An inlet for a flight vehicle engine, such as for a supersonic or hypersonic engine, includes an internal flow diverter to divert boundary layer flow. The flow diverter is configured to minimize disruption to flow outside the diverted boundary by being configured through use of a flow field that is also used to configure the walls of the inlet. The flow field that is used to configure an inlet-creating shape and a diverter-creating shape has the same flow generator, contraction ratio, compression ratio, mass capture ratio, pressure ratio between entrance and exit, and/or Mach number, for example. The internal diverter may be configured so as to allow arbitrary selection of a leading edge shape for the internal diverter, for example to use a shape that helps avoid radar detection.

Passive reduction of internal jet engine component temperature in supersonic and hypersonic vehicles results from use of nanocomposite optical ceramic materials between the heat-generating portions of each jet engine and the ambient environment, allowing heat dissipation from the jet engine components directly to the ambient environment. A propulsion-airframe integrated scramjet aircraft includes a jet engine and an airframe supporting the jet engine, with at least a portion of the airframe between a heat-generating portion of the jet engine and an ambient environment comprising a nanocomposite optical ceramic material in the form of a panel or a grid of windows each supported within a frame. The nanocomposite optical ceramic material portion of the airframe disposed between the heat-generating portion of the jet engine and the ambient environment is infrared-transparent, and may transmit at least 75% of heat energy from the heat-generating portion of the jet engine to the ambient environment.

A regeneration cooler (2) includes a passage forming structure (10) in which a fuel passage (11) is formed for liquid fuel to be supplied. A coating (12, 12A) is formed in the fuel passage (11) to at least partially cover a wall surface of the fuel passage (11). The coating (12, 12A) contains metal particles (13) adhered and fixed to the wall surface (11a) of the fuel passage (11) and a coating material (14, 17).

A scramjet engine includes first and second flow path forming members and first and second fuel injection devices. A flow path formed between the first and second flow path forming members includes a turbulence forming region where compressed air is introduced and a combustion region located downstream thereof. The second flow path forming member is formed with a protrusion in the turbulence formation region. The first fuel injection device is configured to inject fuel into the compressed air via a first fuel nozzle. The second flow path forming member is formed with a cavity located in the combustion region. The second fuel injection device is configured to inject fuel into the compressed air via a second fuel nozzle. The cavity is provided with an inclined surface connected to a bottom surface. An inclination of the inclined surface is adjusted so that a shock wave is generated in the combustion region.

A power module includes a turbo-generator with a propellant selector valve, a stored energy module connected to the propellant selector valve, and bleed air conduit. The bleed air conduit is connected to the propellant selector valve, wherein the propellant selector valve has a first position, wherein the stored energy tank is in fluid communication with the turbo-generator, and a second position, wherein the bleed air conduit is in fluid communication with the turbo-generator. Vehicles and methods of generating electrical power are also described.

An aerospace hybrid hypersonic propulsion system which has a common core airflow path through the engine combining subsonic, transonic, supersonics and hypersonic propulsion system and architecture in such a way that five known engine cycles known in the art are configured and connected to operate seamlessly with a hybrid electric and thermal cycle.

An aircraft includes a fuselage and an engine housed in the fuselage. A ram-air engine inlet is formed on an exterior of the fuselage. The ram-air engine inlet is defined, at least in part, by a cowling door. The cowling door is moveable between a closed position and an open position. An intake duct fluidly couples the ram-air engine inlet to the engine. A filter is disposed across the intake duct. A bypass duct is formed in the ram-air engine inlet aft of the intake duct. The bypass duct is operable to be selectively opened and closed.

Vehicles, such as aircraft, may include turbine-based combined cycle power plants (TBCC) for power to achieve high-mach speeds. Cooling systems for such TBCC may include a turbine-generator arranged to be driven for rotation by ambient air to reduce the temperature of the ambient air while providing electric power for use under cocooning of a primary gas turbine engine in favor of a scramjet engine during high-mach travel.

A rocket motor includes a case and first and second nozzles in the case. The first nozzle is disposed in the second nozzle. The first nozzle includes a forward leg, a rear leg, and an intermediate leg. The intermediate leg has a convex conical geometry, and the forward leg has a forward lip that is spaced from the case. The rear leg has a rear lip that is spaced from the case. The forward leg and the rear leg at least partially define a flow passage through the first nozzle. The first nozzle is exclusively secured by the intermediate leg to at least one of the case or the second nozzle. At least a portion of a fragmentation system is disposed between the first and second nozzles.

An engine system that produces all required thrust for an aerospace vehicle from takeoff through space operation utilizing a turbo ram rocket exhaust nozzle and M-Spike rocket consisting of airbreathing and non-airbreathing propulsion apparatuses. The airbreathing system consists of a turbine engine, a ramjet or scramjet, and the non-airbreathing system is a rocket motor. The turbine engine consists of a turbojet or turbofan configuration. The air breathing turbine, ramjet or scramjet feature an air inlet mechanism, and combustion fuel. The non-airbreathing rocket system includes separate oxidizer system, and either a separate or same source of combustion fuel as the turbine. Airflow velocities in the turbine bypass duct, and burner system, include subsonic and supersonic velocities for ramjet or scramjet operation. The rocket engine utilize either cryogenic or a non-cryogenic fuel and oxidizer system.