Rolls-Royce Corporation

Goodrich Control Systems Limited

Martin Kenneth Yates

Craig W Bemment

Benjamin J Keeler

Andrea Minelli

Peter Swann

Christopher P Madden

Alastair G Hobday

David Michael Beaven

Paul William Ferra

Kevin R McNally

Andrew T Smith

David F Brookes

Daniel James Bickley

Michael Griffiths

Mihir Desai

Glyn Edward Dixon

Philip Laurence Elliott

Luke Ashton

Geir Leivseth

Stephanie E Watkins

Barani P Gunasekaran

Stephanie Watkins

A gear pump assembly includes an inlet configured to receive a fluid, an outlet, a main drive shaft, a secondary gear stage, and a primary gear stage configured to pump the fluid from the inlet to the outlet. The secondary gear stage includes a secondary driver gear driven by the main drive shaft, a secondary driver bearing block disposed adjacent to and supporting the secondary driver gear, a secondary driven gear meshed with and driven by the secondary driver gear, a secondary driven bearing block disposed adjacent to and supporting the secondary driven gear, and a driver bore portion receiving the secondary driver bearing block therein. A fluid passage fluidly communicates the primary gear stage with the driver bore portion.

Gear pump assembly

A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/mof heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Gas turbine performance

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 fuel system for a gas turbine engine configured to combust hydrogen fuel. The fuel system includes a main fuel conduit, a fuel pump configured to operate on hydrogen within fuel conduit to provide pressurised fuel to core combustor of gas turbine engine, an auxiliary combustor downstream in fuel flow of fuel pump, and configured to combust a portion of fuel diverted from main fuel conduit and to heat remainder of fuel in main fuel conduit, and a fuel turbine downstream in fuel flow of the auxiliary combustor. The fuel turbine is configured to be driven by heated fuel from auxiliary combustor and configured to power the fuel pump. The fuel system comprises a turbine inlet valve provided upstream in hydrogen fuel flow of the fuel turbine and configured to control mass flow rate and/or pressure of fuel flowing into fuel turbine. 1. A method of operation is also described.

Hydrogen fuelled gas turbine engine

A bearing block adapted for use in a gear pump assembly. The bearing block includes a bearing surface configured to face one or more gears. The bearing block further includes a bearing block body formed with a recess. The bearing block further includes an insert located within the recess. The insert includes an insert body including an outer surface. The insert body has a tapered shape that tapers outwardly from the bearing surface. The insert further includes a plurality of ridges disposed on the outer surface. The plurality of ridges defines one or more grooves therebetween. The recess includes a tapered shape complementary to the tapered shape of the insert body, such that the insert body engages with the bearing block body.

Bearing block for a gear pump assembly

A method of operating a gas turbine engine having an engine core having a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located 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 heat exchange system having at least one fuel-oil heat exchanger arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located downstream of the at least one fuel-oil heat exchanger. The method comprises controlling the heat exchange system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.

Heat transfer

A gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/mof heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Aircraft heat management

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 operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; at least one bearing supporting the fan shaft; at least one auxiliary system using some of the fuel; an oil loop system to supply oil to the bearing; and a heat exchange system including: a primary fuel-oil heat exchanger through which oil in the oil loop system and substantially all the fuel flow to transfer heat between the oil and the fuel; and a secondary fuel-oil heat exchanger through which oil in the oil loop system and a portion of the fuel flow, to transfer heat between the oil and the fuel; and a fuel pump; wherein, under cruise conditions, fuel temperature on leaving the second heat exchanger is 120° C. to 200° C.

Gas turbine fuel temperature

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 mm/s on entry to the combustor at cruise conditions.

Fuel viscosity

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

A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor to combust fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; a bearing supporting the fan shaft; an oil loop system supplying oil to the bearing; and a heat exchange system including: an air-oil heat exchanger through which oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow to transfer heat between the oil and the fuel; and a bypass pipe to allow a proportion of the oil to flow past one of the air-oil and the fuel-oil heat exchanger; and a bypass valve to allow the proportion of the oil sent through the bypass pipe to be varied.

Growing community of inventors

East Haddon, United Kingdom

Martin Kenneth Yates