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.

The invention relates to a thrust nozzle for a turbofan engine of a supersonic aircraft, wherein the thrust nozzle includes a thrust nozzle wall, and a flow channel that is delimited radially outwards by the thrust nozzle wall, wherein the flow channel has a nozzle throat surface and a central body that is arranged in a flow channel. According to the invention, the central body forms a bypass channel, which extends within the central body, and which is designed for the gas of the flow channels to flow through. The bypass channel has at least one upstream inlet opening, which is arranged upstream of the nozzle throat surface of the flow channel, and at least one downstream outlet opening, which is arranged downstream of the nozzle throat surface of the flow channel

The present invention discloses a supersonic inlet with synchronous adjustment of capture area and throat area, wherein the throat area may be adjusted by providing a movable throat section, while the capture area may be adjusted by providing a movable cowl section at the front end of the cowl lip, thereby realizing the effect that the capture area and the throat area of the inlet may be adjusted in synchronization. Meanwhile, the present invention also provides a control method and a design method of the above-mentioned inlet. The present invention greatly simplifies the actuation system and control system, significantly reduces the weight of the accessory system, and enables the performance of the inlet within a wide envelope range to be maintained in an excellent state.

An inlet duct for use with an engine is presented. The invention includes a duct structure, at least one spike disposed along an interior surface of the duct structure, and an inlet throat formed by one or more apexes disposed along an equal number of spikes. The inlet throat corresponds to the minimum cross-sectional area through which airflow passes as otherwise allowed by the maximal obstruction formed by the apex(es) within the duct structure. Each spike is bounded by a longitudinal ridge and a lateral ridge along an upper end and a base along a lower end. The longitudinal ridge and the lateral ridge intersect at the apex. In preferred embodiments, the longitudinal ridge is at least partially non-linear so as to properly conform to the interior surface of the duct structure. The portion of each spike upstream of the inlet throat functions primarily as a supersonic diffuser. The portion of each spike downstream of the inlet throat functions primarily as a subsonic diffuser. Airflow is isentropically compressed and then expanded within the inlet duct so that greater-than-subsonic flow at an input end is reduced to subsonic flow at an output end.

A supersonic inlet includes a relaxed isentropic compression surface to improve net propulsive force by shaping the compression surface of the inlet to defocus the resulting shocklets away from the cowl lip. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce the cowl lip surface angle, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining overall engine performance.

An engine for mounting in or to an aircraft includes a stage of compression airfoils rotatable about a central axis; a casing surrounding the stage of compression airfoils and defining an inlet; and a low-distortion inlet assembly mounted within the inlet. The inlet assembly includes one or more structural members mounted at predetermined locations around a circumference of the central axis within the inlet, the predetermined locations defining an airflow distortion exceeding a predetermined threshold; and at least one airflow modifying element configured within the inlet so as to reduce airflow distortion entering the stage of compression airfoils.

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.

Disclosed is a supersonic speed attenuator for an air inlet of an aircraft power plant, having a conical external wall with a tapered end upstream and an internal dividing wall that delimits, with the conical external wall, a cavity; the supersonic speed attenuator further having a de-icing device having: an internal wall mounted in the cavity opposite the conical external wall so that, together, the two walls delimit a calibrated de-icing volume; at least one member for supplying a de-icing air flow, opening into the de-icing volume; and at least one member for discharging the de-icing air flow from the de-icing space.

A supersonic jet aircraft and a method of operating the same. The supersonic jet aircraft having at least three turbofan engines and an engine management computer. A first engine of the at least three turbofan engines is configured to be de-activatable during flight to move from an operational state in which it provides thrust to an operational state in which it stops providing thrust. Other engines of the at least three turbofan engines are configured to provide sufficient thrust to the supersonic jet aircraft when the first engine is de-activated such that the aircraft can perform a supersonic climb operation and/or a supersonic cruise operation.

A hybrid propulsion system and methods for cooling the same are provided. The system may comprise a gas turbine and a secondary engine. The gas turbine engine may have a core passage and an engine compartment. The secondary engine may be a supersonic and/or hypersonic engine. The system may comprise a thermal barrier, an inlet and an exhaust. The thermal barrier may longitudinally envelope the gas turbine engine. The thermal barrier may comprise an inner envelope, an outer envelope, an upstream opening, and a downstream opening. The inlet may be in fluid communication with the ambient environment and the gas turbine engine via the upstream opening. The exhaust may be in fluid communication with the ambient environment and the gas turbine engine via the downstream opening. The engine compartment may be located between a boundary of the core passage and the inner envelope.