# US 12,601,486 B2 · Hydrogen fuelled gas turbine engine > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 12,601,486 B2 - Document ID: US12601486B2 - Status: Granted - Kind: B2 - Application number: 18/746,568 - Filed: 2024-06-18 - Published: 2025-07-10 - Granted: 2026-04-14 - Earliest claimed: 2023-07-10 - Assignees: ROLLS-ROYCE plc - Applicants: ROLLS-ROYCE plc - Inventors: Nicholas Grech; Jacopo Tacconi; Chloe J. Palmer - Web page: https://patentgrep.com/patent/US12601486B2 ## CPC - F02C 3/20: using a special fuel, oxidant, or dilution fluid to generate the combustion products - F23R 3/28: characterised by the fuel supply (burners F23D) - F01D 15/10: Adaptations for driving, or combinations with, electric generators - F02C 3/22: the fuel or oxidant being gaseous at standard temperature and pressure (F02C3/28 takes precedence) - F02C 7/224: Heating fuel before feeding to the burner - F02C 9/40: specially adapted to the use of a special fuel or a plurality of fuels - F05D 2220/323: for aircraft propulsion, e.g. jet engines - F05D 2220/76: an electrical generator - F05D 2260/213: by the provision of a heat exchanger within the cooling circuit - F05D 2300/121: Aluminium - F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00) - F23R: GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS (fluidised bed combustion apparatus specially adapted for operation at superatmospheric pressures F23C10/16) - F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS ## Abstract A gas turbine engine is configured to combust hydrogen fuel. The gas turbine engine comprises a fuel system comprising: a fuel pump configured to provide pressurised fuel to a core combustor of the gas turbine engine; a recuperator configured to heat fuel from combustion gases downstream of the core combustor; a fuel turbine downstream in fuel flow of the recuperator. The fuel turbine being configured to be driven by heated fuel and configured to power the fuel pump. Wherein the fuel turbine has an expansion ratio between 1.4 and 2.5. ## Claims ### Claim 1 A method of operating a hydrogen-fueled gas turbine engine, the method comprising: operating a fuel pump to provide pressurised fuel to a core combustor of the gas turbine engine; operating a recuperator to heat fuel using combustion gases downstream of the core combustor; operating a single stage fuel turbine downstream in fuel flow of the recuperator, the single stage fuel turbine being configured to be driven by heated fuel and configured to power the fuel pump, wherein the single stage fuel turbine comprises an aluminium alloy rotor and has an expansion ratio between 1.4 and 1.6; and operating a pre-heater configured to heat hydrogen fuel, wherein: the pre-heater is provided in hydrogen fuel flow between the fuel pump and the recuperator. ### Claim 2 A gas turbine engine configured to combust hydrogen fuel, the gas turbine engine comprising a fuel system comprising: a fuel pump configured to provide pressurised fuel to a core combustor of the gas turbine engine; a recuperator configured to heat fuel using combustion gases downstream of the core combustor; a single stage fuel turbine downstream in fuel flow of the recuperator, the single stage fuel turbine being configured to be driven by heated fuel and configured to power the fuel pump; and a pre-heater configured to heat hydrogen fuel, wherein: the pre-heater is provided in hydrogen fuel flow between the fuel pump and the recuperator; and the single stage fuel turbine comprises an aluminium alloy rotor and has an expansion ratio between 1.4 and 1.6. ### Claim 3 (depends on 2) The gas turbine engine according to claim 2, wherein the single stage fuel turbine is configured to mechanically drive the fuel pump. ### Claim 4 (depends on 2) The gas turbine engine according to claim 2, wherein the single stage fuel turbine is configured to drive an electrical generator. ### Claim 5 (depends on 2) The gas turbine engine according to claim 2, wherein the fuel pump is coupled to an electric motor. ### Claim 6 (depends on 2) The gas turbine engine according to claim 2, wherein the fuel pump is configured to be driven by a core gas turbine engine shaft. ### Claim 7 (depends on 2) The gas turbine engine according to claim 2, wherein the single stage fuel turbine is configured to drive a core gas turbine engine shaft. ### Claim 8 (depends on 2) The gas turbine engine according to claim 2, wherein the fuel system is configured to provide fuel to a fuel injector of the gas turbine engine at a fuel delivery pressure to compressor air delivery pressure ratio of at least 1.5:1. ### Claim 9 (depends on 2) The gas turbine engine according to claim 2, wherein a compressor air delivery pressure ratio at maximum take-off conditions is between 10:1 and 50:1, between 20:1 and 45:1, or approximately 40:1. ### Claim 10 (depends on 2) The gas turbine engine according to claim 2, wherein the fuel system comprises a fuel storage unit configured to store cooled hydrogen, and may be configured to store liquid hydrogen or cryogenically cooled compressed gaseous or supercritical hydrogen. ### Claim 11 (depends on 10) The gas turbine engine according to claim 10, wherein the fuel storage unit is configured to store hydrogen as a liquid at a temperature of less than 30K or at a temperature of less than 25K. ### Claim 12 (depends on 10) The gas turbine engine according to claim 10, wherein the fuel storage unit is configured to store hydrogen at a pressure of between 1 and 4 Bar. ### Claim 13 (depends on 2) The gas turbine engine according to claim 2, wherein the fuel pump is configured to; provide a pressure ratio at maximum take-off conditions of between 10:1 and 60:1, provide a pressure ratio at maximum take-off conditions of between 15:1 and 40:1, or provide a pressure ratio at maximum take-off conditions of approximately 20:1. ### Claim 14 (depends on 2) The gas turbine engine according to claim 2, wherein the recuperator is configured to increase the fuel temperature to between 270K and 600 K or to between 550K and 600K. ### Claim 15 (depends on 2) An aircraft comprising the gas turbine engine of claim 2. ### Claim 16 (depends on 2) The gas turbine engine according to claim 2, wherein the pre-heater comprises an electrical heater. ### Claim 17 (depends on 16) The gas turbine engine according to claim 16, wherein the pre-heater comprises an auxiliary combustor. ### Claim 18 (depends on 2) The gas turbine engine according to claim 2, wherein the pre-heater is configured to: enable starting the gas turbine engine when insufficient heat is generated by the recuperator, avoid overcooling of the recuperator, or control fuel temperature independent of fuel flow rates or engine exhaust gas temperatures.