Rolls-royce Corporation

Paul William Ferra

Craig W Bemment

Benjamin J Keeler

Andrea Minelli

Martin Kenneth Yates

Alastair G Hobday

Kevin R McNally

David Michael Beaven

Andrew T Smith

Peter Swann

Christopher P Madden

John Richard Webster

Alan Robert Maguire

Timothy John Scanlon

Guy David Snowsill

Colin Young

Yi Wang

Lynne H Turner

John A Mylemans

Clive Peter Gravett

Nicolas M P Evanno

Barani P Gunasekaran

John Andrew Mylemans

Lynne Helen Turner

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core and driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators and a fuel supply system. The fuel supply system is arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least ten of the plurality of actuators.

Fueldraulic actuation

A gas turbine engine for an aircraft, comprising: a combustor; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank onboard the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least some fuel that has passed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the flow of fuel along the fuel return line, wherein the modulator valve is arranged to initiate return of fuel to the fuel tank when fuel having passed through the heat exchanger is at a temperature of at least 120° C. A method of operating a gas turbine engine is also disclosed.

Gas turbine fuel temperature management

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and arranged to be driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine includes a plurality of actuators and a fuel supply system arranged to supply fuel to fueldraulically drive at least one actuator of the plurality of actuators. The engine includes two fuel-oil heat exchangers.

Fuel supply system

A gas turbine engine includes engine core, combustor, compressor, and core shaft. The engine includes fan located upstream of engine core and arranged to be driven by core shaft. The engine includes a nacelle surrounding fan and engine core and defining a bypass duct located radially outside of engine core, where a bypass ratio, defined as ratio of mass flow rate of flow through bypass duct to mass flow rate of flow through core at cruise conditions, is at least 4. The engine includes plurality of actuators and fuel supply system arranged to supply fuel for combustion in combustor, and to supply fuel to fueldraulically drive at least one actuator of plurality of actuators. The fuel supply system is arranged to select between: causing fuel to actuate at least one actuator; and causing fuel to bypass the at least one actuator, based on sustainable aviation fuel—SAF—content of fuel.

Gas turbine engine

A gas turbine engine for an aircraft includes an engine core with a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core driven by the core shaft. The engine includes a nacelle surrounding the fan and engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio of the mass flow rate through the bypass duct to the mass flow rate through the core at cruise conditions, is at least 4. The engine includes an engine heat management system and an actuator configured to actuate at least one valve within that system. The engine includes a fuel supply system to supply fuel to fueldraulically drive the actuator. The actuator is configured to actuate said valve to enable non-binary position adjustment between an open and closed valve positions.

Fueldraulic heat management system

A gas turbine engine for an aircraft, comprising: a combustor; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least some fuel that has passed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the flow of fuel along the fuel return line such that a ratio of a temperature, in Kelvin, of fuel in the fuel tank to a temperature, in Kelvin, of fuel being delivered to the combustor is less than 0.56. Methods of operating a gas turbine engine are also disclosed.

Gas turbine fuel return line

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.

Aircraft operation

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.

Aircraft refuelling

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.

Fuel delivery

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.

Method of determining the calorific value of fuel provided to a gas turbine engine

There is provided a method of operating a gas turbine engine. The gas turbine engine comprises a staged combustor comprising an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method comprises providing fuel to the combustor having a calorific value of at least 43.5 MJ/kg. Also disclosed is a gas turbine engine.

Staged combustor

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.

A propulsion system for an aircraft includes a gas turbine engine and a fuel tank, wherein the fuel includes at least a proportion of a sustainable aviation fuel&#x2014;SAF&#x2014;having a density between 90% and 98% of the density, &#x3c1;, of kerosene and a calorific value between 101% and 105% the calorific value CV, of kerosene. The engine includes a combustor; and a fuel pump arranged to supply a fuel thereto at an energy flow rate, C, the pump being arranged to output fuel at a volumetric flow rate, Q, the percentage of fuel passing through the pump not provided to the combustor being referred to as a spill percentage. The fuel include X % SAF, where X % is in the range from 5% to 100%, and has a density, &#x3c1;, and a calorific value CV. The propulsion system is arranged so:  

Aircraft fuel system for gas turbine engine

A method of determining a calorific value of fuel supplied to a gas turbine engine of an aircraft comprises sensing at least one engine parameter during a first time period of aircraft operation during which the gas turbine engine uses the fuel; and, based on the at least one sensed engine parameter, determining a calorific value of the fuel. The sensing may be repeated such that the at least one engine parameter is monitored over time. The gas turbine engine may be a propulsive gas turbine engine of the aircraft or a gas turbine engine of an auxiliary power unit of the aircraft.

Engine parameters

There is disclosed a gas coolerfor providing high-pressure sealing gas to a bearing chamber. The cooler comprises a turbine; a turbine inletarranged to receive gas to drive the turbine and a turbine outletarranged to deliver gas output from the turbine; a compressorarranged to be driven by the turbine; a compressor inletarranged to receive gas to be compressed by the compressor and a compressor outletarranged to deliver gas output from the compressor; and a cooler outletin fluid communication with the turbine outletand the compressor outletso as to deliver high-pressure sealing gas comprising gas merged from the turbine outletand the compressor outlet

Growing community of inventors

Derby, United Kingdom

Paul William Ferra