A gas turbine engine includes a compressor section that provides first and second compressor stages that are configured to respectively provide first and second cooling fluids. The first compressor stage has a higher pressure than the second compressor stage. The gas turbine engine further includes a component that has platform with an internal cooling passage fed by first and second inlets that respectively receive fluid from the first and second cooling sources. The second inlet is downstream from the first inlet.

An assembly for mounting an aircraft engine to an aircraft includes an engine casing flange having a first annular wall extending radially to terminate at an annular rim. A second flange of an additional engine component mounted aft of the engine casing includes a second annular wall. An aft mount bracket has an annular body extending uninterrupted about the center axis and a spigot extending axially from the annular body, the spigot extending circumferentially about an entire circumference of the annular body. The aft mount bracket is axially disposed between the engine casing flange and the additional engine component, with corresponding holes in the first annular wall, second annular wall and aft mount bracket being circumferentially aligned, and the spigot radially abutting the annular rim of the engine casing flange.

An airfoil for a gas turbine engine according to an example of the present disclosure includes a platform section and an airfoil section extending in a spanwise direction from the platform section to a tip portion establishing a tip. The airfoil section has an external wall defining pressure and suction sides extending in a chordwise direction between a leading edge and a trailing edge, and the pressure and suction sides are spaced apart in a thickness direction between the leading edge and the trailing edge. The tip portion includes a tip pocket and a tip shelf extending inwardly from the tip. The tip pocket and tip shelf are on opposite sides of a shelf wall.

An example transmission comprising a lubricant feed passageway to receive lubricant; an input shaft having a cavity that defines a shaft collection trough to receive the lubricant from the lubricant feed passageway, the input shaft defining a radial passageway to enable the lubricant to flow from the shaft collection trough to an exterior surface of the input shaft; and a carrier to receive at least a portion of the input shaft, the carrier defining a carrier collection trough and channels formed in the carrier to distribute the lubricant to at least one of rotating components or non-rotating components in the carrier, the carrier collection trough to receive the lubricant from the shaft collection trough via the radial passageway in response to the input shaft rotating.

There are described methods and systems for operating an aircraft having two or more engines. One method comprises operating the two or more engines of the aircraft in an asymmetric operating regime, wherein a first of the engines is in an active mode to provide motive power to the aircraft and a second of the engines is in a standby mode to provide substantially no motive power to the aircraft; governing the first engine in the active mode using a first governing logic; and governing the second engine in the standby mode using a second governing logic, the second governing logic based on a target compressor speed and variable geometry mechanism (VGM) settings that are adjusted using trim values dependent on at least one parameter of the second engine in the standby mode.

Methods and systems for governing an aircraft propeller of an engine are described. The method comprises obtaining a fluid flow command for speed control of the propeller, determining pulse parameters of a pulse width modulated valve control signal for actuating a two-position solenoid valve in accordance with the fluid flow command based on an average fluid flow through the solenoid valve and an opening and closing time of the solenoid valve, generating the valve control signal with the pulse parameters as determined, and transmitting the valve control signal to the solenoid valve for actuating the solenoid valve, thereby controlling the speed of the propeller.

A hybrid electric propulsion system includes a gas turbine engine having at least one compressor section and at least one turbine section operably coupled to a shaft. The hybrid electric propulsion system includes an electric motor configured to augment rotational power of the shaft of the gas turbine engine. A controller is operable to determine an estimate of hybrid electric propulsion system parameters based on a composite system model and sensor data, determine a model predictive control state and a prediction based on the hybrid electric propulsion system parameters and the composite system model, determine a model predictive control optimization for a plurality of hybrid electric system control effectors based on the model predictive control state and the prediction using a plurality of reduced-order partitions of the composite system model, and actuate the hybrid electric system control effectors based on the model predictive control optimization.

A propulsion system includes a drive shaft movable about an axis of rotation, a fan, a fan shaft that drives the fan, and a reduction device coupling the drive shaft and the fan shaft. The reduction device has first and second reduction stages and includes a sun gear, centered on the axis and driven by the drive shaft, a ring gear that is coaxial with the sun gear and that drives the fan shaft about the axis, and planet gears distributed circumferentially about the axis between the sun gear and the ring gear. Each planet gear includes a first portion meshed with the sun gear and a second portion meshed with the ring gear. A diameter of the first portion is different from a diameter of the second portion, and the first portion of the planet gears extend between the second portion of the planet gears and the fan.

Turbomachine (10) having a contrarotating turbine for aircraft, the turbomachine comprising a contrarotating turbine (22) of which a first rotor (22a) is configured to rotate in a first direction of rotation and is connected to a first turbine shaft (36), and a second rotor (22b) is configured to rotate in an opposite direction of rotation and is connected to a second turbine shaft (38), the first rotor comprising turbine discs that are interleaved between turbine discs of the second rotor, said first shaft (36) being guided by at least two guide bearings (60, 62) mounted between this first shaft and a stator casing, and said second shaft (38) being guided by at least two guide bearings (56, 58) mounted between this second shaft and another stator casing (28).

A method of controlling a gas turbine engine. The gas turbine engine has a compressor, a combustor, and a motor configured to drive the compressor. The method has in a first idle mode, controlling combustor fuel flow to maintain compressor rotational speed at or above a predetermined value; and in a second idle mode, controlling the combustor and the motor to drive the compressor to maintain the compressor rotational speed at or above a second predetermined value, lower than the first predetermined value.