Abstract: A computer system includes a processor controls the computer system to perform a computer-implemented method of determining a minimum humidity required for formation and persistence of contrails produced by an engine of an aircraft. The computer-implemented method includes determining engine performance model parameters of the engine at desired operating conditions with zero humidity; determining additional energy flow out of the engine; and determining an exhaust plume temperature scaling factor. The method further comprises determining a contrail engine efficiency parameter (noverall_?) based on additional energy flow out of the engine and the exhaust plume temperature scaling factor; and generating an improved Schmidt-Appleman equation based on the contrail engine efficiency parameter (noverall_?).

Abstract: A powerplant is provided for an aircraft. This aircraft powerplant includes a compressor section, a fuel-air heat exchanger, a recuperator, a combustor section comprising a combustor, a turbine section, a flowpath and a fuel delivery system. The flowpath extends from an airflow inlet into the flowpath, through the compressor section, the fuel-air heat exchanger, the recuperator, the combustor section and the turbine section, to a combustion products exhaust from the flowpath. The fuel delivery system includes a fuel source, a fuel circuit and a fuel injector. The fuel circuit fluidly couples the fuel source to the fuel injector and extends through the fuel-air heat exchanger. The fuel injector is configured to introduce fuel into the flowpath for combustion within the combustor.

Abstract: An aircraft powerplant includes a core of a gas turbine engine and an exhaust section. The core includes a core flowpath, a compressor section, a combustor section and a turbine section. The core flowpath extends longitudinally through the compressor section, the combustor section and the turbine section. The exhaust section includes an exhaust flowpath and a flow diverter. The flow diverter includes a diverter panel configured to pivot between a first position and a second position. The flow diverter is configured to fluidly couple the diffuser section to the first nozzle section and fluidly decouple the diffuser section from the second nozzle section when the diverter panel is in the first position. The flow diverter is configured to fluidly couple the diffuser section to the second nozzle section and fluidly decouple the diffuser section from the first nozzle section when the diverter panel is in the second position.

Abstract: An impeller for a gas turbine engine includes an impeller hub having an impeller leading edge and an impeller trailing edge and a plurality of bifurcated impeller blades arranged on the impeller hub and extending from the impeller leading edge to the impeller trailing edge, each bifurcated impeller blade having a bifurcated pressure side and a bifurcated suction side. The bifurcated pressure side and the bifurcated suction side of each bifurcated impeller blade separate at a bifurcation point and extend to the impeller trailing edge with a reduced flow channel defined between adjacent bifurcated impeller blades. A blocked zone is defined within each bifurcated impeller blade between the bifurcated pressure side and the bifurcated suction side and each of the bifurcated pressure side and the bifurcated suction side extend tangentially from the bifurcation point at different radii of curvature.

Abstract: An assembly for an aircraft includes a fluid source, a first heat exchanger, a partial-admission turbine, a mechanical load, and a control assembly. The partial-admission turbine includes a rotational assembly and a plurality of partial-admission turbine stages. The rotational assembly includes a bladed turbine rotor. The bladed turbine rotor includes a plurality of rotor blade stages. Each of the plurality of partial-admission turbine stages includes a respective rotor blade stage of the plurality of rotor blade stages. The fluid source, the first heat exchanger, and the partial-admission turbine sequentially form a portion of a fluid flow path through the assembly. The mechanical load is coupled to the rotational assembly.

Abstract: An aircraft propulsion system includes a gas turbine engine and an exhaust nozzle assembly. The gas turbine engine includes a compressor section, a combustor section, a turbine section, and an exhaust section. The compressor section, the combustor section, the turbine section, and the exhaust section form a core flow path through the gas turbine engine. The exhaust nozzle assembly is disposed at the exhaust section. The exhaust nozzle assembly includes a nozzle bypass system and an exhaust treatment system. The nozzle bypass system is selectively configurable in a bypass mode or an exhaust treatment mode. The nozzle bypass system in the bypass mode directs a core combustion gas from the core flow path through the exhaust nozzle assembly bypassing the exhaust treatment system. The nozzle bypass system in the exhaust treatment mode directs the core combustion gas through the exhaust nozzle assembly and exhaust treatment system.

Abstract: An aircraft propulsion system includes a gas turbine engine and an exhaust nozzle assembly. The gas turbine engine includes a compressor section, a combustor section, a turbine section, and an exhaust section. The compressor section, the combustor section, the turbine section, and the exhaust section form a core flow path through the gas turbine engine. The exhaust nozzle assembly is disposed at the exhaust section. The exhaust nozzle assembly includes a nozzle bypass system and an exhaust treatment system. The nozzle bypass system is selectively configurable in a bypass mode or an exhaust treatment mode. The nozzle bypass system in the bypass mode directs a core combustion gas from the core flow path through the exhaust nozzle assembly bypassing the exhaust treatment system. The nozzle bypass system in the exhaust treatment mode directs the core combustion gas through the exhaust nozzle assembly and exhaust treatment system.

Abstract: An engine system is provided that includes an open propulsor rotor, an engine core assembly, a bypass flowpath and a condenser. The engine core assembly is configured to drive rotation of the open propulsor rotor. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The core flowpath extends through the compressor section, the combustor section and the turbine section from an inlet into the core flowpath to an exhaust from the core flowpath. The bypass flowpath bypasses the engine core assembly. The condenser is configured to exchange heat energy between combustion products flowing within the core flowpath downstream of the combustor section and bypass air flowing within the bypass flowpath.

Abstract: An aircraft engine, comprising: a thermal engine having an outlet; a heat exchanger having a first conduit connected to a source of a fluid, and a second conduit connected to the outlet; an inlet conduit connecting the source to the first conduit; an outlet conduit connected to an outlet of the first conduit; a component connected to the outlet conduit; a bypass conduit and a recirculation conduit both connecting the inlet conduit to the outlet conduit while bypassing the heat exchanger; a valve communicating with the recirculation conduit and the bypass conduit, the valve operable to selectively allow the fluid to flow in one of the recirculation conduit and the bypass conduit while limiting the fluid from flowing in the other of the recirculation conduit and the bypass conduit; and a flow inducer communicating with the recirculation conduit for inducing a flow from the outlet conduit to the inlet conduit.

Abstract: A system is provided for an aircraft. This aircraft system includes an aircraft powerplant, a powerplant sensor system, an environment sensor system and a monitoring system. The aircraft powerplant includes a heat engine. The powerplant sensor system is configured to provide engine data indicative of one or more operating parameters of the heat engine. The environment sensor system is configured to provide environment data indicative of one or more environmental parameters of an environment in which the heat engine is operating. The monitoring system is configured to determine formation of a contrail and quantify an impact of the contrail when formed based on the engine data and the environment data.

Abstract: An aircraft propulsion system includes a gas turbine engine, a hydrogen-exhaust heat exchanger, and a hydrogen turboexpander. The hydrogen-exhaust heat exchanger is disposed relative to a turbine exhaust section so that a flow of exhaust gas exiting the engine is in communication with the hydrogen-exhaust heat exchanger. The hydrogen-exhaust heat exchanger changes a flow of liquid hydrogen to gaseous hydrogen. The hydrogen turboexpander has parallel first and second rotor shafts within a housing that includes a first and second rotor shaft cavities. The first rotor shaft and the first rotor shaft cavity define a first rotor shaft fluid flow path (RSFFP). The second rotor shaft and the second rotor shaft cavity define a second RSFFP. The first RSFFP is in series with the second RSFFP between an inlet port and an exit port. The exit port is in fluid communication with the combustor.

Abstract: An aircraft propulsion system includes a gas turbine engine, a water-exhaust heat exchanger, and a fuel cell segment. The gas turbine engine has a compressor, a combustor, and a turbine. The engine is in drive communication with a rotational load component. The water-exhaust heat exchanger is disposed to receive exhaust gas exiting the turbine section. The fuel cell segment includes a liquid hydrogen evaporator, a fuel cell, and a water pump. The liquid hydrogen evaporator is configured to change a flow of liquid hydrogen to a flow of gaseous hydrogen. The fuel cell is configured to receive a flow of air and the flow of gaseous hydrogen and react them to produce electrical power and a flow of water. The flow of water from the fuel cell goes to the water-exhaust heat exchanger which converts the flow of water to a flow of steam.

Abstract: An engine system includes an engine core assembly, a fuel system and a cooling circuit. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The combustor section includes a combustor adjacent a plenum. The fuel system includes a fuel flowpath, a fuel reservoir, a fuel-cooling air heat exchanger and a fuel injector. The fuel flowpath is configured to direct hydrogen fuel, received from the fuel reservoir, through the fuel-cooling air heat exchanger to the fuel injector. The fuel injector is configured to introduce the hydrogen fuel into the combustor. The cooling circuit is configured to direct cooling air, bled from the plenum, through the fuel-cooling air heat exchanger to the turbine section. The fuel-cooling air heat exchanger is configured to exchange heat energy between the hydrogen fuel flowing within the fuel flowpath and the cooling air flowing within the cooling circuit.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an intercooler (IC). The turbine engine includes a source of non-hydrocarbon fuel, and a low pressure compressor having an LPC air inlet and outlet, and a high pressure compressor having an HPC air inlet and outlet. The intercooler has an IC air inlet, an IC air outlet, a U-shaped air passage, and crossflow tubes. The IC air inlet is in fluid communication with the LPC air outlet. The IC air outlet is in fluid communication with the HPC air inlet. The U-shaped air passage is in fluid communication with the IC air inlet and outlet. The crossflow tubes extend through the U-shaped air passage. The crossflow tubes are in fluid communication with the source of non-hydrocarbon fuel, and are configured to contain a flow of the non-hydrocarbon fuel therethrough.

Abstract: A computer system includes a processor controls the computer system to perform a computer-implemented method of determining a minimum humidity required for formation and persistence of contrails produced by an engine of an aircraft. The computer-implemented method includes determining engine performance model parameters of the engine at desired operating conditions with zero humidity; determining additional energy flow out of the engine; and determining an exhaust plume temperature scaling factor. The method further comprises determining a contrail engine efficiency parameter (noverall_?) based on additional energy flow out of the engine and the exhaust plume temperature scaling factor; and generating an improved Schmidt-Appleman equation based on the contrail engine efficiency parameter (noverall_?).

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an intercooler (IC). The turbine engine includes a source of non-hydrocarbon fuel, and a low pressure compressor having an LPC air inlet and outlet, and a high pressure compressor having an HPC air inlet and outlet. The intercooler has an IC air inlet, an IC air outlet, a U-shaped air passage, and crossflow tubes. The IC air inlet is in fluid communication with the LPC air outlet. The IC air outlet is in fluid communication with the HPC air inlet. The U-shaped air passage is in fluid communication with the IC air inlet and outlet. The crossflow tubes extend through the U-shaped air passage. The crossflow tubes are in fluid communication with the source of non-hydrocarbon fuel, and are configured to contain a flow of the non-hydrocarbon fuel therethrough.

Abstract: An engine system includes an engine core assembly, a fuel system and a cooling circuit. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The combustor section includes a combustor adjacent a plenum. The fuel system includes a fuel flowpath, a fuel reservoir, a fuel-cooling air heat exchanger and a fuel injector. The fuel flowpath is configured to direct hydrogen fuel, received from the fuel reservoir, through the fuel-cooling air heat exchanger to the fuel injector. The fuel injector is configured to introduce the hydrogen fuel into the combustor. The cooling circuit is configured to direct cooling air, bled from the plenum, through the fuel-cooling air heat exchanger to the turbine section. The fuel-cooling air heat exchanger is configured to exchange heat energy between the hydrogen fuel flowing within the fuel flowpath and the cooling air flowing within the cooling circuit.

Abstract: An engine system is provided that includes an open propulsor rotor, an engine core assembly, a bypass flowpath and a condenser. The engine core assembly is configured to drive rotation of the open propulsor rotor. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The core flowpath extends through the compressor section, the combustor section and the turbine section from an inlet into the core flowpath to an exhaust from the core flowpath. The bypass flowpath bypasses the engine core assembly. The condenser is configured to exchange heat energy between combustion products flowing within the core flowpath downstream of the combustor section and bypass air flowing within the bypass flowpath.

Abstract: An aircraft propulsion system and method is provided. The system includes compressor, combustor, and turbine sections, a fuel source, a heat exchanger, a compressor bleed air passage, and a cooling air passage. The fuel source is configured to contain a non-hydrocarbon fuel. The heat exchanger has separate air and fuel passages. The air passage permits passage of the compressed air therethrough. The fuel passage is configured to permit a passage of the non-hydrocarbon fuel therethrough. The compressor bleed air passage is configured to receive compressed air bleed off a core flow path upstream of the combustor section. The compressor bleed air passage is in fluid communication with the air passage air inlet. The cooling air passage is configured to receive compressed air exiting the air passage air outlet and configured to direct the compressed bleed air to the vane stage or rotor stage of the turbine section, or both.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, a plurality of air scoops, and a plurality of exhaust gas conduit banks. The housing has an interior cavity, a bottom side, and first and second lateral sides. The exhaust gas conduit banks are disposed in the interior cavity of the housing and are configured to form an interior bypass air chamber. Each exhaust gas conduit bank includes alternating exhaust gas conduits and bypass air passages. The exhaust gas conduits extend axially between the forward and aft ends, and are configured to conduct exhaust gases to the nozzle. The bypass air passages are configured to receive bypass air from the air scoops and direct the bypass air between the exhaust gas conduits and into the interior bypass air chamber.