7 results about "Engine configuration" patented technology

The application discloses a three-engine / two-engine configuration helicopter engine power balancing method and device, which utilizes an engine power control system to realize balancing; the method comprises the following steps: a cockpit power balancing switch is in a torque balancing gear by default; when a pilot observes engine operation, if a difference between gas turbine outlet temperatures ITT of two engines / three engines is less than a preset temperature threshold and a percentage difference between Ngs is less than a preset speed threshold, the cockpit power balancing switch is not operated, and a current torque balancing state is maintained; when the pilot observes that the difference between the ITT of different engines is greater than the preset temperature threshold and the percentage difference between the Ngs is greater than the preset speed threshold, it is considered that the performance difference between the different engines is large, and then the cockpit power balancing switch is manually switched to an Ng balancing gear; after the engine electronic controller receives the Ng balancing gear information, Ng balancing is performed.

A gas turbine engine 10 for an aircraft comprising a rich burn quick quench, lean burn (RQL) combustor 16 having a number of spray nozzles 403, 404, in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range of 2-6, with the engine configured to provide fuel comprising sustainable aviation fuel to the spray nozzles. The engine operation is characterised by specific non-volatile particulate emission ratio beign related to the bypass ratio. The fuel spray nozzles may comprise one or more duplex nozzles and one more single flow nozzles and preferably the combustor comprises 10-14 duplex fuel spray nozzles and 4-8 signle flow fuel spray nozzles. The duplex fuel spray nozzles are arranged in groups about the circumference of the combustor and optionally wherein the groups duplex fuel spray nozzles comprise at least two groups arraned dimametrically opposite each other. A corresponding method is also claimed. Fig 6

Propulsion assembly (100) for an aircraft, particularly for a multi-engine helicopter, comprising at least one first engine (1) and a second engine (2) configured to operate in at least one standby mode, a primary air circulation device (30) configured to bleed air from the first engine (1) via a first bleed channel (310) and / or from the second engine (2) via a second bleed channel (320) to convey it to equipment of the propulsion assembly (100), and a secondary air circulation device (40) configured, when one of the first or the second engine (1, 2) operates in standby mode, to bleed air originating from the other of the first or the second engine (1, 2) not operating in standby mode and to convey it to the first or the second engine (1, 2) operating in standby mode.

A cooling system for an aircraft comprises: an air duct in the aircraft, the air duct comprising an inlet and an outlet; a drive shaft configured to rotate around a first axis and drive a propeller in the aircraft; an engine configured to rotate the drive shaft; a heat exchanger thermally coupled to the engine and mechanically coupled to the air duct; a rotor coupled to the drive shaft; a stator adjacent to the rotor; and a fan located in the air duct and configured to rotate around a second axis, the second axis being noncollinear with the first axis. Rotation of the rotor with respect to the stator generates an electric current sent to the fan through at least one wire to cause the fan to rotate around the second axis. The fan is configured to drive an airflow from the inlet to the outlet.

A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises:a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is defined as:E⁢Iidle×Wf,idleE⁢ImaxTO×Wf,maxTOwhere: EIidle is the nvPM emissions index in mg / kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the nvPM emissions index in mg / kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg / s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg / s at around 100% available thrust for the given operating conditions. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed are methods of operating a gas turbine engine.