Systems and methods for limiting power of a gas turbine engine for an aircraft are described herein. A blade angle of a propeller blade of the engine and a commanded power for the engine are obtained. A thrust transition direction is determined. The commanded power is compared to a selected threshold based on the blade angle and the thrust transition direction. Power to the engine is limited when the commanded power exceeds the selected threshold.

The present disclosure provides a device for controlling a variable section ejection nozzle of a turbojet engine nacelle of an aircraft. The device includes a calculator adapted to determine a position setpoint of the nozzle and a management system of the servo-control of the position of the variable nozzle depending on the flow rate of the fuel supplying the turbojet engine. The management system includes at least one instantaneous flow rate sensor of the fuel and a management unit which is designed to compare the flow rate measured by the flow rate sensor with a theoretical fuel flow rate depending on the parameters of the flight of the aircraft, to determine a correction value of the position of the nozzle depending on the comparison of the measured flow rate and the theoretical fuel flow rate, and to correct the position setpoint of the nozzle according to the correction value.

The present disclosure provides a device for controlling a variable section ejection nozzle of a turbojet engine nacelle of an aircraft. The device includes a calculator adapted to determine a position setpoint of the nozzle and a management system of the servo-control of the position of the variable nozzle depending on the flow rate of the fuel supplying the turbojet engine. The management system includes at least one instantaneous flow rate sensor of the fuel and a management unit which is designed to compare the flow rate measured by the flow rate sensor with a theoretical fuel flow rate depending on the parameters of the flight of the aircraft, to determine a correction value of the position of the nozzle depending on the comparison of the measured flow rate and the theoretical fuel flow rate, and to correct the position setpoint of the nozzle according to the correction value.

A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a nozzle scheduler for determining an exhaust nozzle position goal based on a nozzle schedule of exhaust nozzle positions related to flight conditions. The control system may further include a control module for determining a control command for the gas turbine engine. The control command may include, at least, a fuel flow command and an exhaust nozzle position command and the control command may be based on, at least, the exhaust nozzle position goal and an estimated thrust value. The control system may further include an actuator for controlling the gas turbine engine based on the control command.

A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a nozzle scheduler for determining an exhaust nozzle position goal based on a nozzle schedule of exhaust nozzle positions related to flight conditions. The control system may further include a control module for determining a control command for the gas turbine engine. The control command may include, at least, a fuel flow command and an exhaust nozzle position command and the control command may be based on, at least, the exhaust nozzle position goal and an estimated thrust value. The control system may further include an actuator for controlling the gas turbine engine based on the control command.

A ramjet engine and system and method for operation is generally provided. The ramjet includes a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines an inlet section, a combustion section, and an exhaust section. A fuel nozzle assembly is extended from the longitudinal wall. The fuel nozzle assembly defines a nozzle throat area. The fuel nozzle assembly is moveable along a radial direction to adjust the nozzle throat area based at least on a difference in pressure of a flow of fluid at an inlet of the inlet section and a pressure of the flow of fluid at the fuel nozzle assembly.

A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a nozzle scheduler for determining an exhaust nozzle position goal based on a nozzle schedule of exhaust nozzle positions related to flight conditions. The control system may further include a control module for determining a control command for the gas turbine engine. The control command may include, at least, a fuel flow command and an exhaust nozzle position command and the control command may be based on, at least, the exhaust nozzle position goal and an estimated thrust value. The control system may further include an actuator for controlling the gas turbine engine based on the control command.

A control system for a gas turbine engine, a method for controlling a gas turbine engine, and a gas turbine engine are disclosed. The control system may include a nozzle scheduler for determining an exhaust nozzle position goal based on a nozzle schedule of exhaust nozzle positions related to flight conditions. The control system may further include a control module for determining a control command for the gas turbine engine. The control command may include, at least, a fuel flow command and an exhaust nozzle position command and the control command may be based on, at least, the exhaust nozzle position goal and an estimated thrust value. The control system may further include an actuator for controlling the gas turbine engine based on the control command.

A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

A control system operable includes a modified primary control parameter for an aircraft, the control system includes: a primary control parameter leg configured to output a demand in an aircraft primary control parameter; a primary control parameter compensation leg configured to receive a change in absolute levels and/or spatial distributions of swirl angle and/or fan pressure at a primary control parameter relative to a reference and convert the change into the primary control parameter; a processor adapted to receive the demand in the primary control parameter output from the primary control parameter leg and the change to the primary control parameter output from the primary control parameter compensation leg; compare the demand in the primary control parameter output from the primary control parameter leg and the change to the primary control parameter output from the primary control parameter compensation leg; and generate a modified primary control parameter for the aircraft.