inverter assembly, a gearbox assembly, an electric motor assembly. The inverter assembly, the gearbox assembly, and the electric motor assembly may be substantially aligned along an axis. The inverter assembly, a gearbox assembly, and electric motor assembly may each abut at least one of the others. The electric engine may use a liquid, such as oil, for cooling and lubricating components of the inverter assembly, gearbox assembly, and electric motor assembly. Further, the electric engine may use volumes of oil for cooling and lubricating components below a threshold volume so that the electric engines do not require a fire protective barrier.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly includes a plurality of fan blades defining a fan diameter, and the heat exchanger module is in fluid communication with the fan assembly by an inlet duct. The heat exchanger module includes a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. At full-power condition, the engine produces a maximum thrust T (N), the heat exchanger module transfers a maximum heat rejection H (W) from the first fluid to the airflow, and a Heat Exchanger Performance parameter PEX (W/N) defined as PEX=H/T is 0.4 to 6.0.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a propulsor section, a geared architecture, a high spool and a low spool. The high spool includes a high pressure compressor and a high pressure turbine. The low spool includes a low pressure compressor and a low pressure turbine. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to 1.55. A mechanical tip rotational Mach number of the blades is greater than or equal to 0.5 at an approach speed.

Abstract: A turbomachine, a computing system for a turbomachine, and a method for overspeed protection are provided. The turbomachine includes a first rotor assembly interdigitated with a second rotor assembly together operably coupled to a gear assembly. A plurality of sensors is configured to receive rotor state data indicative of one or more of a speed, geometric dimension, or capacitance, or change thereof, or rate of change thereof, relative to the first rotor assembly or the second rotor assembly. A controller executes operations including receiving rotor state data from the plurality of sensors; comparing rotor state data to one or more rotor state limits; and contacting one or more of the first rotor assembly or the second rotor assembly to a contact surface adjacent to the respective first rotor assembly or the second rotor assembly if the rotor state data exceeds the rotor state limit.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.

Abstract: The present disclosure is directed to providing impeller type oil transportation device including; a transport unit provided such that a mixed fluid including an oil is fed on one side; and an impeller provided in the transport unit, the impeller including a core connected to a rotation axis, and wings extending radially from the core and having hydrophilic teeth on an outer surface to transport the mixed fluid including the oil to the other side of the transport unit by rotation, wherein the impeller is provided in the transport unit such that parts of the wings are exposed above a surface of the mixed fluid, to separate the oil adhered to the teeth while the mixed fluid is fed into a space between the adjacent teeth by capillary flow when the wings exposed above the surface of the mixed fluid move on to the mixed fluid by the rotation.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

Abstract: Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor.

Abstract: A blade pitch control system includes a plurality of serially stacked swashplate assemblies, each having concentric, ring-shaped inner and outer sections, an associated output pitch link coupled to its outer section and an associated input pitch link coupled to its inner section. The inner and outer sections of each swashplate assembly includes pass through holes to accommodate input pitch links and output pitch links of adjacent ones of the stacked swashplate assemblies. The system also includes a plurality of actuators, each coupled to a respective input pitch link of a respective one of the stacked swashplate assemblies. A central static mast accommodates a rotor drive shaft and the stacked swashplate assemblies are configured to slide axially, parallel to a long axis of the static mast.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas pat

Abstract: A method is for making a marine drive for propelling a marine vessel in water. The method includes providing a gearcase; installing a propeller shaft assembly that extends forwardly from the gearcase; coupling front and rear propellers to the propeller shaft assembly, forwardly of the gearcase, such that rotation of the propeller shaft assembly causes rotation of the front and rear propellers, respectively, which thereby propels the marine vessel in the water; and reducing deleterious effects of cavitation on the gearcase by the combination of forming the gearcase with a wide trailing end portion, in particular to maintain pressure alongside the gearcase, and configuring the front and rear propellers so that the front propeller absorbs more torque/thrust load than the rear propeller during said rotation.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub. A blade blockage, which is the ratio of the blade thickness to the product of the circumferential pitch and the cosine of a blade inlet angle (t/s·cos?1), is 0.25 or greater at the 0 percent span position.

Abstract: A coaxial rotor system for a rotorcraft includes a mast, a top rotor assembly and a bottom rotor assembly. The top rotor assembly is coupled to the distal end of the mast. The bottom rotor assembly includes a motor configured to provide rotational energy to the mast, thereby rotating the top rotor assembly. The bottom rotor assembly experiences a torque reaction force responsive to the motor rotating the mast such that the top and bottom rotor assemblies counter rotate.

Abstract: An aircraft propulsion system comprises first and second co-axial propulsors, one of the first and second propulsor being positioned forward of the other propulsor. A first electric motor is configured to drive the first propulsor, and a second electric motor is configured to drive the second propulsor. The first electric motor comprising a rotor radially inwardly of the stator, and the second electric motor comprises a rotor radially outwardly of the stator. The stator of the first electric motor is mounted to the stator of the second electric motor.

Abstract: A turbine engine includes integrated electric machines in the compressor section and the turbine section to supplement power produced from fuel with electric power. The compressor section includes a compressor electric motor that is electrically coupled to a compressor generator. The example turbine section includes a turbine electric motor that is coupled to a geared architecture to supplement power driving the fan section. A turbine generator driven by a portion of the turbine section provides electric power to the turbine electric motor.

Abstract: An unducted thrust producing system has a rotating element with an axis of rotation and a stationary element. The rotating element includes a plurality of blades, each having a blade root proximal to the axis, a blade tip remote from the axis, and a blade span measured between the blade root and the blade tip. The rotating element has a load distribution such that at any location between the blade root and 30% span the value of ?RCu in the air stream is greater than or equal to 60% of the peak ?RCu in the air stream.

Abstract: A mechanical system for transmitting rotary motion between firstly at least one inlet wheel driven in rotation by at least one engine, each inlet wheel being suitable for being driven in rotation by a corresponding dedicated engine, and secondly at least two distinct driven shafts of an accessory box, the at least one engine serving to drive rotation of one of the at least two driven shafts in isolation, or to drive the at least two driven shafts simultaneously, jointly with at least one aircraft rotor, the mechanical system including a selective rotary drive member that is movable axially between two distinct drive positions.

Abstract: Embodiments of the invention relate to advanced propeller assemblies that include a pressure modifying element that modifies its aerodynamic characteristics. Embodiments of the invention may provide benefits including but not limited to: increased thrust per rotation, greater efficiency, modified aerodynamics, and the ability to use such propellers for novel applications. Practical applications for embodiments of propellers of the invention include use with aircraft, drones, watercraft, autogyros, fans, blowers, pumps, generators, compressors, mixers, and blenders.

Abstract: A gear assembly for a turbo machine is provided. The gear assembly includes a first input shaft, a second input shaft, an output shaft, a fixed member, and a spindle. The spindle is extended along a spindle centerline axis. The first input shaft is drivingly connected to the spindle at a first interface. The second input shaft is drivingly connected to the spindle at a second interface. The spindle is connected to the fixed member providing a reactive force at a third interface. The spindle is connected to the output shaft at a fourth interface.

Abstract: A boundary layer ingestion engine includes a fan section configured to extend into a boundary layer of a full annulus of an aft end of a fuselage of an aircraft. The fan section includes a first fan stage and a second fan stage. The boundary layer ingestion engine also includes a differential planetary gear system is operable to transform rotation of an input shaft into counter rotation of a first shaft coupled to the first fan stage and a second shaft coupled to the second fan stage. The boundary layer ingestion engine further includes a motor operable to drive rotation of the input shaft.

Abstract: A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

Abstract: A strut mounted gear box for counter rotating propellers. The gear box is strut mounted for securement to the hull of a boat. A main input shaft is coupled to a propulsion component of a boat with a distal end secured to an idler gear cage assembly located within the gear box. The main input shaft transfers torque and rotation from the propulsion component to an idler gear cage assembly. An inner tail shaft is coupled to the main input shaft and arranged to rotate the inner tail shaft in a first direction. A counter shaft is coupled to the idler gear cage assembly and arranged to rotate the counter shaft in a second direction. A first propeller is secured to the inner tail shaft providing rotation in the first direction; and a second propeller is secured to the counter shaft allowing rotation in the second direction.

Abstract: A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing.

Abstract: The present disclosure is directed to a gas turbine engine including a turbine section including a first rotating component interdigitated along a longitudinal direction with a second rotating component. The first rotating component and the second rotating component are each coupled to a speed reduction assembly in counter-rotating arrangement. The first rotating component comprising an outer shroud and a plurality of outer shroud airfoils extended inward along a radial direction from the outer shroud. A connecting member couples the outer shroud to a radially extended first rotor. The second rotating component comprising an inner shroud and a plurality of inner shroud airfoils extended outward along the radial direction from the inner shroud, the plurality of inner shroud airfoils in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils.

Abstract: The present invention relates to a device for controlling the yaw of an aviation device, such as a helicopter, said aviation device comprising a bearing structure and a rotating member connected to the bearing structure to be mobile in rotation, around an axis of rotation, relative to said bearing structure, wherein the rotating member comprises fixing means for fixing at least one blade, the yaw control device comprising a rotor and a stator which form, in combination, an electrical machine, wherein the bearing structure is connected to the first of this stator and this rotor, and wherein the rotating member is connected to the second of this stator and this rotor, wherein the electrical machine is suitable for generating an electromotive force applied to the rotating member.

Abstract: A turbomachine includes a turbine section including a turbine center frame and a turbine, the turbine including a plurality of low-speed turbine rotor blades and a plurality of high-speed turbine rotor blades alternatingly spaced along the axial direction. The plurality of low-speed turbine rotor blades each extend between a radial inner end and a radial outer end. At least two of the plurality of low-speed turbine rotor blades are spaced from one another along the axial direction and coupled to one another at the radial outer ends. The turbomachine also includes a gearbox aligned with, or positioned forward of, a midpoint of the turbine. The gearbox includes a first gear coupled to the plurality of low-speed turbine rotor blades, a second gear coupled to the plurality of high-speed turbine rotor blades, and a third gear coupled to the turbine center frame.

Abstract: A dual mode turbofan engine includes a jet engine portion having a compressor, a turbine disposed aft of the compressor, and a shaft coupled to the compressor and the turbine. The jet engine portion is configured to produce an exhaust. The system further includes an auxiliary turbine having a plurality of auxiliary turbine blades. The auxiliary turbine is disposed aft of the turbine and decoupled from the shaft. The system also includes a diverter disposed between the turbine and the auxiliary turbine. The diverter is configured to selectively direct the exhaust to an inner flow path bypassing the plurality of auxiliary turbine blades or to an outer flow path engaging the plurality of turbine blades. A plurality of propeller blades is hingedly connected to the auxiliary turbine.

Abstract: A gear system for a gas turbine engine includes a planet gear system which includes an output attached to a carrier for rotating a first fan assembly in a first direction. A star gear system includes an output attached to a ring gear for rotating a second fan assembly in a second direction. A sun gear of the star gear system is mechanically attached to a sun gear of the planet gear system.

Abstract: An aircraft comprising a fuselage and propelled by a turbine engine having two coaxial and contrarotating fans, the turbine engine comprising a power turbine having two contrarotating rotors, one of which drives a fan upstream from the turbine, the other a fan downstream from the turbine, each fan comprising a ring of blades, and the assembly of the fans and the power turbine being incorporated at the rear of the fuselage, in the extension of same. The aircraft comprises, for at least one of the fans, a device for blocking the rotation of the fan and a device configured to modify the pitch of the blades of the fan in such a way as to make it operate as a flow straightener with respect to the other fan.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a low pressure (LP) turbine defining an outer flowpath. The outer flowpath defines a first outer flowpath radius at an upstream-most end of the LP turbine, a last outer flowpath radius disposed at a downstream-most end of the LP turbine, a middle outer flowpath radius disposed therebetween along the longitudinal direction. The middle outer flowpath radius is greater than the last outer flowpath radius.

Abstract: A pitch actuation system for a turbine engine propeller, comprising an actuator, a movable part of which is designed to be connected to blades of the propeller so as to rotate said blades relative to blade pitch-setting axes, wherein the actuator is an electromechanical actuator and comprises at least two electric motors for driving a common rotor, and a transmission screw rotated by the common rotor, and in that the system further comprises a nut, through which the transmission screw passes and which is designed to cooperate with the blades so as to move them.

Abstract: A rotor hub assembly is provided including an open rotor hub having a base plate and a plurality of rotor members mounted about a circumference of the base plate. The rotor members are configured to form a plurality of rotor blade openings between adjacent rotor members, and a central opening between all of the rotor members. A plate includes a central body portion and a plurality of arms. Each of the plurality of arms is configured to mount to one of the plurality of rotor members such that the central body portion is generally positioned within the central opening of the open rotor hub.

Abstract: A fan system includes a rotor having plurality of blades and a ring airfoil, the plurality of blades being rotatably joined to a hub and the ring airfoil. The fan system may include a second contra-rotationally disposed rotor having a plurality of blades and a ring airfoil. The first and second ring airfoils having a cambered shape and an angle of attack between about ?5 degrees and about 45 degrees, more preferably between about 5 degrees and about 30 degrees. Optionally, an outlet guide vane may be mounted rearward of the one or more rotors having a ring airfoil.

Abstract: A gas turbine engine includes first and second shafts rotatable about a common axis. A first turbine section is supported on the first shaft. Second compressor and turbine sections are supported on the second shaft. The gas turbine engine includes a fan. A first compressor section is arranged in an axial flow relationship with the second compressor and the first and second turbines. A geared architecture operatively connects the first shaft and the fan. An inducer operative couples to the gear train.

Abstract: A drive mechanism (10) for driving a first adjustment member (12) for adjusting the orientation of the blades (14) of a first turbomachine rectifier stage (16) and a second adjustment member (18) for adjusting the orientation of the blades (20) of a second turbomachine rectifier stage (22), which comprises a mechanism for simultaneously moving the two adjustment members (12, 18) in the turbomachine, characterised in that it comprises a single drive wheel (24) that simultaneously drives the first adjustment member (12) and the second adjustment member (18) and comprises two gear stages (26, 28) that are arranged between the drive wheel (24) and one or the other of the first adjustment member (12) and the second adjustment member (18) and that have different transmission ratios.

Abstract: A control system for controlling the pitch of a propeller, the system including a propeller shaft, a blade swivel device having a rotary control element suitable for placing the blades in an angular position corresponding to a desired propeller pitch, and a transmission presenting an outlet member coupled in rotation with the rotary control element of the blade swivel device. The transmission includes a variable speed drive having drive, control, and outlet rotors that are coupled in rotation respectively with the propeller shaft, with a control member, and with the outlet member of the transmission. With the variable speed drive, the speed of rotation of the outlet member of the transmission is a predetermined function of the speeds of rotation not only of the propeller shaft, but also of the control member.

Abstract: An actuation system is provided for varying the pitch of the blades of a variable pitch fan or propeller. The actuation system includes a first actuator and a second actuator. The actuation system further includes a first linkage which operably connects the first actuator to the blades such that operation of the first actuator varies the pitch of the blades via the first linkage. The actuation system further includes a second linkage which operably connects the second actuator to the first actuator such that operation of the second actuator varies the pitch of the blades by series connection of the second linkage to the first linkage.

Abstract: An airplane unducted propeller turboshaft engine having a gas generator and a receiver including a propulsion assembly carrying least one propeller, the engine including a first casing, a second casing, and a third casing, the third casing being provided between the first and second casings and surrounding at least a portion of the gas generator, a reinforcing shell presenting a first attachment zone mounted on the first casing, second attachment zone mounted on the second casing, and a wall provided between the first and second attachment zones and surrounding the third casing, wherein the reinforcing shell further includes at least one pipe segment integrated in the wall.