The invention relates to a device for assisting a propulsion system of a single-engine helicopter, comprising an engine connected to a power transmission gearbox (15) suitable for rotating a rotor (88) of the helicopter, characterised in that it comprises: a turbine (18) for driving in rotation an output shaft (34) that is mechanically connected to said power transmission gearbox (15); and controlled means (16) for supplying said drive turbine (18) with pressurised fluid in order to allow said turbine (18) to transform the power from said pressurised fluid into mechanical power for rotating said output shaft (34).

A method for eliminating rotational stall in a compressor of a turbine engine, includes automatically detecting surge in the turbine engine; automatically shutting-down the turbine engine; in the event surge is detected, automatically restoring a surge margin; and automatically re-igniting the turbine machine.

An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.

A direct metering architecture is provided having a metering pump and a servo pump wherein the metering and servo pumps are driven by an engine shaft by way of a gearbox transmission. The system reduces wasted horsepower previously occurring with oversized fixed displacement pumps, reduces instances of engine flameout and reduces the amount of heat added to fuel which ultimately improves engine oil cooling.

The invention deals with the restriction on the mode change of an engine due to latching in the case of restarting the engine that has stopped in flight. A FADEC that controls an engine of an aircraft in accordance with a mode, includes a latch operating section that latches the mode, and an in-parking/in-flight determining section that determines whether the aircraft is on the ground or flying off the ground. The latch operating section operates when the in-parking/in-flight determining section determines the aircraft is on the ground, and does not operate when the in-parking/in-flight determining section determines that the aircraft is in fight.

A method of controlling fuel flow for combustion in a gas turbine engine comprising a low pressure spool, a high pressure spool and a fuel metering valve is disclosed. The method comprises scheduling the fuel metering valve in dependence upon the speed of the high pressure spool.

Systems and methods for conditionally performing engine operational tests for a turbine engine are provided. A system comprising at least one processor can be configured to obtain sensor data associated with at least one sensor for a turbine engine. The sensor data identifies a current fuel flow associated with the turbine engine. The system can determine a predicted fuel flow of the turbine engine based at least in part on the current fuel flow and a fuel flow reduction associated with an engine operational test. The system can compare the predicted fuel flow to at least one threshold. The system can selectively initiate the engine operational test based on comparing the predicted fuel flow to the at least one threshold.

A propulsion system includes a gas turbine designed so that a combustion chamber can be ignited in a first ignition range of rotational speeds of a compressor shaft. The system further includes a control device designed to control an electric starter to accelerate the compressor shaft and, when the compressor shaft is accelerated, to control an attempt to ignite the combustion chamber. The gas turbine is designed so that the combustion chamber can be ignited in a second ignition range which is higher than the first ignition range, but not between these two ignition ranges, and the ignition attempt is carried out in the second ignition range.