Abstract: A gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to about 8 at cruise power. A gear arrangement is configured to drive the fan section. A compressor section includes both a first compressor section and a second compressor section. A turbine section is configured to drive the gear arrangement, and may have a low pressure turbine with four stages and a low pressure turbine pressure ratio greater than about 5:1, and a high pressure turbine with two stages. An overall pressure ratio is provided by the combination of a pressure ratio across the first compressor section and a pressure ratio across the second compressor section, and greater than about 40, measured at sea level and at a static, full-rated takeoff power.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A gas turbine engine according to the present disclosure includes a first compressor and a first turbine for driving the first compressor. A core section includes a second compressor and a second turbine for driving the second compressor. A third turbine is arranged fluidly downstream of the first turbine and the second turbine and configured to drive a power take-off. A first duct system is arranged fluidly between the low-pressure compressor and the core section. The first duct system is arranged to reverse fluid flow before entry into the core section.

Abstract: A gas turbine engine includes a primary flowpath fluidly connecting a compressor section, a combustor section, and a turbine section. A heat exchanger is disposed in the primary flowpath downstream of the turbine section. The heat exchanger includes a first inlet for receiving fluid from the primary flowpath and a first outlet for expelling fluid received at the first inlet. The heat exchanger further includes a second inlet fluidly connected to a supercharged CO2 (sCO2) bottoming cycle and a second outlet connected to the sCO2 bottoming cycle. The sCO2 bottoming cycle is an overexpanded, recuperated Brayton cycle.

Abstract: A gas turbine engine includes a primary flowpath fluidly connecting a compressor section, a combustor section, and a turbine section. A heat exchanger includes an first inlet connected to a high pressure compressor bleed, a first outlet connected to a high pressure turbine inlet. The heat exchanger further includes a second inlet fluidly connected to a supercharged CO2 (sCO2) work recovery cycle and a second outlet connected to the sCO2 work recovery cycle. The sCO2 work recovery cycle is an overexpanded, recuperated work recovery cycle.

Abstract: A gas turbine engine includes a primary flowpath fluidly connecting a compressor section, a combustor section, and a turbine section. A heat exchanger is disposed in the primary flowpath downstream of the turbine section. The heat exchanger includes a first inlet for receiving fluid from the primary flowpath and a first outlet for expelling fluid received at the first inlet. The heat exchanger further includes a second inlet fluidly connected to a supercritical CO2 (sCO2) bottoming cycle and a second outlet connected to the sCO2 coolant circuit. The sCO2 bottoming cycle is a recuperated Brayton cycle.

Abstract: A gas turbine engine includes a primary flowpath fluidly connecting a compressor section, a combustor section, and a turbine section. A heat exchanger includes an first inlet connected to a high pressure compressor bleed, a first outlet connected to a high pressure turbine inlet. The heat exchanger further includes a second inlet fluidly connected to a supercharged CO2 (sCO2) coolant circuit and a second outlet connected to the sCO2 work recovery cycle.

Abstract: A recuperated gas turbine engine includes an engine core that has a compressor section, a combustor section, and a turbine section. An exhaust duct is located downstream of the turbine section for receiving a hot turbine exhaust stream from the turbine section. The exhaust duct includes a heat exchanger and a temperature-control module upstream of the heat exchanger. A first compressor bleed line portion leads into the heat exchanger, and a second compressor bleed lie portion leads into the exhaust duct upstream of the heat exchanger. A compressor return line leads from the heat exchanger into the engine core upstream of the combustor section. The compressor bleed line is operable to selectively feed compressed air to the heat exchanger, and the temperature-control module is operable to selectively modulate at least one of temperature and flow of the hot turbine exhaust stream with respect to the heat exchanger.

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage is fluidly coupled to a bleed port of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system includes a pump fluidly coupled to the first cooling stage to receive and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air. The intercooled cooling system includes an air cycle machine in fluid communication to outputs of the cooling stages to selectively receive the cool bleed air or the intercooled cooling air.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including an array of fan blades rotatable about an engine axis, a compressor including a first compressor section and a second compressor section, the second compressor section including a second compressor section inlet with a compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, and a turbine having a first turbine section driving the first compressor section, a second turbine section driving the fan through an epicyclic gearbox, the second turbine section including blades and vanes, and wherein the second turbine section defines a maximum gas path radius and the fan blades define a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is less than 0.6.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a first compressor section and a second compressor, the second compressor section including a second compressor section inlet with a second compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the second compressor section inlet, a shaft assembly having a first portion and a second portion, a turbine in fluid communication with the combustor, the turbine having a first turbine section coupled to the first portion of the shaft assembly to drive the first compressor section, and a second turbine section coupled to the second portion of the shaft assembly to drive the fan, an epicyclic transmission coupled to the fan and rotatable by the second turbine section through the second portion of the shaft assembly to allow the second turbine to turn fa

Abstract: A gas turbine engine includes, among other things, a fan, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10. A gear arrangement drives the fan. A compressor section includes both a low pressure compressor and a high pressure compressor. A turbine section drives the gear arrangement. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor and a pressure ratio across the high pressure compressor, and greater than 50, measured at sea level and at a static, full-rated takeoff power. The pressure ratio across the high pressure compressor is greater than 7.

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes a plurality of cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage of the plurality of cooling stages is fluidly coupled to a bleed port of a compressor of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system also includes a pump fluidly coupled to the first cooling stage to receive the cool bleed air and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage of the plurality of cooling stages is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air, which is provided to the gas turbine engine.

Abstract: An intercooled cooling system for a gas turbine engine is provided. The intercooled cooling system includes cooling stages in fluid communication with an air stream utilized for cooling. A first cooling stage is fluidly coupled to a bleed port of the gas turbine engine to receive and cool bleed air with the air stream to produce a cool bleed air. The intercooled cooling system includes a pump fluidly coupled to the first cooling stage to receive and increase a pressure of the cool bleed air to produce a pressurized cool bleed air. A second cooling stage is fluidly coupled to the pump to receive and cool the pressurized cool bleed air to produce an intercooled cooling air. The intercooled cooling system includes an air cycle machine in fluid communication to outputs of the cooling stages to selectively receive the cool bleed air or the intercooled cooling air.

Abstract: A graphene heat pipe for a gas turbine engine includes a body of graphene. The body has a hot side to accept heat from the gas turbine engine, a cold side to reject heat from the body, and an adiabatic portion to flow heat within the body between the hot side and the cold side.

Abstract: A compressor section of a gas turbine engine includes a bleed port having a flow splitter therein so as to define a downstream bleed channel having a downstream inlet and an upstream bleed channel having an upstream inlet that is positioned radially outward from the downstream inlet.

Abstract: A gas turbine engine includes a core having a compressor section with a first compressor and a second compressor, a turbine section with a first turbine and a second turbine, and a primary flowpath fluidly connecting the compressor section and the turbine section. The first compressor is connected to the first turbine via a first shaft, the second compressor is connected to the second turbine via a second shaft, and a motor is connected to the first shaft such that rotational energy generated by the motor is translated to the first shaft. The gas turbine engine includes a takeoff mode of operation, a top of climb mode of operation, and at least one additional mode of operation. The gas turbine engine is undersized relative to a thrust requirement in at least one of the takeoff mode of operation and the top of climb mode of operation, and a controller is configured to control the mode of operation of the gas turbine engine.

Abstract: A gas generator has at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine is positioned downstream of a path of the products of combustion having passed over the at least one gas generator turbine rotor. The fan drive turbine drives a shaft and the shaft engages gears to drive at least three fan rotors.