829 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 relates to the technical field of flight control, in particular to a fixed-wing unmanned aerial vehicle attitude correction and precise flight control method fused with visual navigation, which comprises the following steps: extracting ORB feature points and constructing a dynamic feature point cloud; simultaneously acquiring angular velocity and acceleration data of the inertial measurement unit; performing tight coupling fusion processing on the dynamic feature point cloud and the inertial measurement data to generate real-time pose data; inputting the real-time pose data into an attitude correction parameter calculation module, and generating an attitude correction parameter through a dynamic weight distribution algorithm in combination with a preset flight path planning model and GPS positioning data; and a control surface deflection instruction and an accelerator control instruction are generated through the multi-constraint optimization model, and unmanned aerial vehicle flight attitude closed-loop correction is completed. According to the method, aerodynamic parameters such as the air speed, the wind field lateral disturbance and the attack angle are introduced to serve as coupling input, the control surface deflection amount and limiting conditions are dynamically adjusted, and self-adaptive modeling and control over the nonlinear flight characteristics are achieved.

Disclosed herein is are systems, devices, and methods for an autorotating aerial device which includes a housing member having disposed thereon an actuator and a controller configured to control the actuator. The device also includes a wing member coupled to the actuator, the wing member including a main wing portion and a flap portion adjustable with respect to the main wing portion. The controller is configured to control the actuator to switch the autorotating aerial device between a diving mode of operation and an autorotating mode of operation based on adjusting an angle of attack of the flap portion, the angle of attack being with respect to a lateral axis along the main wing portion.

The invention provides a gust-oriented wind generating set anti-overspeed control method and device and a medium, and belongs to the field of wind generating set control. The method comprises the following steps: acquiring real-time wind speed data and generator speed data; determining whether to enter a gust mode of the wind generating set according to the real-time wind speed data; under the condition that the wind generating set enters the gust mode, the minimum pitch angle of operation of the wind generating set is adjusted to be the minimum pitch angle of operation of the gust mode; calculating generator speed acceleration according to the generator speed data; determining whether to enter an acceleration anti-overspeed control strategy according to the generator rotating speed data and the generator rotating speed acceleration; and under the condition that it is determined that the acceleration anti-overspeed control strategy is entered, the preset variable-pitch rate is superposed on the normal variable-pitch rate. The pitch angle of the wind generating set in operation is increased when the wind generating set enters the gust mode, so that the attack angle margin of the blade can be kept in a safe range, and after the acceleration overspeed prevention control strategy is entered, the preset variable pitch rate is superposed on the normal variable pitch rate.

This disclosure provides an adaptive STSMC hierarchical control method for depth tracking of an underwater vehicle. The method includes: using an adaptive line-of-sight guidance algorithm to estimate the vehicle's angle of attack in real time and obtain an expected pitch angle; employing a nonlinear interference observer based on a sliding mode surface error to detect unknown interference; and applying an STSMC controller with adaptive saturation compensation to determine the expected control rudder angle from the sliding mode surface error, the expected pitch angle, and the observed interference. The adaptive saturation compensation corrects the rudder angle based on the saturation deviation, which is determined by the previously calculated expected control rudder angle and the actual output rudder angle at the previous moment. By introducing an adaptive saturation compensator in the dynamic control layer, the issue of rudder angle saturation is mitigated, thereby enhancing both control performance and stability.

A fixed-wing unmanned aerial vehicle large-attack-angle longitudinal aerodynamic characteristic test device and a test method relate to the technical field of aircraft aerodynamic characteristic test, and mainly comprise a wind tunnel simulation device, an unmanned aerial vehicle bearing and motion simulation device, a sensor arranged on an unmanned aerial vehicle or a model, a data acquisition system and the like. The simulation module is used for simulating longitudinal motion states and attitudes of the unmanned aerial vehicle in various pneumatic environments; the device is used for experiments, nonlinear longitudinal aerodynamic characteristics of the unmanned aerial vehicle under various attack angles are obtained, a phase diagram reflecting the attack angle-pitch angle speed change relation of the system state is obtained, a balance point under the normal working condition and the limit working condition is determined, and a basis is provided for flight control of the unmanned aerial vehicle under the large attack angle limit working condition.

The invention discloses an online efficient load shedding control method based on a forecast wind field. The method comprises the following steps: firstly, acquiring a load attack angle and a sideslip angle which are calculated in real time on a rocket; amplitude limiting processing is carried out on the load attack angle and the sideslip angle, an attack angle control instruction is generated after gain control network calculation is carried out, and an attitude control system carries out direct attack angle control in a load sensitive area in combination with the load shedding requirement and the guidance deviation correction capacity; and pitching and yaw feedback control instructions are generated, active load shedding control is implemented, and errors generated by attack angle control are compensated. The on-line active and passive load shedding mode covers the real deviation or interference influence during flight, the generated load shedding instruction is better matched with the actual flight, the universality is good, the implementation is easy, and the load shedding effect is better.

The invention provides a multi-airfoil combined axial flow fan blade design method and system based on CST parameterization, and relates to the technical field of axial flow fan blade optimizing.The multi-airfoil combined axial flow fan blade design method comprises the following steps that S1, a plurality of characteristic sections are arranged in the radial direction of a blade, and the airfoil corresponding to each characteristic section is determined; s2, geometric modeling is conducted on the airfoils of all the characteristic sections through a CST parameterization method, and description parameter sets corresponding to all the airfoils are obtained; s3, performing aerodynamic performance simulation on the airfoils through the description parameter group to determine the attack angle of each airfoil; s4, the airfoils are combined at the corresponding attack angles, smooth transition cambered surfaces are generated between the airfoils, and an initial blade model is obtained; and S5, optimizing the initial blade model to obtain a target blade model by adjusting the description parameter group of each airfoil by taking the maximization of the lift-drag ratio as a target. Compared with a traditional single arc thin plate blade profile design, the pneumatic performance of the blade is more optimized, and the problem that a single blade profile is difficult to cope with complex pneumatic conditions where different sections are located is effectively solved.

The invention provides a multi-degree-of-freedom coupling execution mechanism for deflection of a water tunnel test model. The multi-degree-of-freedom coupling execution mechanism comprises a shell, an attack angle motion module, a yaw motion module and a rolling motion module, wherein the attack angle motion module, the yaw motion module and the rolling motion module are arranged in the shell; according to the mechanism, in the aspect of attack angle motion, a pitching motor drives a pitching lead screw through a mechanical transmission assembly, and control over the attack angle of a model balance component can be achieved by combining a translation sliding block and a pitching sliding block which slide in a matched mode in the height direction. The yaw movement module drives a yaw lead screw through a yaw motor, a pitching arc-shaped guide rail is stably pushed to move along the yaw arc-shaped guide rail through a yaw sliding block, and accurate adjustment of the yaw angle is achieved. The rolling motion module effectively realizes the rolling action of the model balance component through a rolling motor; the pitching, yawing and rolling motion modules can be independently controlled and can also be controlled in a coupling mode, the unique design greatly improves the flexibility and adaptability of mechanism motion control, and various complex working condition requirements can be accurately met.

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The first wing can be coupled to the boom by a (first) articulated joint. The propulsion assembly can be coupled to the boom by a (second) articulated joint. The articulated joints can provide various adjustments between the wing(s) and the boom (e.g., stowed, deployed, anhedral, dihedral, sweep, angle of attack, collective pitch control, the cyclic pitch control, etc.) and between the propulsion system and the boom (e.g., to move between vertical flight and horizontal flight). A payload can be connected to the boom with one of the articulated joints whereby the articulated joint adjusts a position of the payload during flight to maintain desired flight characteristics and stabilize the payload.

The invention discloses a reinforcement learning driven hypersonic variant aircraft reentry trajectory optimization method, and belongs to the field of hypersonic aircraft and trajectory optimization. The implementation method comprises the following steps: establishing a waverider-like hypersonic variant aircraft parameterized model comprising a waverider fuselage and two deformed wings based on a category shape conversion (CST) method; the influence of factors such as flight height H, flight Mach number Ma, attack angle alpha, inner wing section sweepback x1, outer wing section sweepback x2 and the like is considered, a Kriging agent model is established, and the aerodynamic performance of the hypersonic speed variant aircraft is obtained. Based on a DDPG reinforcement learning algorithm, according to flight characteristics of a hypersonic speed variant aircraft in a reentry section flight process, and in combination with expert experience and dynamics, start and end states, heat flow, overload and dynamic pressure constraints in the flight process, training of a reentry section trajectory optimization model driven by the DDPG algorithm is realized. Through flight simulation prediction of different hypersonic speed variant aircrafts, variant decisions are given in real time, and reentry section trajectory optimization of the hypersonic speed variant aircrafts is realized.

The invention discloses a balance centering measurement and correction method for a wind tunnel force measurement test, and belongs to the technical field of wind tunnel force measurement. The problem that in the prior art, a traditional balance centering measurement and correction method is large in test result error due to the fact that the actual centering position and the theoretical centering position of a balance are inconsistent is solved. According to the invention, a wind tunnel test model is designed and processed, the positions of loading points are determined, each loading point is loaded, the angle of attack of a balance is adjusted through a support system, and then a normal force and a pitching moment are measured through the balance; according to the normal force and the pitching moment, force arms of the loading points at the fixed angle of attack are obtained, and the actual distance between each loading point and the balance calibration center in the flow direction is obtained through fitting; and obtaining the actual position of the balance center according to the difference between the actual distance and the theoretical distance from each loading point to the balance center along the flow direction. The method effectively improves the precision of the obtained balance center calibration position, and can be applied to a wind tunnel force measurement test.

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 belongs to the technical field of aircraft design, and particularly relates to a static pressure source position error correction method, which comprises the following steps: performing theoretical appearance static pressure source position error correction analysis based on current numerical simulation and flight test state points to obtain a static pressure source pressure coefficient Cp, a flight test static pressure source pressure coefficient Cp0 and an actually measured static pressure source pressure coefficient Cp1; static pressure source position error correction data optimization is carried out, a static pressure source position error correction relation Cp'during real flight of the theoretical appearance is obtained, and static pressure source position error correction is carried out; appearance measurement, CFD simulation and test flight data comprehensive application are realized; correcting the position error of the static pressure source into a function of Mach number, angle of attack and height; and not only can the requirement of the CCAR25 part for the air pressure height be met, but also the requirement of RVSM operation in AC-21-13 for the error of a height measurement system can be met. The method can be used for atmospheric data correction of transportation aircrafts, and atmospheric data correction of other aircrafts can be used for reference.

The invention discloses a non-linear state rapid fitting method for an aircraft, and belongs to the field of aircraft control. The implementation method comprises the following steps: constructing a two-loop attitude pilot and a three-loop attitude pilot, and discretizing to obtain a guidance instruction with a time-delay effect; according to coordinate conversion, forward deriving each aerodynamic parameter of the aircraft and an aircraft overload instruction equation; the method comprises the following steps: calculating all pneumatic data according to part of pneumatic data obtained by a blowing experiment of an aircraft, and fitting a guidance instruction with a balance attack angle, a sideslip angle and a rudder deflection angle of the aircraft to obtain a corresponding interpolation table; all aerodynamic parameters in the flight process of the aircraft are obtained according to an aircraft guidance instruction and an interpolation table in the actual situation, namely, nonlinear state rapid fitting of the aircraft is achieved, and simulation reproduction of the aircraft as a mass point in the whole guidance process is achieved according to the nonlinear state rapid fitting result. And the precision and efficiency of aircraft guidance control are improved.

Disclosed in the present invention are a method and system for estimating an airspeed, an angle of attack, and a sideslip angle of an aircraft. The method comprises: modeling the full six-degree-of-freedom dynamics, the aerodynamics, a force, and a moment of an aircraft in a nonlinear state-space form; using an inertial measurement unit to measure the body angular rate and body acceleration of the aircraft; using a global navigation satellite system to measure a ground speed; using data of systems of the aircraft other than an air data system to determine air data measurement; and using a plurality of Kalman filters to estimate the airspeed and / or the angle of attack and / or the sideslip angle from all measured data. The method and the system provided by the present invention solve the problem in the prior art of inability to improve the safety and reliability by installing a redundant air data computer / sensor on a size-, weight-, and power-constrained aircraft, and the problem in the prior art of being not independent of a conventional air data system at the measurement principle level.

The invention provides a displacement vortex generator and a position regulation and control method, and belongs to the field of active flow control equipment. The VG comprises a belt, a belt guide rail, a belt bearing, a sliding block, a supporting column, a triangular VG body, a guide rail supporting block, a limiting groove, a body connecting plate and a motor. The belt is arranged on the belt guide rail and is driven by the motor; the sliding block is arranged on the belt, is in an inverted concave shape and is connected with the body connecting plate through the supporting column, and the VG bodies are fixed to the two ends of the body connecting plate. The belt guide rail is arranged on the guide rail supporting block, and the sliding block can penetrate through the limiting groove. The motor is controlled to drive the belt to drive the sliding block and the VG body to move, and accurate adjustment of the position of the vortex generator body is achieved. The position of the vortex generator is automatically adjusted according to changes of the incoming flow wind speed and the attack angle, the optimal pneumatic effect is kept under different working conditions, the pneumatic performance and the power generation efficiency of the wind turbine blade are remarkably improved, and the negative influence of the front edge roughness can be effectively weakened.

The present invention discloses a dynamic composite distortion simulator and a working method thereof in the technical field of aero-engine air inlet and engine matching. The simulator comprises a process air inlet, a spoiler column, a rotary valve and a high-speed motor. The spoiler column is installed inside the process air inlet, and the rotary valve is installed in the cavity of the spoiler column. One end of the rotary valve extending out of the spoiler column is connected to the high-speed motor. Dynamic distortion vortices are formed by utilizing the flow around the column and the rotary valve, and parameters such as the intensity and scale frequency of the vortex change with changes in the motor speed and the angle of attack and height of the spoiler column. This can effectively simulate the dynamic distortion flow field structure generated in the air inlet due to various reasons, overcome the shortcomings of existing simulators in simulating dynamic distortion, form a flow structure similar to the distortion flow field at the outlet of the real air inlet, so as to more realistically simulate the air inlet matching process, significantly improve the experimental efficiency and accuracy, and do not require an external air source.

The invention belongs to the technical field of aerodynamic parameter identification, and relates to an intelligent fitting optimization method for aerodynamic parameters of a fixed-wing aircraft based on direct lift force. According to the invention, a BP network is used to carry out fitting of a lift coefficient and an aerodynamic derivative of the lift coefficient to an attack angle on noiseless data. The input of the BP network comprises an attack angle, an elevator, a flap, a speed and a pitch angle speed, and is input into the BP network to obtain a prediction result when no noise influence exists, a loss function value is calculated, a gradient is calculated according to the loss function value, and then network parameters are updated according to a gradient descent algorithm to reduce the loss function value. When the loss function value is small, the difference between the prediction result of the network and the actual result is small, that is, the lift coefficient identification value output by the network is accurate in the BP network. The aerodynamic derivative obtained through calculation can be applied to a controller and has good control quality.

The invention relates to the technical field of unmanned aerial vehicles, and discloses an unmanned aerial vehicle attitude control method and system based on flight aerodynamic stability, and the method comprises the steps: constructing a flight control model, collecting a system attitude angle, setting an ideal attitude angle, calculating an attitude angle error, setting a proportion parameter, an integral parameter and a differential parameter, generating an attitude control signal, and setting collection time. The method comprises the steps of continuously collecting flight wind speed and flight angle of attack, setting aerodynamic resistance parameters and aerodynamic moment parameters, calculating aerodynamic moment disturbance, obtaining the surface area of the unmanned aerial vehicle, calculating aerodynamic disturbance, setting a proportion upper limit, an integral upper limit and a differential upper limit, and dynamically adjusting to obtain adjustment proportion parameters, adjustment integral parameters and adjustment differential parameters. And generating an update control signal based on the adjustment proportion parameter, the adjustment integral parameter and the adjustment differential parameter to complete unmanned aerial vehicle attitude control. The precision of unmanned aerial vehicle attitude control can be improved.

The invention relates to a multi-aircraft cooperative guidance method based on an MAPPO algorithm, and the method comprises the steps: building a multi-aircraft cooperative guidance model which comprises an aircraft dynamics model, a process constraint model, a terminal constraint model, and a heeling angle amplitude constraint model; when the remaining flight range is larger than or equal to 500 km, a cooperative guidance law based on improved prediction correction is designed, and an aircraft attack angle and heeling angle guidance instruction is solved, and when the remaining flight range is smaller than 500 km, the cooperative guidance law based on the MAPPO algorithm is switched to, and the aircraft attack angle and heeling angle guidance instruction is solved. The method can accurately arrive at the middle-end shift change point in a relatively large time difference, and smoothly realizes a time cooperation task. The method has high robustness for state disturbance, and has a larger time cooperation range and a better time cooperation guidance capability.

The invention discloses an incremental model prediction control method and device for variant aircraft information, a medium and a product, and relates to the field of aircraft control, and the method comprises the steps: building a nonlinear model of a variant aircraft based on the variant aircraft information; the information of the variant aircraft comprises speed, angle of attack, pitch angle, height, thrust, resistance, lift force, pitch moment, mass and rotational inertia; classifying and simplifying the nonlinear model of the morphing aircraft according to a time scale separation principle to obtain an incremental equation; based on the incremental equation and an MPC framework, constructing an incremental model prediction controller under input and state constraints; and completing predictive control of information of the variant aircraft based on the incremental model predictive controller. The real-time control precision of the morphing aircraft can be improved while the calculation amount is reduced and the calculation efficiency is improved.

The invention belongs to the technical field of aircraft simulation, and particularly relates to an aircraft longitudinal maneuvering load simulation method and device. The method comprises the following steps: S1, carrying out 1g level flight balancing on the aircraft, and determining a balancing attack angle of the aircraft and a balancing elevator skewness of the aircraft; s2, determining the aerodynamic force of the aircraft; s3, dynamic response solving is carried out on aerodynamic force of the aircraft, and the aerodynamic attack angle, normal overload, the flight speed and the pitch angle speed are obtained; s4, acquiring a first elevator skewness calculated by the stability augmentation system according to the pneumatic attack angle and the normal overload, and superposing a second elevator skewness given by the elevator instruction to obtain an actually required elevator skewness; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angle speed, the aerodynamic angle of attack and the actually required elevator skewness until the simulation is finished. The effect of the stability augmentation system in the process of determining the longitudinal maneuvering load of the aircraft is considered, the simulation process is simple, and the precision is high.

The intelligent bionic ornithopter with the body and the self-adaptive attack angle adjusting joints comprises an ornithopter body, wings and an empennage, the wings and the empennage are arranged on the ornithopter body, driving motors are symmetrically arranged on the flapping portions of the wings, and each driving motor is provided with a limiting joint; the limiting joint comprises an outer ring, an inner ring and an elastic cushion block, the outer ring is composed of a plane end and first protruding parts, the plane end and the driving motor are connected and rotate synchronously, the first protruding parts are located on the two sides of the plane end and provided with penetrating holes, the inner ring is a cylinder, and the inner ring is arranged between the first protruding parts and swings with the middle of the plane end as the center. The inner ring and the penetrating hole are matched and jointly sleeved and fixedly connected with the edge of the wing, a second protruding part is arranged on the inner ring, and elastic cushion blocks are arranged at the upper end and the lower end of the second protruding part. The invention aims to provide the bionic ornithopter which can fully simulate flapping wing flapping actions of flying organisms and can adaptively adjust the attack angle in the flapping process so as to improve the flapping lift force and reduce the resistance.

The invention belongs to the technical field of aircraft simulation, and relates to a flight simulation method and device under gust load considering a downwash effect. The method comprises the following steps: S1, carrying out 1g level flight balancing on the aircraft, and determining a balancing attack angle of the aircraft and a balancing elevator skewness of the aircraft; s2, superposing the aerodynamic coefficient increment brought by the downwash angle of attack to the aerodynamic coefficient of the aircraft body, and determining the aerodynamic force of the aircraft; s3, dynamic response solving is conducted on aerodynamic force of the aircraft, the flight speed is obtained, the corrected speed of the aircraft is determined after the flight speed is superposed with the wind speed of the vertical gust field, and the aerodynamic attack angle and the flight Mach number are determined according to the corrected speed; and S4, re-determining the aerodynamic force of the aircraft according to the flight Mach number, the pitch angle speed, the aerodynamic angle of attack, the given wind speed of the vertical gust field and the elevator skewness, and circulating the steps until the simulation is finished. The method is simple in simulation process and high in precision, and meets the requirements of engineering practice.

Some implementations relate to methods, systems and computer readable media to provide an aerodynamic force model for real-time distributed physics simulation. According to one aspect, a computer-implemented method includes receiving a description of a mechanism that includes physically coupled geometric assemblies within a virtual experience, with each geometric assembly defining a surface mesh. The description includes motion data of the mechanism. The method further includes, for each geometric assembly, identifying exposed surface areas of the surface mesh. The method further includes evaluating an aerodynamic force model based on the exposed surface areas and the motion data of the mechanism, where the aerodynamic force model includes a pressure coefficient that varies based on an angle of attack and based on windward and leeward facing surfaces. The method further includes integrating the aerodynamic force models into a physics simulation to refine the motion data of the mechanism.

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 belongs to the technical field of aircraft simulation, and particularly relates to an aircraft maneuvering load simulation method and device considering a downwash effect. The method comprises the following steps: S1, carrying out 1g level flight balancing on the aircraft, and determining a balancing attack angle of the aircraft and a balancing elevator skewness of the aircraft; s2, superposing the aerodynamic coefficient increment brought by the downwash angle of attack to the aerodynamic coefficient of the aircraft body, and determining the aerodynamic force of the aircraft; s3, dynamic response solving is carried out on aerodynamic force of the aircraft, and the aerodynamic attack angle, the flight speed and the pitch angle speed are obtained; s4, calculating elevator skewness at the current moment according to elevator instruction input; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angle speed, the aerodynamic angle of attack and the elevator skewness until the simulation is finished. According to the method, the effect of the downwash influence in the aircraft maneuvering load determination process is reasonably considered, the simulation process is simple, the precision is high, and the engineering practice requirement is met.

The retractable lift force type wind-assisted propulsion device comprises a sail assembly, a retractable device and a control system, and the sail assembly can rotate relative to a ship body; the sail assembly comprises a main wing sail, a first aileron sail, a second aileron sail and a lower end plate; wherein all the sails are wing-shaped sails, and the main wing sails are elliptical wing-shaped sails; each sail is rotatably connected with the lower end plate; and the rotating axes of the sails are parallel to each other and are not coplanar. Through the arrangement of the multi-body wing-shaped sail, the multi-body wing-shaped sail can adapt to different wind directions to change the attack angle, and the wind power boosting effect is improved; through the arrangement of the folding and unfolding device, the sail can be freely folded and unfolded in different working states.