A control method for controlling an air intake system for an engine of a vehicle; the intake system has a main air intake coupled to an air filter provided with a heating device. The control method comprises the steps of: determining a pressure difference between upstream and downstream of the air filter; determining a variation speed of the pressure difference between upstream and downstream of the air filter by calculating the first derivative in time of the pressure difference between upstream and downstream of the air filter; and turning on and/or turning off the heating device based on the variation speed of the pressure difference between upstream and downstream of the air filter.

Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor inlet to receive airflow from a first air supply source. An air mixing device having a first mixer inlet to receive compressed air from the compressor outlet and a second mixer inlet to receive bleed air from a bleed air system. The bleed air to provide a motive fluid to enable the air mixing device to mix the bleed air and the compressed air to produce mixed air for the anti-icing system.

A gas turbine control device includes a detection value acquisition unit that acquires a detection value of at least one of a supply amount of fuel, pressure of compressed air, and electric power generated by a generator; a flue gas temperature acquisition unit that acquires a flue gas temperature detection value; a combustion gas temperature estimate value calculation unit that calculates a combustion gas temperature estimate value based on the detection value; a correction term acquisition unit that calculates a correction term based on a ratio between the combustion gas temperature estimate value and the flue gas temperature detection value; a corrected combustion gas temperature estimate value calculation unit that corrects the combustion gas temperature estimate value using the correction term to calculate a corrected combustion gas temperature estimate value; and a gas turbine controller that controls the gas turbine based on the corrected combustion gas temperature estimate value.

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.

In a compressed air energy storage and power generation device, a compressed air energy storage and power generation method defines, as a reference storage value, a storage value indicating that a storage amount of air in an accumulator tank is in a predetermined intermediate state. At the reference storage value, at least one of a motor and a generator rotates at a rated rotation speed. When a storage value indicating a current storage amount in the accumulator tank is larger than the reference storage value, at least one of the motor and the generator is controlled to rotate at equal to or less than the rated rotation speed. When the storage value indicating the current storage amount in the accumulator tank is smaller than the reference storage value, at least one of the motor and the generator is controlled to rotate at equal to or more than the rated rotation speed and equal to or less than a maximum permissible rotation speed.

Flow multiplier systems for aircraft are described herein. A flow multiplier system includes a turbo-compressor having a compressor, a turbine, and a drive shaft coupled between the compressor and the turbine. A compressor outlet of the compressor is fluidly coupled to an ejector in a gas turbine engine. The system also includes a supply line fluidly coupling a compressed air tank and a turbine inlet and a valve coupled to the supply line. The system includes a controller configured to, based on an input signal requesting to increase output power of the gas turbine engine, send a command signal to open the valve to enable a flow of pressurized air from the compressed air tank to the turbine inlet. The turbine drives the compressor to create high pressure air at the compressor outlet, which is provided into the gas turbine engine to increase the output power.

An auxiliary power unit (APU) control system for an aircraft is disclosed. The APU control system includes an APU, one or more processors, and a memory coupled to the one or more processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the APU control system to receive a one or more ambient signals indicative of an air density value and one or more power signals indicative of a specific amount of power generated by the APU. The APU control system is further caused to determine a variable rotational speed of the APU based on the air density value and instruct the APU to operate at the variable rotational speed. The APU continues to generate the specific amount of power when operating at the variable rotational speed.

A flutter control system for a turbine includes a processor. The processor is configured to detect blade flutter of a turbine. The blade flutter indicates that blades of the turbine are in a deflected position different from a nominal operating position. The processor is configured to control operational parameters of the turbine that reduce or eliminate the blade flutter to improve the reliability and efficiency of the turbine.

A method for determining an engine heath of an aircraft engine includes determining, by one or more control devices, the aircraft engine is operating in a bleed off condition; determining, by the one or more control devices, a first engine health modifier value while the aircraft engine is operating in the bleed off condition, the first of engine health modifier value including a compressor leakage flow value; determining, by the one or more control devices, a second plurality of engine health modifier values while the aircraft engine is operating in a bleed on condition; and determining, by the one or more control devices, an engine health parameter using at least one of the second plurality of engine health modifier values determined while the aircraft engine is operating in the bleed on condition and the compressor leakage flow value determined while the aircraft engine was operating in the bleed off condition.

A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.