An aircraft system is provided that includes a first propulsor rotor, a first transmission, a second propulsor rotor, a second transmission and an intermittent combustion engine. The first propulsor rotor is rotatable about a first propulsor axis. The first transmission is coupled to the first propulsor rotor. The second propulsor rotor is rotatable about a second propulsor axis. The second transmission is coupled to the second propulsor rotor. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor through the first transmission. The intermittent combustion engine is configured to drive rotation of the second propulsor rotor through the second transmission.

An aircraft power plant, has: a thermal engine having a combustion chamber inside a housing defining a coolant passage, and an oil passage; and a heat management system including: an oil-to-coolant heat exchanger having an oil conduit in fluid communication with the oil passage and a first coolant conduit in heat exchange relationship with the oil conduit and in fluid communication with the coolant passage; and a coolant-to-air heat exchanger having a second coolant conduit in fluid communication with the coolant passage and an air conduit in heat exchange relationship with the second coolant conduit and in fluid communication with an environment outside the aircraft power plant, wherein the coolant passage inlet and the coolant passage outlet of the coolant passage of the housing are fluidly connected to one another via the first coolant conduit of the oil-to-coolant heat exchanger and the second coolant conduit of the coolant-to-air heat exchanger.

A rotating detonation combustion system includes a detonation channel including an inner wall and an outer wall and extend in a longitudinal direction from an inlet of the detonation channel to an outlet of the detonation channel. A first mixing element and a second mixing element are disposed in the rotating detonation combustion system. The first mixing element forming a ring on the inner wall and the second mixing element forming a ring on the outer wall of the detonation channel adjacent the inlet. Each of the first mixing element and the second mixing element comprise a plurality of protrusions disposed circumferentially along the inner wall and the outer wall and extend into the detonation channel such that the plurality of protrusions affects vectors of at least a portion of the fuel and at least a portion of the fluid passing through recesses between the plurality of protrusions.

A rotating detonation combustion system includes a detonation channel including an inner wall and an outer wall and extend in a longitudinal direction from an inlet of the detonation channel to an outlet of the detonation channel. A first mixing element and a second mixing element are disposed in the rotating detonation combustion system. The first mixing element forming a ring on the inner wall and the second mixing element forming a ring on the outer wall of the detonation channel adjacent the inlet. Each of the first mixing element and the second mixing element comprise a plurality of protrusions disposed circumferentially along the inner wall and the outer wall and extend into the detonation channel such that the plurality of protrusions affects vectors of at least a portion of the fuel and at least a portion of the fluid passing through recesses between the plurality of protrusions.

A propulsion system for an aircraft includes an engine assembly and a turbocompressor. The engine assembly includes an engine and an interburner. The engine includes an engine output shaft. The engine is configured to drive rotation of a propulsor with the engine output shaft. The interburner is configured to mix and burn an exhaust gas from the engine with fuel to form a combustion gas. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The turbine is connected in fluid communication with the interburner to receive the combustion gas. The compressor is connected in fluid communication with the engine to direct a compressed air to the engine. The rotational assembly is mechanically independent of the engine output shaft.

A propulsion system for an aircraft includes an engine assembly and a turbocompressor. The engine assembly includes an engine and an interburner. The engine includes an engine output shaft. The engine is configured to drive rotation of a propulsor with the engine output shaft. The interburner is configured to mix and burn an exhaust gas from the engine with fuel to form a combustion gas. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The turbine is connected in fluid communication with the interburner to receive the combustion gas. The compressor is connected in fluid communication with the engine to direct a compressed air to the engine. The rotational assembly is mechanically independent of the engine output shaft.