292results about "Boundary layer controls" patented technology

The invention relates to an air system for an aircraft, that includes air consumers; air sources and a network of ducts and associated control valves controlled by a control unit. The air system is characterized in that: the network of ducts and associated valves includes at least one isolation valve, arranged between an air bleed device and an air duct connecting an air conditioning pack and an auxiliary power unit; the control unit is configured to be able to determine an ideal configuration of the control valves according to the identified requirements of each consumer and a degraded configuration that makes it possible to supply air to predetermined air consumers from the available air sources when the ideal configuration is not attainable.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

The invention relates to the technical field of vehicle active control, in particular to a vehicle stabilizing system based on active airflow compensation and a control method, and the system comprises a distributed airflow generation array unit, a multi-source sensing system unit, an intelligent control unit and an energy management unit. Directional compensation airflow is generated through the distributed airflow generation array unit, the multi-source sensing system unit monitors a flow field and motion parameters in real time, the intelligent control unit fuses a hybrid control strategy and reinforcement learning to achieve dynamic compensation and algorithm iteration, and the energy management unit recycles braking energy and optimizes distribution. The stability and response speed of the vehicle in a complex environment are remarkably improved, meanwhile, the airflow utilization rate is increased through the gradient porous material, energy consumption is reduced through reverse airflow power generation, and efficient and energy-saving active stable control is achieved.

A raised structure for reducing frictional drag due to viscosity of a flow toward an object in a direction defining an acute angle with a leading edge of the object. The raised structure includes raised bodies configured to be provided on a surface of the object at the leading edge on a downstream side of a stagnation point of the flow. A height of each raised body changes along a convex curve, and the raised bodies are arranged in an array to define a first uneven shape changing periodically in a first cross section configured to have a constant distance from the stagnation point and be orthogonal to the surface, and a second uneven shape changing in a second cross section configured to be orthogonal to a line composed of the stagnation point and the first cross section, the second uneven shape having concave and convex portions that change periodically.

The invention relates to a blade (7) made of composite material for a turbine engine (10), in particular of an aircraft, the blade (7) comprising an airfoil (70) formed by a fibrous preform embedded in a resin and extending along an axis (A) of elongation, the airfoil (70) comprising a leading edge (73), a trailing edge (74), a pressure side (71) and a suction side (72) connected to the pressure side (71) by the leading edge (73) and the trailing edge (74), the airfoil (70) further comprising at least one serrated element (9) formed by a plurality of teeth (92) and covering at least part of the leading edge (73) and / or the trailing edge (74), wherein the at least one serrated element (9) is made of a thermosetting or thermoplastic material.

Provided is a moving body capable of moving in a fluid, wherein a surface of the moving body includes a first surface to which a first sheet member having a first riblet structure formed thereon is attached, and a second surface to which a second sheet member having a second riblet structure formed thereon is attached, a characteristic of the first riblet structure being different from a characteristic of the second riblet structure.

A propulsion system for an aircraft can include an electric power source and an propulsion assembly having a propulsor. An electric power bus can electrically connect the electric power source to the 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 apparatus and system are provided to provide transpiration cooling for an airframe. An airframe includes a body that contains circuitry, a phase change material (PCM) that changes phase to a gas when cooling the circuitry, a fluid reservoir that retains the PCM in liquid and gaseous form, a vapor reservoir that retains the gas, a thermal sensor that detects a temperature of an external surface of the body, and a controller that control ejection of the gas from the vapor reservoir based on the temperature to control a thickness of a boundary air layer at the external surface. The controller also controls flow of the gas from the fluid reservoir to the vapor reservoir via a valve based on pressure in the vapor reservoir detected by a pressure sensor.

An aircraft may comprise propulsors. The propulsors may be integrated into an array of propulsors. An array of propulsors may be integrated in an airfoil of the aircraft. The aircraft may include a flow turning control system. The flow turning control system may include an end wall and one or more movable flaps. The flow turning control system may enable selective flow turning of the flow of propelled by the array of propulsors. The selective flow turning may include actuating the one or more movable flaps. The selective flow turning may achieve flow turning of over a trailing edge of a movable flap.

The present invention discloses an electroactive smart material extensionally deformable drag reduction surface, comprising a dielectric elastomer film and a flexible electrode layer, wherein the dielectric elastomer film is processed with straight micro-ribs with cross-sectional geometric dimensions of 2-20 μm in width, 10-100 μm in height and 20-200 μm in spacing in the middle of the upper surface, and flexible electrode arrays are distributed on both sides. According to the electroactive smart material extensionally deformable drag reduction surface of the present invention, the drag reduction surface is deformed in the extension direction by turning on and off the power of the flexible electrode array, which can: (1) change the extensional spacing of the straight ribs, (2) form a sinusoidal structure, and (3) simultaneously change the spacing and form a sinusoidal structure. The beneficial effect achieved is that the flexible electrode layer changes with the flow field environment, turns on and off the power, or changes the size of the received voltage signal, thereby changing the spacing between the ribs, the wavelength and amplitude of the sinusoidal structure alone or simultaneously, so that the turbulent friction resistance near the wall can be significantly reduced under different flow field environments, and a better drag reduction effect can be provided.

The embodiment of the present invention discloses a vortex generator for reducing fuselage drag and delaying airflow separation, which relates to the technical field of rotorcraft manufacturing, and can realize the passive flow control of the flow field near the fuselage of the rotorcraft when encountering strong crosswind or sideways flight, thereby achieving the purpose of reducing drag of the rotorcraft, while reducing the implementation cost and failure rate. The present invention includes: the value of the installation angle (2) of the vortex generator is greater than or equal to 10° and less than or equal to 45°; the second corner (20) of the single-piece vortex generator is the angle between the upper surface (7) of the second section and the connection section (19) between the vortex generator and the object surface, the value of the first corner (9) of the single-piece vortex generator is less than 30°, and the value of the second corner (20) is greater than or equal to 40° and less than or equal to 60°. The present invention is suitable for the passive flow control of the flow field near the fuselage of the rotorcraft when encountering strong crosswind or sideways flight.

An air injection assembly for an aircraft is provided. The aircraft includes a fuselage extending longitudinally between a front end and a rear end, and a boundary layer intake fan mounted to the fuselage at the rear end of the fuselage. The air injection assembly includes a plurality of injection ports defined on a surface of the fuselage at locations upstream of the boundary layer intake fan. Supplemental airflow is provided to the injection ports through a fluid passageway, and the supplemental airflow is injected into the injection ports to displace at least a portion of the relatively high-velocity boundary layer airflow. In this manner, the airflow entering the boundary layer intake fan is more uniform, has less vortex distortion, and has a lower average velocity.

A panel for an active laminar flow control arrangement may comprise a longitudinal wall (434) and one or more division walls (428) extending from the longitudinal wall (434) and extending between the first end (491) and the second end (492). The panel may be coupled to a nacelle outer skin (312) via a plurality of adhesive fasteners (326, 322) preinstalled onto an inner surface (314) of the outer skin (312) and onto stiffeners (382) and / or stringers (384) associated with the outer skin (312).

The invention relates to an improved deflection flap for a vehicle with a fuselage featuring an airfoil profile. The deflection flap at the bow is two-part. It includes a device for raising and lowering it.

There is provided a flight vehicle including: a wing unit; a battery that is arranged in the wing unit; an air intake unit that is formed at a position corresponding to the battery on a front side of the wing unit; a heat sink unit that is arranged for the battery and cools the battery by air which flows in from the air intake unit and that includes a ventilation unit having a shape widening from the front side toward a rear side; and an exhaust unit that is formed at a position corresponding to the battery on the rear side of the wing unit and that exhausts air which flows out from the heat sink unit.

The invention relates to an improved hinged deflection panel for a vehicle comprising a fuselage having a wing profile. The hinged deflection panel at the front is a two-part piece. Also provided is a device for lifting and lowering.

An intentional fluid manipulation apparatus (IFMA) assembly that includes an upstream intentional momentum shedding apparatus (IMSA) configured to impart a first induced velocity to a local free stream flow during a nominal operation requirement. The upstream IMSA creates a streamtube. The IFMA includes a downstream IMSA, with some or all of the downstream IMSA being located in a downstream portion of the streamtube. The downstream IMSA imparts a second induced velocity to the local free stream flow within the streamtube. The second induced velocity at the location of the downstream IMSA has a component in a direction opposite to the direction of the first induced velocity at the location of the downstream IMSA.

The invention discloses a device capable of reducing local airflow deposition of a wing of a tilt rotorcraft in the technical field of rotorcrafts, and the device comprises a covering cap frame which is installed on the wing of the tilt rotorcraft and is provided with a covering cap hole; the covering cap plate is mounted on the covering cap frame and used for closing or opening the covering cap hole; the connecting assembly is connected with the opening cover frame and the opening cover plate and drives the opening cover plate to close or open the opening cover hole; the driving assembly is used for driving the connecting assembly to drive the covering cap plate to close or open the covering cap hole, the covering cap is arranged on the wing, 90-degree opening and closing are achieved, and therefore the pulling force loss is reduced.

A device to reduce turbulent flow skin friction uses carbon dioxide, riblets, and suction holes to reduce the overall drag of a body moving in a fluid. A deflector (2) pushes up the boundary layer of the incoming flow on a surface (1). Carbon dioxide is injected into a plenum (3) and ejected through an exhaust slot (4). Riblets of specific dimensions (5) are placed to interact with the carbon dioxide flow that is now the sublayer viscous flow. After passing the riblets, the flow is sucked partly through suction holes (6), preventing flow separation that could increase the form and wave drag.

A laser processing device for generating a plurality of grooves in a surface comprises an optical diffraction arrangement adapted to receive a laser radiation and to generate an output radiation hereupon, the output radiation having a plurality of intensity maxima. An actuator arrangement is provided for generating a relative movement between the output radiation and the surface, wherein each intensity maximum generates a groove of the plurality of grooves.

A vehicle, includes a main body. A fluid generator is coupled to the main body and produces a fluid stream. At least one tail conduit is fluidly coupled to the generator. First and second fore ejectors are coupled to the main body and respectively coupled to a starboard side and port side of the vehicle. The fore ejectors respectively comprise an outlet structure out of which fluid flows. At least one tail ejector is fluidly coupled to the tail conduit. The tail ejector comprises an outlet structure out of which fluid flows. A primary airfoil element includes a closed wing having a leading edge and a trailing edge. The leading and trailing edges of the closed wing define an interior region. The at least one propulsion device is at least partially disposed within the interior region.

The invention relates to an aircraft air system comprising: a plurality of air consumers (6, 16); a plurality of air sources (5); and a network of ducts (23, 25, 26, 27) and associated control valves (PRV, OPV, APUCV, IPCV, FAV) operated by a control unit. The air system is characterized in that: said network of ducts and associated valves includes at least one isolation valve (APUIV), arranged between an air intake device and an air duct connecting an air conditioning pack (6) and an auxiliary power unit (5); said control unit (14) is configured to determine an ideal configuration of the control valves according to the identified needs of each consumer and a degraded configuration that allows predetermined air consumers to be supplied with air from said available air sources when said ideal configuration is not attainable.

A propulsion system 300 for an aircraft 10 includes an electric power source 302 and an electric propulsion assembly 304 having an electric motor 312 and a propulsor 314, the propulsor 314 powered by the electric motor 312. The propulsion system 300 also includes an electric power bus 306 electrically connecting the electric power source 302 to the electric propulsion assembly 304. The electric power source 302 is configured to provide electrical power to the electric power bus 306, and the electric power bus 306 is configured to transfer the electric power to the electric propulsion assembly 304 at a voltage exceeding 800 volts.

A propulsion system coupled to a vehicle. The system includes a convex surface, a diffusing structure coupled to the convex surface, and at least one conduit coupled to the convex surface. The conduit is configured to introduce to the convex surface a primary fluid produced by the vehicle. The system further includes an intake structure coupled to the convex surface and configured to introduce to the diffusing structure a secondary fluid accessible to the vehicle. The diffusing structure comprises a terminal end configured to provide egress from the system for the introduced primary fluid and secondary fluid.

Propulsion assembly for an aircraft, comprising a turbomachine (1), an attachment strut (40) and a structural element (10) carrying the turbomachine (1) by the attachment strut (40), the turbomachine having at least one rotating part (30) rotating about an axis of rotation (A), the rotating part (30) being arranged upstream of the structural element (10) and of the attachment strut (40) so that the air jet exiting from the rotating part (30) impacts the structural element (10) and the attachment strut (40) in the wake thereof, the leading edge of the structural element (10) and / or of the attachment strut (40) locally comprising at least one acoustic attenuation device (50, 60, 70, 80) arranged at least partially in the wake of the rotating part (30), the acoustic attenuation device being a partial modification of the structure and / or of the profile of the leading edge (11).

The present application belongs to the technical field of aircraft skin, and relates to an aircraft elastic porous medium skin, a plurality of hole bodies are prefabricated on the skin, an organ pipe type nozzle is installed in the hole body, the organ pipe type nozzle is connected with a jet control system for providing a heat medium for the organ pipe type nozzle; the organ pipe type nozzle comprises an inlet section, a resonance section, an intermediate section and an outlet section connected in sequence; the inlet section is provided with an incidence hole, the resonance section is provided with a resonance hole, the intermediate section is provided with an intermediate hole, and the outlet section is provided with an emission hole which is a diverging hole; the incidence hole, the resonance hole and the intermediate hole are all cylindrical holes, and the diameters of the incidence hole, the resonance hole and the intermediate hole gradually decrease. The elastic porous medium skin is used to induce the slip and micro-cavitation effect of unsteady boundary layer fluid, and the geometric parameters of the control hole are controlled, so that the self-adaptive boundary layer flow state is greatly reduced, the frictional resistance is reduced, and the outer lining has the characteristics of noise reduction, and has the scattering and absorption effect on sound waves.

A leading edge for an aircraft designed for active laminar flow control comprising a removable microperforated skin (10) for boundary layer suction, such microperforated skin comprising stiffeners (20) on its inner face that also act as spacers when the skin is installed on the supporting structure of the leading edge to form suction chambers, characterized in that the microperforated skin with its associated stiffeners is joined to the supporting structure (1) by means of non-permanent fasteners (30) attaching the stiffeners to the supporting structure and not visible on the outer surface of the skin, and the leading edge includes access to its interior through its lower surface that allows the installation of the non-permanent fasteners joining the skin to the leading edge structure. The invention also includes the process and the tool for an efficient and easy installation of the microperforated skin on the leading edge structure.

This application relates to a laser processing device which comprises the following: an optical arrangement; wherein the optical arrangement comprises an input for receiving a laser beam; wherein the optical arrangement comprises a beam splitter, which splits the laser beam into at least two partial beams; wherein the optical arrangement recombines the partial beams to form a laser spot in order to generate an interference pattern in the laser spot; wherein a first state of the laser beam at the input generates a first interference pattern and a second state of the laser beam generates a second interference pattern; wherein the first state and the second state differ in at least one of (i) a position of the laser beam at the input and (ii) an angle of incidence of the laser beam with respect to the input; and wherein the optical arrangement is configured such that the second interference pattern continues the first interference pattern in phase.