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 gas turbine engine for an aircraft includes: an engine core; a fan; turbomachinery bearings; a power gearbox; and a heat management system for providing lubrication and cooling to the power gearbox and turbomachinery bearings, at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the heat management system is configured to provide a first amount of heat and a second amount of heat such that a ratio of a first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed at an environment temperature of ISA +40° C. to the first proportion at an environment temperature of ISA ?69° C. is in the range of from 1.5 to 4.5.

Abstract: A gas turbine engine for an aircraft includes: an engine core including a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor, wherein the core shaft has a core shaft maximum take-off speed in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm; a fan; turbomachinery bearings; a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the gearbox and turbomachinery bearings, at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink.

Abstract: A geared gas turbine engine includes a heat management system configured to provide lubrication and cooling to a power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, an air-lubricant heat exchanger to dissipate a first amount of heat, and a fuel-lubricant heat exchanger to dissipate a second amount of heat wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that at cruise conditions a proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is in the range of from 0.35 to 0.80.

Abstract: A gas turbine engine for an aircraft includes: an engine core including a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor; a fan including a plurality of fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings.

Abstract: An aircraft gas turbine engine includes: an engine core including a compressor, combustor, turbine, and a core shaft connecting the turbine to the compressor; a fan including fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox to drive the fan at a lower rotation speed than the turbine; and a heat management system providing lubrication and cooling to the power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel.

Abstract: A method of operating a gas turbine engine for an aircraft including: a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor; a fan; turbomachinery bearings; a power gearbox; and a heat management system configured to provide lubrication and cooling to the gearbox and turbomachinery bearings. The method includes operating the heat management system to provide a first amount of heat and a second amount of heat such that a first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed is in the range of from 0.25 to 0.70; and operating the fan at cruise condition to provide a fan pressure ratio in the range of from 1.35 to 1.43.

Abstract: A combination of a gas turbine engine and power electronics, includes an engine core and oil circuit to cool and lubricate bearings of the engine core, and a fuel circuit for supplying fuel to the combustor. The fuel circuit includes a low pressure pump for pressurising the fuel to a low pressure, and a high pressure pump to receive the low pressure fuel and increase the pressure to a high pressure for supply to a fuel metering system and the combustor. The engine includes a fuel-oil heat exchanger having a fuel side on the fuel circuit between an outlet of the low pressure pump and an inlet of the high pressure pump, and an oil side on the oil circuit to transfer heat from the oil circuit to the fuel circuit. The power electronics transfers heat to a cooling flow formed by a portion of the low pressure fuel.

Abstract: A gas turbine engine for an aircraft includes: an engine core; fan; turbomachinery bearings; power gearbox; and a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings and including at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the heat management system is configured to provide a first proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed in the range of 0.25 to 0.70, and a second proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 65% of the core shaft maximum take-off speed in the range of 0.60 to 1.

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 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: The invention relates to a planetary gearbox device with a static ring gear, a rotatable planet carrier, a plurality of planet gears, each planet gear being connected to a bearing device, wherein at least one of the bearing devices is coupled to at least one lubricant scooping device and/or at least one lubricant reservoir device for collecting lubricant, in particular oil coming from the at least one bearing device, and at least one lubricant channel is for allowing a flow of the collected lubricant towards a lubrication location and/or a lubricant supply for the planetary gearbox device.

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 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: 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 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: A combination of a gas turbine engine and power electronics, includes an engine core and oil circuit to cool and lubricate bearings of the engine core, and a fuel circuit for supplying fuel to the combustor. The fuel circuit includes a low pressure pump for pressurising the fuel to a low pressure, and a high pressure pump to receive the low pressure fuel and increase the pressure to a high pressure for supply to a fuel metering system and the combustor. The engine includes a fuel-oil heat exchanger having a fuel side on the fuel circuit between an outlet of the low pressure pump and an inlet of the high pressure pump, and an oil side on the oil circuit to transfer heat from the oil circuit to the fuel circuit. The power electronics transfers heat to a cooling flow formed by a portion of the low pressure fuel.

Abstract: A combination of a gas turbine engine and a power electronics for powering aircraft and/or engine systems. The engine includes an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor, and a fuel circuit for supplying a fuel flow to the combustor. The power electronics is configured to transfer heat produced by the power electronics to a cooling flow formed by a portion of the fuel flow. The fuel circuit is configured to circulate the cooling flow in a loop during selected engine conditions such that the cooling flow transfers heat from the power electronics to a phase change material located on the loop. The phase change material has a phase change temperature at a predetermined limiting temperature whereby the phase change material stores heat from the cooling flow to prevent the power electronics exceeding the limiting temperature.

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: A method of controlling a propulsion system of an aircraft, the propulsion system comprising a gas turbine engine arranged to be powered by a fuel and at least one variable inlet guide vane—VIGV, comprises obtaining at least one fuel characteristic of the fuel being provided to the gas turbine engine; and making a change to scheduling of the at least one VIGV based on the at least one obtained fuel characteristic.