Abstract: A heat exchanger for heat exchange between a first fluid and a second fluid has a plurality of tube sections, each comprising; an interior for passing the first fluid; an exterior for exposure to the second fluid; a first leg; a second leg; a turn joining the first leg to the second leg; and a first face and a second face. A support has at least one carbon member engaging the plurality of tube sections.

Abstract: A heat exchanger for heat exchange between a first fluid and a second fluid has a plurality of tube sections, each having; an interior for passing the first fluid; an exterior for exposure to the second fluid; a first leg; a second leg; and a turn joining the first leg to the second leg. A has: fiber members passing between legs of the tube sections; and an end plate.

Abstract: An aircraft propulsion systems and aircraft having the same are described. The aircraft propulsion systems have one or more aircraft systems including at least one hydrogen tank and a first heat exchanger and one or more engine systems including at least a main engine core, a second heat exchanger, and a third heat exchanger. The main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is configured to be supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the first heat exchanger of the aircraft systems, the second heat exchanger of the engine systems, and the third heat exchanger of the engine systems, and then supplied into the burner for combustion.

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 heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

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 high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor inlet annulus area, a fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and a turbine having a high pressure turbine section and a low pressure turbine section driving the fan through a speed reduction mechanism, wherein the low pressure 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: Gas turbine engines are described. The gas turbine engines include a compressor section, a combustor section, a turbine section, a nozzle section, wherein the compressor section, the combustor section, the turbine section, and the nozzle section define a core flow path that expels through the nozzle section, and a waste heat recovery system. The waste heat recovery system includes a heat recovery heat exchanger arranged at the nozzle section, wherein the heat recovery heat exchanger is arranged within the nozzle section such that the heat recovery heat exchanger occupies less than an entire area of an exhaust area of the nozzle section and a heat rejection heat exchanger arranged to reduce a temperature of a working fluid of the waste heat recovery system.

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 low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas pat

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: Gas turbine engines are described. The gas turbine engines includes a compressor section, a combustor section, a turbine section, and a nozzle section. The compressor section, the combustor section, the turbine section, and the nozzle section define a core flow path that expels through the nozzle section. A cooling duct is provided that is separate from the core flow path. A waste heat recovery system is arranged with a heat rejection heat exchanger arranged within the cooling duct and a blower is arranged within the cooling duct and configured to generate a pressure drop across the heat rejection heat exchanger.

Abstract: A heat exchanger has arcuate inlet and outlet manifolds and a plurality of tube banks, each tube bank coupling one of the inlet manifold outlets to an associated one of the outlet manifold inlets. Each tube bank partially nests with one or more others of the tube banks and has: a first header coupled to the associated inlet manifold outlet and the associated the outlet manifold inlet; a second header; and a plurality of tube bundles each having a first end coupled to the associated first header and a second end coupled to the associated second header. A flowpath from the each inlet manifold outlet passes sequentially through flowpath legs formed by each of the tube bundles in the associated tube bank to exit the tube bank to the associated outlet manifold inlet.

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: A gas turbine engine and engine mount structure includes a core engine including a compressor section, a combustor section and a turbine section mounted within a core engine housing. The fan, the compressor section and the turbine section rotate about an axis of rotation. An outer nacelle surrounds the fan, and is spaced from the core engine housing to define a bypass duct. The fan delivers air into the bypass duct and into the core engine housing. The nacelle is formed with camber so as to be curved in a first plane away from the axis of rotation in a first lateral direction. An engine mount structure extends from the nacelle at an angle that is non-parallel and non-perpendicular to the first plane, and has a component in a lateral direction that is opposed to the first lateral direction. An aircraft is also disclosed.

Abstract: A gas turbine engine and engine mount structure includes a core engine including a compressor section, a combustor section and a turbine section mounted within a core engine housing. The fan, the compressor section and the turbine section rotate about an axis of rotation. An outer nacelle surrounds the fan, and is spaced from the core engine housing to define a bypass duct. The fan delivers air into the bypass duct and into the core engine housing. The nacelle is formed with camber so as to be curved in a first plane away from the axis of rotation in a first lateral direction. An engine mount structure extends from the nacelle at an angle that is non-parallel and non-perpendicular to the first plane, and has a component in a lateral direction that is opposed to the first lateral direction. An aircraft is also disclosed.

Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

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 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 turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and a fan duct annulus area outboard of the low pressure compressor section inlet, and a fan drive turbine section. The fan drive turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35, and is less than 0.55.