An external cooling system for hydraulic fluid of an aircraft hydraulic system. The external cooling system includes a heat exchanger, where an input side of the heat exchanger is connected to a hydraulic fluid reservoir of the aircraft hydraulic system and an output side of the heat exchanger is connected to suction ports of a return side of an electric motor driven pump (EMDP) of the aircraft hydraulic system. The external cooling system operates on 120 VAC power and the hydraulic fluid does not exceed a maximum pressure of 50 pounds per square inch gauge. The EMDP pumps hydraulic fluid through the hydraulic system under conditions wherein fuel tanks in the aircraft are empty, and the external cooling system cools the hydraulic fluid as the EMDP pumps the hydraulic fluid, wherein the hydraulic fluid passes from the hydraulic fluid reservoir and through the external cooling system before entering the EMDP.

A high efficiency hydrogen fuel system for an aircraft at high altitude which utilizes compressors to compress air to a sufficiently high pressure for the fuel cell. Liquid hydrogen is compressed and then utilized in heat exchangers to cool the compressed air, maintaining the air at a temperature low enough for the fuel cell. The hydrogen is also used to cool the fuel cell as it is also depressurized prior to its entry in the fuel cell cycle. A water condensation system allows for water removal from the airstream to reduce impacts to the atmosphere. The hydrogen fuel system may be used with VTOL aircraft, which may allow them to fly at higher elevations. The hydrogen fuel system may be used with other subsonic and supersonic aircraft, such as with asymmetric wing aircraft.

A nacelle for an engine includes an exterior housing a least partially surrounding the engine. The nacelle further includes a front portion proximate an engine face of the engine. The front portion of the nacelle defines an opening into an interior of the nacelle. The nacelle further includes a seal disposed proximate the opening. The seal is configured to selectively allow air into the interior of the nacelle. The nacelle further includes a heat exchanger disposed within the interior of the nacelle. The heat exchanger is configured to exchange heat between a fluid flowing within the heat exchanger and air at the interior of the nacelle.

A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. The leading edge assembly includes a plurality of structural layers and a plurality compliant layers alternately stacked with each other to facilitate thermal expansion and movement between the plurality of structural layers, while also providing a thermal break between the plurality of structural layers.

A hypersonic aircraft includes one or more leading edge assemblies that are designed to cool the leading edge of certain portions of the hypersonic aircraft that are exposed to high thermal loads, such as extremely high temperatures and/or thermal gradients. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. A coolant supply may be in fluid communication with at least one fluid passageway that passes through the outer wall to deliver a flow of cooling fluid, such as liquid metal, to the stagnation point. The liquid metal vaporizes when the leading edge experiences a high heat load, thereby transpiration cooling the leading edge and/or facilitating a magnetohydrodynamic process for generating thrust or electricity.

A hypersonic aircraft includes one or more leading edge assemblies that are designed to cool the leading edge of certain portions of the hypersonic aircraft that are exposed to high thermal loads, such as extremely high temperatures and/or thermal gradients. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. A coolant supply may be in fluid communication with at least one fluid passageway that passes through the outer wall to deliver a flow of cooling fluid to the stagnation point. In addition, a nose cover is positioned at least partially over or within the at least one fluid passageway and is formed from a material that ablates or melts when the leading edge is exposed to a predetermined critical temperature, the nose cover being configured for restricting the flow of coolant until the nose cover is ablated or melted away.

A centrifugal separator for removing particles from a fluid stream includes an angular velocity increaser configured to increase the angular velocity of a fluid stream, a flow splitter configured to split the fluid stream to form a concentrated-particle stream and a reduced-particle stream, and an exit conduit configured to receive the reduced-particle stream. An inducer assembly for a turbine engine includes an inducer with a flow passage having an inducer inlet and an inducer outlet in fluid communication with a turbine section of the engine, and a particle separator, which includes a particle concentrator that receives a compressed stream from a compressor section of the engine and a flow splitter. A turbine engine includes a cooling air flow circuit which supplies a fluid stream to a turbine section of the engine for cooling, a particle separator located within the cooling air flow circuit, and an inducer forming a portion of the cooling air flow circuit in fluid communication with the particle separator. A method of cooling a rotating blade of a turbine engine having an inducer includes directing a cooling fluid stream from a portion of turbine engine toward the rotating blade, separating particles from the cooling fluid stream by passing the cooling fluid stream through a inertial separator, accelerating a reduced-particle stream emitted from the inertial separator to the speed of the rotating blade, and orienting the reduced-particle stream by emitting the reduced-particle stream from the inertial separator into a cooling passage in the inducer.

A propulsion system for an aircraft is disclosed, and includes a first propeller, a second propeller, a first hybrid propulsion system, a second hybrid propulsion system, and across-connecting clutch. The first hybrid propulsion system includes a first motor coupled to a first engine by a first overrunning clutch, where the first hybrid propulsion system is operably coupled to drive the first propeller. The second hybrid propulsion system includes a second motor coupled to a second engine by a second overrunning clutch, where the second hybrid propulsion system is operably coupled to drive the second propeller. The cross-connecting clutch is operably coupled to both the first hybrid propulsion system and the second hybrid propulsion system and configured to actuate into an engaged position.

Systems and methods for cooling bleed air from an aircraft engine are described. An example system includes a housing to bifurcate airflow exiting a fan of an engine fan system, a pre-cooler assembly disposed within the housing to remove heat from the bleed air, and an inlet duct within the housing to direct the airflow to the pre-cooler assembly. The example system also includes at least one diverter duct within the housing and coupled to the inlet duct to divert the airflow around the pre-cooler assembly. The at least one diverter duct includes an exit directing the airflow into a fan duct of the engine fan system. The example system also includes at least one valve to control an amount of airflow through the inlet duct and an amount of airflow through the at least one diverter duct.

An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inner structure and a deflection limiter. The nacelle inner structure includes an internal compartment and a cowl. The internal compartment is configured to house a core of a gas turbine engine. The cowl is configured to form an outer radial periphery of the internal compartment. The cowl is also configured to form an outer radial periphery of a compartment exhaust to the internal compartment at an aft end of the cowl. The deflection limiter is attached to the cowl. The deflection limiter is configured to limit radial outward movement of the cowl.