An aircraft propulsion system with a drag reduction portion adapted to reduce skin friction on at least a portion of the external surface of an aircraft. The drag reduction portion may include an inlet to ingest airflow. The aircraft may also have an internally cooled electric motor adapted for use in an aerial vehicle. The motor may have its stator towards the center and have an external rotor. The rotor structure may be air cooled and may be a complex structure with an internal lattice adapted for airflow. The stator structure may be liquid cooled and may be a complex structure with an internal lattice adapted for liquid to flow through. A fluid pump may pump a liquid coolant through non-rotating portions of the motor stator and then through heat exchangers cooled in part by air which has flowed through the rotating portions of the motor rotor. The drag reduction portion and the cooled electric motor portion may share the same inlet.

An aircraft propulsion system with a drag reduction portion adapted to reduce skin friction on at least a portion of the external surface of an aircraft. The drag reduction portion may include an inlet to ingest airflow. The aircraft may also have an internally cooled electric motor adapted for use in an aerial vehicle. The motor may have its stator towards the center and have an external rotor. The rotor structure may be air cooled and may be a complex structure with an internal lattice adapted for airflow. The stator structure may be liquid cooled and may be a complex structure with an internal lattice adapted for liquid to flow through. A fluid pump may pump a liquid coolant through non-rotating portions of the motor stator and then through heat exchangers cooled in part by air which has flowed through the rotating portions of the motor rotor. The drag reduction portion and the cooled electric motor portion may share the same inlet.

A thermal management system for transferring heat from a heat load includes a composite structural member that supports a heat load source and a heat transfer member in thermal contact with the composite structural member, and in thermal contact with a heat sink. The system further includes at least one thermally-conductive first fastener that is in thermal contact with the heat transfer member, couples the heat load source to the composite structural member, and conducts heat from the heat load source into the heat transfer member. The heat transfer member conducts heat from the thermally-conductive first fastener to the heat sink.

A rectification structure body 100 of a flying vehicle is provided with a rectification section 30, a heat input control section 20 and a vacuum thermal insulation section 10. The rectification section 30 has a rectification surface 30a and a back surface 30b. The rectification surface 30a rectifies airflow 5 from a travelling direction. The back surface 30b is arranged opposite to the rectification surface 30a. The heat input control section 20 is connected to the back surface 30b. The vacuum thermal insulation section 10 is connected to the heat input control section 20 and its surface is formed of rigid body. In addition, the heat input control section 20 is sandwiched between the back surface 30b and the vacuum thermal insulation section 10.

A rectification structure body 100 of a flying vehicle is provided with a rectification section 30, a heat input control section 20 and a vacuum thermal insulation section 10. The rectification section 30 has a rectification surface 30a and a back surface 30b. The rectification surface 30a rectifies airflow 5 from a travelling direction. The back surface 30b is arranged opposite to the rectification surface 30a. The heat input control section 20 is connected to the back surface 30b. The vacuum thermal insulation section 10 is connected to the heat input control section 20 and its surface is formed of rigid body. In addition, the heat input control section 20 is sandwiched between the back surface 30b and the vacuum thermal insulation section 10.

A gas turbine engine is provided. The gas turbine engine defines a radial direction. The engine includes: an airfoil positioned within an airflow and extending between a root end and a tip along the radial direction; and a mount coupled to or formed integrally with the root end of the airfoil for mounting the airfoil to the engine, the mount including an outer surface along the radial direction exposed to the airflow and defining an air-cooling channel extending between an inlet and an outlet, the inlet positioned on the outer surface of the mount.

A gas turbine engine is provided. The gas turbine engine defines a radial direction. The engine includes: an airfoil positioned within an airflow and extending between a root end and a tip along the radial direction; and a mount coupled to or formed integrally with the root end of the airfoil for mounting the airfoil to the engine, the mount including an outer surface along the radial direction exposed to the airflow and defining an air-cooling channel extending between an inlet and an outlet, the inlet positioned on the outer surface of the mount.

An ultrasonic airflow severing resistance reducing device for vehicles, including an airflow severing blade assembly, shock-absorbing installation assemblies, ultrasonic vibration components, and a rubber sealing ring assembly; a bottom portion of the airflow severing assembly is provided with an accommodating cavity; a top end of each shock-absorbing installation assembly is fixedly connected to an inner cavity wall of the accommodating cavity; a bottom end of each shock-absorbing installation assembly is formed with a fixed installation bottom portion; the ultrasonic vibration components are installed on the inner cavity wall; the rubber sealing ring assembly is arranged and installed on the airflow severing blade assembly. The ultrasonic vibration components drive the airflow severing blade assembly to generate high frequency ultrasonic vibration to sever airflow.

A computing system for an unducted rotor engine with a variable pitch vane assembly in aerodynamic relationship with an unducted rotor assembly, including a sensor-based controller configured to execute a first set of operations and a model-based controller configured to execute a second set of operations. The first set of operations includes obtaining a first signal corresponding to a commanded low spool speed; obtaining a second signal indicative of a pitch angle corresponding to thrust output from the unducted rotor assembly and variable pitch vane assembly; generating a pitch feedback signal corresponding to a commanded adjustment to the pitch angle based at least on one or both of a variable blade pitch angle or a variable vane pitch angle. The second set of operations include obtaining a desired thrust output via a throttle input; determining, at least via a power management block, a commanded thrust output signal; receiving the commanded thrust output signal; and generating an output signal.

A computing system for an unducted rotor engine with a variable pitch vane assembly in aerodynamic relationship with an unducted rotor assembly, including a sensor-based controller configured to execute a first set of operations and a model-based controller configured to execute a second set of operations. The first set of operations includes obtaining a first signal corresponding to a commanded low spool speed; obtaining a second signal indicative of a pitch angle corresponding to thrust output from the unducted rotor assembly and variable pitch vane assembly; generating a pitch feedback signal corresponding to a commanded adjustment to the pitch angle based at least on one or both of a variable blade pitch angle or a variable vane pitch angle. The second set of operations include obtaining a desired thrust output via a throttle input; determining, at least via a power management block, a commanded thrust output signal; receiving the commanded thrust output signal; and generating an output signal.