A method to predict an onset of flow separation from a surface of an inner barrel of a nacelle is disclosed. In various embodiments, the method includes determining a static pressure distribution about the inner barrel surface of the nacelle; determining a mean static pressure value and a minimum static pressure value using the static pressure distribution; determining a separation indicator value using the mean static pressure value and the minimum static pressure value; and comparing the separation indicator value against a separation threshold value.

A tile assembly (22) which, in use, is fitted to a base structure to form at least part of a fluid washed surface. The tile assembly comprises a housing (42) with at least one plenum (45) being provided within the housing (42). A wall (44) of the housing (42) is provided with a plurality of flow passages (46) which extend from the plenum side of the wall (44) to an outer surface (48) of the wall. Flow passage closures (50) are provided which are operable to open and close at least some of the flow passages (46).

An actuator assembly is capable of manipulating a fluid flowing around a flow body, the fluid being received or able to be received in a volume of at least one cavity arranged in the flow body, and the fluid passing through at least one opening in the at least one cavity during manipulation of the fluid. In this process, the volume of the at least one cavity can be changed by moving a wall portion delimiting or defining the cavity. The actuator assembly has a drive unit with at least one actuator, which executes a periodic movement over time when actuated, causing a translational movement of the wall portion delimiting or defining the cavity and the wall portion being shaped in terms the topology thereof in such a way that it is adapted to the shape of the at least one cavity with the at least one opening thereof.

An aerodynamic surface of an aircraft comprises a main part having a leading and a trailing edges and having an airfoil section. The aerodynamic surface also having at least two vortex generators in the form of teeth having edges along the length thereof. The teeth are mounted on the leading edge of the main part so as to be capable of generating two vortex cores on one tooth. The edges of a tooth adjoin the leading edge of the main part of the aerodynamic surface. The radius of an edge of each tooth along the length of the vortex generator is five times less than the radius of the leading edge of the main part. The main part of the aerodynamic surface has a cambered airfoil section, wherein the teeth are mounted with a deflection towards the smallest degree of curvature of the airfoil section of the main part. The invention is intended for reducing an aerodynamic drag at low angles of attack while maintaining an increased load hearing capacity of the aerodynamic surface by generating vortex cores adjoining one of the sides thereof.

A vehicle includes a main body and a gas generator producing a gas stream. At least one fore conduit and tail conduit are fluidly coupled to the generator. First and second fore ejectors are fluidly coupled to the at least one fore conduit. At least one tail ejector is fluidly coupled to the at least one tail conduit. The fore ejectors respectively include an outlet structure out of which gas from the at least one fore conduit flows. The at least one tail ejector includes an outlet structure out of which gas from the at least one tail conduit flows. First and second primary airfoil elements have leading edges respectively located directly downstream of the first and second fore ejectors. At least one secondary airfoil element has a leading edge located directly downstream of the outlet structure of the at least one tail ejector.

The invention discloses a lift source for an aircraft comprising a fuselage and wings, wherein first channels are formed in the wings, a plurality of first inlets are formed in upper surfaces of the wings, a plurality of first pressure ports are formed in lower surfaces of the wings and are communicated with the first inlets via the first channels; and spoiler devices are arranged in the first channels and under the effect of the spoiler devices, form high-speed fluid layers on the upper surfaces of the wings, thereby generating a pressure difference from the lower surfaces of the wings which counteracts an external fluid pressure on the upper surfaces of the wings in the opposite direction, so a lift is generated by reduction of fluid resistance when fluid flows through the upper and lower surfaces of the wings, thereby developing a high-speed aircraft with a larger lift and thrust.

Active fluid control systems and related methods are disclosed. A disclosed example active fluid control system includes a plurality of plenums coupled together to define a fluid flow passageway, and a plurality of fluidic actuators coupled to outer surfaces of respective ones of the plenums. The fluidic actuators define actuator inlets and actuator outlets. The fluid flow passageway defined by the plenums to fluidly couple the fluidic actuators and a pressurized fluid supply source. The plenums are configured to couple to an aircraft structure supporting an aerodynamic surface to enable the actuator outlets to be mounted to the aerodynamic surface. The fluidic actuators are configured to provide the pressurized fluid to the aerodynamic surface to modify an aerodynamic characteristic of the aerodynamic surface.

Example active flow control systems and methods for aircraft are described herein. An example active flow control system includes a plurality of nozzles arranged in an array across a surface of an aircraft. The nozzles are oriented to eject air across the surface to reduce airflow separation. The active flow control system also includes an air source coupled to the nozzles and a controller to activate the nozzles to eject air from the air source in sequence from outboard to inboard and then from inboard to outboard to create a wave of air moving from outboard to inboard and then from inboard to outboard across the surface.

A propulsion system for an aircraft can include an electric power source and an electric propulsion assembly having an electric motor and a propulsor. The propulsor can be powered by the electric motor. An electric power bus can electrically connect the electric power source to the electric propulsion assembly. The electric power source can be configured to provide electrical power to the electric power bus. An inverter converter controller can be positioned along the electric power bus and can be electrically connected to the electric power source at a location downstream of the electric power source and upstream of the electric propulsion assembly.

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.