167results 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 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.

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).

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.

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 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.

Multilayer rib ornaments and methods of producing the same are described herein. One disclosed example method includes applying a first high elongation polymeric material to a web tool, where the web tool is provided from a first roller; and heating the first high elongation polymer material via a first heating process. The disclosed example method further includes applying a second high elongation polymeric material to the first high elongation polymeric material; and heating the second high elongation polymer material via a second heating process. The disclosed example method also includes applying a support layer to the second high elongation polymeric material via a lamination roller.

The present invention discloses a rotor based on flexible blades. A plurality of blades are evenly spaced on a rotor shaft. The blades are made of flexible materials, and high-strength tensile ropes are arranged therein. The blades reduce the interference of lift changes in the advancing area and the retreating area of the blades during forward flight and gust interference through deformation; a winglet mechanism is inserted and fixed at the wingtip of the blade, and a winglet is installed on the winglet mechanism. The winglet is arranged parallel to the blade; the winglet has an angle of attack relative to the blade chord line, which is used to provide a torsional moment for the blade to have an angle of attack. By adopting the above rotor, the unmanned aerial vehicle is less affected by the wind, flies more stably, has clear imaging during shooting, and is also more convenient for shipboard operation and storage.

A laser processing device comprising: an optical arrangement; wherein the optical arrangement comprises an input for receiving a laser beam; wherein the optical arrangement comprises a beam splitter that splits the laser beam into at least two partial beams; wherein the optical arrangement recombines the partial beams into a laser spot for generating 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.

An aircraft propulsion system with 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 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.

In order to improve de-icing systems for an aircraft (10), a de-icing system (26) is proposed. The de-icing system (26) comprises a thin film heating foil (28) that is adhesively bonded to the leading edge portion (24). The heating foil (28) comprises a plurality of heating elements (36). In addition, an ice detecting device (30) is provided, which is able to detect the presence of ice on the leading edge portion (24) by optical, acoustical or thermal means. The de-icing system (26) has a control device (34), which is configured for periodically and selectively energizing the heating elements (36). To further improve the performance of the de-icing system (26), a silicone nanofilament network (SNF) is added as a superhydropho-bic coating (98) on a microperforated skin (100) and within the microperforations (94).

An aircraft includes a fuselage and at least one primary wing having an upper surface, at least one recess in the upper surface and at least one conduit in fluid communication with the at least one recess. At least one ejector is disposed within the at least one recess and is configured to receive compressed air via the at least one conduit.

The invention relates to a ridged foil (10) for reducing the air resistance of an aircraft, comprising a suspension (20) with magnetic particles contained therein, wherein depending on the course of a magnetic field acting on the ridged foil (10), the acting magnetic field can be made visible at least locally by changing the orientation of the magnetic particles. The ridged foil (10) according to the invention allows the inspection of the aircraft structure located below the ridged foil (10) through the ridged foil.

An aircraft structure (10) having a fuselage (20), an engine (70) and an air inlet (100) for receiving air for the engine is specified. The air inlet has a cowl which surrounds an inlet opening. The air inlet furthermore has a compression surface (102) and a transition shoulder (104). The air inlet (100) additionally has a flow guide surface (110), which is arranged as a local elevation in a region of the transition shoulder (104) and extends into the inlet opening. This flow guide surface designed as a local elevation contributes to improving the displacement of the boundary layer of a fluid flowing along the aircraft structure at the air inlet.

An assembly is provided that includes a regulator manifold and a regulator element. The regulator manifold includes an inlet passage, a first outlet passage and a second outlet passage. The inlet passage is configured to receive bleed gas from a core flowpath. The regulator element is within the regulator manifold and configured to pivot to a first position, a second position and an intermediate position between the first position and the second position. The regulator element is configured to fluidly couple the inlet passage with the first outlet passage in the first position. The regulator element is configured to fluidly couple the inlet passage with the second outlet passage in the second position. The regulator element is configured to fluidly decouple the inlet passage from the first outlet passage and the second outlet passage in the intermediate position.

A hybrid electric gas turbine propulsion system (100) may comprise: a first propulsion system (401), a second propulsion system (402), and a third propulsion system (403). The first propulsion system (401) may comprise a first fan (212), a first turbine (252), a first compressor (232), and a first electric motor (190), the first fan (212) operably coupled to the first turbine (252) and the first compressor (232) by a first shaft (126), the first shaft (126) coupled to the first electric motor (190), the first shaft (126) configured to be disposed radially inward of a fuselage (12) of an aircraft (10). The second propulsion system (402) and the third propulsion system (403) may be in accordance with the first propulsion system (401). The hybrid electric gas turbine propulsion system (100) may be symmetric about a vertical plane extending through a neutral aerodynamic axis (104).

An air injection assembly for an aircraft is provided. The aircraft includes a fuselage extending between a forward end and an aft end along a longitudinal direction and a boundary layer ingestion fan mounted to the fuselage at the aft end of the fuselage. The air injection assembly includes at least one injection port defined on a surface of the fuselage at a location upstream of the boundary layer ingestion fan. A supplemental airflow is provided through a fluid passageway to the injection port where it is ejected to displace at least a portion of relatively higher velocity boundary layer airflow. In this manner, the airflow entering boundary layer ingestion fan is more uniform, has less swirl distortion, and has a lower average velocity.