This gas turbine power generation system includes: a two-shaft gas turbine; a power generator that is driven by the two-shaft gas turbine and that generates power; an electric motor that is driven by power generated by the power generator; and a frequency converter that converts a frequency of the power that is transmitted between the power generator and the electric motor. The frequency converter is disposed within a predetermined range near the electric motor at the side opposite to the two-shaft gas turbine side relative to the electric motor so as to face the plane perpendicular to a rotation shaft of the electric motor.

There is described a method of operating a multi-engine system of an helicopter. The multi-engine system has a first turboshaft engine having a first shaft, a second turboshaft engine having a second shaft, a gearbox having a clutch system, and a range of rotation speeds defined as a placarded zone. The method generally has: rotating the first and second shafts at a flight rotation speed above the placarded zone when clutched to a load; decreasing a rotation speed of the first shaft from the flight rotation speed to a first idle rotation speed above the placarded zone; unclutching the first shaft from the load during the decreasing; and subsequently to the decreasing and the unclutching, simultaneously decreasing the rotation speeds of the first shaft and of the second shaft to a second idle rotation speed below the placarded zone, the simultaneously decreasing including clutching the first shaft to the load.

There is described a method of operating a multi-engine system of an helicopter. The multi-engine system has a first turboshaft engine having a first shaft, a second turboshaft engine having a second shaft, a gearbox having a clutch system, and a range of rotation speeds defined as a placarded zone. The method generally has: rotating the first and second shafts at a first idle rotation speed below the placarded zone when clutched to a load; increasing a rotation speed of the first shaft from the first idle rotation speed to a flight rotation speed above the placarded zone; unclutching the second shaft from the load during the increasing; and increasing a rotation speed of the second shaft to a second idle rotation speed when the second shaft is unclutched from the load, the second idle rotation speed above the placarded zone and below the flight rotation speed.

There is described a method of operating a multi-engine system of an helicopter. The multi-engine system has a first turboshaft engine having a first shaft, a second turboshaft engine having a second shaft, a gearbox having a clutch system, and a range of rotation speeds defined as a placarded zone. The method generally has rotating the first shaft at a flight rotation speed when clutched and rotating the second shaft at a first idle rotation speed when unclutched, the first idle rotation speed above the placarded zone; decreasing a rotation speed of the first shaft from the flight rotation speed to a given rotation speed within the placarded zone; decreasing a rotation speed of the second shaft to the given rotation speed; clutching the second shaft; and decreasing the rotation speeds of the first and second shafts to a second idle rotation speed below the placarded zone.

A method for controlling a clearance control valve of a turbomachine wherein, during a maneuver to increase engine speed in cruise phase, a command to reduce the opening of the clearance control valve is actuated by a Full Authority Digital Engine Control based on a change in the state of a step-climb signal provided by a flight management system in order to increase clearances at the tips of the turbomachine blades and an increase in the opening of the clearance control valve follows its reduction, at the expiry of either of the following two time limits: a first time limit starting at the change in the state of the step-climb signal and determined not to penalize the performance of the engine for too long and a second time limit starting at the end of the maneuver and determined as a function of a thermal time constant of the casing.

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.

A gas turbine engine comprises a relatively high pressure compressor coupled to a relatively high pressure turbine by a relatively high pressure shaft; a relatively low pressure compressor coupled to a relatively low pressure turbine by a relatively low pressure shaft rotatable independently of the high pressure shaft; a first combustor located downstream of the high pressure compressor and upstream of the high pressure turbine; and a second combustor located downstream of the high pressure turbine, and upstream of the low pressure turbine. The engine further comprises a coupling arrangement configured to selectively transfer torque between the high pressure shaft and the low pressure shaft.

In an asymmetric operating regime, a first engine is operating in an active mode to provide motive power to an aircraft while a second engine is operating in a standby mode and de-clutched from a gearbox of the aircraft. In response to an emergency exit request, the second engine's speed is increased, at a maximum permissible rate, to a re-clutching speed while increasing the first engine's power output at a maximum permissible rate. When the re-clutching speed is reached, the second engine's power output is increased at a maximum permissible rate. In response to a normal exit request, the second engine's speed is increased to the re-clutching speed at a rate lower than the maximum permissible rate. When the re-clutching speed is reached, the second engine's power output is increased at a rate lower than the maximum permissible rate.

A system and method for providing feedback for an aircraft-bladed rotor about a longitudinal axis and having an adjustable blade pitch angle. At least one position marker is provided at the rotor, extends along an axial direction, from a first end to a second end, and has varying magnetic permeability from the first end to the second end. At least one sensor is coupled to the rotor and configured for producing, as the rotor rotates about the longitudinal axis, at least one sensor signal in response to detecting passage of the at least one position marker. A control unit is communicatively coupled to the at least one sensor and configured to generate a feedback signal indicative of the blade pitch angle in response to the at least one sensor signal received from the at least one sensor.

A method of bowed rotor start response damping for a gas turbine engine is provided. A spring rate and a damping characteristic of one or more bearing supports in the gas turbine engine are selectively modified while a shaft of the gas turbine engine rotates below a speed which is adversely affected by a bowed rotor condition of the gas turbine engine.