41results about How to "Increase stall angle of attack" patented technology

The invention provides a high-efficiency pneumatic layout structure of synergistic jet for a high-altitude propeller and a control method. The high-efficiency pneumatic layout structure is characterized in that a plurality of synergistic jet devices are distributed along propeller blades of the propeller in a spanwise sectional manner; each synergistic jet device comprises an air blowing opening formed in a negative-pressure area of the front edge of the upper surface of the propeller, an air suctioning opening formed in a high-pressure area of the rear edge of the upper surface of the propeller, an airflow pipeline arranged in the paddle and an air pump arranged in the airflow pipeline; the air blowing opening and the air suctioning opening are communicated by the airflow pipeline to form an air-blowing and air-suctioning loop; the air pump is used for driving air suctioning and air spraying to be implemented simultaneously, and by control of the air pump, the air suctioning amount and the air spraying amount are same. The high-efficiency pneumatic layout structure provided by the invention has the advantages that the defects of the traditional propeller layout and the common flow control technology can be made up, and the working efficiency of a propelling system of the high-altitude propeller can be improved.

The invention relates to a wind machine blade airfoil profile capable of controlling flow stalling through standing vortex. An airfoil profile between a shaping starting point and an airfoil profile trailing edge point is modified, thus a concave pit is formed; the shaping starting point and the airfoil profile trailing edge point are connected through a straight line, a concave pit starting point and a concave ending point are connected by virtue of a 1 / 4 circular arc, and the concave ending point, a first transition point, a second transition point and the airfoil profile trailing edge point are connected through a B spline; and the airfoil profile trailing edge is provided with a Gurney wing flap. According to the invention, the amplitude of the thickness reduction of the upper surface is slowed down to delay the separation of the flow field of the upper airfoil, vortex with a fixed position is formed by the air flow at the concave pit of the upper surface, under the action of the vortex, the separation of the upper surface of the airfoil can be controlled effectively, the irregular falling of separated vortex is avoided, and vortex lifting force is formed at the upper surface of the airfoil, and the stalling of the airfoil is slowed down, so that the stalling incidence of the airfoil is improved.

The invention which aims at close-coupled canard type configuration aircrafts provides an efficient indirect vortex control method and a device thereof. Places over a canard leading edge and at a certain height from the canard surface are respectively symmetrically provided with a jet tube, nozzle directions of the jet tubes parallel to the canard leading edge, high-speed pulse jets are jetted through the nozzles to generate a fluid vortex ring so as to increase the intensity of canard leading edge vortex and increase the angle of attack of canard configuration during canard leading edge vortex breakdown, and beneficial interference of the canard leading edge vortex and wing leading edge vortex is utilized to generate an indirect vortex control effect so as to enhance the intensity of the wing leading edge vortex and delay wing leading edge vortex breakdown when the angle of attach is large, so the configuration lift of the canard configuration aircraft is improved, the stall angle of attack is increased; and simultaneously the pulse frequency and the pulse width of pulse blowing are controlled through a solenoid valve, so the air entrainment amount is reduced to realize high efficiency and energy saving.

The invention discloses a canard aerodynamic configuration of a subsonic-velocity high-maneuver drone aircraft. A pair of canards is symmetrically arranged on the two sides of an aircraft body, the pair of canards utilizes symmetrical wing type with the maximum thickness of 10% to 14%, the sweepback angle of each leading edge is 30 to 32 DEG C, and the area of the canards is 8% to 12% of the area of a wing; a pair of wings is symmetrically arranged on the two sides of the back part of the aircraft body, ailerons are arranged on the outer sides of the wings and can be taken as flaps for lift enhancement according to the flight state, the wings utilize spanwise variable camber reverse-bending wing type with the maximum thickness of 8% to 12%, the sweepback angle of each leading edge of the wings is 38 to 42 DEG C, the wings are provided with twin vertical fins, the twin vertical fins utilize symmetric wing type with the maximum thickness of 6% to 10%, the area of the twin vertical fins is 15% to 25% of the area of the wings, and rudders are arranged on the twin vertical fins. According to the duck-type aerodynamic layout, the preposition double-canard configuration can remarkably increase the stalling angle of attack of an aircraft so as to provide conditions for high maneuver; a wing body fusion layout can also remarkably reduce the resistance of the aircraft and increase the speed; an air inlet of an inlet is formed in the ventral, and the structure is in an embedded type semi-crescent shape, so that the collision damage of the air inlet in a drone aircraft recycling process can be avoided.

The invention discloses a novel tailstock type axisymmetric multi-propeller vertical take-off and landing unmanned aerial vehicle. According to the scheme, the novel tailstock type axisymmetric multi-propeller vertical take-off and landing unmanned aerial vehicle comprises an unmanned aerial vehicle body assembly, a head cover assembly, a tail cover assembly, power propeller assemblies and a symmetric wing assembly. The novel tailstock type axisymmetric multi-propeller vertical take-off and landing unmanned aerial vehicle is characterized in that four wings are symmetrically arranged relativeto an unmanned aerial vehicle body in a crossed shape; the power propeller assemblies are correspondingly contained on the four wings; landing gears with shock absorbing devices are used at the lowerparts of the wings; in addition, the head cover assembly and the tail cover assembly are positioned at the two ends of the unmanned aerial vehicle body to reduce the air resistance in the cruise levelflight process. According to the novel tailstock type axisymmetric multi-propeller vertical take-off and landing unmanned aerial vehicle, through differential control of thrust forces of the four propellers in the vertical taking-off and landing or hovering stage, the attitude stabilization of the unmanned aerial vehicle is guaranteed; maneuvering actions, such as rolling and yawing, of the unmanned aerial vehicle are controlled by using trailing edge flaps and ailerons during horizontal flight; and the novel tailstock type axisymmetric multi-propeller vertical take-off and landing unmanned aerial vehicle is especially suitable for being applied to the fields such as carrier-based unmanned aerial vehicles and power line inspection or places with higher requirements on maneuverability, higher speeds and higher safeties.

The invention provides a method for designing a UAV (unmanned aerial vehicle) provided with a deflectable winglet. The problems that most UAVs nowadays have big height loss when performing dive and climb maneuvers, cannot guarantee cruising property and has poor stalling characteristics are solved by respective design of a main wing, inverted V empennages and the deflectable winglet of the UAV. The inverted V empennage layout and a rear propeller power mode are adopted. According to the layout, interference drag between the empennages as well as between the empennages and the fuselage can be reduced with fewer parts, and downwash of the empennages can be weakened to enhance course stability. Meanwhile, propeller slipstream can be avoided by the layout, so that pitching stability is improved.

The invention discloses a vortex generator, a fan blade and a manufacturing method of the fan blade. The vortex generator comprises at least one pair of grooves formed in the blade, and when airflow flows nearby the groove, high-energy vortex is generated under the function of the groove and flows downstream to disturb airflow on the upper surface of the blade. The vortex generator has the main functions of a current vortex generator, can generate enough vortex disturbance to restrain the flowing separation of the wing section surface, improves the lift-drag ratio of a wing section, improves the pneumatic performance of the blade, reduces the flowing noise of the blade, increases the annual energy output of a wind driven generator, and has the characteristics of being low in cost, high inreliability and the like.

The invention discloses a wind turbine blade capable of improving stalling characteristics based on fractal optimization. The wind turbine blade comprises a blade body positioned at the front part corresponding to a flow incoming direction, and a blade tail edge positioned at the back part corresponding to the flow incoming direction; the blade tail edge is distributed with multiple elastic pieces and fractal holes with fractal structures; the elastic pieces are arranged on suction surface sizes of the fractal holes, and are suffered from a pressure difference effect of a blade pressure surface and a suction furnace to locate in a closing or opening state; and the fractal holes penetrate through the pressure surface and the suction furnace. When the blade incoming flow attack angle dose not reach stalling conditions, as the pressure difference between the blade pressure surface and the suction surface is smaller, and at this time, the elastic pieces are located in the closing state to stop passing of a fluid to stop small-scale turbulence so as to guarantee the blade pneumatic performances; when the blade incoming flow attack angle is increased to reach the stalling conditions, the elastic pieces are located in the opening state due to the effect of pressure difference, and the fluid can pass through the fractal holes to generate small-scale turbulence on the suction furnace to delay separation of a blade surface boundary layer, so that the formation of stalling clusters is delayed, and the purpose of improving the blade stalling characteristics is achieved.

The invention discloses a novel configuration vertical take-off and landing unmanned aerial vehicle belonging to the field of aircraft transmission system and lifting rotor, which comprises a fuselage, a rolling wing system, a vertical tail, a rudder, an aileron, a flap, a detection pod and an take-off and landing actuating cylinder. A rolling wing system is installed in the duct of the wing, andthe rolling wing system is composed of two groups of rolling wings with the same size and opposite rotating directions. The propeller shaft of each roller section is coaxial, and the full vector propulsion system of cycloidal propeller is composed of two sections of each group. The invention discloses a novel configuration vertical take-off and landing unmanned aerial vehicle, which can realize vertical take-off and landing on land-based platform and sea-based platform, and has better flying quality. The invention combines the characteristics of a rotor aircraft and a fixed-wing aircraft, andhas low dependence on take-off and landing site and strong environmental adaptability. At the same time, the whole structure of cycloidal propeller system is light and simple, which has broad militaryand civilian prospects.

Provided are a deformable aircraft leading edge slat and a method for designing the same. An outer section of the leading edge slat is divided at the 20-40% position of the leading edge slat to form a movable section and a fixed section. The movable section is located at the position of the leading edge of the slat, the fixed section is connected with a main wing section through a leading edge slat sliding rail, and the movable section is connected with the fixed section through a hinge, so that the movable section can be deflected through delta degrees around the fixed section. The air flow separation of the leading edge slat can be effectively delayed, the stall attack angle of an aircraft is increased, the flight border is expanded, and the flight safety is improved. The deformable aircraft leading edge slat improves maximum lifting power of the aircraft at a take-off and landing stage, and accordingly the loading capacity of the aircraft is improved. The aerodynamic load limitation borne by the leading edge slat is reduced too, the design difficulty of the leading edge slat is reduced, the deflection angle of the movable section of the leading edge slat is flexibly adjusted according to the specific requirements of different flight phases, and the deformable aircraft leading edge slat has high task adaptive capability.

The invention provides an efficient control method of high-altitude airscrew collaborative jet. The efficient control method comprises the following steps: setting configuration parameters and working parameters of collaborative jet devices; and configuring each collaborative jet device for propeller blades in a spanwise sectional type according to the configuration parameters; enabling each collaborative jet device to work according to the configured working parameters, so that the airfoil profile of each station reaches the best lift-drag ratio and the minimum energy loss. The operation process of each collaborative jet device is that an air pump simultaneously drives a front-edge negative-pressure region to spray air and a rear-edge high-pressure region to suck air, and actively controls the air stream on the surface of the airfoil profile in a flow mode, wherein the counter acting force of jet flow generated by air spraying and suction is decomposed to two directions, one is the rotation direction of a propeller, and further the rotation of the propeller is boosted, and the resistance is reduced; and the other one is the thrust direction of the propeller, the thrust of the propeller is enhanced, and finally the pneumatic efficiency of the propeller is improved. According to the invention, the operating efficiency of the propelling system of the high-altitude airscrew can be improved.

The invention belongs to the field of design of horizontal axis wind power wings, and in particular, relates to a thick wing group of a large turbine blade. The thick wing group comprises eight wings with relative thicknesses of 0.35, 0.40, 0.45, 0.50, 0.55, 0.60, 0.65 and 0.70; the eight wings all have thicker blunt tail edges, and are suitable for the inner side and the root part of an above-MW-grade large turbine blade working under high reynolds number; the thick wing group is higher in rise coefficient, lift-to-drag ratio and stall attack angle compared with a traditional contrast wing, so that higher rise and torque are provided to the blade, the start air speed of a turbine is lower, the power coefficient is higher, and the output power is higher; the designed larger attack angle facilitates reduction of the twisting angle of the blade, and facilitates the structural design of the blade and the rigidity, strength and manufacturing of a blade girder, so that the turbine blade of the thick wing group is lighter; and meanwhile, the blunt tail edge design of the wing improves the blade manufacturability, and effectively reduces the production cost.

The invention provides a notch airfoil shape of a stamping parafoil and an optimum design method thereof. The optimum design method comprises the following steps of selecting a high aerodynamic efficiency base airfoil shape, designing a front edge notch and designing front edge notch parameter optimization. Compared with a conventional traditional CLARK-Y notch airfoil shape, the notch airfoil shape has the advantages that the max thickness of the notch airfoil shape is large, the thickness distribution is uniform, the space of a parafoil air chamber is large, and the notch airfoil shape is beneficial to improving of the charge efficiency and the charge rigidity of the parafoil; as the distribution of surface pressure along a thread direction is uniform, the notch airfoil shape is beneficial to arrangement of a parafoil parachute cord, and the pitching driving efficiency is improved; the stalling angle is large, the usable aerodynamic efficiency is high, and the notch airfoil shape has the large stalling angle range for smooth and stable control.

The invention discloses a vortex generator, a fan blade and a manufacturing method thereof. The vortex generator comprises a groove formed in a blade, wherein the groove gradually reduces; and when gas flow flows through near the groove, high-energy vortex is generated under action of the groove, and develops towards the downstream to disturb gas flow on the upper surface of the blade. According to the vortex generator disclosed by the invention, blade quality is reduced, the producing and mounting period of the vortex generator is reduced in process, and certain cost reducing meaning is achieved; besides, the vortex generator is integrally formed along with the blade main body, and attachments do not need to paste to the surface of the blade, so that compared with a conventional mountingtype vortex generator, the vortex generator disclosed by the invention has extremely high reliability, is convenient in later-stage operation and maintenance, can meet main functions of an existing vortex generator, can generate enough disturbance vortex to inhibit downstream flow separation, and delays speed reduction.

The invention discloses a wing applied to a pop-up type glider model. The wing is composed of a wing middle section and two wing tips, wherein the two wing tips are fixedly installed at the two ends of the wing middle section respectively, the dihedral angle of the wing tips is 15-25 degrees, the rear edge of the wing tip located at the left end of the wing middle section has misalignment by 0.8-1 mm in the vertical direction, and the rear edge of the wing tip at the right end of the wing middle section has misalignment by 0.8-1 mm in the vertical direction; a stepped table is formed on the upper surface of the wing in the length direction of the wing and located at the position of 45-55% of the length of a wing chord, and the step height of the stepped table accounts for 4-6% of the length of the wing chord. Due to the misaligned installation of the wing tips, the pop-up type glider model ascends perpendicularly with stable ascending trails, and the climbing height can exceed 30 meters. The stepped table makes the wing stepped, the life-drag ratio of the wing is increased, and the stalling attack angle of the wing is increased.

The invention discloses an awning wing for increasing the lift force and the stall angle of attack, which is a thin fin structure placed on top or obliquely above the upper surface of the wing. A ventilation duct is formed in front of the canopy and the upper surface of the wing, which delays the separation of the airflow, increases the stall angle of attack of the entire wing, and greatly increases the maximum lift coefficient of the whole aircraft. Comparing the wind tunnel test data of the aircraft model with awnings with the situation without awnings, it is found that the maximum lift coefficient has increased by more than 70%, and the stall angle of attack has increased by more than 8°-18°. When flying, the wing can be released or retracted as required. When retracted, the awning is stored on the upper surface of the basic airfoil to keep the profile of the original airfoil unchanged. The canopy wing is novel in principle and simple in structure, and can be widely used in civil and military aircraft or other machines moving in fluid.

The invention discloses a high-lift wind turbine airfoil. The airfoil is formed by the following steps of: taking an S809 airfoil as a basic profile, multiplying the chord length of an NACA4412 airfoil by a scale factor alpha, shrinking the NACA4412 airfoil, and rotating the shrunk airfoil at a fixed angle beta based on a trailing edge point; and at the trailing edge, when the vertical coordinate of the upper surface of the S809 airfoil is smaller than the vertical coordinate of the NACA4412 airfoil, replacing the upper surface of the part of the S809 airfoil with the surface of the NACA4412 airfoil. According to the high-lift wind turbine airfoil, the middle section position of the upper surface of the airfoil is in smooth transition, and a small adverse pressure gradient of the airfoil under a large angle of attack is guaranteed to inhibit airflow separation, the airfoil has a larger lift coefficient, a larger stalling angle and smaller resistance, and the wind energy absorption efficiency of a wind wheel of a wind generating set is improved.

The invention belongs to the technical field of aviation aircrafts, and discloses a high-aspect-ratio high-strength double-layer wing solar unmanned aerial vehicle which comprises a fuselage, wings and an empennage. The wings are transverse through wings, the number of the wings is two, the wings are an upper layer wing and a lower layer wing respectively, the fuselage is arranged between the upper layer wing and the lower layer wing, and the upper layer wing and the lower layer wing are connected through a wingtip supporting wall; solar panels are arranged on the upper surfaces of the upper-layer wing and the lower-layer wing; the empennage is arranged at tail of the fuselage. According to the double-layer united wing layout, on the premise that structural strength limitation is met, the aspect-ratio of the wings is increased to 30 or above, and the double-layer united wing layout has great significance in designing aerodynamic drag reduction of airplanes with large lift coefficients; the wingspan is reduced by 29%, the structural weight is reduced by 22%, and the mission load and the continuous flight time of the aircraft can be remarkably improved; the wings are high in bending resistance and torsion resistance; and satisfactory three-axis stability and control capability can be provided.

The invention discloses a horizontal shaft tidal current energy water turbine and a using method. The horizontal shaft tidal current energy water turbine comprises blades and a supporting structure, each blade comprises a main wing and a flap, each flap is arranged at the tail edge of the corresponding main wing, and a water-permeable channel with a variable area exists between each main wing and the corresponding flap. Due to the adoption of the structural design that the water-permeable channels exist between the main wings and the flaps, the flow speed at the tail edges of the suction surfaces of the main wings can be increased through the ejection effect of the water-permeable channels between the main wings and the flaps on water flow, and the aim of improving boundary layer separation is achieved. Meanwhile, the flow separation phenomenon and energy loss on the blade surfaces are reduced, and therefore the lift force of the blades and the energy conversion efficiency of the water turbine are improved. Meanwhile, through the structural design that the area of the water-permeable channels is variable, when the speed of water flow around the blades is increased, the area of the water-permeable channels is adjusted, then the overall camber of the blades is increased, the stall attack angle is increased, the occurrence of blade stall is delayed, and therefore the operation range of the tidal current energy water turbine is widened.

A split type linked wing layout aircraft comprises a main wing, the main wing is a double-sweepback flying wing, a sailboard type wingtip winglet is arranged at each of wing tips of the two sides of the main wing, two empennage connecting rods are symmetrically arranged on the upper surface, close to the rear edge, of the middle section of the main wing, the empennage connecting rods are hollow variable-section straight rods, and a power device is arranged at the tail of the middle section of the main wing. The trailing edge of the wingtip of the main wing is provided with a flaperon rudder, the linked wing is a forward swept wing and adopts a symmetrical wing type, one end of the linked wing is fixedly connected with the top end of the wingtip winglet, the other end of the linked wing is fixedly connected with the top end of the V-shaped empennage, the trailing edge of the linked wing is provided with an elevator, the bottom end of the V-shaped empennage is fixed on an empennage connecting rod, and the trailing edge of the V-shaped empennage is provided with a rudder. The closed structure formed by mutually connecting all the parts improves the overall structural strength of the aircraft, the pneumatic flutter problem of the aircraft during high-speed flight is solved, direct force control of the aircraft with the split-type linked-wing layout is achieved through the split-type linked-wing layout form and the single-engine power arrangement higher than the main wings, and the flight quality of the aircraft is improved.

The invention discloses an ejection type glider model comprising a glider body, a wing and an empennage mounted on the glider body. The wing is formed by a wing middle section and two wing tips; the two wing tips are respectively fixed on two ends of the wing middle section; an anhedral of each wing tip is 15 to 25 degrees; the back edge of the wing tip on the left end of the wing middle section is 0.8 to 1mm upwardly staggered in a vertical direction; the back edge of the wing tip on the right end of the wing middle section is 0.8 to 1mm downwardly staggered in the vertical direction; the upper surface of the wing forms a step bench along a length direction of the wing; the empennage is formed by a first empennage and a vertical second empennage; the first empennage is installed on the glider body in a way of forming a rotation angle of 2-3 DEG by taking the glider body as a center shaft and the horizontal plane as a starting point; the lower end of the second empennage is fixed on the upper surface of the first empennage; and the side surface of the second empennage is fixed on the glider body. The wing tip staggered installation provides a stable rising path when the ejection glider model vertically rises; and the climbing height can surpass 30 meters.

The invention discloses a step wing used for an ejection type glider model. The step wing comprises a wing body, a stair is formed on the upper surface of the wing body in the length direction and located at the position of 45-55% wing chord length, and the step height of the stair is 4-6% of the wing chord length. The step wing has the advantages that the stair makes the wing into the step wing, the step wing is easy to manufacture, time and labor are saved, airfoil processing and control are precise, and the weight of the wing is reduced. If the step wing is installed on the ejection type glider model, the lift-drag ratio of the wing can be improved, and the stalling attack angle of the wing is increased.

The invention discloses an ejection type glider model, which comprises a glider body, and a wing and an empennage which are installed on the glider body. The wing comprises a middle section and two wing ends that are installed on two ends of the middle section, respectively. A reverse angle on the wing ends is 15-25 degrees, wherein a stepped edge is formed on the upper surface of the wing in the length direction of the wing and located at a point of 45-55% of the length of a wing chord, and the height of the stepped edge is 4-6% of the length of the wing chord. The empennage comprises a first empennage and a second vertical empennage. The first empennage is installed on the glider body at an angel of 2-3 degrees that is formed by rotating the first empennage with the glider body as a central axis and with a horizontal level as a starting point. The lower end of the second empennage is fixed on the upper surface of the first empennage. Both sides of the second empennage are fixed on the glider body. The stepped edge makes the wing become a stepped wing, increasing lift-drag ratios and stalling angles of the wing.