A multivariable fuel control and estimator (MFCE) of a gas turbine engine for preventing combustor blowout is provided. The MFCE includes a first input port that receives controller requests and provide system usage commands, a second input port that receives measured disturbance values, a third input that receives system and component limits, a fourth input port that receives sensed parameters, a fuel system model of the fuel system of the gas turbine engine and an engine model of the engine system that includes the combustor of the gas turbine engine, a processor that generates a control signal for controlling the fuel valve and generates a control signal for controlling the actuator using the fuel system and engine model based on the controller requests, the measured disturbance values, the system and component limits, and the sensed parameters, and an output port that transmits the control signals to the fuel system.

This invention concerns an aircraft propulsion system in which an engine has an engine core comprising a compressor, a combustor and a turbine driven by a flow of combustion products of the combustor. At least one propulsive fan generates a mass flow of air to propel the aircraft. An electrical energy store is provided on board the aircraft. At least one electric motor is arranged to drive the propulsive fan and the engine core compressor. A controller controls the at least one electric motor to mitigate the creation of a contrail caused by the engine combustion products by altering the ratio of the mass flow of air by the propulsive fan to the flow of combustion products of the combustor. The at least one electric motor is controlled so as to selectively drive both the propulsive fan and engine core compressor.

This invention concerns an aircraft propulsion system in which an engine has an engine core comprising a compressor, a combustor and a turbine driven by a flow of combustion products of the combustor. At least one propulsive fan generates a mass flow of air to propel the aircraft. An electrical energy store is provided on board the aircraft. At least one electric motor is arranged to drive the propulsive fan and the engine core compressor. A controller controls the at least one electric motor to mitigate the creation of a contrail caused by the engine combustion products by altering the ratio of the mass flow of air by the propulsive fan to the flow of combustion products of the combustor. The at least one electric motor is controlled so as to selectively drive both the propulsive fan and engine core compressor.

A speed control system and a power load balance detector for a turbine is provided. The speed control system includes a speed wheel with a plurality of teeth. A timer stores a time stamp when each of the teeth passes by a speed probe. A first speed estimate is determined for overspeed protection, and a second speed estimate is determined for operational speed control. The power load balance detector trips or shuts down the turbine when an unbalance is above a first threshold and the speed of the turbine is above a second threshold.

The invention relates to a method for starting up a gas turbine engine of a combined cycle power plant. The method includes applying load to the gas turbine engine and increasing the load until a predetermined combustor firing temperature is reached, while keeping the adjustable inlet guide vanes in a start position adapted to reduce the mass flow of air into the compressor; further increasing the load of the gas turbine engine while opening the adjustable inlet guide vanes and keeping the predetermined combustor firing temperature constant until the inlet guide vanes reach an end position adapted to increase the mass flow of air into the compressor; further increasing the load of the gas turbine engine while keeping the adjustable inlet guide vanes in the end position until a predetermined load of the gas turbine engine is reached.

The invention relates to a method for starting up a gas turbine engine of a combined cycle power plant. The method includes applying load to the gas turbine engine and increasing the load until a predetermined combustor firing temperature is reached, while keeping the adjustable inlet guide vanes in a start position adapted to reduce the mass flow of air into the compressor; further increasing the load of the gas turbine engine while opening the adjustable inlet guide vanes and keeping the predetermined combustor firing temperature constant until the inlet guide vanes reach an end position adapted to increase the mass flow of air into the compressor; further increasing the load of the gas turbine engine while keeping the adjustable inlet guide vanes in the end position until a predetermined load of the gas turbine engine is reached.

A method to reduce aerodynamic drag of a engine exhaust/engine nozzle includes collecting data that is indicative of an instant flight condition, entering the data into a decision algorithm that, based on the data, outputs at least first and second drag control parameters corresponding, respectively, to an angle of one or more variable area turbines of a turbine engine and a position of a variable area exhaust nozzle of the turbine engine, and adjusting the angle of the one or more variable area turbines and the position of the variable area exhaust nozzle according to, respectively, the first and second drag control parameters to reduce aerodynamic drag of an engine exhaust/engine nozzle of the turbine engine.

A method to reduce aerodynamic drag of a engine exhaust/engine nozzle includes collecting data that is indicative of an instant flight condition, entering the data into a decision algorithm that, based on the data, outputs at least first and second drag control parameters corresponding, respectively, to an angle of one or more variable area turbines of a turbine engine and a position of a variable area exhaust nozzle of the turbine engine, and adjusting the angle of the one or more variable area turbines and the position of the variable area exhaust nozzle according to, respectively, the first and second drag control parameters to reduce aerodynamic drag of an engine exhaust/engine nozzle of the turbine engine.

A 2-shaft gas turbine has a controller which controls the opening degree of an air inlet guide vane to adjust the inlet mass flow rate to a compressor. The air inlet guide vane control unit includes a first control unit that adjusts the opening degree of the inlet guide vane to keep the speed of a high pressure turbine shaft constant; a control status confirmation unit that confirms the actual speed and the opening degree of the inlet guide vane; and a low ambient temperature correction unit that reduces the actual speed in a case where the actual speed is equal to or greater than a predetermined threshold value, the opening degree of the inlet guide vane is equal to or greater than a predetermined threshold value, and the ambient temperature is equal to or less than a predetermined threshold value.

The present disclosure relates to systems and methods that are useful in control of one or more aspects of a power production plant. More particularly, the disclosure relates to power production plants, methods of starting power production plants, and methods of generating power with a power production plant wherein one or more control paths are utilized for automated control of at least one action. The present disclosure more particularly relates to power production plants, control systems for power production plants, and methods for startup of a power production plant.