440 results about "Angle of attack" patented technology

An attitude display system for use in an aircraft has at least one display unit visible by a pilot within a cockpit. A symbology generator is configured for receiving input from at least one attitude sensor and generating symbology image data representative of a pear-shaped attitude indicator, distorted vertically to represent a pitch of the aircraft and rotated to represent a roll angle of the aircraft. The symbology generator may further be configured for adding a horizontal line that represents a horizon, a regular trapezoid representing a downward gravity vector at a base of the trapezoid, a pitch angle of the aircraft, split wings representing an angle of attack of the aircraft, and a stall warning in a sawtooth crack shape. A tail portion of a T-shape with the curved top edge is itself curved in a direction opposite that of an aircraft spin direction.

The invention provides a wind tunnel experiment-based wing high-attack-angle flow separation combined intelligent control method, which comprises the following steps of: arranging active flow control devices at 10% chord length and trailing edge flap of a wing to inhibit wing high-attack-angle flow separation; the active flow control device at the 10% chord length of the wing adopts direct jet flow, the jet flow direction is perpendicular to the chord line of the wing, three jet flow exciters are uniformly arranged along the spanwise direction, the active flow control device at the trailing edge flap adopts sweeping jet flow, the jet flow direction is parallel to the surface of the wing at the position, and 18 sweeping jet flow actuators are uniformly arranged along the spanwise direction; the jet flow of the two sets of flow control devices is adjusted through corresponding mass flow controllers, jet flow momentum distribution is dynamically optimized in combination with a reinforcement learning algorithm, the lift coefficient is remarkably increased, and energy consumption is reduced. The flow field state is fed back in real time through a pressure measuring system and a force measuring system, a closed-loop control system is constructed, and an execution action is output according to the flow field information around the wing to control an active flow control device arranged on the wing so as to inhibit flow separation on the surface of the wing, so that the wing gets rid of the stall state and the lift force is improved.

The invention relates to the technical field of aircraft control, and discloses a flight program planning method and device for an aircraft, equipment and a medium, and the method comprises the steps: obtaining a nominal flight program and a corresponding flight envelope function which is generated in a preset profile domain based on a process constraint condition and a terminal constraint condition of the aircraft; key nodes are selected in the speed domain of the nominal flight program, random quantities are overlaid on all calculation nodes based on a random walk strategy, and an overlaid result is generated; performing constraint correction on the superposition result through a flight envelope function to generate a corrected attack angle node value; and performing polynomial interpolation fitting on the corrected attack angle node value to generate a random walk flight program. According to the method, the random walk flight program with randomness and constraint conformity can be generated, so that the track unpredictability of the hypersonic gliding aircraft is enhanced, and diversified combat requirements of the aircraft such as high maneuverability and strong penetration are met.

The invention relates to the technical field of flight safety of civil aircrafts, in particular to a flight trajectory prediction method under the condition of air high-speed target hood perforation, which comprises the following steps: acquiring calculation coefficient parameter data of a target flight dynamics model, establishing a trajectory prediction framework based on a three-degree-of-freedom target flight dynamics model and target real-time flight data; the method comprises the following steps: performing numerical simulation on a flow field of a hood perforation area through fluid mechanics analysis software to obtain speed field and pressure field data, establishing a hood perforation flow field pneumatic analysis model by adopting a high-order polynomial fitting method, and calculating Mach number and pressure distribution after hood perforation; dynamically updating aerodynamic coefficient data and an aerodynamic parameter database; the speed and pressure distribution data are introduced into the three-degree-of-freedom target flight dynamics model for registration updating, a prediction track is obtained, the prediction precision is improved by 30%-50%, and especially under the condition of a large angle of attack, the prediction error is controlled to be within 5% from 20% or above.

The invention discloses a roll angle measuring method for a continuous variable roll test of a transient wind tunnel, and belongs to the technical field of wind tunnel test. Comprising the following steps: step S1, early-stage preparation: selecting two angle measurement sensors to carry out measurement work, and installing the two angle measurement sensors and an installation platform according to a preset angle; s2, carrying out field calibration operation; carrying out field calibration on the coefficient of the angle measurement sensor by adopting a fixed attack angle and variable roll angle mode; and S3, data interception, discretization and application during a step variable attack angle continuous variable roll angle test: in a wind tunnel test process, continuously collecting output voltage values of the two angle measurement sensors, calculating roll angles corresponding to the output voltage values, intercepting roll angle range values corresponding to the two angle measurement sensors according to needs, and calculating the roll angle range values of the two angle measurement sensors according to the intercepted roll angle range values. Splicing into a target measurement range; compared with a traditional step variable rolling and fixed rolling variable attack angle test mode, the scheme has the advantage that the test efficiency is greatly improved.

The invention relates to the field of wind driven generators, and particularly discloses a floating type vertical axis wind turbine multi-partition variable pitch control strategy which comprises the steps that S1, basic parameters and operating condition parameters of a unit are obtained; s2, establishing an aerodynamic model based on a double-disk multi-flow management theory and a superposition wind speed method; s3, measuring parameters through a sensor, and calculating a theoretical attack angle of the blade; s4, establishing a multi-partition pitch angle change function based on the theoretical attack angle; s5, constructing an NSGA-II multi-target optimization model, and taking maximization of an average power coefficient and minimization of a blade thrust standard deviation as targets; s6, substituting a function to obtain an optimal variable pitch control factor and an optimal pitch angle; and S7, the optimal pitch angle is converted into a variable pitch actuator instruction. The problems that an existing strategy is limited in partition, the wind direction influence is ignored, and optimal parameters for working condition adaptation are lacked are solved, the wind energy capture efficiency is improved, the load fluctuation is reduced, and the variable pitch control accuracy is improved.

The invention belongs to the technical field of hypersonic flight vehicle control, and particularly relates to an INDI-based flight and thrust integrated control method for a hypersonic flight vehicle. The method comprises the following specific steps: designing a vertical-direction-controlled flight-thrust integrated longitudinal dynamic model of the hypersonic aircraft, designing a trajectory guiding outer ring to realize height and speed tracking control, designing an attitude stabilizing outer ring to realize attack angle tracking control, constructing an attitude control inner ring to realize flight-thrust control of an acceleration level, and designing a flight-thrust model of the hypersonic aircraft. And designing an engine compensation layer to process dynamic delay of an actuator. Through the MIMO-INDI mechanism, the pneumatic / propulsion coupling effect is actively utilized, collaborative distribution of the fuel equivalence ratio of the elevator and the engine is achieved, and the comprehensive control performance of the aircraft is improved. The method is a flight / propulsion integrated control method and has a wide application prospect.

The invention relates to the technical field of aircraft electric heating deicing, in particular to a pulse heating aircraft deicing method and system.The method comprises the following steps that the ice layer thickness, the ice layer type, the airspeed and the attack angle of each aircraft surface area are collected, the inherent pneumatic sensitivity coefficient of each area is called, the aerodynamic penalty index of each area is established, and the aerodynamic penalty index of each area is calculated; and according to the values of the aerodynamic penalty indexes of the areas, performing descending order arrangement to generate a deicing area priority sequence. Through multi-parameter joint evaluation of the ice layer thickness, the ice layer type, the airspeed and the incidence angle, the heating resource distribution sequence is optimized, energy waste is avoided, in the local heating execution stage, temperature change slope real-time monitoring is adopted to replace fixed time control, the actual thermal response behavior of the ice layer can be dynamically responded, and the real-time thermal response performance of the ice layer is improved. The initial phase change process of the ice crystal is identified through a phase change starting slope threshold value, and then a medium-power pulse control mode is switched, so that a steady-state regulation heating path is formed, and the phase change control continuity is strengthened.

The invention discloses a hydrofoil with an adjustable attack angle for an electric control type boat and a control system, relates to the technical field of attack angle control, is used for solving the hidden danger of grounding caused by insufficient underwater distance and limited lift force compensation during navigation in a shallow water area, and comprises a working condition judgment module, a backstepping fitting module, a risk grade judgment module and an attack angle execution module, the modules are in signal connection; the working condition judgment module obtains a water depth value, a draught value and a water leaving height value through a sensor, calculates a bottom gap and compares the bottom gap with a safe minimum gap, and judges shallow water working condition triggering; the backstepping fitting module calculates the actual lift force by combining the gravity, the buoyancy and the vertical acceleration of the boat body when the shallow water working condition is triggered, and fits the shallow water corrected lift force slope; the risk grade judgment module calculates the residual available lifting force and the lifting force required by lifting, and sets a risk flag bit through a dual-condition criterion; and the attack angle execution module adjusts the attack angle according to the risk flag bit to realize closed-loop control.

The invention discloses a ship air sail navigation aiding energy efficiency dynamic optimization method and system and a ship, and relates to the technical field of ship energy conservation. Dynamic parameters in the ship navigation process are collected in real time; establishing a dynamic coupling model; the coupling model is utilized to quantify the dynamic interference of the hull attitude parameters on the air sail state parameters into attack angle offset; a reinforcement learning algorithm is adopted to simulate the attack angle offset, and the optimal attack angle of the air sail is obtained; according to the optimal attack angle and the environmental parameters, target adjustment parameters of the air sail are calculated; and driving the air sail to adjust the sail body rotation angle and the sail surface inclination angle in real time according to the target adjustment parameters. According to the method, the air sails always work in the efficient interval with the maximum lift coefficient and the minimum resistance coefficient, the pneumatic interference model is introduced, the total auxiliary thrust of the multiple sails is improved, the problem that the energy-saving effect is reduced due to dynamic deviation of the attack angle is effectively solved, and the situation that the fuel consumption of a main engine is reduced is stably achieved.

The invention relates to an active gravity center adjustment control method for an advanced aircraft variant process, and the method comprises the steps: collecting the variant state information of an aircraft in real time, synchronously obtaining the flight state information, and obtaining an input vector; wherein the variant state information comprises a wing sweep angle, a span-chord ratio and a flap deflection angle, and the flight state information comprises an attack angle, a flight speed, a height, a pitch angle and an acceleration; according to the input vector, an initial pneumatic center position is calculated through a pneumatic center prediction function, and the pneumatic center position is determined according to the pneumatic center position variable quantity; calculating a target gravity center position according to the longitudinal static stability through the average pneumatic chord length of the aircraft and the pneumatic center position, and obtaining a target gravity center position adjustment amount; and adjusting the center of gravity of the aircraft to the target center-of-gravity position through a center-of-gravity adjustment executing mechanism. Active gravity center adjustment can be provided for the advanced aircraft variant process, attitude disturbance is restrained through information collection, pneumatic center calculation and the like, and flight safety is guaranteed.

The invention belongs to the technical field of high-speed train aerodynamic optimization, and particularly relates to a train pneumatic drag reduction and wake flow regulation and control device and method based on bionic shark fins. The device comprises a base, the base is arranged on the lower surface of the vehicle body fairing through a plurality of bolts, a fin plate is arranged on the base, and a fixing block is arranged at the bottom of the fin plate; the base is provided with a plurality of limiting and locking mechanisms, the fin plate is limited through the limiting and locking mechanisms, the limiting and locking mechanisms are provided with driving mechanisms, the driving mechanisms are connected with rotating shafts, and the rotating shafts are connected with the fixing blocks and used for driving the fin plate to rotate; a lateral reinforcing plate is arranged on the outer side of the base, a pneumatic compensation plate is arranged between the lateral reinforcing plate and the fin plate, and the pneumatic compensation plate is fixedly connected with the base plate. By dynamically adjusting the attack angle and clearance compensation of an additional device on the surface of the vehicle body, aerodynamic resistance is reduced, wake flow stability is improved, and self-adaptive optimization under complex working conditions is achieved.

The invention belongs to the related technical field of underwater communication equipment, and discloses a variable-depth pull-type buoy body, the overall appearance of the variable-depth pull-type buoy body is of a wing-shaped structure, and the variable-depth pull-type buoy body comprises flow stop plate assemblies, an antenna assembly, a watertight instrument cabin, a pull sliding block and a horizontal rudder assembly, the antenna assembly is arranged in a groove in the upper surface of the wing-shaped structure, and the whole watertight instrument cabin is pre-buried in the wing-shaped structure; the dragging sliding block is arranged on the lower surface of the wing-shaped structure and can move front and back, the diving rudder assembly can change the attack angle of the diving rudder, and therefore the underwater depth and posture of the buoy body can be adjusted through cooperation of the diving rudder assembly and the diving rudder assembly. According to the invention, rapid switching between a water surface working state and an underwater working state can be realized, and a function of real-time depth control in an underwater towing process is realized; in addition, when the buoy body sails underwater along with the platform, no obvious signal characteristic exists on the water surface, good working concealment is achieved, and the communication capacity of the deep-sea movable underwater platform can be effectively improved.

The invention discloses a laser guidance unmanned aerial vehicle false attack angle feedback stability augmentation control method comprising the following steps that a three-loop autopilot frame with a false attack angle as a feedback signal is set, and the false attack angle is a calculated value of the attack angle; three-loop autopilot control parameters are set, and an autopilot is obtained; and an autopilot is adopted to control the flight state of the laser guidance unmanned aerial vehicle. According to the false attack angle feedback stability augmentation control method for the laser guidance unmanned aerial vehicle, stability augmentation control of the system is achieved through a three-loop autopilot frame with the false attack angle as a feedback signal, higher robustness is obtained, and an expected instruction can still be tracked under interference.

The invention provides a method and device for calculating rudder deviation required by unit normal overload of a statically unstable aircraft, and the method comprises the steps: selecting a plurality of calculation state points, and obtaining a small disturbance linearization equation at each calculation state point; a stability augmentation control law is constructed, the pitching moment characteristic of the aircraft is changed by feeding back the attack angle, the pitching damping characteristic of the aircraft is improved by feeding back the pitching angle rate, and therefore the aircraft is not divergent under the control of the stability augmentation control law and has the expected stability and damping characteristic; establishing a calculation model of rudder deflection required by unit normal overload based on a stability augmentation control law and a small disturbance linearization equation at a calculation state point, calculating normal overload by the calculation model through a pitch angle rate, an attack angle and a vacuum speed, and processing the calculation model to obtain a transfer function from a control channel elevator instruction to the normal overload; and calculating a normal overload value generated by unit elevator skewness based on the transfer function, wherein the reciprocal of the normal overload value is the rudder skewness required by unit normal overload.

The invention discloses a synthetic attack angle control method during aircraft-rocket separation, and relates to the technical field of spaceflight, and the method comprises the following steps: controlling a deviation value between a pitching program angle and a trajectory inclination angle to be smaller than a preset deviation threshold value, and adjusting the pitching program angle in real time according to the trajectory inclination angle, thereby reducing a trajectory attack angle; when the ballistic attack angle is within a preset ballistic attack angle threshold value, an overload difference value between the actual normal overload at the current moment and the normal overload at the corresponding current height is calculated, and the correction amount of the pitching program angle is calculated based on the overload difference value; the correction amount of the pitching program angle is compensated into the pitching program angle, and the synthetic attack angle of the rocket is converged, so that the wind attack angle is reduced; before device-rocket separation, transverse normal acceleration data are continuously collected, and if the transverse normal acceleration data exceed a change threshold value, device-rocket separation is carried out after the acceleration data are recovered to be normal. According to the method, potential safety hazards caused by consideration of a single factor can be effectively avoided, and the safety and reliability of the separation process are remarkably improved.

The invention discloses a deformable empennage applied to a bionic aircraft, and relates to the technical field of bionic flying robots. Through efficient cooperation of five-degree-of-freedom motion of the empennage and deformation motion of the wings, the maneuverability and adaptability of the aircraft are remarkably improved. The bionic aircraft comprises a motor, a fuselage, a deformation wing, an empennage and an empennage deformation driving mechanism. And flexible switching among a horizontal tail state, a positive V tail state and an inverted V tail state can be realized. When the state is switched to the positive V tail state, parasitic resistance can be effectively reduced, wingtip vortex interference is reduced, and the aerodynamic efficiency of the cruising stage is improved; when the inverted V-tail state is switched, the influence of wing wake flow can be avoided, the aerodynamic stability is obviously enhanced under the working condition of a large angle of attack, and the stall risk is effectively inhibited; and when the horizontal tail state is switched, aerodynamic lift force distribution in the low-speed stage can be optimized, and the aerodynamic response of take-off and landing control is improved. The flexible switching of the empennage among the three states provides better pneumatic suitability for the aircraft in different flight stages and task scenes.

The present application belongs to the technical field of unmanned sailboat, and particularly relates to a rigid wing sail self-balancing turning sail system of unmanned sailboat, which comprises a main sail skeleton, a main sail shell, a turning sail control box, a tail wing connecting rod assembly, a tail wing main body, a steering engine and a tail wing rope driving assembly. The driving end of the steering engine is controlled to move, and the tail wing main body is driven to rotate through the tail wing rope driving assembly, so as to adjust the deflection angle of the tail wing main body relative to the main sail as a whole, adjust the included angle (attack angle) between the main sail as a whole and the wind, and make the main sail as a whole reach a predetermined attack angle and keep a stable balance state. The mast of the main sail as a whole can rotate freely around the unmanned sailboat body, realizes the rotation decoupling of the rigid wing sail and the ship body, the bow shaking movement of the unmanned sailboat does not affect the attack angle of the rigid wing sail, the tail wing can be regarded as the air rudder of the main sail by adjusting the attack angle of the main sail through the tail wing, and the method makes the main sail attack angle adjustment more efficient and rapid, and reduces the adjustment power consumption.

The invention provides an aircraft control method and a rotorcraft. The method comprises the steps that the current attack angle of a high lift wing is obtained through an angle-of-attack device; calculating a target included angle between the high lift wing and the rotor wing according to the relation between the current attack angle and a set angle; the set angle relation is an incidence relation between an included angle between the high lift wing and the rotor wing and the attack angle; adjusting the included angle between a high lift wing and a rotor wing of the rotor craft to be the target included angle through the angle adjuster; when the included angle between the high lift wing and the target included angle is the target included angle, the attack angle of the high lift wing is correspondingly adjusted to be the optimal attack angle. The current attack angle of the high-lift wing is measured through the angle-of-attack device, the relative angle between the high-lift wing and the rotor wing is adjusted accordingly, the optimal lift-drag ratio working condition of the high-lift wing is achieved, power consumption of the rotor wing is reduced, and the loading capacity, the endurance time and the endurance mileage of the rotorcraft are improved.

The invention provides a fine adjustment device for the meshing depth of a transmission gear, which belongs to the technical field of attack angle sensors and comprises a flange bracket, a mounting bracket, a rotary transformer assembly, a damper assembly and a weather vane rotating shaft assembly, the rotary transformer assembly, the damper assembly and the weather vane rotating shaft assembly are fixed in the circular mounting groove through an eccentric mounting flange and are meshed with one another, and mounting bracket scale marks are arranged on the eccentric mounting flange and the circular mounting groove. According to the device, a linear relation is established between the meshing depth and the scale marks, and the gear meshing depth debugging process of all parts of the attack angle sensor, which is adjusted according to artificial experience, is converted into an indication value which can be directly visually observed. The requirement for experience of operators is lowered, unmeasurable artificial experience is solidified into measurable indication data, the debugging efficiency is improved, and the number of times of wind tunnel test verification caused by repeated debugging is reduced.

The application belongs to the technical field of aircraft simulation, and particularly relates to a flight simulation method and device under gust load. The method comprises the following steps: S1, performing aircraft 1g level flight trimming to determine the aircraft trimming angle of attack and the aircraft trimming elevator deflection; S2, determining the aerodynamic force of the aircraft; S3, performing dynamic response solution on the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, the flight speed and the pitch angular velocity, and determining the normal overload of the aircraft after superimposing the flight speed on the wind speed of the vertical gust field; S4, obtaining the elevator deflection calculated by the stability augmentation system, superimposing the elevator deflection at a simulation moment to obtain the actual required elevator deflection; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angular velocity, the aerodynamic angle of attack and the actual required elevator deflection until the simulation is completed. The application considers the role of the stability augmentation system in the determination of the aircraft gust load, and the simulation process is simple and has high precision.

The application discloses a kind of high-speed low-energy consumption autonomous control method of underwater vehicle based on prediction model optimization, belongs to underwater vehicle autonomous control technical field.The application establishes the longitudinal prediction model of vehicle, in each control cycle, with current navigation state as initial condition, based on the prediction model is carried out rolling optimization.Optimization function not only includes depth tracking error, more critically, introduce the navigation resistance estimation term directly related with angle of attack and rudder angle, to minimize total resistance as target, solve the optimal angle of attack sequence in future time domain and corresponding rudder angle, thrust control instruction.The application embeds hydrodynamic optimization in real-time closed-loop control, so that the vehicle can actively maintain the optimal drag-reducing attitude navigation, rather than passive maintenance level, thereby significantly reducing the total resistance and energy consumption of underwater vehicle high-speed navigation, effectively suppresses the phenomenon of buried bow, improves navigation stability.

The application relates to a stall-avoiding angle-of-attack limit design method, and belongs to the technical field of unmanned aerial vehicle control. The method comprises the following steps: acquiring aircraft parameters related to the angle of attack, acquiring the corresponding maximum pitching-down moment according to an angle-of-attack limiter starting threshold, and acquiring an actual crossing angle of attack according to the maximum pitching-down moment; designing an angle-of-attack integral formula according to a pitching axis mass moment of inertia, a wing area, an average aerodynamic chord length and a point dynamic pressure, and calculating an expected crossing angle of attack according to the angle-of-attack integral formula; comparing a difference value between the actual crossing angle of attack and the expected crossing angle of attack with a margin threshold, and judging whether the angle-of-attack limiter meets the requirements according to the comparison size. The application is based on the constraint relationship between the maximum pitching-down moment curve slope and the zero-crossing angle of attack, relies on the most original mechanical equation, designs the angle-of-attack limit design criterion, can be evaluated only by using a few simple parameters of the aircraft, is universal, and can be used for precise evaluation of a new round of aircraft layout in a short time under various states.

The invention provides an aircraft dynamic derivative test device in a strong interference environment, which relates to the technical field of aircraft aerodynamic characteristic tests and comprises an interference flow field establishing device, a dynamic derivative measuring device and a synchronous control module. The disturbing flow field establishing device is composed of a shock wave generator, a supporting connecting piece and a wind tunnel attack angle mechanism, the shock wave generator is connected with the wind tunnel attack angle mechanism through the supporting connecting piece, + / -15-degree pitching and + / -10-degree yawing motion is achieved based on the wind tunnel attack angle mechanism, and multiple types of strong disturbing flow fields can be established by replacing the shock wave generator; the dynamic derivative measuring device is installed in a plug-in mechanism of a wind tunnel, based on a forced vibration test method, a model is driven by a high-precision servo motor to vibrate slightly, precise acquisition is realized by matching with a high-frequency dynamic balance and a data filtering unit, and the displacement of the model relative to a shock wave generator can be adjusted through the plug-in mechanism. According to the scheme, a real strong interference environment can be accurately reproduced, and the dynamic derivative measurement accuracy is improved.

To provide a simulator capable of appropriately reproducing acceleration acting on a user.SOLUTION: The control unit of the simulator receives the squaring angle detected by the first detecting unit or the squaring angle calculated from the angular velocity detected by the first detecting unit and the angle of attack detected by the second detecting unit, calculating an inertial force generated in the support plate in the turn action from an equation of motion related to a gravity of the user acting on the support plate, a frictional force between the support plate and the reference surface, and an air resistance acting on the user in the turn action, using a friction coefficient between the support plate and the reference surface in the turn action as a sum of a first friction coefficient between the support plate and the reference surface in a carving turn action in which the turn action is performed in a state of being inclined in the roll direction and a second friction coefficient between the support plate and the reference surface in a squatting turn action in which the turn action is performed so as to rotate in the yaw direction; The turning radius of the support plate in the turn operation is calculated, and the driving part is controlled based on the calculation result.SELECTED DRAWING: Figure 1

The utility model provides a three-blade curved surface axial flow type impeller. The three-blade curved surface axial flow type impeller comprises a hub, a mounting plate and blades, an inner cavity of the hub is provided with an internal thread; the mounting plates are uniformly distributed along the periphery of the hub, the mounting plates are provided with locking pieces, and the blades are connected with the mounting plates through the locking pieces; the blade wing is connected with the hub through the mounting plate, the attack angle of the blade wing is alpha, and the torsion angle of the blade wing is Beta. According to the utility model, the hub, the mounting plate, the locking piece, the screw rod, the screw cap, the blade wing and the like are matched for use, and the blade wing and the mounting plate are connected through the locking piece, so that the blade wing and the mounting plate are connected through the screw rod and the screw cap; the design that the three blades are matched with the mounting plate and the hub is changed into an axial flow type impeller, so that the shearing force of the blades in the using process is weakened, the output performance of axial flow is enhanced, and the energy-saving effect is achieved by reducing the dynamic coefficient of the whole mechanism during rotation.

The invention discloses a multi-rotor aircraft with lift-enhanced wings, which comprises rotor power devices, a fuselage for mounting airborne equipment and connecting the rotor power devices, and a pair of independent and controllable lift-enhanced wings, the high-lift wings are mounted on the fuselage; when the multi-rotor aircraft with the high-lift wings flies, a flight control algorithm controls the tilting angles of the high-lift wings according to the attitude and the flight speed of the aircraft, so that the high-lift wings are in a proper attack angle state, and proper lift force can be provided for the aircraft. The flying speed can be increased, the cruising ability is improved, and transportation and storage are convenient.

The application designs a high-speed aircraft cluster cooperative guidance method based on space-time synchronous attraction domain. An artificial potential field is designed to keep a proper distance between aircrafts, avoid collision while gathering; a speed matching item is designed based on consistency theory to realize dynamic consistency of aircraft speed; a guidance law based on resistance acceleration-speed profile is designed to meet the cooperative requirements of terminal range, height, speed and other aircrafts, and the combination of the three realizes efficient autonomous clustering and guidance of the aircraft. The application can make the aircraft cluster reach the target point with high precision under the conditions of meeting complex constraints such as heat flow, overload, dynamic pressure, collision avoidance, clustering, dynamic consistency of speed and the like through the design of attraction domain, angle of attack and roll angle control.

The invention provides a dual-channel attitude control retrimming method based on an interference source, which comprises the following steps: establishing an attitude dynamic model, and calculating to obtain a static attack swing angle; taking the calculated static attack swing angle as the input of load calculation, and drawing a segmented limit line by using the limit value of each working condition load qalpha meeting the strength requirement of the calibration structure; according to the structural bearing capacity on a pitching channel and a yawing channel of the rocket, the total load q alpha limiting value is decomposed to the pitching channel and the yawing channel, the wind attack angle and the wind sideslip angle of shear wind are kept unchanged, and the attack angle and the sideslip angle of stable wind needing to be retrimmed are obtained through calculation; according to the obtained attack angle and sideslip angle after rebalancing, according to different Ma number intervals, the wind speed and the wind speed direction at different moments or under Ma are adjusted to achieve the balancing condition, matching of the needed balancing attack angle and sideslip angle with an interference source is completed, and a corresponding matched swing angle after rebalancing is obtained; and taking a re-trimmed attack swing angle calculation result as a final condition of load design.

Watercraft (1) with a keel-shaped hydrofoil (25) arranged on a keel (20) and a rudder-shaped hydrofoil (35) arranged on a rudder (30), wherein the sword and / or the rudder wing (25, 35) are electrically adjustable, - with an electronic distance sensor (60) arranged on the bow or mast of the watercraft (1) for non-contact detection of a distance to a water surface, - with a control unit (70) for controlling an angle of attack of the sword and / or rudder wing (25, 35), wherein the control unit (70) is designed such that the angles of attack of the sword and / or rudder wings (25, 35) are controlled depending on the distance detected by the electronic distance sensor (60), characterized in that that the water level sensor (60) is arranged on the bow or mast in such a way that a distance to a water surface in front of the watercraft (1) is detected, and that the sword (20) has a capacitive water level sensor (80) for detecting the sword wing immersion depth (h) KT ) and / or distance (h R ) of the hull (10) to the water surface.