Rolls-Royce Patente

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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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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).

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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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A combustion system (120) comprising: a combustion chamber (122) extending in an axial direction between an inlet (124) and an outlet (126), the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port (128) configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system (130) for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points (132) disposed downstream of the fuel injection port (128), and distributed radially and circumferentially around the combustion chamber (122), wherein each electrical plasma initiation point (132) comprises a pair of electrodes (132a;132b) configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

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A fuel system (205) for an aircraft propulsion system (103) comprising a gas turbine engine. The fuel system includes a fuel tank (104) configured to store compressed gaseous hydrogen pressures greater than 100 Bar at a temperature below 200 Kelvin and a tank heating system (224, 232, 234, 242, 302, 406) configured to raise the temperature of hydrogen fuel within the hydrogen tank (104). The tank heating system includes a primary tank heater comprising an oil to fuel heat exchanger (242) and an auxiliary tank heater comprising one or more of a recuperator (224), an electric heater (232), a gas turbine engine core compressor bleed-air powered heater (302, 402), and an auxiliary combustor (306).

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