Abstract: A fuel system for a hydrogen fueled aircraft propulsion system comprises first and second hydrogen fuel tanks configured to store liquid hydrogen, first fuel line and second fuel lines configured to supply hydrogen from the first hydrogen fuel tank to a combustor of a first gas turbine engine and from the second hydrogen fuel tank to a combustor of a second gas turbine engine respectively. First and second fuel pumps are provided, each being configured to pump fuel in a respective first and second fuel line. First and second fuel heaters are provided, each being configured to heat fuel in a respective first and second fuel line. A fuel cross-feed fuel line is provided, which is configured to transfer fuel between the first and second fuel lines. The fuel cross-feed line is provided downstream in fuel flow of the first and second fuel heaters.

Abstract: A propulsion system comprises a propulsive hydrogen-burning gas turbine engine, a fuel cell stack auxiliary power unit (APU) and a first tank arranged to store liquid hydrogen with an ullage. A first fuel line includes a first pump and a first vaporiser and transports hydrogen from the first tank to combustion apparatus of the engine during operation of the propulsion system. A second fuel line includes a second fuel pump and a second vaporiser and transports hydrogen from the first tank to the fuel cell stack APU. A duct connects the second fuel line at a position thereon between the second vaporiser and the fuel cell stack to the ullage of the first tank, providing for pressure in the first tank to be maintained therein as liquid hydrogen within the first tank is depleted, thus avoiding cavitation of liquid hydrogen within the first fuel pump.

Abstract: A propulsion system comprises a propulsive hydrogen-burning gas turbine engine, a first tank storing liquid hydrogen with an ullage and a first fuel line including a fuel pump and a vaporiser, the first fuel line providing gaseous hydrogen to the engine during operation of the system. A second tank storing gaseous hydrogen is coupled by a second fuel line to the first fuel line at a position thereon between the vaporiser and the engine, providing for engine start-up (when the vaporiser is inoperative) and power-boosting during operation of the system. A duct connects gaseous hydrogen within the second tank to the ullage in the first tank in order maintain pressure in the first tank as liquid hydrogen within it is depleted, preventing cavitation of liquid hydrogen within the fuel pump.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? exhaust ? nozzle ? pressure ? ratio core ? exhaust ? nozzle ? pressure ? ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? exhaust ? nozzle ? pressure ? ratio core ? exhaust ? nozzle ? pressure ? ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.

Abstract: A gas turbine engine for an aircraft having an engine core configured with a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, the fan comprising a plurality of fan blades, with a nacelle surrounding the gas turbine engine, and a bypass duct outlet guide vane extending radially across the bypass duct between an outer surface of the engine core and the inner surface of the nacelle. An outer wall axis is defined joining the radially outer tip of the trailing edge of the bypass duct outlet guide vane and the rearmost tip of the inner surface of the nacelle. An outer bypass duct wall angle is defined as the angle between the outer wall axis and the centreline, and the outer bypass duct wall angle is in a range from ?15 to ?2.5 degrees.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? exhaust ? nozzle ? pressure ? ratio core ? exhaust ? nozzle ? pressure ? ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: bypass ? ? exhaust ? ? nozzle ? ? pressure ? ? ratio core ? ? exhaust ? ? nozzle ? ? pressure ? ? ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

Abstract: There is provided an air intake system for providing air to a tip clearance control system. The air intake system comprises a ram-air intake having a scoop portion and a body portion. The body portion of the ram-air intake houses a heat exchanger.

Abstract: A gas turbine engine for an aircraft includes an engine core having a core length and comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius, wherein a ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in a range from 1.2 to 2.0; and wherein the engine core length is in a range from 150 cm to 320 cm.

Abstract: A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

Abstract: A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the ? ? maximum ? ? take - off ? ? rotational ? ? speed ? ? of ? ? the ? ? fan fan - turbine ? ? radius ? ? difference ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the ? ? maximum ? ? take ? - ? off ? ? rotational ? ? speed ? ? of ? ? the ? ? fan fan ? - ? turbine ? ? radius ? ? difference ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan defining a fan tip radius at the fan face; wherein a downstream blockage ratio is defined as: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial ? ? location of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage a ? ? distance ? ? f ? rom ? ? a ? ? ground ? ? plane ? ? to ? ? the ? ? wing ? and wherein an engine blockage ratio of: ( 2 × the fan tip radius/the engine length ) the ? ? downstream ? ? blockage ? ? ratio is in the range from 2.5 to 4.

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: the ? ? turbine ? ? diameter ? ? at ? ? an ? ? axial location ? ? of ? ? the ? ? lowest ? ? pressure ? ? rotor ? ? stage a ? ? distance ? ? from ? ? a ? ? ground ? ? plane ? ? to ? ? the ? ? wing is in the range from 0.2 to 0.3.