The invention relates to a method for controlling the speed and the power of a turbine engine propeller, wherein at least two operating modes are implemented: —one operating mode, called “speed mode”, in which the pitch (β) of the propeller is controlled as a function of the desired propeller speed, while the fuel flow is controlled as a function of the desired torque; the other operating mode, called “β mode”, in which the fuel flow is controlled as a function of the desired propeller speed, the pitch (β) of the propeller being set to a limit angle (βmin) that limits the pitch of the propeller in the two operating modes, the pitch angle (βmin(t)) being continuously computed and updated during a flight on the basis of parameters relating to the flight conditions estimated in real time.

A fuel pump system can include a motor and a pump connected to the motor. The pump can be configured to receive an inlet flow from an inlet line, to pressurize the inlet flow, and to output a pressurized flow to an output line for an engine. The system can include a bypass line disposed between the outlet line and the inlet line, and a bypass valve disposed on the bypass line and configured to allow pressurized flow to flow to the inlet line in an open state, and to prevent pressurized flow from flowing to the inlet line in a closed state. The bypass valve can be configured to allow pressurized flow to flow to the inlet line to circulate flow and to maintain a constant pressure on the output line.

A method and a system for providing in-flight reverse thrust for an aircraft are provided. The aircraft comprises an engine having a rotor, a compressor mechanically coupled to the rotor, and a variable geometry mechanism provided upstream of the compressor and configured to modulate an amount of compression work performed by the compressor. The method comprises operating the rotor with the variable geometry mechanism in a first position, receiving a request to increase reverse thrust for the rotor, in response to the request, adjusting the variable geometry mechanism from the first position towards a second position, the variable geometry mechanism having a greater opening angle in the second position than in the first position, and operating the rotor with the variable geometry mechanism in the second position for causing an increase in the amount of compression work performed by the compressor and an increase in reverse thrust for the rotor.

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 invention relates to a device for starting a gas turbine, comprising an initiation system able, on command, to rotate said gas turbine.

The starting device comprises means for receiving information representing the rotation speed of the gas turbine, means for calculating a set torque value according to said information representing the rotation speed of the gas turbine, and means for transmitting said set torque value to the initiation system, and the means for calculating the set torque value are suitable for making a comparison of the information representing the rotation speed of the gas turbine with a predetermined speed profile and calculating the set torque value from said comparison.

To provide an aircraft engine in which an accessory can be directly supported with no accessory gearbox intervening to thereby suppress an increase in size thereof, the aircraft engine (E) includes a take-out shaft (11) having a first end portion, connected with an engine rotary shaft (9) and extending in a radially outward direction, and also having a second end portion connected with an accessory (1), and a mounting pad (12) provided in an engine main body (EB) and to which the accessory (1) is fitted. The mounting pad (12) forms an outer perimeter of an opening (48) through which the take-out shaft (11) extends. The opening (48) is sealed by a covering (47), through which the take-out shaft (1) extends, and a sealing member (49) to seal between the covering (47) and the take-out shaft (11).

A power extraction system for a gas turbine engine may comprise a low spool transmission and a low spool accessory gearbox. The low spool accessory gearbox may comprise a generator and a dual clutch transmission. The dual clutch transmission may be coupled between the low spool transmission and the generator.

The invention relates to a method for regulating an acceleration of a turbomachine, the turbomachine comprising a combustion chamber, a high-pressure turbine located downstream of the combustion chamber and which drives a high-pressure shaft in rotation, and a low-pressure turbine located downstream of the high-pressure turbine and which drives a low-pressure shaft in rotation, characterized in that the method comprises the following steps: Injecting (100) mechanical power onto the high-pressure shaft until a speed of the high-pressure shaft attains a target value, then Extracting (200) mechanical power from the high-pressure shaft so as to maintain the speed at the target value.

A two-gas shaft turbine control system (31) is disclosed. The gas turbine control system comprises a fuel controller (35), which receives a speed error signal, indicating whether the low-pressure turbine wheel (11) of the gas turbine (1) is rotating at the desired target speed. The gas turbine control system (31) further comprises an NGV controller (41), which receives a speed error signal, indicating whether the high-pressure turbine wheel (9) of the gas turbine engine (1) is rotating at the desired target speed. Two cross channel controllers are further provided. On the basis of a gas turbine model, a first cross channel controller (43) provides a fuel control modification signal, which is added to a control signal generated by the fuel controller (35). A second cross channel controller (45) provides an NGV control modification signal. The modification signals are aimed at reducing or canceling the effect of mutual interaction between fuel control and NGV control. A two-shaft gas turbine engine system and a method of operation are further disclosed.

A method and a system for providing in-flight reverse thrust for an aircraft are provided. The aircraft comprises an engine having a rotor, a compressor mechanically coupled to the rotor, and a variable geometry mechanism provided upstream of the compressor and configured to modulate an amount of compression work performed by the compressor. The method comprises operating the rotor with the variable geometry mechanism in a first position, receiving a request to increase reverse thrust for the rotor, in response to the request, adjusting the variable geometry mechanism from the first position towards a second position, the variable geometry mechanism having a greater opening angle in the second position than in the first position, and operating the rotor with the variable geometry mechanism in the second position for causing an increase in the amount of compression work performed by the compressor and an increase in reverse thrust for the rotor.