In an aspect of the present disclosure, an aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; a single unducted rotor engine mounted at a location spaced from the fuselage of the aircraft, the single unducted rotor engine comprising an unducted rotor assembly having a single stage of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the single stage of rotor blades of the unducted rotor assembly along the lateral direction.

A single unducted rotor engine is provided. The single unducted rotor engine defines an axial direction and a radial direction, the engine including: a power source; a casing surrounding the power source; an unducted rotor assembly driven by the power source comprising a single row of rotor blades; and an outlet guide vane assembly comprising a plurality of pairs of outlet guide vanes, each pair of the plurality of pairs of outlet guide vanes including first outlet guide vane extending from the casing at a location downstream from the single row of rotor blades of the unducted rotor assembly and a second outlet guide vane also positioned downstream from the single row of rotor blades of the unducted rotor assembly; wherein the first outlet guide vane of each pair of outlet guide vanes defines a first geometry, wherein the second outlet guide vane of each pair of outlet guide vanes defines a second geometry, and wherein the first geometry is not equal to the second geometry.

A control device includes a suction amount controller and a bleed flow controller. When an actual output is in a second output range from a second output to a third output less than the second output, the suction amount controller controls a suction amount so that the suction amount becomes constant even if the actual output decreases. The bleed flow controller controls a bleed flow so that the bleed flow gradually increases as the actual output decreases.

A method for selectively modulating bleed air used for cooling a downstream turbine section in a gas turbine engine. The method including: measuring an engine and/or aircraft performance parameter by an engine sensor device; comparing the engine and/or aircraft performance parameter to a performance threshold; determining a bleed trigger condition, if the engine and/or aircraft performance parameter crosses the performance threshold; determining a non-cooling condition, if the engine and/or aircraft performance parameter is below the performance threshold; actuating a flow control valve to an open position, in response to the bleed trigger condition, so that bleed air is extracted from the compressor section and flowed to the downstream turbine section; and terminating, in response to the non-cooling condition, the flow of the bleed air to the downstream turbine section of the engine by actuating the flow control valve to a closed position.

Methods and systems for controlling an engine having a variable geometry mechanism are described. An output power of the engine is determined. A speed of the engine is determined. A temperature-independent position control signal for the variable geometry mechanism is generated based on a power-to-speed ratio, the power-to-speed ratio obtained by dividing the output power by the speed. The position control signal is output to a controller of the engine to control the variable geometry mechanism.

A method and a system for pre-shutdown motoring of an aircraft engine are provided. A period of time during which the engine, prior to a shutdown thereof, was operating at a power level below a predetermined threshold is tracked. Based on the tracked period of time, a motoring duration for the engine is determined. The engine is then ran, prior to the shutdown thereof, at a low power setting for the motoring duration.

A gas turbine plant includes a connection line configured to connect an outlet of a compressor high-pressure stage and an inlet of a turbine via a combustor, a bypass line configured to cause some or all of air compressed at a compressor low-pressure stage to bypass the compressor high-pressure stage and to be supplied to the connection line, and an adjustment device configured to adjust a flow rate of the air flowing through the bypass line. A plurality of types of fluid are supplied to the connection line in addition to the air compressed by the compressor, and during operation of the gas turbine plant, supply of at least one type of fluid of the plurality of types of fluid to the connection line is stopped according to an operating state of the gas turbine plant.

A control device selects and executes any of normal control of outputting a fuel flow rate command value through feedback control based on a deviation between a target output and an actual output of a gas turbine and load decreasing control of outputting the fuel flow rate command value for reducing an output of the gas turbine to a predetermined target output during a predetermined first time without performing feedback control, and performs control of reducing a flow rate of air flowing into a compressor of the gas turbine such that an air-fuel ratio settles within a predetermined range, in parallel with execution of the load decreasing control when the load decreasing control is selected.

Methods and systems for operating an auxiliary power unit (APU) of an aircraft are described. The method comprises obtaining external environment parameters of the aircraft, determining an available output power for the APU as a function of the external environment parameters, and setting an overcurrent protection threshold of the APU to a level associated with the available output power.

A method includes receiving a command to operate a power load of a power system at a command output power while operating the power load at a reference output power; operating a gas turbine engine of the power system in a maximum regulator mode to increase a power generation of the gas turbine engine when the command output power is greater than the reference output power or in a minimum regulator mode to decrease the power generation of the gas turbine engine when the command output power is less than the reference output power; and coordinating an electric machine power draw from the gas turbine engine with a change in power generation of the gas turbine engine to maintain a rotational speed parameter of the gas turbine engine substantially constant while operating the gas turbine engine in the maximum regulator mode or in the minimum regulator mode.