A method of reducing low-energy flow in a flight vehicle engine includes an isolator of the engine having a swept-back wedge to improve flow mixing. The wedge includes forward shock-anchoring locations, such as edges or rapidly-curved portions, that anchor oblique shocks in situations where the isolator has sufficient back pressure. The swept-back wedge may also create swept oblique shocks along its length. Boundary layer flow streamlines are diverted running parallel to or parallel but moving outward conically to the swept-wedge leading edge moving outboard and upward. The non-viscous flow outside the boundary layer is processed through the swept-back ramp shock and diverted outboard and upward as well. The outboard aft portion of the wedge at the sidewall intersection may also induce shocks and divert flow near the walls closer toward the walls and upward, and/or improve flow mixing.

A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

A motor and fuel-powered hybrid system of a rocket thruster is disclosed, which mainly provides power through a motor and a fluid fuel injector. In particular, at the beginning stage of the rocket lift-off, the motor drives the compressor to provide power to send the rocket into air. When the speed and height of the rocket gradually increase, the fuel is ignited to give power to keep propelling the rocket, thereby reducing the fluid fuel that needs to be carried on the rocket, increasing the rocket's loading space and enhancing the carrying capacity.

A motor and fuel-powered hybrid system of a rocket thruster is disclosed, which mainly provides power through a motor and a fluid fuel injector. In particular, at the beginning stage of the rocket lift-off, the motor drives the compressor to provide power to send the rocket into air. When the speed and height of the rocket gradually increase, the fuel is ignited to give power to keep propelling the rocket, thereby reducing the fluid fuel that needs to be carried on the rocket, increasing the rocket's loading space and enhancing the carrying capacity.

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.

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.

Airframe integrated scramjet engines are disclosed. Scramjet engines within the scope of this disclosure may be configured to integrate smoothly with an airframe of a hypersonic flight aircraft or vehicle. The scramjet engine may include capture shape of an inlet configured to capture airflow, a combustor configured for combustion of fuel and air, and an exit shape of a nozzle configured for expansion of the combusted fuel and air to provide hypersonic thrust. In some embodiments, the scramjet engine has a fixed geometry and a transitioning cross-sectional shape over its full length. The scramjet engine is configured to be a component of launch vehicle system.

Airframe integrated scramjet engines are disclosed. Scramjet engines within the scope of this disclosure may be configured to integrate smoothly with an airframe of a hypersonic flight aircraft or vehicle. The scramjet engine may include capture shape of an inlet configured to capture airflow, a combustor configured for combustion of fuel and air, and an exit shape of a nozzle configured for expansion of the combusted fuel and air to provide hypersonic thrust. In some embodiments, the scramjet engine has a fixed geometry and a transitioning cross-sectional shape over its full length. The scramjet engine is configured to be a component of launch vehicle system.

A system for bistatic radar target detection employs an unmanned aerial vehicle (UAV) having a ramjet providing supersonic cruise of the UAV. Deployable antenna arms support a passive radar receiver for bistatic reception of reflected radar pulses. The UAV operates with a UAV flight profile in airspace beyond a radar range limit. The deployable antenna arms have a first retracted position for supersonic cruise and are adapted for deployment to a second extended position acting as an airbrake and providing boresight alignment of the radar receiver. A mothership aircraft has a radar transmitter for transmitting radar pulses and operates with an aircraft flight profile outside the radar range limit. A communications data link operably interconnects the UAV and the tactical mothership aircraft, transmitting data produced by the bistatic reception of reflected radar pulses in the UAV radar antenna to the mothership aircraft.