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 = P E ? M ? 1 ) ( Maximum ? Dry ? Thrust = T ) is in a range of between 2.0 and 10.0.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Abstract: An embodiment of a method for supplying refrigerants to a liquefied natural gas (LNG) facility includes: advancing a first refrigerant from a first storage device to a heat exchanger, the first refrigerant having a first temperature; advancing a second refrigerant from a second storage device to the heat exchanger, the second refrigerant having a second temperature different than the first temperature; flowing the first refrigerant and the second refrigerant through the heat exchanger; adjusting the second temperature based on at least a transfer of heat between the first refrigerant and the second refrigerant in the heat exchanger; and transferring the first refrigerant and the second refrigerant to the LNG facility.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, first thermal bus and heat exchanger. The first thermal bus includes a first heat transfer fluid in a closed loop flow sequence, between the first gas turbine engine, or each first electric machine, and the first heat exchanger. Waste heat energy transfers to the first heat transfer fluid. The first heat exchanger configures to transfer waste heat energy from the first heat transfer fluid to a dissipation medium. During steady-state operation of the first gas turbine engine, the first heat transfer fluid entering the first heat exchanger has a temperature of TFLUID(° C.), and a temperature of an inlet air flow entering the first gas turbine engine has a temperature TAIR(° C.) and a ratio B is in a range of between 5.0-18.0.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger. Waste heat energy generated by at least one first gas turbine engine, and first electric machine, transfers to the first heat transfer fluid. The first heat exchanger directs a first proportion of the first heat transfer fluid through a first heat dissipation portion wherein a first proportion of the waste heat energy transfers to a first dissipation medium dependent on the first dissipation medium temperature and mass flow rate. The first heat exchanger directs a second proportion of the first heat transfer fluid through a second heat dissipation portion wherein the second proportion of waste heat energy transfers to a second dissipation medium dependent on the second dissipation medium temperature and mass flow rate.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat is exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, a first heat exchanger, and one or more first ancillary systems. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the or each first electric machine, the first gas turbine engine, the first heat exchanger, and the or each first ancillary system. Waste heat energy generated by at least one of the first gas turbine engine, the or each first electric machine, and the or each first ancillary system, is transferred to the first heat transfer fluid. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, a first thermal bus, and a first heat exchanger. The first thermal bus includes a first heat transfer fluid in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to first heat transfer fluid. When airspeed of aircraft is less than Mn0.6, the first heat exchanger transfers the waste heat energy from the first heat transfer fluid to a first dissipation medium. When the airspeed of the aircraft is greater than Mn0.6, the first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a second dissipation medium.

Abstract: A thermal management system for an aircraft includes a first thermal bus including one or more first heat sources, a heat sink, a vapour compression system, and one or more second heat sources. The vapour compression system includes a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor. A first heat flow (Q1) of waste heat energy generated by the first heat sources is transferred via the first heat transfer fluid to the heat sink. A second heat flow (Q2) of waste heat energy generated by the second heat source(s) being transferred via the evaporator to a refrigerant. A third heat flow (Q3) of heat energy in the refrigerant is transferred via the condenser to the first heat transfer fluid. The controller is configured to ensure that: 1.1*Q2<Q3<3.0*Q2.

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. 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: An embodiment of a method for supplying refrigerants to a liquefied natural gas (LNG) facility includes: advancing a first refrigerant from a first storage device to a heat exchanger, the first refrigerant having a first temperature; advancing a second refrigerant from a second storage device to the heat exchanger, the second refrigerant having a second temperature different than the first temperature; flowing the first refrigerant and the second refrigerant through the heat exchanger; adjusting the second temperature based on at least a transfer of heat between the first refrigerant and the second refrigerant in the heat exchanger; and transferring the first refrigerant and the second refrigerant to the LNG facility.

Abstract: The invention relates to various nonlimiting embodiments that include methods, apparatuses or systems for processing natural gas comprising a heavies removal column processing natural gas and light oil reflux. The overhead stream goes to heavies treated natural gas storage. The heavies removal column reboiler bottoms stream product is input to a debutanizer column. The debutanizer column overhead lights are input to a flash drum where the bottoms is pumped through a heat exchanger as a light oil reflux input to the heavies removal column, while the debutanizer reboiler bottoms product is stored as stabilized condensate. Alternatively, debutanizer column overhead lights are sent to heavies treated gas storage and the bottoms stream product goes to a depentanizer column, the overhead lights are pumped through a heat exchanger as a light oil reflux input to the heavies removal column, while the depentanizer reboiler bottoms product is stabilized condensate.

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: The invention relates to various nonlimiting embodiments that include methods, apparatuses or systems for processing natural gas comprising a heavies removal column processing natural gas and light oil reflux. The overhead stream goes to heavies treated natural gas storage. The heavies removal column reboiler bottoms stream product is input to a debutanizer column. The debutanizer column overhead lights are input to a flash drum where the bottoms is pumped through a heat exchanger as a light oil reflux input to the heavies removal column, while the debutanizer reboiler bottoms product is stored as stabilized condensate. Alternatively, debutanizer column overhead lights are sent to heavies treated gas storage and the bottoms stream product goes to a depentanizer column, the overhead lights are pumped through a heat exchanger as a light oil reflux input to the heavies removal column, while the depentanizer reboiler bottoms product is stabilized condensate.

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: A twin-engine aircraft power and propulsion system including first and second propulsive gas turbine engines, each having combustion equipment and a first and second spool; first, second, third, and fourth electrical power generation sub-systems including electric machines respectively mechanically coupled with the first spool of the first propulsive gas turbine engine, the second spool of the first propulsive gas turbine engine, the first spool of the second propulsive gas turbine engine, and the second spool of the second propulsive gas turbine engine; and first, second, third, and fourth power channels respectively connected with distribution sides of the first electric machine, second electric machine, third electric machine, and fourth electric machine.

Abstract: Multi-engine aircraft power and propulsion systems and methods of starting the engines of multi-engine aircraft disclosed, including supplying electrical power from an electrical power source to electric machines of the a first gas turbine engine and operating electric machines as motors to drive rotation of spools of the first gas turbine engine; starting the first gas turbine engine by lighting combustion equipment of the first gas turbine engine; operating the electric machines of the first gas turbine engine as generators to extract mechanical power and generate electrical power from spools of the first gas turbine engine; transferring the electrical power to electric machines of a second gas turbine engine and operating the electric machines as motors to drive rotation of spools of the second gas turbine engine; and starting the second gas turbine engine by lighting combustion equipment of the second gas turbine engine.

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.