The invention relates to a method for controlling a pitch angle of the vanes or blades of a propellant body of a turbine engine, comprising generating a pitch command (i.sub.final) according to a rotational speed of the propeller (XN.sub.mes) and a speed setpoint (XN.sub.cons), the method comprises a nominal regulating chain (13), wherein the pitch command is further generated according to a value of a pitch angle (βmes) of the vanes or blades of the propellant body, and an off-nominal regulating chain (16), wherein the pitch command is generated independently of a value of a pitch angle of the vanes or blades of the propellant body.

This single-shaft combined cycle plant comprises: a power generator; a gas turbine; a steam turbine that is driven by using waste heat from the gas turbine, and is connected to the power generator by a clutch when the rotational speed syncs with the rotational speed of the gas turbine; a steam turbine over-rotation prevention device; a gas turbine over-rotation prevention device; and a control device. The control device sets the power generator to an unloaded state and, whilst maintaining the rotational speed Ng of the gas turbine so as to be higher than the rotational speed Ns of the steam turbine and lower than the maximum rotational speed Nglim of the gas turbine, increases the rotational speed Ns of the steam turbine to the maximum rotational speed Nslim of the steam turbine (time t2-t4) and tests whether or not the steam turbine over-rotation prevention device operates normally.

A system and method for adaptively controlling a cooldown cycle of an auxiliary power unit (APU) that is operating and rotating at a rotational speed includes reducing the rotational speed of the APU to a predetermined cooldown speed magnitude that ensures combustor inlet temperature has reached a predetermined temperature value, determining, based on one or more of operational parameters of the APU, when a lean blowout of the APU is either imminent or has occurred, and when a lean blowout is imminent or has occurred, varying one or more parameters associated with the shutdown/cooldown cycle.

An assembly for a turbine engine turbine includes a turbine rotor disc centered on a longitudinal axis, a turbine shaft centered on the longitudinal axis and driven in rotation by the rotor disc. The assembly further includes first means of transmitting torque from the rotor disc to the shaft, wherein the rotor disc is locked in translation relative to the shaft in the direction of the longitudinal axis by a screwed member on the said shaft and second means of transmitting torque from the rotor disc to the screwed member. The screwed member has an unscrewing direction identical to the direction of rotation of the rotor disc in operation, and the second means of transmitting torque are configured to transmit the rotational torque from the rotor disc to the screwed member when the first means of transmitting torque cease to transmit torque from the rotor disc to the shaft.

A method for stopping an engine of a rotorcraft in overspeed, the rotorcraft comprising at least one engine, the engine comprising a gas generator and a power assembly, the power assembly comprising at least one power turbine rotated by gases originating from the gas generator, the power assembly comprising at least one power shaft rotationally secured to the power turbine, the power assembly rotating about a longitudinal axis at a speed referred to as the “speed of rotation”. The method comprises steps consisting in measuring a current value of the speed of rotation, determining a time derivative of the current value of the speed of rotation, referred to as the “current derivative $\left(\frac{dN2i}{dt}\right)”,$
and automatically stopping the engine when the current derivative $\left(\frac{dN2i}{dt}\right)$
changes sign.

A method (600) for testing control logic for a propeller driven by a gas turbine engine of an aircraft includes overriding (602) a signal indicating the aircraft is operating in a ground mode. The method can further include testing (604) minimum pitch protection logic when the signal is overridden; determining (606) the gas turbine engine is operating at a ground fine setting; restoring (608) the signal to an original state in which the signal indicates the aircraft is operating in the ground mode; modifying (610) pitch protection logic; determining (614) the propeller is operating at an overspeed condition; and testing (616) the propeller overspeed protection logic. In addition, the method can also determine (612) the propeller is operating at a low pitch condition when the gas turbine engine is operating at the ground fine setting.

A flight control computer (FCC) for a rotorcraft includes a processor and a non-transitory computer-readable storage medium storing a program to be executed by the processor, with the program including instructions for providing main rotor overspeed protection. The instructions for providing the main rotor overspeed protection include instructions for monitoring sensor signals indicating a main rotor RPM, determining a target operating parameter, determining one or more flight parameters in response to a relationship between the main rotor RPM and the target operating parameter indicating a main rotor overspeed condition. Determining the one or more flight parameters includes determining a setting for a flight control device of the rotorcraft that changes the main rotor RPM, controlling positioning of a pilot control according to the flight parameters, and controlling the flight control device of the rotorcraft according to positioning of the pilot control.

A method of controlling a gas turbine engine includes the steps of: detecting a shaft break event in a shaft connecting a compressor of the gas turbine engine to a turbine of the gas turbine engine; and in response to this detection, activating a shaft break mitigation system which introduces a fluid into a gas flow of the gas turbine engine downstream of the turbine, or increases an amount of a fluid being provided into the gas flow of the gas turbine engine downstream of the turbine, whereby the fluid reduces an effective area of a nozzle for the gas flow so as to reduce the mass flow rate of the gas flow through the turbine.

A gas turbine engine includes, among other things, a fan, a fan drive gear system that is coupled with the fan and a fan drive input shaft, a compressor section that includes a first compressor and a second compressor, and a turbine section. The turbine section includes a first turbine coupled with a first shaft and a second turbine coupled through a second shaft to the second compressor. A bearing supports the fan drive input shaft. The bearing is located proximal to, and radially spaced from, a forward end of the first shaft. The bearing includes a speed sensor target that is rotatable with the forward end and that defines a rotation path. A speed sensor probe is situated proximal to the rotation path and is operable to read the speed sensor target.

A control system (10) for a gas turbine engine (1) having a gas generator (4) and a turbine (6) driven by the gas generator (4), is provided with: a control unit (12) to control a forward operating mode or a reverse operating mode of the gas turbine engine (1); and a supervising unit (14), operatively coupled to the control unit (12), to receive an input signal (PLA) indicative of a forward, or reverse, power request and to cause the control unit (12) to control the forward, or reverse, operating mode based on the input signal (PLA). The supervising unit (14) has an enabling stage (20) to enable a transition between the forward and reverse operating modes based on a check that a safety condition is satisfied.