439 results about "Flight test" patented technology

The invention relates to the technical field of unmanned aerial vehicle sensing, and discloses an unmanned aerial vehicle complex vector wind field sensing method and system, and the method comprises the steps: respectively recording the original data of a wind field sensing device in a wind tunnel simulation wind field test and the original data of an unmanned aerial vehicle carrying the wind field sensing device for a flight test, and constructing a feature data set; using the constructed feature data set to train an LSTM model, so that the LSTM model can predict wind field information at the current moment in real time according to historical wind field data, installing a wind field sensing device array on the unmanned aerial vehicle, collecting wind field information data in real time, and using the trained LSTM model to predict the wind field information in real time; and performing data fusion on prediction results of the plurality of sensors by adopting a Kalman filtering algorithm, and calculating wind field information such as three-dimensional wind speed, wind-induced air pressure change and turbulence. According to the method, real-time efficient prediction of the complex wind field environment in the air of the unmanned aerial vehicle is realized, and an effective means is provided for improving wind field detection and reliable operation of various unmanned aerial vehicles in the complex wind field environment.

The invention belongs to the technical field of hypersonic flight vehicle test, and relates to a hypersonic flight vehicle simulation flight test system based on a multi-modal large model, which comprises a data acquisition and fusion system and a control system. Through a data acquisition and fusion system and a cross-modal attention mechanism, efficient alignment and feature representation of time sequence data, environmental data and a historical case library acquired by a ground test system are realized, and the data acquisition and analysis precision is remarkably improved. In combination with a hardware system, effectiveness of test parameters is ensured, and the aircraft design and verification period is shortened. The control system dynamically adjusts test parameters through a natural language instruction, the model generates a structured control parameter vector through semantic understanding, the performance and stability of the aircraft control system are improved, the test operation complexity is reduced, and the user experience is improved.

The invention belongs to the technical field of flight test photoelectric testing, and particularly relates to a cloud particle parameter measuring device and method based on digital holographic imaging. The device comprises a device pneumatic shell (100), a signal transmitting end arm (300), a signal receiving end arm (200), a measuring module and a data processing module, the device pneumatic shell (100) is integrally provided with a columnar pneumatic appearance, the signal transmitting end arm (300) and the signal receiving end arm (200) are arranged at the same end of the device pneumatic shell in parallel, and the opposite sides of the signal transmitting end arm (300) and the signal receiving end arm (200) are provided with a transparent transmitting light window (301) and a transparent receiving light window (201) respectively. The measuring module is arranged in the pneumatic shell of the device, the measuring module generates an emission light beam, the emission light beam is emitted through the signal emission end arm and then received through the signal receiving end arm to form holographic image information, and the holographic image information is transmitted to the data processing module; and the data processing module analyzes the image and calculates to obtain a particle median volume diameter MVD and an atmosphere liquid water content LWC. Accurate measurement of geometric quantities such as three-dimensional morphology, particle size distribution and the like, motion quantities such as three-dimensional speed, rotation characteristics and the like, and statistical quantity information such as median volume diameter (median volume diameter), liquid water content (liquid water content) and the like of ice crystal / water drop particles in multi-scale complex-form cloud can be realized; and autonomous control of the core test equipment for the icing flight test of the aircraft is realized.

The invention provides an aircraft test flight optimization method and system and a storage medium, and belongs to the technical field of aircraft test flight, and the method comprises the steps: S1, generating an initial scheme; s2, test flight execution and data acquisition; s3, dynamic safety monitoring and early warning; s4, dynamically adjusting parameters and generating a control instruction; s5, comprehensive performance evaluation; s6, performing multi-objective optimization; s7, verifying the difference degree; and S8, performing closed-loop iterative optimization. According to the method, objective feedback optimization is adopted, and adverse effects caused by subjective feedback optimization are eliminated, so that the optimization efficiency and the optimization effect are 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 relates to the field of aviation flight tests, and discloses a flight test airborne data unified recording and processing method, system and device and a computer readable storage medium. The method comprises the following steps: receiving an external time code time service signal, and generating a synchronization timestamp for performing time synchronization on each module in a system; collecting multi-source data streams from a network interface, an analog quantity PCM interface, a video interface and a 1553B interface; performing protocol conversion and format standardization processing on the multi-source data stream to obtain a data stream with a uniform format; based on a data fusion algorithm, carrying out fusion processing on the standardized data streams to generate a fusion data sequence in a unified format; performing grading processing on the fused data sequence according to the importance of the data; performing adaptive compression processing on the graded data sequence to generate a compressed storage data sequence; and transmitting the compressed data storage sequence to a ground station through a wireless transmission protocol. According to the invention, the flight test efficiency is effectively improved.

The invention discloses a fairing pulsating pressure scaling wind tunnel test method capable of reducing space-ground difference. The method comprises the following steps: determining initial parameters; based on the determined initial parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test, arranging pulsating pressure measuring points, and carrying out transonic speed state fairing pulsating pressure environment measurement to obtain a measurement result; performing parameter optimization according to a comparison result of a measurement result and a standard parameter under a real flight test condition to obtain an optimal parameter of manual transition; and based on the optimal parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test to obtain test results of the configuration in different free incoming flow Mach number, attack angle and sideslip angle states. According to the method, the free incoming flow local shock wave-boundary layer interference state and the boundary layer separation state of the ground wind tunnel test and the flight test are ensured to be closer through a manual transition mode, and the ground test prediction precision of the carrier rocket fairing pulsating pressure environment is improved.

The utility model provides a rope-driven flexible robot for a helicopter test flight test. The rope-driven flexible robot comprises a test fixing table, a flexible periodic pitch changing device and a high-precision total pitch adjusting device. Wherein the test fixing table, the flexible periodic pitch changing device and the high-precision total pitch adjusting device are all fixed in a helicopter cockpit, and the flexible periodic pitch changing device is used for simulating human hands to execute helicopter blade periodic pitch changing operation; the high-precision collective pitch adjusting device is used for simulating a human hand to execute collective pitch changing operation; an output shaft gear of the high-precision collective pitch adjusting device is connected with a collective pitch variable rod, a flexible belt is connected with a driving motor and a variable-size gear disc, and the output shaft gear and the variable-size gear disc are coaxially fixed to be used for outputting variable rotating speed. According to the rope-driven flexible robot for the helicopter test flight test, the test precision and reliability are improved, and meanwhile the personnel safety in the test process is guaranteed.

The invention belongs to the technical field of aircraft test data processing, and relates to an aircraft ground test and flight test consistency evaluation method and device. The method comprises the following steps: acquiring strain data of each first measuring point of a ground test, and acquiring strain data of each second measuring point of a flight test; fusing the finite element simulation data with the strain data of each first measuring point of the ground test, calculating the strain data of each second measuring point, reconstructing a ground test strain field containing the strain data of each first measuring point and each second measuring point, and fusing the finite element simulation data with the strain data of each second measuring point of the flight test to obtain a ground test strain field containing the strain data of each first measuring point and each second measuring point. Calculating strain data of each first measuring point, and reconstructing a flight test strain field containing the strain data of each first measuring point and each second measuring point; and for the reconstructed ground test strain field and the flight test strain field, counting the proportion of the test points of which the strain data difference value of the same test point is smaller than a set value, and taking the proportion as a consistency index of the ground test and the flight test. The utilization rate of test data is improved.

The invention discloses an attitude-elasticity-pneumatic coupling flutter control method of a flying wing layout aircraft, and relates to the technical fields of mechanics, aeroelasticity, control and the like. The control method comprises the steps that a Lagrangian equation under a general body shafting is synthesized, a full-quantity coupling dynamic model is established, and a disturbance motion dynamic equation used for flight control system design is established through small disturbance linearization; correcting the disturbance motion kinetic equation based on a system identification flight test; designing an attitude-elastic-pneumatic coupling flutter suppression controller based on a traditional or modern control theory, and embedding the attitude-elastic-pneumatic coupling flutter suppression controller into a flight control system; the effectiveness of the flutter suppression controller is verified through a flight test. The problems that existing numerical analysis and wind tunnel experiments cannot accurately reflect the complex coupling effect in flight, flutter cannot be effectively restrained in flight, and the flight envelope cannot be expanded are solved, and the flight safety and stability of the flying wing layout aircraft can be effectively improved.

The invention belongs to the technical field of flight testing, and discloses an aircraft testing method and system for simulating a flight environment, and the method comprises the steps: obtaining design parameters of an aircraft and environment data of a real scene, and constructing a simulation test model; in the simulation test model, generating a simulation scene based on a preset test task; collecting multi-source simulation data of a simulation scene where the virtual object is located; inputting the multi-source simulation data into a pre-constructed state perception prediction model, and predicting the state perception of the virtual object to obtain a state prediction result; the state prediction result is input into a pre-constructed test control model, a navigation control instruction is output, and the virtual object sails along a test track when responding to the navigation control instruction; the test trajectory is generated based on a preset trajectory planning algorithm; and recording test data of the virtual object in the process of sailing along the test track to obtain a simulation test result. The method has the advantages of low test cost, few limited conditions, high controllability and the like.

The invention provides an unmanned aerial vehicle laboratory testing method and device, and belongs to the technical field of unmanned aerial vehicle testing. The scheme comprises the following steps: determining a global reference point based on the position of a laboratory, and further determining a conversion relation between a local northeast high coordinate system and a global coordinate system and a plurality of fixed reference points; according to the fixed reference point, sequentially executing stand setting and starting of an automatic measurement tracker in a laboratory, and further obtaining real-time coordinates of the to-be-tested unmanned aerial vehicle by using the automatic measurement tracker; the GNSS signal generator generates a simulation navigation signal according to the real-time coordinates of the to-be-tested unmanned aerial vehicle and records a simulation track; the to-be-tested unmanned aerial vehicle performs indoor flight test based on the simulation navigation signal; and comparing the simulation track with a flight track recorded by a flight control system of the to-be-tested unmanned aerial vehicle so as to output a laboratory test result. The scheme does not depend on hardware modification of the unmanned aerial vehicle, the implementation difficulty is reduced, the implementation cost is saved, and controllable testing of medium and large unmanned aerial vehicles in the indoor environment is facilitated.

The invention discloses a rocket aircraft aerodynamic performance prediction method based on heaven and earth test and simulation data, and relates to the technical field of aerodynamic design. The method mainly comprises the following steps: obtaining a large number of aerodynamic performance parameters by using a non-viscous rapid design theory to construct a low-precision flow field training data set; performing heaven-earth correlation error correction on ground wind tunnel test data through a flight test; performing simulation error correction on a large number of aerodynamic performance parameters obtained by CFD numerical simulation through test data to obtain a high-precision training data set; and performing data fusion on the high-precision data set and the low-precision data set, and constructing a pneumatic prediction model. The invention provides the method for rapidly predicting the aerodynamic performance of the rocket aircraft, and an efficient and feasible solution is provided for rapid prediction of the aerodynamic performance.

The embodiment of the invention provides an equivalent model design method and system for an unmanned aerial vehicle carrying mission load flight test, and the method comprises the steps: obtaining mission load demand parameters and test condition parameters, the demand parameters comprise a long side A1 and a short side A2 of a rectangular airspace, A1 > = A2, and a demand speed v0; the test parameters comprise a test airspace long side B1 and a test airspace short side B2, B1 is greater than or equal to B2, the test machine minimum speed v1 and the test machine maximum speed v2, and v2 is greater than or equal to v1; when A1 < = B1, A2 < = B2 and v1 < = v0 < = v2 are not satisfied, calculating an equivalent coefficient p, and if the airspace condition A1 < = B1 and A2 < = B2 are not satisfied, calculating and determining the equivalent coefficient p; if the airspace condition A1 < = B1 and A2 < = B2 are satisfied, but the speed v1 < = v0 < = v2 is not satisfied, determining an equivalent coefficient p; based on the reference point coordinates and the equivalent coefficient p, longitude and latitude are achieved through geocentric coordinate system conversion, and horizontal navigational speed mapping is carried out. According to the technical scheme, by simulating the flight profile required by the mission load, the problem that the unmanned aerial vehicle cannot fly out of the flight profile required by the mission load in a flight test carrying the mission load can be solved.

The invention belongs to the technical field of airworthiness verification of aircrafts, and particularly relates to a test flight method for verifying longitudinal static stability of a civil helicopter instrument. Comprising the following steps: S1, determining a relation between a longitudinal control lever force and a longitudinal control position of a control lever; s2, verifying the longitudinal joystick force gradient of the acceleration section of the helicopter through a flight test; s3, verifying the longitudinal joystick force gradient of the helicopter deceleration section through a flight test; s4, verifying the balancing speed recovery characteristic of the acceleration section of the helicopter through a flight test; and S5, verifying the balancing speed recovery characteristic of the deceleration section of the helicopter through a flight test.

The invention relates to a high-speed aircraft transition identification method based on a flight test high-frequency vibration signal. The method comprises the following steps: selecting high-quality measurement point data suitable for analysis from flight test high-frequency vibration data as a vibration signal to be analyzed; performing time-frequency analysis on the to-be-analyzed vibration signal, outputting a time-frequency diagram, and representing the change condition of energy in a frequency domain by using a centroid frequency and an envelope difference of the centroid frequency; carrying out preliminary identification on the transition occurrence moment; decomposing a to-be-analyzed vibration signal into a plurality of intrinsic mode functions with different center frequencies through variational mode decomposition; quantizing the change condition of the energy of each modal component along with time; and aiming at the change condition of the energy of each modal component along with the time, the transition is identified according to the energy of the mode excited by the transition, and the energy sudden increase point is the moment when the transition begins to occur. According to the method, the transition occurrence time can be effectively identified under the condition that no hole is formed in the surface of the aircraft.

The invention belongs to the technical field of wind tunnel test and flight control crossing, and discloses a control device and a control method of a multivariable hypersonic variable configuration aircraft model. The control device comprises a main control computer; the main control computer is used for project management and development, data storage and display, flight and control simulation; the system also comprises a real-time simulation machine which is used for simulating the real-time environment of the flight control system, and completing the data collection of the sensor, and the generation and transmission of a control instruction. The control method comprises the steps of establishing a digital simulation model; verifying the digital simulation model; compiling a control law program and testing; performing a virtual flight test; and analyzing and processing data. According to the control device and the control method of the multivariable hypersonic-speed variable-configuration aircraft model, simulation of a multivariable hypersonic-speed variable-configuration aircraft digital model and a control law is realized, and a control algorithm operated in a real-time simulation machine can be used for carrying out a wind tunnel pneumatic, motion and control coupling virtual flight test; the engineering practical value is realized.

The invention discloses a flight test platform parametric modeling method based on a CST method, and relates to the field of aerospace aircraft modeling design, and the method comprises the steps: S1, carrying out the function division according to the appearance structure of a flight test platform aircraft, and dividing each divided part into a plurality of different section shapes according to the appearance structure; s2, setting a corresponding control point on each section shape; s3, through a five-parameter mode line design method, adjacent circumferential control points on each section shape are connected one by one to generate a transverse section curve, then curve control points are correspondingly connected one by one based on adjacent transverse sections to generate a longitudinal curve, finally a curved surface is obtained, and parametric modeling of each part is completed; and S4, assembling each part which independently completes parametric modeling, and completing final parametric modeling of the aircraft. According to the flight test platform parametric modeling method based on the CST method, the comprehensive performance, the reliability and the design iteration efficiency are remarkably improved.

The invention relates to a tilt rotorcraft semi-physical simulation system and method, and relates to the technical field of wind tunnel flight test semi-physical simulation. Based on the initialized system with the test parameters set, six-component balance load, wind speed and attitude data are collected in real time, and the collected key physical data are filtered and then sent to the flight simulation subsystem; the flight simulation subsystem carries out calculation based on a flight dynamics model, a rotor wing power model, a control instruction preprocessing model and the received data and outputs an actuation instruction; the actuator receives an actuation instruction and drives the aircraft model, and the airborne sensor feeds back new attitude data to form a control closed loop; the flight simulation subsystem sends state data to the 3D visual simulation module in real time to realize synchronization of control and visual; in this way, the stable state under the combination of different tilting angles and flight speeds is recorded, and a tilting transition corridor is obtained through fitting. Therefore, simulation precision, test safety and iteration efficiency are taken into consideration, and fused semi-physical simulation is formed.

The invention relates to the field of GNSS simulator testing, and discloses an unmanned aerial vehicle formation flight testing method and device based on a GNSS simulator, and the method comprises the steps: determining a GNSS signal, corresponding to each unmanned aerial vehicle, of the GNSS simulator according to the flight parameter information of a plurality of unmanned aerial vehicles; the flight parameter information of the multiple unmanned aerial vehicles comprises the flight path, the flight height and the flight speed of each unmanned aerial vehicle and the relative positions of the multiple unmanned aerial vehicles; the GNSS simulator and each unmanned aerial vehicle are connected, time synchronization is carried out on the GNSS signals of all the unmanned aerial vehicles, and the GNSS simulator sends the GNSS signals to the corresponding unmanned aerial vehicle; and the multiple unmanned aerial vehicles simulate formation flight according to a preset formation control algorithm and the GNSS signal corresponding to each unmanned aerial vehicle, obtain flight state information of each unmanned aerial vehicle, and analyze and debug the flight state information of each unmanned aerial vehicle. According to the invention, flight testing can be efficiently and accurately carried out on the unmanned aerial vehicle formation.

The invention belongs to the technical field of flight test, and discloses a cockpit vision field evaluation test flight method suitable for a troposphere manned airship, which comprises the following steps: selecting an airship test condition; the airship climbs to a preset field height and keeps level flight; a preset flight test action is completed; and commenting by a flight tester, and filling in a cockpit vision field commenting card. According to the method, the blank of the flight test method for evaluating the visual field of the cockpit of the troposphere manned airship with the vector propulsion system is filled, and the flight test efficiency and the flight test quality of the manned airship are improved.

The invention discloses a multi-modal unmanned aerial vehicle flight control system and a switching method, and relates to the technical field of flight control systems, the multi-modal unmanned aerial vehicle flight control system comprises a parameter generation module for performing real flight test on an unmanned aerial vehicle to form a parameter combination and a modal mapping table; the mode switching module is used for collecting current multi-dimensional data of the unmanned aerial vehicle for preprocessing, generating an undetermined control mode in combination with the mode mapping table for verification, obtaining a target mode and executing mode switching; the instruction processing module is used for executing smooth operation to obtain an initial control quantity during mode switching, calculating a final control law and an inverse matrix, executing matrix multiplication based on the final control law and the inverse matrix, generating a torque instruction and taking the torque instruction as an execution instruction of motor thrust distribution or steering engine deflection in the unmanned aerial vehicle; according to the invention, the synchronism and consistency of mode switching and control law updating are realized, and the real-time performance and accuracy of the unmanned aerial vehicle during flight task execution are improved.

The invention belongs to the field of flight tests of unmanned aerial vehicles, and particularly relates to a calibration method for an airspeed system of a round-trip flight unmanned aerial vehicle with a fixed course angle. Comprising the following steps: step 1, selecting a test point in an aircraft flight envelope; step 2, the aircraft flies back and forth under the condition corresponding to the test point, and flight data of the going path and the returning path in the stable level flight stage are obtained; step 3, respectively selecting a plurality of flight data points of a going path and a returning path from the flight data, acquiring a course angle and a track angle of each flight data point, and according to the course angle and the track angle, judging whether the non-uniformity of the wind field in the flight process meets the requirement; if so, entering step 4; if not, returning to the step 2, and re-executing the test point; 4, acquiring the ground speed of each flight data point, and calculating the wind speed according to the ground speed; 5, calculating the vacuum speed in the flight process according to the wind speed and the ground speed; and 6, calibrating the air speed of the test point according to the vacuum speed.

The invention discloses a fixed-wing aircraft non-airport take-off and landing adaptability evaluation method based on test flight performance data, and belongs to the field of aviation aircraft design. Comprising typical difference analysis of a non-airport environment and an airport environment, takeoff and landing sliding tolerance analysis of the unmanned aerial vehicle in the non-airport environment and takeoff and landing lateral deviation tolerance analysis of the unmanned aerial vehicle in the non-airport environment. The non-airport take-off and landing adaptive capacity of the airplane is analyzed through take-off and landing data of previous test flights. Key differences and possible problems faced by taking off and landing in the non-airport environment are found out by comparing the test flight airport environment with the non-airport environment, and key constraints of taking off and landing in the non-airport environment are formed; taking-off and landing taxiing section flight parameters in the test flight data are extracted, conformity analysis is carried out from all constraint dimensions, and the conformity analysis comprises analysis of taxiing tolerance based on grounding point deviation and analysis of taxiing feasibility of all levels of roads based on taxiing side offset distance probability distribution; whether field cleaning work needs to be carried out or not during taking-off and landing is evaluated by combining airplane height and highway guardrail design standard comparison.

The utility model discloses an unmanned aerial vehicle hosting test device which comprises an upper clamp and a lower clamp which are used for fixing an unmanned aerial vehicle, the clamps are connected with a suspender, the upper end of the suspender is connected with the front end of a safety rope, the safety rope has a safety length, and the safety rope passes through a pulley installed on a cross beam and the rear end of the safety rope is connected with a take-up and pay-off wheel on a motor. The motor rotates to control winding and unwinding of the safety rope, the upper end of the suspender is connected with the linear displacement sensor, the linear displacement sensor monitors the moving distance and speed of the suspender, and the motor is controlled to wind and unwind the safety rope through the linear displacement sensor. The device is provided with an emergency button, when the emergency button is pressed down, the unmanned aerial vehicle ground control station receives a signal to control the unmanned aerial vehicle to stop propellers, and the motor rapidly retracts the safety rope with the safe length. The safety rope slowly lifts the unmanned aerial vehicle to a test initial position.

The invention discloses a light environment design method combining engineering and numerical values. The light environment design method comprises the following steps: carrying out numerical calculation of a three-dimensional thermochemical unbalanced flow field; selecting a typical observation window, and generating observation rays penetrating through the flow field area; interpolating the numerical calculation result of the three-dimensional thermochemical unbalanced flow field to the observation rays to obtain density and temperature data on different observation rays; calculating a high-temperature gas emission coefficient at each point on the observation ray based on a high-temperature gas visible light wave band radiation engineering calculation formula; integrating along an observation ray path to obtain a background flow field light radiation intensity value Isim of each point on the light path reaching the optical window; light environment measurement data of a previous flight test are used, a background flow field light radiation intensity simulation value Isim0 of each point on a light path reaching the optical window is obtained through the method, the simulation value Isim0 is compared with a measured value Iexp, and a light environment simulation correction factor f = Iexp / Isim0 is obtained; and multiplying the light environment simulation correction factor by a background flow field light radiation intensity value Isim of the optical window to obtain a light environment design result.

The invention relates to a closed-loop virtual test flight pre-evaluation method and a medium, and relates to the field of flight tests. The method comprises the steps of constructing a digital twin module which comprises a pneumatic sub-model, a flight control sub-model, a power sub-model and a structure sub-model, defining a tilt transition specific high-risk subject and generating a test case, executing parallel Monte Carlo simulation and hardware-in-the-loop test, performing flight safety risk identification and envelope boundary quantification, and determining a tilt transition specific high-risk subject according to the tilt transition specific high-risk subject. And generating a risk assessment report and optimizing a real aircraft test flight outline. Wherein the aerodynamic sub-model adopts a dynamic vortex grid correction method based on CFD offline calibration and online query; the flight control sub-model comprises a configurable fault injection module based on a script engine, and the configurable fault injection module is used for defining a coupling fault sequence of sequential logic and a condition triggering relation. Compared with the prior art, the accuracy and the applicability of the pre-evaluation of the virtual test flight are improved.

The present invention discloses a variable-pitch propeller design verification method and a dedicated simulation test device. The method adopts a variable-pitch propeller design verification method based on blade engineering multi-segment three-dimensional surface elements and PSP surface pressure distribution measurement technology. There is no need to use a large-scale wind tunnel to test the propeller, nor does it require a high-power engine / motor to drive the propeller to rotate. Only the surface pressure distribution of a series of three-dimensional surface elements needs to be measured in the test device proposed by the present invention. The method has the advantages of being applicable to a wide range of propellers, a relatively simple test system, low cost and high efficiency. It has more prominent advantages than full-scale wind tunnel tests and flight tests, is easier to implement, and has higher safety.

The invention discloses a coaxial double-propeller multi-rotor unmanned aerial vehicle stability control method based on aerodynamic disturbance optimization, and the method comprises the following steps: S1, collecting flight data, preprocessing the data, and generating an aerodynamic disturbance training data set; s2, based on the aerodynamic disturbance training data set, constructing an improved Fourier neural operator network model; s3, optimizing parameters of the model by adopting an improved chaos grey wolf optimization algorithm; s4, embedding the optimized model into an unmanned aerial vehicle flight control system; s5, according to the predicted aerodynamic disturbance data, a model prediction control algorithm is adopted to adjust the thrust and the attitude; and S6, verifying the stability of the control method through a simulation test and an actual flight test. The optimization method for improving the flight stability of the unmanned aerial vehicle in the complex aerodynamic environment is described in detail, and the method has the advantages of being high in calculation efficiency, accurate in aerodynamic disturbance prediction and high in control system robustness.

The invention discloses an atmospheric environment simulation method and device for aerospace flight simulation, and relates to the related field of data processing, and the method comprises the steps: carrying out the multi-stage environment construction of a flight test scheme of an aerospace vehicle according to an atmospheric data system tester, and obtaining a multi-stage flight simulation environment; performing a simulation test on the aerospace vehicle to obtain a first test sequence; performing simulation scene variation compensation to obtain a second test sequence; performing multi-dimensional risk prediction on the test scheme to obtain a multi-stage test risk prediction result; and carrying out sensitivity association to obtain a multi-order risk environment sensitive factor, and optimizing a test scheme. The technical problem that the test result is incomplete and inaccurate due to the fact that the environment simulation scene is limited and uncertain factor consideration is lacked in the adaptability test of the aerospace vehicle in different atmospheric environments in the prior art is solved, the flight state of the aerospace vehicle in the complex atmospheric environment is truly and comprehensively simulated, and the test efficiency is improved. And the comprehensiveness and the accuracy of a test result are improved.