375results about How to "Improve the lift-to-drag ratio" patented technology

A fluid dynamic section provides one or more fixed size escapelets through a foil body to reduce the induced and interference drag caused by trailing vortices and similar wake turbulence. The escapelets, which can be provided in both aerodynamic and hydrodynamic structures, such as wings, tail sections; rotary blades, guy wire frames, wing sails, and various underwater keels and wing keels. The escapelets transfer energy from an inlet located in the high-pressure surface of the foil or foil body to an outlet located in the lower-pressure surface, allowing energy that would normally form a vortex at the tip of the foil to be redirected and dissipated in a beneficial way. As a result, drag is reduced and fuel economy is increased, while at the same time increasing the authority of ailerons and similar flight control surfaces, allowing aircraft that were not previously spin recovery rated to become spin recoverable.

Unlike conventional winglets that attempt to block out wingtip vortices, these winglets use multiple airfoils to recycle much of the energy of these vortices back into useful lift and thrust (drag reduction). This will improve the lift to drag ratio at high angles of attack. These airfoils extend outward in the plane of the wing from each wingtip and are specially shaped to make them easy to extend and retract. Not only can they function over a continuum of airspeeds and angles of attack, but they can also be fully retracted at very high speeds to avoid parasite drag penalties.

The invention relates to an airfoil flow control device. The airfoil flow control device is arranged on the upper surface of the main airfoil of an airplane near the front edge of the airfoil. The device consists of a fundamental frequency Gershgorim band and a subharmonic frequency Gershgorim band which are arranged in parallel; both the fundamental frequency Gershgorim band and the subharmonic frequency Gershgorim band are in a sawtooth shape; during the taking off stage, a front flap extends the airfoil flow control device, thereby controlling the boundary layer of the airfoil to fast transit and form an onflow boundary layer used for restraining or eliminating the separation of the boundary layers, thereby effectively improving the life force of the airfoil; during cruising flying, the front flap of the airfoil resets to cover the device, thus leading the device to be out of use and leading the boundary layer of the airfoil to maintain a layer flow during the cruising stage, so as to reduce the friction resistance at the airfoil surface and improve the lift-drag ratio. The airfoil flow control device has the advantages of simple structure, light weight and convenient for being maintained, can solve the separation problem of the boundary layers of the airfoil surfaces during the taking off state of the airplane under the situations of not increasing the weight of the airplane and the power; besides, the airfoil flow control device also avoids the increasing problem of the friction resistance brought by the transition of the boundary layers of the airfoil surfaces during the cruising flying stage.

Embodiments of the present invention provide a gliding submersible transport system. Exemplary submersible gliders have wings capable of providing sufficiently high lift-to-drag ratios such that the submersible gliders of may be used for transporting large volumes of military or commercial hardware, equipment, personnel, or the like. According to one exemplary embodiment of the present invention, a submersible glider has a step-wise glider range. The glider includes a substantially cylindrical hull having a bow and a stern. A generally planar lifting surface is disposed toward the stern. The lifting surface has a pair of generally planar stabilizer surfaces that extend generally perpendicular to a plane of the lifting surface from ends of the lifting surface. A nose cone and at least one steering device are disposed toward the bow.

The invention discloses a foldable wing extensible in wingspan .The foldable wing comprises a front wing, a rear wing, a base and an actuating system, wherein the front wing comprises a left front inner wing and a right front inner wing which are arranged symmetrically, the rear ring comprises a left rear wing and a right rear wing which are arranged symmetrically, the end of the left front inner wing and the end of the right front inner wing are hinged and installed to one end of the base though front wing installing shafts, the actuating system is arranged at the other end of the base, a left front outer wing and a right front outer wing which slide in the length directions of the inner wings are respectively arranged on the left front inner wing and the right front inner wing, one end of the left rear wing and one end of the right rear wing are hinged and installed to a sliding block, the other end of the left rear wing and the other end of the right rear wing are in hinged connection with the left front outer wing and the right front outer wing, and the sliding block is driven by the actuating system to longitudinally move along the base back and forward .The foldable wing is large in wingspan variation amplitude and small in occupied space, leans against the base when being folded, the left front outer wing and the right front outer wing are rotatably unfolded around the front wing installing shaft, meanwhile the left front outer wing and the right front outer wing increase the wingspan width, a lift force and the lift-drag ratio are increased, and the flight efficiency is improved .

The invention discloses a waverider designing method based on a reference flow field of a revolution body of a cuspidal Von Karman curve. The method includes: patching a generatrix of a revolution body of a Von Karman curve to a generatrix of the revolution body of the cuspidal Von Karman curve; designing the revolution body of the cuspidal Von Karman curve; solving a supersonic-velocity axisymmetric reference flow field of the zero-incidence cuspidal Von Karman curve revolution body; performing streamline tracing in the supersonic-velocity axisymmetric reference flow field; and generating a waverider aerodynamic configuration. By adoption of the technical scheme of the method, the volume and the lift-drag ratio of a waverider are increased.

The invention relates to a revolving-ring underwater glider which comprises a machine body and an internal platform mounting structure, wherein a comprehensive control mechanism, a control communication and power system, a revolving wing system and a measurement sensor are mounted in the machine body. The revolving-ring underwater glider is compact and reasonable in structure and convenient to operate; by integration of net buoyance and a gravity center adjustment mechanism into a comprehensive attitude adjustment mechanism, the internal load-bearing space and the system weight are effectively reduced, and the effective load can be increased by about 80 percent, which accounts for 10 to 20 percent of the total load; by the design of revolving wings, the lift-drag ratio of the glider can be effectively increased, and the gliding efficiency can be improved; meanwhile, the glider has steering capacity and the capacity of moving on a horizontal plane; and the maneuverability and the flexibility of the glider are effectively improved.

The invention relates to a layout scheme of a tilt rotor transport aircraft. According to the layout scheme, the tilt rotor transport aircraft comprises a body, front wings, rear wings, front-rear-wing connecting end plates, a vertical tail wing, rotor wings, front wing auxiliary wings, rear wing auxiliary wings, a rotatable main shaft, engines, engine small-angle rotating mechanisms, a slip-flow rudder, a slip-flow rudder surface and a rudder; the body is connected with the front wings, the rear wings, the rotatable main shaft and the vertical tail wing; the front wings and the rear wings are arranged on the front end and the rear end of the body, and the front wing auxiliary wings and the rear wing auxiliary wings are respectively installed on the front wings and the rear wings; the vertical tail wing is installed on the tail end of the body, the rudder is installed on the vertical tail wing, and the rotatable main shaft is connected with the engine small-angle rotating mechanism in a rotating manner; the rotor wing is installed on an output shaft of the engine; the slip-flow rudder is installed on the rotatable main shaft, and the slip-flow rudder surface is installed on the slip-flow rudder. The invention relates to a control method of the tilt rotor transport aircraft. The control method comprises five steps. By adopting the layout and the control method of the tilt rotor transport aircraft, the pneumatic interference and mode transformation between the rotor wings and the wings of the aircraft can be solved, the transport capacity and the vertical take-off and landing capacity of the connected-wing layout can be reasonably combined, and a novel type of transport aircraft is formed.

The invention discloses an inclined fixed wing unmanned plane. The inclined fixed wing unmanned plane comprises a plane body and a wing inclined mechanism which can incline the wings, wherein the two sides of the plane body are provided with at least one pair of wings, and the tail of the plane body is also provided with a tail wing; and each wing is fixedly provided with a power device, each power device is connected with a screw propeller, the wings and the power devices are fixedly arranged together and can simultaneously move, and the rotation surfaces of the screw propellers are perpendicular to the horizontal planes of the wings all the time. The inclined fixed wing unmanned plane adopting the technical scheme can vertically take off and land and can realize high the cruising speed and power utilization efficiency and reasonable structure of a fixed wing plane; as the preferred technical scheme, the wings are both provided with subsidiary wings which are controlled by the wing inclined mechanism, and the power devices are all fixedly arranged on the middle parts of the wings; and the wings have enough rigidity and intensity, the wings are prevented from deforming because the power devices are fixedly arranged, and the localization of the wings is exact in revolution and fixation.

The invention provides a waverider integration design method for a hypersonic aircraft forebody, an air inlet and a wing. On the basis of waverider forebody/air inlet integration design, the theory of characteristic lines is applied for establishing a hypersonic aircraft forebody-air inlet integration axial symmetrical benchmark flow field and a wing symmetrical benchmark flow field, and a hypersonic aircraft forebody-air inlet integration configuration and a wing configuration are generated by applying the streamline-tracing technique in the two benchmark flow fields correspondingly. The two configurations together form a hypersonic aircraft forebody-air inlet-wing waverider integration configuration. In the design state, a whole hypersonic aircraft external flow field has waverider-like characteristics, the waverider forebody is used as a precompression face of the air inlet, and precompressed airflow is efficiently captured and provided for the air inlet, so that the waverider wing provides a high lift-drag ratio for the aircraft.

The invention discloses a method for designing a special wind turbine airfoil profile and a special wind turbine wing profile obtained by the designing method. In the invention, a corresponding airfoil profile can be obtained by expressing the geometrical shape of an airfoil profile with a simple high-order polynomial shape function and altering the coefficient of the shape function polynomial according to a special constraint condition. The special wind turbine airfoil profile is in a Reynolds number range with Re=1*106-3*106, has low-airfoil profile surface roughness sensitivity, higher lift coefficient and maximum lift drag ratio, as well as high lift drag ratio and favorable structural property in a larger attack angle range, and has favorable compatibility and acoustical property compared with other wind turbine airfoil profiles.

The invention provides a low-Reynolds-number wing section matched with a full-wing solar unmanned aerial vehicle. The low-Reynolds-number wing section is characterized in that the relative thickness of the wing section is 11%-13%, the maximum camber of the wind section is 2%-4%, and a chordwise position with the maximum camber is 25%-30%; the wing section has a single bent outline in a 70% chordwise range from a front edge. With a full-aircraft matching design, the checking of pneumatic characteristics of a pair of wing sections is finished by checking the corresponding pneumatic characteristics of the wing sections and applying the two wing sections to the full-wing solar unmanned aerial vehicle, and the low-Reynolds-number wing section shows good pneumatic characteristics and engineering enforceability and can meet application demands on low Reynolds number, high lift force and lift-to-drag ratio and large relative thickness of the full-wing solar unmanned aerial vehicle.

The invention provides a variant stealth aircraft and a variant method and application thereof. According to stealth requirements, the aircraft is designed by adopting a force-lifting body edge-wing variable-configuration pneumatic layout scheme, and comprises two retractable full-motion canards and two deflectable full-motion empennages. The retractable full-motion canards can be retracted into amachine body during cruising of an aircraft or flying at a high speed, so that the scattering cross section of a radar is reduced, thus improving the invisibility. The deflectable full-motion empennages can be deflected to different angles in different flight states so as to serve as V-shaped tails, flat tails or ventral fins. Thus, transformation of the pneumatic layout of the aircraft is achieved, and the layout comprises a canard-type layout, a tail-free layout and a three-wing-surface layout. Through the switching of the three pneumatic layouts, the advantages of the three pneumatic layouts can be combined, thus better satisfying the requirements for the maneuverability and invisibility of the aircraft under different flight profiles and flight states. As a new-generation manned and unmanned aerial combat platform with high lift-drag ratio, high maneuverability and good invisibility, the variant stealth aircraft can be used for executing air dominant combat tasks.

The invention discloses a hypersonic flight vehicle aerodynamic configuration based on pre-cooling type combined powder and a working method. The hypersonic flight vehicle aerodynamic configuration based on pre-cooling type combined powder comprises a flight vehicle aerodynamic configuration and a pre-cooling type air turbine rocket engine, wherein the pre-cooling type air turbine rocket engine comprises air inlet channels, a pre-cooling system, an air compressor, an annular combustion chamber, a turbine, an afterburner and exhaust nozzles which are successively connected; the air inlet channels and the lower surface of the front body of the flight vehicle aerodynamic configuration are designed integrally; an adjustable hemicone is arranged at an air inlet end of each air inlet channel; the exhaust nozzles and the lower surface of the rear body of the flight vehicle aerodynamic configuration are designed integrally; and an adjustable cone is arranged in each exhaust nozzle. Medium-altitude and low-altitude working performance advantages of turbine engines are retained, by application of a pre-cooler, it can be ensured that large residual thrust can be provided for a flight vehiclein a transonic/supersonic climbing acceleration process, and therefore, the flight vehicle can climb to a cruise altitude in an accelerated manner within a short time.

The invention discloses a hypersonic aircraft aero-propulsive integrated layout method based on combined power. According to the hypersonic aircraft aero-propulsive integrated layout method, a hypersonic air inlet channel of a dual-mode ramjet passage and a tailpipe nozzle are integrally designed with the front body and the rear body of an aircraft respectively; the dual-mode ramjet passage is taken as a basis; a turbine engine passage, a rocket ejector ramjet passage and the dual-mode ramjet passage are parallelly arranged, so that the three passages can coordinate with one another to work according to the own working conditions under the flight Mach number; the dual-mode ramjet passage and the rocket ejector ramjet passage are parallelly connected in an up-down manner; the turbine engine passage is parallelly connected with the dual-mode ramjet passage and the rocket ejector ramjet passage in a left-right manner; and a dual-engine mode under the left-right side-by-side layout is designed. By adoption of the hypersonic aircraft aero-propulsive integrated layout method disclosed by the invention, the middle-altitude and low-altitude working performance advantages of a turbine engine can be remained; through a rocket ejector ramjet, relatively high residual thrust can be provided in a transonic and supersonic climbing acceleration process of the aircraft; and the climbing acceleration of the aircraft to a cruising altitude within a short time can be achieved.

The invention discloses a vane airfoil profile of a low power wind driven generator, which is designed by adopting biological airfoil profile simulation technology and fluent software numerical simulation. The ratio of the maximum thickness d of the vane airfoil profile and the chord length c of the airfoil profile is d/c =9.83 percent. The maximum thickness position is x d /c =23.4 percent. The ratio of the maximum camber f of the airfoil profile and the chord length c of the airfoil profile is f/c=9.91 percent. The maximum camber position is the chord length part of x f/c=43.6 percent. Xd is the horizontal ordinate value of the maximum thickness part of the airfoil profile. Xf is the horizontal ordinate value of the maximum camber part of the airfoil profile. The coordinate axis origin is the front edge point of the airfoil profile. The horizontal ordinate x axis is superposed with chord length c and the direction points to the back edge of the airfoil profile from the front edge of the airfoil profile. Compared with the traditional conventional airfoil profile, the maximum thickness position xd and the maximum camber position xf of the invention are moved back and the invention has better lift efficiency and higher lift drag ratio. That is to say, the performance of the invention is improved on the whole.

The invention discloses an aerodynamic layout of a transatmospheric reusable aircraft, which comprises a head, a fuselage, a horizontal wing, a vertical tail, a body flap and a liquid hydrogen/liquid oxygen linear plug-type nozzle rocket engine. The invention has no booster, and adopts a single stage to vertically launch into orbit from the ground; after completing the mission, it can return to the atmosphere and glide horizontally to land, and can be completely reused after recovery. Through reasonable aerodynamic configuration and optimized design, the aircraft has a high lift-to-drag ratio and good stability and maneuverability in a wide range of altitudes and speeds. At the same time, the fuselage has a large effective volume and needs to carry propulsion Dose less. The invention has low launching cost, short launching preparation period and high reliability. It can overcome the shortcomings of current space vehicles that cannot be completely reused, expensive to launch, and cumbersome to use and maintain. It can be used not only for military space missions, but also for commercial and civil space services.

The invention discloses a small wind turbine wing type aerodynamic robust optimization design method suitable for turbulent working conditions, belonging to the technical field of wind turbines. The invention adopts the non-embedded probability collocation point method in the design to realize the quantitative characterization of the turbulence intensity of the uncertain parameter. Through the improved Hick-Henne type function parameterization method, Latin hypercube test method and CFD numerical calculation method, the kriging agent model between design variables, uncertain parameters and themaximum lift-to-drag ratio of the airfoil is established. On this basis, a robust optimization mathematical model is established to maximize the mean value of the maximum lift-to-drag ratio and minimize the standard deviation of the wind turbine wing under uncertain turbulent conditions. The non-embedded probabilistic configuration point method, the kriging proxy model, and the NSGA-II optimization algorithm are used to optimize the wind turbine airfoil. The invention improves the maximum lift-drag ratio of the airfoil, reduces the fluctuation range, improves the wind energy capture efficiency, and enhances the aerodynamic robustness under the turbulent condition. At the same time, the computational workload of robust optimization design is reduced, and the optimization efficiency is improved, which provides an important reference for the optimization design of wind turbine wing under turbulent conditions.