Abstract: A gas turbine engine for an aircraft comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module. The gas turbine engine further comprises a first electric machine that is rotationally connected to the turbine module, and an electrical energy storage unit. The gas turbine engine is configured to generate a maximum dry thrust T (N). The first electric machine is configured to generate a maximum electrical power PEM1 (W). The electrical energy storage unit has an energy storage capacity E (Wh), a maximum charge rate C (h?1), and a maximum discharge rate D (h?1). The electrical energy storage unit is configured to store electrical energy that may be generated by the first electric machine.

Abstract: A gas turbine engine for an aircraft 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 a maximum electrical power PEM1 (W), and the gas turbine engine is configured to generate a maximum shaft power PSHAFT (W); and a ratio R of: R = ( Maximum ? Electrical ? Power ? Generated = P E ? M ? 1 ) ( Maximum ? Shaft ? Power = P S ? H ? A ? F ? T ) is in a range of between 0.005 and 0.020.

Abstract: Aircraft power and propulsion systems, aircraft comprising such power and propulsion systems, and methods of restarting a gas turbine engine of such power and propulsion systems during flight are provided. One such aircraft power and propulsion system comprises: a propulsive gas turbine engine comprising a plurality of spools, combustion equipment, one or more electric machines mechanically coupled with one or more of the spools and an electrically-powered fuel pump for delivering fuel to the combustion equipment; an electrical system connected with the one or more electric machines and the electrically-powered fuel pump, the electrical system comprising an energy storage system; and a control system configured to: responsive to a determination to the effect that a flame in the combustion equipment has been extinguished, control the electrical system to supply electrical power from the energy storage system to the fuel pump during an engine restart attempt.

Abstract: Aircraft power and propulsion systems and methods of operating the systems, one method includes: operating electric machines of gas turbine engines as generators to extract mechanical power from spools and generating electrical power therefrom; meeting an electrical power demand of a plurality of electrical loads connected with an electrical system by supplying the plurality of electrical loads with electrical power generated by the electric machines; determining when there is an electrical system and/or the gas turbine engine fault and an amount of electrical power generated by the power and propulsion system is reduced to a lower level; and responsive to the determination: during a time period ?T, controlling the plurality of electrical loads reducing the electrical power demand; and during the time period ?T, meeting at least part of the electrical power demand of the plurality of electrical loads by discharging the electrical energy storage system.

Abstract: Multi-engine aircraft power and propulsion systems and methods of restarting an engine of a multi-engine aircraft during fight are provided. One such method comprises: determining a condition to the effect that a flame in the combustion equipment of the second gas turbine engine has been extinguished; responsive to the determination, supplying electrical power from the electrical energy storage system to one or more of the electric machines of the second gas turbine engine and operating said one or more electric machines as motors to limit a reduction in a speed of the one or more spools of the second gas turbine engine following extinguishment of the flame in its combustion equipment; and restarting the second gas turbine engine by relighting the combustion equipment of the second gas turbine engine.

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 gas turbine engine includes a heat exchanger module, and a core engine. The core engine includes an intermediate-pressure compressor, high-pressure compressor, and high and low-pressure turbines. The high-pressure compressor rotationally connects to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor rotationally connects to the low-pressure turbine by a second shaft. The heat exchanger module fluidly communicates with the core engine by an inlet duct. The heat exchanger module includes a central hub and multiple heat transfer elements extending radially from the hub and spaced in a circumferential array, for heat energy transfer from a first fluid within the elements to an inlet airflow passing over a surface of the elements prior to airflow entry into an inlet to the core engine. The gas turbine engine further includes a first electric machine rotationally connected to the first shaft, and positioned downstream of the heat exchanger module.

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 gas turbine engine for an aircraft 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 a maximum electrical power PEM1 (W), and the gas turbine engine is configured to generate a maximum dry thrust T (N); and a ratio S of: S=(Maximum Electrical Power Generated=PEM1)/(Maximum Dry Thrust=T) is in a range of between 2.0 and 10.0.

Abstract: Implementations described and claimed herein provide systems and methods for processing liquefied natural gas (LNG). In one implementation, a solvent is injected into a feed of natural gas at a solvent injection point. A mixed feed is produced from a dispersal of the solvent into the feed of natural gas. The mixed feed contains heavy components. A chilled feed is produced by chilling the mixed feed. The chilled feed includes a vapor and a condensed liquid. The condensed liquid contains a fouling portion of the heavy components condensed by the solvent during chilling. The liquid containing the fouling portion of the heavy components is separated from the vapor. The vapor is directed into a feed chiller heat exchanger following separation of the liquid containing the fouling portion of the heavy components from the vapor, such that the vapor being directed into feed chiller heat exchanger is free of freezing components.

Abstract: A heat exchanger system includes a core-in-shell heat exchanger and a liquid/gas separator. The liquid/gas separator is configured to receive a liquid/gas mixture and to separate the gas from the liquid. The liquid/gas separator is connected to the core-in-shell heat exchanger via a first line for transmitting gas from the liquid/gas separator to a first region in the core-in-shell heat exchanger and connected to the core-in-shell heat exchanger via a second line for transmitting liquid from the liquid/gas separator to a second region of the core-in-shell heat exchanger.

Abstract: The invention relates to a system, method and apparatus for processing natural gas in an LNG facility. A natural gas feed is introduced into a heavies removal unit. The heavies removal system includes a heavies removal column and a distillation column. The heavies removal column and the distillation column are connected via a purge/recovery line. One or more components of the natural gas feed is purged from the heavies removal column to the distillation column via the purge/recovery line to obtain a specified concentration or concentration range of heavy components feeding into the distillation column.

Abstract: An aircraft gas turbine engine comprises at least a first main engine shaft an electric machine coupled to the main engine shaft, and a lubrication pump configured to provide oil to at least a part of the electric machine. The lubrication pump is controlled by a controller. The controller is configured to operate the lubrication pump in a main gas turbine engine running mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a first flow rate relative to the rotational rate of the main engine shaft, and a fault mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a second, higher flow rate relative to the rotational rate of the main engine shaft.

Abstract: The invention relates to a system, method and apparatus for removing heavies from natural gas. Natural gas and an external rich reflux gas feed are processed in a single column refluxed absorber. A bottoms stream is routed to a first heat exchanger and then to a stabilizer column where an overhead stream from the stabilizer column is routed through a condenser for partial separation into an overhead stream. A rich solvent may be introduced to the stabilizer column. The overhead stream is routed through a condenser for partial separation into a stabilizer reflux and a second overhead stream lights. The second overhead stream lights is routed to a heat exchanger and then routed to a partial condenser where the stream is separated into a heavies rich reflux stream, a distillate stream and heavies treated natural gas stream.

Abstract: An aircraft gas turbine engine (10) comprises a main engine shaft (22, 23), a main engine shaft bearing arrangement (36, 44, 49, 50) configured to rotatably support the main engine shaft (22, 23) and an electric machine (30) comprising a rotor (34) and a stator (32). The rotor (34) is mounted to the main engine shaft (22, 23) and is rotatably supported by the main engine shaft bearing arrangement (36, 44, 49, 50), and the stator (32) is mounted to static structure (46) of the gas turbine engine (10).

Abstract: The invention relates to a system, method and apparatus for removing heavies from natural gas. Natural gas and an external rich reflux gas feed are processed in a single column refluxed absorber. A bottoms stream is routed to a first heat exchanger and then to a stabilizer column where an overhead stream from the stabilizer column is routed through a condenser for partial separation into an overhead stream. A rich solvent may be introduced to the stabilizer column. The overhead stream is routed through a condenser for partial separation into a stabilizer reflux and a second overhead stream lights. The second overhead stream lights is routed to a heat exchanger and then routed to a partial condenser where the stream is separated into a heavies rich reflux stream, a distillate stream and heavies treated natural gas stream.

Abstract: An aircraft gas turbine engine includes a main engine shaft, a main engine shaft bearing arrangement for rotatably supporting the main engine shaft, and an electric machine including a rotor and a stator. The rotor is mounted to the main engine shaft and is rotatably supported by a further electric machine bearing arrangement extending between the rotor and the stator, and the stator is mounted to static structure of the gas turbine engine.

Abstract: A core-in-shell heat exchanger, a method of fabricating the core-in-shell heat exchanger, and a method of exchanging heat in a core-in-shell heat exchanger disposed on a slosh-inducing moving platform are described. The method of exchanging heat includes introducing a shell-side fluid into a shell of the core-in-shell heat exchanger and introducing a fluid to be cooled into each of one or more cores of the core-in-shell heat exchanger, the one or more cores being arranged along an axial length of the shell with a plurality of baffles disposed on either side of the one or more cores along the axial length of the shell to reduce slosh of the shell-side fluid. The method also includes draining excess shell-side fluid using a plurality of drains, at least two of the plurality of drains being disposed on opposite sides of one of the plurality of baffles.

Abstract: A gas turbine engine includes a main gas flow exhaust nozzle having an annular inner surface which, in use, bounds a flow of exhaust gas. The gas turbine engine further includes cooling passages having respective outlets therefrom to provide a flow of cooling air over a surface of the engine or an adjacent airframe component, thereby protecting the cooled surface from the exhaust gas flow. Adjacent cooling passages of the or each pair of the nested cooling passages are separated from each other by a respective dividing wall. The outlets from the nested cooling passages are staggered in the axial direction of the exhaust nozzle such that cooling air flowing out of an inner one of the adjacent cooling passages of the or each pair of the nested cooling passages flows over the dividing wall separating the adjacent passages.