A speed reducer includes two intermediate lines for a turboprop engine, the reducer including: an input line including an input shaft bearing an input gear wheel, a first intermediate line including a first intermediate shaft bearing a first intermediate gear wheel, a second intermediate line including a second intermediate shaft bearing a second intermediate gear wheel, a spring held by a frame, the spring surrounding a longitudinal portion of the input shaft in order to allow a movement of the input gear wheel towards an equilibrium position corresponding to an equal distribution of the transmission of the power coming from the input shaft to the first intermediate shaft of the first intermediate line and the second intermediate shaft of the second intermediate line.

A gas turbine engine comprises a gearbox including a first epicyclic gearbox and a second epicyclic gearbox. The first epicyclic gearbox comprises a first sun gear meshing with the first planet gears and the first planet gears meshing with a first annulus gear. The second epicyclic gearbox comprises a second sun gear meshing with the second planet gears and the second planet gears meshing with a second annulus gear. An input shaft is arranged to drive the first sun gear and a first planet gear carrier is arranged to drive a propulsor. The first and second annulus gears are fixed to a static structure. A first clutch is arranged between the first planet gear carrier and the second planet gear carrier and a second clutch is arranged between the second sun gear and the input shaft. The gearbox is more efficient at cruise conditions.

A system, including: a linear actuator axially arranged in a first propeller and rotatably secured thereto, a first linking mechanism connecting a rod of the linear actuator to a second propeller to change a setting of blades thereof, and including an intermediate bearing between the propellers, which is capable of transmitting translation of the rod of the linear actuator and disconnecting a rotatable link to the first propeller to make it possible to change the setting of the blades of the second propeller rotated in an opposite direction to the first propeller, and a second linking mechanism combined with blades of the first propeller to change a setting thereof.

A suspension structure for suspending a turboprop including two unducted propellers from a structural element of an aircraft, the structure including a suspension frame for fastening to a structural element of the aircraft and including a main beam extending parallel to a longitudinal axis of the turboprop, the suspension frame being fastened to an air intake structure of the turboprop by an attachment frame.

There is described a method and system for in-flight start of an engine. The method comprises rotating a propeller; generating electrical power at an electric generator embedded inside a propeller hub from rotation of the propeller; transmitting the electrical power from the electric generator to an engine starter mounted on a core of the engine via an electric power link; and driving the engine with the engine starter to a sufficient speed while providing fuel to a combustor to light the engine to achieve self-sustaining operation of the engine.

A shock absorber assembly for a non-streamlined blower includes an extension collar having a first interface suitable for being rigidly attached on a first module of the blower; a connector support housing including a third interface suitable for being rigidly attached on a second module of the blower; and a fourth interface allowing a mechanical connection to be made between the housing and a second interface of the collar; a plurality of elastic devices arranged between the collar and the housing; a flange element rigidly attached on the third interface of the housing comprising a stop bearing on the collar in such a way as to define a space between the collar and the flange element; the elastic means are arranged so that a residual clearance exists between the flange element and the housing.

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and an overall engine efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include a low-pressure turbine. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine of the core engine and comprises a first rotational speed, the output of the gearbox is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-4.0.

There are described methods and systems for providing an autothrottle mode in a propeller-driven aircraft. A thrust change is obtained corresponding to a difference between an actual thrust and a desired thrust for an engine. When greater than a pre-determined threshold, a setting change to one or more control input(s) of the engine is determined. One or more commands is output to cause the setting change of the control input(s).

A gas turbine engine has an engine core including a low pressure compressor, a high pressure compressor, a high pressure turbine and a low pressure turbine coaxially mounted about an engine axis and fluidly connected in series by a core gaspath. The low pressure turbine is drivingly connected to the low pressure compressor via an external shaft disposed radially outwardly of the core gaspath.

A three-stream engine is provided. The three-stream engine includes a fan section, a core engine disposed downstream of the fan section, and a core cowl annularly encasing the core engine and at least partially defining a core duct. A fan cowl is disposed radially outward from the core cowl and annularly encasing at least a portion of the core cowl. The fan cowl at least partially defining an inlet duct and a fan duct. The fan duct and the core duct at least partially co-extending axially on opposite sides of the core cowl. A heat exchanger disposed within the fan duct. The heat exchanger provides for thermal communication between a fluid flowing through fan duct and a motive fluid flowing through the heat exchanger.