Abstract: A method of determining at least one fuel characteristic of a fuel provided to a gas turbine engine of an aircraft includes making an operational change, the operational change being effected by a controllable component of a propulsion system of which the gas turbine engine forms a part, and being arranged to affect operation of the gas turbine engine, sensing a response to the operational change; and determining the at least one fuel characteristic based on the response to the operational change.

Abstract: A method of operating an aircraft including a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and proposing or initiating a change to a flight profile of the aircraft based on the at least one fuel characteristic. For example, intended route and/or altitude of the aircraft may be changed based on the one or more fuel characteristics.

Abstract: A gas turbine engine including a fuel delivery system arranged to provide fuel; a combustor arranged to combust at least a proportion of the fuel; a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough; and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. The method includes using at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger to heat the fuel so as to adjust the fuel viscosity to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: A gas turbine engine includes: a combustor that combust the fuel and having an exit, a combustor exit temperature (T40) is the average temperature of flow and a combustor exit pressure (P40) is the total pressure there; a turbine including a rotor having a leading edge and a trailing edge, and wherein a turbine rotor entry temperature (T41) is an average temperature of flow at the leading edge and a turbine rotor entry pressure (P41) is the total pressure there; and a compressor having an exit, wherein a compressor exit temperature (T30) is the average temperature of flow at the exit from the compressor and a compressor exit pressure (P30) is the total pressure there (all at cruise conditions). A method of determining at least one fuel characteristic includes changing a fuel supplied to the engine; and determining a change in a relationship between T30 or P30, T40 and T41, or of P40 and P41, respectively.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

Abstract: A method of operating an aircraft including a gas turbine engine and a plurality of fuel tanks arranged to provide fuel to the gas turbine engine, where at least two of the fuel tanks contain fuels with different fuel characteristics. The method includes obtaining a flight profile for a flight of the aircraft; and determining a fuelling schedule for the flight based on the flight profile and the fuel characteristics. The fuelling schedule governs the variation with time of how much fuel is drawn from each tank. Fuel input to the gas turbine engine may then be controlled in operation in accordance with the fuelling schedule.

Abstract: A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: A gas turbine engine for an aircraft includes a staged combustion system having pilot fuel injectors and main fuel injectors. The gas turbine engine further includes a fuel delivery regulator arranged to control delivery of fuel to the pilot and main fuel injectors, and a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being supplied to the staged combustion system. A controller is configured to determine a staging point defining the point at which the staged combustion system is switched between pilot-only operation and pilot-and-main operation, the staging point being determined based on the determined one or more fuel characteristics, the controller being configured to control the staged combustion system according to the determined staging point.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

Abstract: A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.

Abstract: A method of refuelling an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises obtaining an amount of energy required for an intended flight profile; obtaining a calorific value of fuel available to the aircraft for refuelling; calculating the amount of the available fuel needed to provide the required energy; and refuelling the aircraft with the calculated amount of the available fuel. The calculating the amount of the available fuel needed to provide the required energy may comprise obtaining an energy content of fuel already in the fuel tank and subtracting that from the determined amount of energy required for the intended flight profile.

Abstract: A gas turbine engine for an aircraft, including: staged combustion system having pilot fuel injectors and main fuel injectors, staged combustion system being operable in pilot-only range of operation and pilot-and-main range of operation; and fuel delivery regulator arranged to control delivery of fuel to pilot and main fuel injectors. Fuel delivery regulator arranged to receive fuel from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a different second fuel characteristic. The fuel delivery regulator is arranged to deliver fuel to the pilot fuel injectors during at least part of the pilot-only range of operation having a different fuel characteristic from fuel delivered to one or both of the pilot and main fuel injectors during at least part of the pilot-and-main range of operation. A method of operating a gas turbine engine is also disclosed.

Abstract: A power system for an aircraft includes one or more gas turbine engines arranged to burn a fuel so as to provide power to the aircraft; a plurality of fuel tanks each arranged to contain a fuel to be used to provide power to the aircraft; and a fuel manager. At least two of the fuel tanks contain different fuels, which have different proportions of a sustainable aviation fuel. The fuel manager is arranged to store information on the fuel contained in each fuel tank; and to control fuel supply so as to select a specific fuel accordingly to power at least the majority of operations on the ground. The fuel manager may additionally identify which tank contains the fuel with the highest proportion of a sustainable aviation fuel; and that fuel may be used to power at least the majority of operations on the ground.

Abstract: A method of checking refuelling of an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises: receiving an input of calorific value data for fuel provided to the aircraft on refuelling; independently determining at least one of: (i) the calorific value of fuel supplied to the gas turbine engine in use; and (ii) the calorific value of the fuel provided to the aircraft on refuelling; and providing an alert if the determined calorific value is inconsistent with the calorific value data input received

Abstract: The present application discloses a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft, the gas turbine engine having a combustor supplied with fuel from a fuel system, the method comprising: determining a mass of the fuel being supplied to the combustor; determining a corresponding volume of the fuel being supplied to the combustor; and determining one or more fuel characteristics based on the determined mass and volume. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A propulsion system for an aircraft comprises a gas turbine engine; a plurality of fuel tanks arranged to contain different fuels to be used to power the gas turbine engine, wherein the fuels have different calorific values; and a fuel manager. The fuel manager is arranged to store information on the fuel contained in each fuel tank and to control fuel input to the gas turbine engine in operation by selection of a specific fuel or fuel combination from one or more of the plurality of fuel tanks based on thrust demand of the gas turbine engine such that a fuel with a lower calorific value is supplied to the gas turbine engine at lower thrust demand.

Abstract: The present application provides a method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: passing UV-visual spectrum light through the fuel; measuring a transmittance parameter indicating the transmittance of light through the fuel; determining one or more fuel characteristics of the fuel based on the transmittance parameter; and communicating the one or more fuel characteristic to a control module of the gas turbine engine or the aircraft. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.