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, 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 semi-passive cooling system for a component exposed to a fluid flow utilizes a hierarchical vasculature and a sacrificial transpirant to cool the component. The component includes a body that defines a transpirant reservoir and the hierarchical vasculature. The transpirant is configured to transition between a solid phase and a vapor phase over an operating temperature range of the component.

An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising rotor blades; and at least one fuselage shield removably coupled to the fuselage at a location in alignment with the rotor blades along the lateral direction.

An aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes a fuselage; an engine mounted at a location spaced from the fuselage of the aircraft, the engine comprising rotor blades; and at least one fuselage shield removably coupled to the fuselage at a location in alignment with the rotor blades along the lateral direction.

A method of generating a pressure wave proximate an airflow surface and altering airflow to promote a localized lowering of skin friction over the airflow surface is described herein. A series of pressure waves may be configured to create a virtual riblet to control turbulent vortices in a boundary layer adjacent to the airflow surface creating a virtual riblet. The pressure waves may be configured to prevent disruption of the flow of air relative to at least one of a step or a gap associated with the airflow surface. The pressure wave generating system may be comprised of at least one of a thermoacoustic material, a piezoelectric material and a semiconductor material, and a microelectric circuit.

A method of generating a pressure wave proximate an airflow surface and altering airflow to promote a localized lowering of skin friction over the airflow surface is described herein. A series of pressure waves may be configured to create a virtual riblet to control turbulent vortices in a boundary layer adjacent to the airflow surface creating a virtual riblet. The pressure waves may be configured to prevent disruption of the flow of air relative to at least one of a step or a gap associated with the airflow surface. The pressure wave generating system may be comprised of at least one of a thermoacoustic material, a piezoelectric material and a semiconductor material, and a microelectric circuit.

High-pressure fan duct bleed air is used to pressurize a cavity between the fan duct inner wall and the inner wall thermal insulation blankets. The cavity is pressurized to prevent hot air from the nacelle core compartment from flowing under the insulation blankets and degrading the fan duct inner wall. Pressure regulating valves (PRV) allow better control of the cavity pressure during different phases of the flight profile and under different levels of insulation blanket seal degradation by passively controlling exit area from the cavity based on an established pressure limit. Moreover, the pressurization system can be implemented as a passive cooling system by increasing the mass flow rate into the cavity and then the core compartment to a level suitable for core compartment cooling. The cooling air can be vented at the forward end of the insulation blanket assembly to provide core compartment ventilation flow, or vented through dedicated ports in the insulation blanket for targeted core compartment component cooling.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a low speed spool and a high speed spool; a rotor assembly coupled to the low speed spool; an electric machine rotatable with the low speed spool for extracting power from the low speed spool, for adding power to the low speed spool, or both; and an inter-spool clutch positioned between the low speed spool and the high speed spool for selectively coupling the low speed spool to the high speed spool.

In an aspect of the present disclosure, an aircraft defining a longitudinal direction and a lateral direction is provided. The aircraft includes: a fuselage; a single unducted rotor engine mounted at a location spaced from the fuselage of the aircraft, the single unducted rotor engine comprising an unducted rotor assembly having a single stage of rotor blades; and a fuselage shield attached to or formed integrally with the fuselage at a location in alignment with the single stage of rotor blades of the unducted rotor assembly along the lateral direction.