Inlet systems are described. An example inlet system has a first wall, a second wall, and a compressor. The first wall has a first end, a second end, and defines a passageway, a first passageway opening, and a second passageway opening. The passageway extends from the first passageway opening to the second passageway opening. The first wall and the second wall cooperatively define an inlet, an outlet, and a channel. The channel extends from the inlet to the outlet. The compressor is disposed within the passageway and is configured to pressurize fluid that passes through the passageway. The first end of the first wall is disposed outside of the channel. The passageway is in fluid communication with the channel.

An aerospace plane having an elongate body supporting a pair of wings each having at least two angled, trailing edge portions. The pair of wings are adapted to extend away from the elongated body in opposing directions. A landing gear assembly is operatively associated with the elongated body to be moveable from a retracted position where the landing gear assembly is substantially locatable within the elongated body and an extended position where the landing gear assembly extends at least partially away from the elongated body. At least one engine adapted to generate thrust. At least one stabilizer adapted to assist with movement of the aerospace plane during flight. The at least one engine is located at least partially within an intake housing adapted to direct air into said at least one engine. The intake housing having at least one inlet door adapted to move from a fully open position, which allows air to pass into the engine, to a sealingly closed position which prevents air from flowing into the engine, when the engine is shut down during flight.

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.

The present invention provides a hypersonic large internal contraction ratio air inlet channel having stepless adjustable air release valve, including an air inlet channel front body, an air-discharging slit cover plate, sidewalls, a lip cover, air-discharging cavities, valve plates, partition plates, a rotatable shaft, an expansion section and a driver. The valve plates are rotated through the driver according to the actual working conditions of air inlet channel, the minimum cross-section of the air discharging flow path is thus changed, and a stepless dynamic adjustment of the air discharging flow of the air inlet channel can be realized, so that the aerodynamic performance of the air inlet channel is improved, and the air discharging resistance of the air vehicle is reduced.

A flight vehicle engine includes an isolator with 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 method of operating a gas turbine engine includes generating a flow of combustion products from a gas turbine generator that has a gas generator axis of rotation. A duct is oriented in a first position to direct the flow of combustion products that have passed over at least one gas generator turbine rotor through a fan drive turbine in response to a first desired flight condition. An axis of rotation of the fan drive turbine is transverse to a gas generator axis of rotation. The duct is oriented in a second position to direct the flow of combustion products that have passed over at least one gas generator turbine rotor through an augmentor section in response to a second desired flight condition.

A nozzle arrangement is disclosed herein for use with a supersonic jet engine that is configured to produce a plume of exhaust gases. The nozzle arrangement includes, but is not limited to, a nozzle having a trailing edge and a plug body partially positioned within the nozzle. The plug body has an expansion surface and a compression surface downstream of the expansion surface. A protruding portion of the plug body extends downstream of the trailing edge for a length greater than a conventional plug body length. The plug body is configured to shape the exhaust gases to flow substantially parallel to a free stream of air flowing off of the trailing edge of the nozzle and to cause the plume of exhaust gases to isentropically turn the free stream of air to move in a direction parallel to a longitudinal axis of the plug body.

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.

An assembly is provided for an aircraft propulsion system. This assembly includes a variable area inlet, and the variable area inlet includes an inlet structure, an inlet lip, an inlet orifice and an inlet passage extending within the variable area inlet from the inlet orifice. The inlet lip is configured to pivot about a pivot axis between a first position and a second position. The inlet orifice is formed by and between the inlet structure and the inlet lip. The inlet orifice has a first area when the inlet lip is in the first position. The inlet orifice has a second area when the inlet lip is in the second position. The second area is different than the first area.