Abstract: A novel configuration for an axial flow gas turbine engine for aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and combustor volume (in litres) which when divided by fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit an electric motor to be located within fan hub diameter.

Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine is positioned downstream of the fan assembly and is connected to the turbine module. The fan assembly includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The turbine module includes a lowest pressure turbine stage having a row of rotor blades. The gas turbine engine has a fan tip axis that joins a radially outer tip of the leading edge of one of the plurality of fan blades of the highest pressure fan stage, and the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage. The fan tip axis lies in a longitudinal plane which contains a centreline of the gas turbine engine. The fan axis angle is between 11 and 20 degrees.

Abstract: A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit electric motor to be located within fan hub diameter.

Abstract: A novel configuration for axial flow gas turbine engine for aircraft has an engine core having a core length and including first turbine, axial compressor, and drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.

Abstract: A cooling system for an aircraft comprises a gas turbine engine, an ancillary apparatus, and a heat exchanger. The gas turbine engine comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate an electrical power PEM1 (W). The heat exchanger is configured to transfer a total waste heat energy Q (W) generated by the gas turbine engine and the ancillary apparatus, to an airflow passing through the heat exchanger, and a ratio S of: S = ( Total ? Electrical ? Power ? Generated = P EM ? 1 ) ( Total ? Heat ? Energy ? Rejected ? to ? Airflow = Q ) is in a range of between 0.50 and 5.00.

Abstract: An gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, together with a first electrical machine rotationally connected to the turbine module. The combustor module has a combustor volume V (cm3). In use, at a full power condition, the gas turbine engine has a maximum corrected core flow Q (m3/sec), and a ratio T of: T = ( Maximum ? Corrected ? Core ? Flow = Q ) ( Combustor ? volume = V ) is in a range of between 450 and 2,500.

Abstract: An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.

Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine is positioned downstream of the fan assembly and is connected to the turbine module. The fan assembly includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The turbine module includes a lowest pressure turbine stage having a row of rotor blades. The gas turbine engine has a fan tip axis that joins a radially outer tip of the leading edge of one of the plurality of fan blades of the highest pressure fan stage, and the radially outer tip of the trailing edge of one of the rotor blades of the lowest pressure turbine stage. The fan tip axis lies in a longitudinal plane which contains a centreline of the gas turbine engine. The fan axis angle is between 11 and 20 degrees.

Abstract: A gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a total electrical power PEM1 (W), and the gas turbine engine is configured to generate a total shaft power PSHAFT (W); and a ratio R of: R = Total Shaft Power = P S H A F T Total Electrical Power Generated = P E M 1 is in a range of between 0.005 and 0.020.

Abstract: A cooling system for an aircraft comprises an apparatus, a heat exchanger, and a vapour cycle machine. The apparatus comprises a first fluid being circulated to provide cooling to the apparatus, with the heat exchanger being configured to transfer waste heat energy from the first fluid to a second fluid. The second fluid has a temperature T2 (°C), and the vapour cycle machine is configured to increase a temperature T1 (°C) of the first fluid, to a temperature greater than T2 (°C).

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes fan blades defining a fan diameter. The heat exchanger module is in communication with the fan assembly by an inlet duct, and the heat exchanger module further includes radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between hollow vanes. Each hollow vane accommodates at least one heat transfer element to transfer heat from a first fluid contained within the or each heat transfer element to a corresponding vane airflow passing through the hollow vane and over a surface of the or each heat transfer element. Each hollow vane further includes a flow modulator configured to regulate airflow in proportion to total airflow entering the heat exchanger module in response to a user requirement.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

Abstract: A turbofan gas turbine engine includes heat exchanger module, fan assembly, compressor, turbine and exhaust modules. The fan includes a plurality of fan blades. The heat exchanger in fluid communicates with the fan assembly by an inlet duct, and the heat exchanger includes a plurality of radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger is divided between a set of vane airflows and a set of channel airflows. Each vane airflow has a vane mass flow rate FlowVane, and each channel air flow has a channel mass flow rate FlowChan. Each hollow vane includes, an inlet, heat transfer, and exhaust portions, with the inlet portion comprising a diffuser element and the heat transfer portion including at least one heat transfer element. The diffuser element causes FlowVane to be lower than FlowChan.

Abstract: A gas turbine engine comprising: a combustor configured to initiate combustion; and a turbine comprising a stator vane ring defining a plurality of passageways between adjacent vanes; wherein at least one of the passageways is provided with a restrictor which defines a temporary gas washed surface for the stator vane ring and is configured to be ablated upon initiation of combustion to reveal an operational gas washed surface of the stator vane ring. A method of starting a gas turbine engine is also described.

Abstract: A supersonic jet aircraft and a method of operating the same. The supersonic jet aircraft having at least three turbofan engines and an engine management computer. A first engine of the at least three turbofan engines is configured to be de-activatable during flight to move from an operational state in which it provides thrust to an operational state in which it stops providing thrust. Other engines of the at least three turbofan engines are configured to provide sufficient thrust to the supersonic jet aircraft when the first engine is de-activated such that the aircraft can perform a supersonic climb operation and/or a supersonic cruise operation.

Abstract: A propulsion system includes a main gas turbine engine adapted for generating propulsive thrust during subsonic and supersonic flight operations and a supplementary propulsion unit adapted for generating additional thrust. The supplementary propulsion unit has an air intake and an exhaust for gas accelerated by the supplementary propulsion unit to provide the additional thrust and is adapted to generate the additional thrust during a limited range of subsonic flight operations, and to be dormant during other flight operations. The propulsion system has housing for the supplementary propulsion unit, including intake and exhaust covers which are moveable between deployed and stowed configurations. During the limited range of subsonic flight operations the intake and exhaust cover are moved to the deployed configuration to open the intake and the exhaust. During other flight operations the intake and exhaust cover are moved to the stowed configuration to close the intake and the exhaust.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes fan blades defining a fan diameter. The heat exchanger module is in communication with the fan assembly by an inlet duct, and the heat exchanger module further includes radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between hollow vanes. Each hollow vane accommodates at least one heat transfer element to transfer heat from a first fluid contained within the or each heat transfer element to a corresponding vane airflow passing through the hollow vane and over a surface of the or each heat transfer element. Each hollow vane further includes a flow modulator configured to regulate airflow in proportion to total airflow entering the heat exchanger module in response to a user requirement.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module including a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the fan assembly. Each heat transfer element may be individually and independently fluidly isolated from the remaining heat transfer elements.

Abstract: A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, and an exhaust module. The fan assembly includes a plurality of fan blades defining a fan diameter (D). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and the heat exchanger module includes a plurality of radially-extending hollow vanes arranged in a circumferential array with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger module is divided between a set of vane airflows through each of the hollow vanes and a set of channel airflows through each of the channels.