Rolls-Royce Patente

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A gas turbine engine for an aircraft. The gas turbine engine (10) comprises: a combustor (16), comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustor (16) is operable in a condition in which each of the fuel spray nozzles of the first subset (124A) of fuel spray nozzles (124) is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of fuel spray nozzles (124), wherein a ratio of the number of fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:3 to 1:6. An MTO nvPM emissions index ratio-modified fuel flow is defined as: EImaxTO,SAFEImaxTO,FF×Wf,maxTO where: EI_maxTO,SAF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine (10) when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles (124) comprises a sustainable aviation fuel (SAF); EI_maxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine (10) when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles (124) is a fossil-based hydrocarbon fuel; and W_f,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles (124) in kg/s when the gas turbine engine (10) is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine (10) in kg/s is less than 4. The gas turbine engine (10) is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles (124). Also disclosed are methods of operating the gas turbine engine.

A gas turbine engine (10) for an aircraft is disclosed. The gas turbine engine comprising: a combustor (16), comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustor (16) is operable in a condition in which each of the fuel spray nozzles of the first subset (124A) of fuel spray nozzles (124) is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of fuel spray nozzles (124), wherein a ratio of the number of fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:2 to 1:5. A MTO nvPM emissions index ratio is defined as: EImaxTO,SAFEImaxTO,FF where: EI_maxTO,SAF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine (10) when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles (124) comprises a sustainable aviation fuel (SAF); and EI_maxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine (10) when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles (124) is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine (10) is less than 1. The gas turbine engine (10) is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles (124). Also disclosed is a method of operating the gas turbine engine.

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An annular brush seal (100) comprises a bristle pack (114) formed of bristles (112) and a bristle pack support (116), which supports the bristle pack (114). The bristle pack (114) has an upstream face (117), a downstream face (118) and a sealing face (119). The sealing face (119) comprises an upstream portion (119a) and a downstream portion (119b) that is adjacent to the upstream portion (119a). A length of at least a fraction of the bristles (112) of the upstream portion (119a) is shorter than a length of the bristles (112) of the downstream portion (119b) to form a circumferentially repeating pattern (160) at an interface between the upstream portion (119a) of the sealing surface (119) and the upstream surface (117).

An aircraft comprises a turbofan engine 201A, a cabin blower system 210 and a PEM fuel cell stack 242. The cabin blower system 210 includes a cabin blower compressor 216 arranged to be driven by mechanical power derived from a shaft 205 of the turbofan engine 201A. A ducting system 220 delivers respective portions of the mass flow rate of compressed air output by the cabin blower system to a cabin space 230 of the aircraft via an air-conditioning unit 228 and to the cathode input 241 of the fuel cell stack 242. An electric motor 246 receives electrical power from the fuel cell stack 242 and provides mechanical power to the compressor 216 via a drive arrangement 211, 214, 215 of the cabin blower system. Power associated with excess capacity of the cabin blower system is recovered to the cabin blower compressor, thereby mitigating or eliminating the fuel consumption penalty on the turbofan engine associated with the excess capacity.

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There is provided an air pressurisation system 400 comprising: a blower compressor 420 configured to be mechanically coupled to a spool 440 of a gas turbine engine and configured to receive a first inlet flow of air from a bypass duct 22 of the gas turbine engine; an air treatment apparatus 460 comprising an air cycle line 470 configured to condition air received from the blower compressor 420, and a bypass line 480 for receiving air from the blower compressor, and configured to bypass the air cycle line 470, wherein the air treatment apparatus 460 is configured to mix air from the air cycle line 470 with air from the bypass line 480 to control a temperature of air for discharge to an airframe system 450; and a core bleed line 410 configured to provide a second inlet flow of air from a compressor 412 of the gas turbine engine to the air treatment apparatus 460 in an augmented air supply mode of the air pressurisation system 400; wherein the core bleed line 410 is configured to provide the second inlet flow of air directly to the bypass line 480 or mix the second inlet flow of air with air from the blower compressor 420 upstream of the bypass line 480.

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A method of operating a fuel system (18) of a hydrogen fuelled aircraft propulsion system (20) comprises exposing one or more fuel system components (12, 26, 34, 36, 38, 40, 48, 50, 52, 54) to a hydrogen embrittlement inhibiting gas comprising one of oxygen and carbon monoxide at a concentration above an embrittlement inhibition concentration.

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