219results about How to "Improve aerodynamics" patented technology

The invention discloses a tilt ducted unmanned aerial vehicle. A central duct is arranged in an inner cavity of a center of a body, and is connected with the body through a tilting axis; the part of the body connected with the tilting axis is provided with a duct tiling mechanism; wings are symmetrically arranged on two sides of the body; a nose and wingtips of the wings are provided with posture control ducts; the lateral rear of the body is provided with extraverted double vertical tails; the rear of the extraverted double vertical tails is provided with a rudder; the lower side of the body is provided with a front three-point wheel landing gear; and the central duct, the duct tilting mechanism, the posture control duct and the rubber are respectively connected with a flight control system. The tilt ducted unmanned aerial vehicle has the advantages that: 1, a tail-free and large-triangle wing fusion layout is adopted, and the pneumatic property is good; 2, a unique tilt central duct design is adopted; 3, a unique three-duct posture control design is adopted; and 4, a duct rotor wing design is adopted, namely all paddles are embedded in the ducts, and the flight safety in an complex environment is effectively improved.

An aircraft, preferably a large capacity passenger aircraft with two passenger decks, has a double-T tail assembly arranged on its tail section. The double-T tail assembly includes a double fin and rudder assembly and a tailplane and elevator assembly, and provides a significant increase in the usable passenger cabin space in the tail section of the aircraft. The double fin and rudder assembly has a U-shaped structure and comprises two upright stabilizer fins connected to each other by a horizontal torsion box that fixedly extends crosswise through the fuselage tail section within the floor of a cabin deck. The tailplane and elevator assembly is pivotably mounted on the upper tips of the two upright stabilizer fins in a double-T arrangement.

The invention discloses a skin for morphing wings, belonging to the technical field of design of morphing aircrafts. The skin for the morphing wings is made from fibre reinforced composites, and the cross section of the skin is in an isosceles trapezoid ripple structure. The invention also carries out analysis aiming at the chordwise and spanwise equivalent elastic modulus and flexural modulus ofthe skin structure with isosceles trapezoid ripples, provides a corresponding mathematical model and can further optimize the skin structure according to the mathematical model. The invention meets the requirements of the skin of the morphing wings of the morphing aircrafts for large deformation aerodynamic load bearing, can realize the design optimization of the skin structure under different load operating conditions according to the mathematical analysis model of the equivalent elastic modulus and the flexural modulus in two different directions of the skin structure, also has structure health monitoring and self-adapting deformation capability and can realize optimal aerodynamic characteristics by controlling the shape and the vibration of the skin.

A blade for use in a wind turbine comprises a pressure side and suction side meeting at a trailing edge and leading edge, the pressure side having a pressure side surface and the suction side having a suction side surface, the pressure side surface and suction side surface for providing lift to the turbine blade upon the flow of air from the leading edge to the trailing edge and over the pressure side and suction side surfaces, the pressure side and suction side extending from a root portion to a tip portion of the turbine blade. In some situations, the root portion is non-aerodynamic. The aerodynamics of such a blade is improved with the aid of pneumatic blowing through one or more blown passages for providing pressurized air (or other fluid) to a suction side and/or pressure side of the blade.

Described herein are embodiments of fastening systems that can provide easily removed and reapplied fasteners and that provide a hindrance to those that might tamper with the fasteners. In some embodiments, fasteners are also described that are configured to reduce the accumulation of dust and dirt inside the fasteners. In further embodiments, fasteners are provided for improving the aerodynamics of the fastener, and some embodiments describe provision of commercial advertising or other messages on the fastener.

The invention relates to a canard forward-sweep telescoping wing aerodynamic configuration with variable span wing area. The canard forward-sweep telescoping wing aerodynamic configuration comprises a fuselage, canards, wings, and a vertical fin; the wings comprise forward-sweep inner wings and forward-sweep telescopic outer wings having internal connections with the forward-sweep inner wings by a telescoping mechanism; and the telescoping mechanism extends to allow the forward-sweep telescopic outer wings to extend to the outside of the forward-sweep inner wings when the mach number (Ma) of an unmanned aerial vehicle is 0.2, and the telescoping mechanism contracts to allow the forward-sweep telescopic outer wings to contract to the inside of the forward-sweep inner wings when the Ma of the unmanned aerial vehicle is 0.4, wherein the area ratio of the forward-sweep telescopic outer wings to the forward-sweep inner wings is 0.25-0.45. The telescoping wing configuration with variable span wing area adopted in the invention allows the unmanned aerial vehicle to have good aerodynamic performances in scopes of different spatial domains and different speeds and the mobility and the flexibility of the unmanned aerial vehicle to be improved. According to the canard forward-sweep telescoping wing aerodynamic configuration of the invention, when Ma is equal to 0.2 or 0.4, the cruise lift-to-drag ratio of the high aspect ratio aerodynamic configuration is over 20% higher than the cruise lift-to-drag ratio of the low aspect ratio aerodynamic configuration when the unmanned aerial vehicle flies at a low speed, and the cruise lift-to-drag ratio of the low aspect ratio aerodynamic configuration is about 15% higher than the cruise lift-to-drag ratio of the high aspect ratio aerodynamic configuration when the unmanned aerial vehicle flies at a high speed, so the capacities of the cruise flight with the Ma of 0.2 and 0.4 are possessed.

The invention discloses an aerodynamic optimization method of a wind turbine airfoil profile considering a high turbulence free flow effect. Based on the high turbulence characteristics of incoming flow from a blade airfoil in an actual wind field, the effect of inflow turbulence intensity on the airfoil profile aerodynamic efficiency and load are parameterized and the sensitivity parameters of the maximum lift-drag ratio and the maximum lift coefficient are established, and then the aerodynamic optimization model of the airfoil is formed by combining the objective parameters of airfoil efficiency and load, surface roughness sensitivity and Reynolds number effect with the weight coefficient method, and the aerodynamic optimization method of the airfoil is constructed by combining the genetic algorithm. As shown by the numerical simulation and evaluation, the aerofoil profile aerodynamic optimization method provided by the invention not only maintains a high design lift coefficient andimprove the maximum lift-drag ratio, but also remarkably reduces the sensitive parameters of the aerofoil aerodynamic performance varying with the turbulence intensity of the free flow; in addition, the maximum lift coefficient of the airfoil is effectively limited, and the stall process is more smooth.

The invention relates to a centrifugal fan blade, belonging to the technical field of fluid machinery. The blade comprises a front edge, a pressure surface, a suction surface and a rear edge, wherein the front edge is tangent to the inner diameter of a fan; the rear edge is positioned on the outer diameter of the fan; one end of the pressure surface and one end of the suction surface are respectively connected with the two ends of the front edge, and the other ends of the pressure surface and the suction surface are respectively connected with the two ends of the rear edge; and the blade is formed by the area surrounded by the front edge, the pressure surface, the suction surface and the rear edge. The blade adopts variable thickness forward bending structure, and the pressure surface and the suction surface of blade are controlled by high-order bezier curve, so that higher pressure rise and working efficiency can be ensured under the condition of low flow rate, the situation of a flow channel is improved, and the energy loss is reduced.

The invention provides a phased-array antenna subarray for circularly polarized wide-angle scanning, and aims at providing a phased-array antenna subarray which has a good wide-angle scanning gain characteristic and a good wide-angle scanning circular polarization characteristic, and can avoid derotation properly. The phased-array antenna subarray is achieved by the technical scheme as follows: three circularly polarized microstrip antenna radiation patches are distributed on a microstrip antenna dielectric plate (10) in a ternary rotary inversion subarray form; a metal layer (6) at the lower part of the microstrip antenna dielectric plate serves as a metal place for three microstrip antenna units; each coaxial joint inner core (3) and each coaxial joint outer core (4) form a coaxial feeding port; each coaxial outer core is connected with the metal place; each coaxial joint inner core is connected with the corresponding radiation patch located in the microstrip antenna dielectric plate in blind-inserting and non-contact manners; and the coaxial feeding ports perform electromagnetic coupling feeding and implement phase compensation on the microstrip antenna units to obtain a subarray radiating left-hand circular polarization electromagnetic waves or a subarray radiating right-hand circular polarization electromagnetic waves.

The invention discloses a flapping wing capable of automatically folding and unfolding for a flapping wing type micro aerial vehicle. The flapping wing consists of a power device, two hydraulic pipes, a plurality of creases, a plurality of supporting rods, a wing membrane and two folding devices, and the folding and unfolding of the flapping wing are realized by adopting a hydraulic principle. According to the flapping wing, the two hydraulic pipes and the plurality of supporting rods form wing veins of a bionic flapping wing, and the wing membrane is adhered to the bionic wing veins to form a wing surface. The flapping wing pumps liquid into different liquid chambers by virtue of a micro flow pump to generate a pressure so as to drive the flapping wing to fold up or unfold. The flapping wing disclosed by the invention has the beneficial effects of reducing the transverse size of flapping wing type micro aerial vehicle when the flapping wing type micro aerial vehicle does not fly, reducing the spaces required for carrying, transportation and storage, reducing transportation requirements, conveniently protecting the flapping wing of the aerial vehicle from being damaged, favorably replacing a damaged flapping wing to improve the service life of the micro aerial vehicle.

The invention discloses an aerodynamic layout design method based on a variable-configuration aerospace vehicle. The aerodynamic layout design method comprises the steps that the low-speed datum layout and the high-speed datum layout of the variable-configuration aerospace vehicle are determined correspondingly; the geometrical shape of the variable-configuration aerospace vehicle is described ina parameterization mode through a state-type function method; a corresponding parameterization kinetic model is established through an engineering estimation method, models of different speed domainsare integrated based on the similarity criterion, and a corresponding agent model is established; the influence of the flight condition and aerodynamic layout change on the performance of a flight control system is determined through a sensitivity analysis method, and the stability boundary and system requirements of the variable-configuration aerospace vehicle are determined; and through full-speed-domain performance matching analysis, a variable-layout scheme of the variable-configuration aerospace vehicle is determined. By properly changing the wing configuration of the aerospace vehicle, it is guaranteed that the aerospace vehicle has the optimal aerodynamic layout at each mission stage, large envelope continuous flight under the minimum energy consumption is achieved, and the need ofintegrated design of the aerospace vehicle is met.

The invention belongs to the field of overall design of aviation aircrafts and particularly relates to a solar aircraft. The solar aircraft comprises a wing which is a rectangular wing. An energy storage battery is fixedly distributed in the wing. Solar cells are laid on the upper surface of the wing. Two aircraft bodies are fixed to the positions 1/4 and 3/4 of the wing in the tangential direction. Flat tails are parallel to the wing, perpendicular tails are perpendicular to the flat tails, and the flat tails and the perpendicular tails are fixed to the rear ends of the aircraft bodies. Two power systems are fixed to the front ends of the two aircraft bodies respectively. By the adoption of the solar aircraft, the arrangement of the two aircraft bodies is adopted and the energy storage battery is distributed in the wing, so that loads are evenly distributed, lifting weight balance and thrust and resistance balance are considered, energy balance needs to be further considered, it is guaranteed that the energy consumed for maintaining lifting force of the aircraft is equivalent to the energy obtained by the solar cells, and long-time high-altitude surveillance and monitoring can be achieved; and meanwhile, the solar aircraft of the structure has the good aerodynamic performance and is convenient to install.

A wheel for a vehicle includes an inner hub portion that is configured to receive an axle therethrough and which defines an axis of rotation at a center thereof. The wheel includes an outer rim portion. A plurality of wheel spoke portions extend generally between the inner hub portion and the outer rim portion. The inner hub portion, the outer rim portion, and the plurality of spoke portions cooperate to form a wheel outer surface having a plurality of turbine openings. The outer surface of the wheel defines a wheel reference plane that is oriented generally perpendicular to the axis of rotation. The wheel outer surface is the side that face away from a vehicle. Each of the plurality of wheel spoke portions has an outer side adjacent the wheel outer surface, an inner side opposite the outer side, a leading side surface and a trailing side surface. The inner side of each of the plurality of spoke portions has a first edge adjacent the leading side surface that is spaced a first distance apart from the wheel reference plane and a second edge adjacent the trailing side surface that is spaced a second distance apart from the wheel reference plane. The inner side of the spoke is configured such that the first distance is less than the second distance.

A headgear apparatus for delivering temperature conditioned air including passively or actively cooled, heated, or ventilated air to the interior of a shell is disclosed. A device housing which enshrouds a device for providing temperature conditioned air is placed on a rear portion of the shell. The device housing has air inlets and air rejector vents for delivering fresh air to and removing heated air from the device, and is configured to reduce the heated air from re-entering the device through the air inlets. The headgear apparatus is designed to withstand impact forces without injuring the wearer.

The invention relates to the field of aero-engine nozzles, in particular to a two-dimensional vector convergent-divergent nozzle capable of achieving pitching and yawing. The two-dimensional vector convergent-divergent nozzle comprises a barrel, two pitching adjusting sheets, two yawing adjusting sheets, four sealing sheets, first actuating cylinders, first pulling rods, second actuating cylinders and second pulling rods. The outer side face of each pitching adjusting sheet is connected with the corresponding first pulling rod and the corresponding first actuating cylinder so that the pitching adjusting sheet can rotate in the pitching direction. The outer side face of each yawing adjusting sheet is connected with the corresponding second pulling rod and the corresponding second actuating cylinder so that the yawing adjusting sheet can rotate in the yawing direction. Sealing sheets are located between the pitching adjusting sheets and the yawing adjusting sheets. The two-dimensional vector convergent-divergent nozzle capable of achieving pitching and yawing is simple in structure; deflecting, converging and diverging in the pitching and yawing directions are achieved through deformation combination of the adjusting sheets on the four sides; good aerodynamic performance is also achieved, the control law is simple, and the reliability is high.

The invention discloses an outer rectifying cover of a scramjet engine and a method for designing the same. The main design thought comprises the following steps: firstly carrying out non-adhesion numerical analysis on a flow field generated on a front body of a hypersonic speed aerocraft; then designing the lower wall face shape of the outer rectifying cover in the flow field in a flown line tracing mode, and adopting a plane vertical to the lower wall on the side wall of the outer rectifying cover; and deflecting the outer rectifying cover by certain angle to obtain the outer rectifying cover of the scramjet engine according to the information such as flight attack angle, flight mach number and the like of the aerocraft by taking the lip opening position of the outer rectifying cover asa rotating point. The outer rectifying cover of the scramjet engine adopts the design principle of a waverider body, makes full use of a high pressure area after shocking waves to generate additionallift force, and reduces incident flow area and pressure difference resistance by deflecting the rectifying cover to make the whole aerodynamic performance of the aerocraft obviously improved. The outer rectifying cover can be directly applied to designing the hypersonic speed aerocraft taking various scramjet engines as power.

The invention provides automatic garment label sewing equipment. The automatic garment label sewing equipment has the advantages that a label sewing bar tacking machine, pneumatic and electric control systems and related mechanical mechanisms are combined with one another, accordingly, labels can be automatically conveyed and folded and can be automatically accurately placed, the production efficiency of label sewing work can be improved, the labor cost can be reduced, and the integral automatic level of the garment label sewing work can be upgraded.

The invention discloses a design method for reusable sky-earth aircraft configuration. According to the method, an up-down contour control line and a left-right width contour line of a fuselage are determined, a quadratic curve and a CST method based on a type function and a shape function are adopted to generate a corresponding section shape, and therefore a key stance shape of the whole fuselage is obtained; the shape of the whole fuselage is obtained through a surface lofting technology; a numerical computation method is adopted to compute a flow field; matching design is performed on a control plane; and aerodynamic configuration is optimized. Compared with the prior art, the method has the advantages that high supersonic velocity lift-to-drag ratio characteristics and sub-cross-ultralow velocity entrance characteristics are comprehensively considered, the reusable sky-earth aircraft configuration with all velocity domains meeting design demands is designed, and an aircraft obtained through optimization has excellent aerodynamic performance and good heat resistance as well as good practicability.

A Mach number adjusting mechanism suitable for a trans-supersonic wind tunnel belongs to the technical field of aviation aerodynamic force wind tunnel test, and aims at solving the problem that a present Mach number adjusting mechanism is high in the adjusting obstruction degree of an air flow cross section. The adjusting mechanism comprises a front tail plate, a rhombic adjusting sheet and a rear tail plate, the front tail plate and the rear tail plate are mounted in a cavity formed by an upper double-throat wallboard and a lower double-throat wallboard, the rhombic adjusting sheet is hinged between the front tail plate and the rear tail plate, an auxiliary adjusting mechanism is arranged in a rhombic frame of the rhombic adjusting sheet, a main electric cylinder drives the rear tail plate to slide in a slide rail soleplate, and the rear tail plate is pulled to adjust opening and closing of the rhombic adjusting sheet. The Mach number adjusting mechanism uses the rhombic adjusting sheet, the rhombic adjusting sheet is continuous with and has no order difference from an air flow profile formed by the front and rear tail plates, and can be folded completely, the mechanism is low in obstruction degree, and a pneumatic performance is high.

The invention provides a sea and air amphibious propeller. The propeller comprises a ducted propeller stator, air propellers, air propeller rotors, ducted propellers and ducted propeller rotors, wherein a group of magnets is fixed on the air propeller rotors and the ducted propeller rotors respectively, the air propeller rotors and the ducted propeller rotors rotate coaxially and are nested to thesame ducted propeller stator, and two coils corresponding to the air propeller rotors and the ducted rotors respectively are embedded into the ducted propeller stator. The air propellers are used foraerial propulsion, water propellers are used for underwater propulsion, and the both can work independently. During underwater propulsion, the air propellers are folded back and rotated at a lower speed to eliminate the resistance to the air propellers, and the shaftless ducted propellers work to generate propulsion. In the air, the air propellers are unfolded and rotate to generate propulsion, and the ducted propellers assist the operation. The shaftless ducted propellers have good propulsion performance under water, the air propellers are suitable for working in the air, and the air propellers are combined with the shaftless ducted propellers to realize efficient propulsion and free switching under water and in the air.

The invention discloses a wave rider concept gliding aircraft profile design method satisfying filling demands; the method comprises the following steps: determining fuselage up and down contour control lines according to dimension constraints of the fuselage length and different cross sections, wherein the contour lines are spliced by linear segments; calculating an airflow deflection angle, calculating a two dimension tangent wedge flow field pressure, and calculating a FCT arc curve radius; determining an arc curve according to the FCT arc curve radius, and using the two dimension tangent wedge flow field to form a lower surface wave rider surface; using a secondary curve to form a leeside cross section shape; carrying out shrink ratio design for the lower surface wave rider surface according to the fuselage width and height dimension constraints and the upper surface shape, thus obtaining the whole fuselage shape; designing a passivation rounding with the radius of 15mm on the head and the wave rider tip leaning edge; match-designing a control surface. The aircraft has excellent pneumatic performance, and is high in practicality.

A system for deploying a deployable structure in an aircraft, comprising a drogue parachute, at least one deployable structure, a riser line attached to the drogue parachute and the at least one deployable structure, wherein when the drogue parachute is deployed, a tension is applied to the riser line, and a parachute deployment system, wherein the parachute deployment system is configured to utilize the applied tension to deploy the at least one deployable structure.