59 results about "Flight experiment" patented technology

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 present invention discloses a method, device and computer-readable storage medium for calibrating the drag coefficient of a multi-rotor. The method includes: performing flight tests at multiple set airspeeds along a straight and level flight route in a windless environment and a windy environment, and recording flight control logs according to a measurement sequence composed of airspeed, inclination angle and acceleration during the flight tests; combining each flight control log, respectively selecting corresponding measurement combinations according to the acceleration, endurance and acceleration corresponding stages; taking the least square solution obtained from a preset number of the measurement combinations as the calibration result, and applying the calibration result to a preset real-time airspeed estimation algorithm to obtain an estimation result. The present invention realizes a low-cost experimental calibration algorithm, efficiently and accurately calculates the multi-rotor drag coefficient function, and effectively verifies the accuracy of the calibration result and the estimation algorithm.

The present invention relates to a new generation of information technology, and discloses a method and a system for controlling the flight area of an aircraft to improve the adaptability to the test environment. The method includes: calculating the transformation matrix between the associated coordinate systems; during the flight of the aircraft, in real time calculating the coordinates of the current point and the corresponding safety zone boundary in the launch coordinate system according to the transformation matrix between the associated coordinate systems, the longitude and latitude information of the current point, and the azimuth data of the flight test reference trajectory; calculating the longitude and latitude corresponding to the safety zone boundary according to the coordinates of the safety zone boundary in the launch coordinate system, the angular velocity of the earth's rotation in the launch coordinate system, and the radius vector of the origin of the launch coordinate system; judging whether the current point is within the safety zone range according to the longitude and latitude of the current point and the longitude and latitude corresponding to the safety zone boundary. If the current point is within the safety zone, calculating the normal flight control program; otherwise, calculating the safety control program to make the aircraft dive and land.

The invention discloses an intelligent self-adaptive control bionic flapping wing flight experiment data acquisition system and method, and belongs to the technical field of bionic aircraft testing and intelligent control, and the system comprises a hardware system integration module, a wind tunnel environment simulation module, a three-degree-of-freedom motion control module, a flexible sensor array module and a reinforcement learning interaction module. According to the intelligent self-adaptive control bionic flapping wing flight experiment data acquisition system, the upper computer is used as a control center, and a multifunctional visual interface is integrated; the multi-fan wind tunnel is linked with the three-degree-of-freedom platform to reproduce an airflow-attitude coupling relationship and improve the suitability of the two; data-training-optimization integration is realized, and system compatibility is improved.

The application discloses a transition measurement method based on a plasma probe, wherein the plasma probe is arranged at different positions of an undisturbed part of an aircraft surface; the probe arrangement position is subjected to simulation analysis of aerodynamic heat and a temperature field in a flight process, and a probe material is selected according to a simulation result; three-dimensional plasma flow field simulation results are obtained for typical states in a transition window interval of a flight profile; three-dimensional plasma flow field electron density data are extracted, and a range of the plasma probe is determined according to the three-dimensional plasma flow field electron density data; a probe circuit is designed, and a plasma probe measurement device is obtained; the plasma probe measurement device is used to obtain a measured current-voltage curve and a saturated ion current of the probe, and electron density data are calculated according to the measured current-voltage curve and the saturated ion current, that is, a transition measurement signal of the probe plasma in the flight test, and a step change time of the transition measurement signal is taken as a transition occurrence time of measurement; according to the measured transition occurrence time, combined with telemetry data of the flight test, a transition measurement result is obtained, and validity of the transition measurement result is confirmed.

The invention discloses an envelope extension type flight test state point risk level assessment method, which comprises the steps of designing a test flight state point of an envelope extension test flight subject according to information such as an initial safety value and a limit value of a to-be-extended envelope (speed, overload, angle of attack and the like) involved in the test flight subject; establishing a risk measurement value calculation method of a single envelope extension test flight state point in a flight test; based on the risk measurement value of the single envelope extension test flight state point, establishing a calculation method related to the risk measurement value of a plurality of envelope extension test flight state points in the flight test; designing a risk grading scale, and visually representing the risk grade of the test flight state point through the grade and the color tape; obtaining the risk level of the test flight state point through the position of the test flight state point risk measurement value in the risk analysis scale; and the envelope extension type flight test flight test state point optimization method based on the flight test state point risk level is formed. According to the embodiment of the invention, the problem of evaluation of the risk level of the test flight state point in envelope extension type test flight design is solved, and test flight design personnel can carry out test flight design optimization according to the visual risk level, so that the test flight progress, the test flight cost and the test flight risk are balanced.

The invention relates to a high-frequency-response high-temperature heat flux sensor for a flight test, which is characterized in that the type of a thermocouple is selected according to the maximum value of the surface temperature of an aircraft in the flight process, a heat transfer element is of a cylindrical structure, the end part is polished and ground, and two mounting through holes are symmetrically formed in the same distance from the center line of the cylindrical structure; positive and negative electrodes of the thermocouple are processed into wires with corresponding diameters and are fixed in the two mounting through holes of the heat transfer element, and one ends of the two electrodes are flush with the end surface of the heat transfer element; the end face of the heat transfer element provided with the thermocouple is sequentially plated with a copper film and a silicon dioxide film in a magnetron sputtering mode, and the high-frequency response characteristic of the heat flow sensor is achieved. The coated heat transfer element is fixed in the high-temperature alloy packaging shell, the tail parts of the positive electrode and the negative electrode of the thermocouple are welded with the compensating leads with the same polarity, and a cavity in the tail part of the high-temperature alloy packaging shell is filled with high-temperature glue. The sensor has the advantages of being small in size, resistant to temperature and high in frequency response.

The present application belongs to the field of flight test technology, and relates to a kind of indication recording system test flight method suitable for troposphere manned airship.The method comprises: hovering 360 in the three time periods of sunrise, strongest sunlight and sunset respectively;In the above hovering process, it is observed whether the display data displayed by the display is easy to see under different angle illumination conditions;If there is an abnormal working condition in the above process, optimization improvement suggestions are proposed for it, and then the above steps are repeated until the qualified criterion is met.

The invention provides a multi-dynamic infrared simulation method, system and equipment based on a spacecraft test. The multi-dynamic infrared simulation system comprises a control system, a horizontal sliding rail system, a rotating mechanism system, a lifting mechanism system, a target heat source and a communication network system. The control system controls the horizontal sliding rail system to achieve horizontal movement of a target, drives the rotating mechanism system to complete composite rotation and adjusts the height of the lifting mechanism system so as to keep position synchronization with a target heat source, and during the period, the communication network system guarantees bidirectional data transmission and communication between all the systems and an aircraft. According to the invention, the problem of simulating the characteristics and motion characteristics of the infrared target under the real flight condition of the external field airspace is solved, the index requirement of high-precision tracking is met, the visual display effect of the ground suspension / flight test is obviously enhanced, and the stability and reliability of target control are effectively improved.

The application provides a five-degree-of-freedom virtual flight test device based on motion compensation, wherein a test model of the device can complete three-axis rotation motion with the help of a three-degree-of-freedom hinge, the test model and an abdominal support strut can move up and down along the axial direction of a three-pronged suspension support to complete the release of the heave degree of freedom, and the test model and the three-pronged suspension support can realize lateral undamped motion under the support of a pneumatic suspension bearing, so as to achieve the purpose of releasing the lateral translation degree of freedom. A parallel six-degree-of-freedom mechanism base at the bottom is responsible for motion compensation of the lateral translation, heave, roll and pitch degrees of freedom, and based on the above structure, the device can realize the five-degree-of-freedom coupled motion test capability except the forward and backward motion. The realization of the above functions has important significance for improving the restoration capability of the virtual flight test of a wind tunnel to the real dynamic characteristics of an aircraft model.

The present invention relates to the technical field of wind tunnel dynamic simulation, and discloses a dynamic similarity simulation device for virtual flight tests in a wind tunnel, including: a first connecting member rotatably arranged on a gas source device along a first direction; a second connecting member rotatably arranged on the first connecting member along a second direction; a third connecting member connected to the second connecting member; a fourth connecting member rotatably arranged on the third connecting member along a third direction; the first connecting member has a plurality of gas source inlets, and the fourth connecting member has a plurality of gas source outlets. Airflow enters from the gas source inlets, flows along the first connecting member, the second connecting member, the third connecting member and the fourth connecting member, and is discharged from the gas source outlets; wherein, along the flowing direction of the airflow, the cross-sectional area of the inner cavity of the second connecting member is larger than the cross-sectional area of the inner cavity of the first connecting member. By the above method, the present invention solves the technical problem that it is difficult to simultaneously simulate the thrust-to-weight ratio and the pressure drop ratio on a scaled model in the wind tunnel simulation of aircraft tests in the related art.

The invention belongs to the technical field of flight test real-time monitoring and data processing, and discloses a flight test telemetering monitoring picture integrated design method, which comprises the following steps: establishing a front-end and rear-end separation mode, forming that a user is responsible for front-end interface development and parameter association, and a developer is responsible for rear-end parameter calculation and transmission; a digital management model of the monitoring parameters is constructed, and full-process management of the monitoring parameters is achieved; researching monitoring picture development common components and display modes thereof, and designing a component digital model; and a four-level management mechanism of design uploading, auditing and publishing, data confirmation and online approval is constructed. According to the invention, the problem that the description of the real-time monitoring demand of the existing flight test is not accurate enough is solved, and the implementation efficiency of the monitoring task is improved.

The invention belongs to the technical field of flight-launch matching design, and particularly relates to a nested nacelle for a scaling flight test and a design method. The nested nacelle is characterized in that the jacketed nacelle is obtained by directly scaling an actual nacelle of an aircraft scheme; the embedded nacelle is embedded in the outer sleeve nacelle, an airflow flowing channel is formed between the embedded nacelle and the outer sleeve nacelle, an embedded nacelle air inlet channel is formed in the embedded nacelle, and the embedded nacelle comprises an embedded nacelle front section and an embedded nacelle rear section. An opening for communicating the airflow flowing channel with the spray pipe cooling channel is formed between the embedded nacelle front section and the embedded nacelle rear section; the engine is installed in the embedded nacelle and connected with an air inlet channel of the embedded nacelle, and a spray pipe is arranged at the tail of the engine. The method is used for the scaling flight test of the wing hanging type large aircraft, and the flight-launch matching problem faced by the scaling flight test can be solved.

This application belongs to the field of aircraft design technology, and particularly relates to a static pressure source position error correction method. Based on current numerical simulation and flight test state points, a static pressure source position error correction analysis is performed on the theoretical configuration, obtaining the static pressure source pressure coefficient Cp, the flight test static pressure source pressure coefficient Cp0, and the measured static pressure source pressure coefficient Cp1. The static pressure source position error correction data is optimized to obtain the static pressure source position error correction relationship Cp′ during actual flight of the theoretical configuration, and then the static pressure source position error is corrected. The application of configuration measurement, CFD simulation, and flight test data is achieved. The static pressure source position error is corrected as a function of Mach number, angle of attack, and altitude. This method not only meets the pressure altitude requirements of CCAR Part 25, but also meets the altitude measurement system error requirements of RVSM operations in AC‑21‑13. This method can be used for atmospheric data correction on transport aircraft and can be used as a reference for atmospheric data correction on other types of aircraft.

The application provides a multi-target identification method based on real-time data analysis. In some flight test tasks, there are several flight targets flying simultaneously in the air, and there are many measurement devices on the ground simultaneously tracking and measuring the flight targets. The task of some precise measurement devices only needs to capture, track and measure one flight target. Since the multiple flight targets have intersection in time and space, the precise measurement devices inevitably have identification and tracking confusion. The reasons for the identification confusion of the flight targets tracked and locked by the measurement devices under the background of multi-target flight are discussed. The internal relationship between the data of the same target obtained by the measurement devices is discussed by using mathematical methods. The process of the spatial coordinates of the flight target changing with time is further revealed, and the properties of the trajectory, data continuity and smooth curve are established. A mathematical model for the measurement devices to avoid identification confusion of the flight targets in a short time is established.

The invention provides an application system and an application method of a flight simulation training data packet, and belongs to the technical field of aviation simulation training. The system comprises a data acquisition module, an engineering data packet processing module, a training data packet generation module, an error verification module and an adaptive output module. According to the method, an engineering data packet is constructed by arranging aircraft design data and wind tunnel data, a training data packet meeting the standard of a CAAC full-motion D-level simulator is formed after flight experiment correction, and the error is controlled within 5%. The system realizes localization of flight simulation training data packets, breaks through foreign technical monopoly, guarantees national defense safety, improves training fidelity and economical efficiency, and can be widely applied to the field of civil aviation and military aircraft simulation training.

The invention relates to the technical field of flight tests, and discloses a test flight planning method and system for verifying a communication distance index of an airborne ultra-short wave radio station, and the method comprises the steps: obtaining an optimized test flight route through the implementation of spatial geographical environment analysis, communication coverage and communication link simulation analysis; the test route planning of airborne ultra-short wave radio air-ground communication distance index verification in a limited airspace and complex terrain environment is realized, the problems of low flight test success rate and difficult organization and implementation of airborne ultra-short wave radio air-ground communication distance index verification in the limited airspace and complex terrain environment are solved, the flight test quality can be improved, and the flight test efficiency is improved. And the flight cost consumption is reduced.

The invention relates to a flight test platform structure modal parameter prediction method based on a graph neural network, and the method comprises the steps: obtaining a geometric parameter, a modal parameter, a stiffness matrix, a mass matrix and the like of a flight test platform structure through Abaqus finite element simulation calculation, converting the obtained parameters into graph data, taking the graph data as training data of the graph neural network, and carrying out the prediction of the modal parameters of the flight test platform structure. And then constructing a modal parameter prediction model based on the graph neural network, and predicting the modal parameters of the new structure by using the trained prediction model. Based on topological structure information of the graph neural network and in combination with node features of the finite element mesh model, efficient prediction of the new structure modal parameters of the flight test platform is realized, and the high calculation cost for modal analysis of the new structure can be saved.

The invention provides a flight test task dynamic optimization method based on a model, and the method comprises the steps: S10, carrying out the statistical analysis of the execution data of a test flight task, and constructing an uncertainty mathematical model according to the influence factors which affect the plan change of the test flight task; the uncertainty mathematical model comprises a task uncertainty model, a task execution time uncertainty model and a resource uncertainty model, S20, according to the uncertainty mathematical model, constructing a proactive test flight task optimization model to obtain a proactive test flight task planning scheme, and performing task scheduling on the proactive test flight task. The proactive test flight task optimization model comprises a proactive target model and a proactive constraint model; and S30, according to the uncertainty mathematical model, a reactivity scheduling test flight task optimization model is constructed, a reactivity test flight task plan is obtained, and the reactivity scheduling test flight task optimization model comprises a reactivity target model and a reactivity constraint model.

This invention discloses a method and system for planning multi-aircraft flight test subjects considering meteorological window constraints, belonging to the field of aircraft flight test management technology. The method includes: constructing a multi-aircraft flight test mission planning model incorporating immediate logic constraints, resource exclusivity constraints, and meteorological time window constraints; solving the model using an improved gold mining optimization algorithm (I-GRO), representing the joint decision of "mission priority-aircraft allocation" through a two-layer hybrid encoding strategy; introducing a dynamic topology sorting and meteorological window alignment mechanism in the decoding stage to map continuous position vectors into feasible scheduling schemes that satisfy temporal logic; and in the search stage, adaptively adjusting the execution probabilities of migration, gold mining, and cooperation operators through a dynamic strategy selection mechanism, and introducing a linearly decaying elite cooperation and perturbation mechanism. This invention can effectively solve the search stagnation problem under fragmented feasible domains, significantly shorten the total flight test duration, and improve the robustness of the planning scheme.

The invention provides a blunt cone virtual flight test system and method with a variable head deflection angle and a variable mass center, and belongs to the technical field of wind tunnel dynamic tests. The system comprises a variable elbow mechanism, a variable centroid mechanism and a virtual flight mechanism. The variable elbow mechanism is matched with an electric push cylinder through an upper push rod and a lower push rod, so that the switching of a blunt cone head part between zero-position locking and rapid deflection is realized; the variable centroid mechanism drives a balancing weight to move axially through an electric push cylinder, and the whole centroid position is changed. The virtual flight mechanism supports the blunt cone through the supporting rod and the low-damping bearing set to achieve pitching motion. According to the invention, head deflection and centroid change can be independently or cooperatively controlled, the pneumatic / motion coupling effect of the intelligent deformable aircraft in the dynamic process is truly simulated, the limitation that the shape and the centroid are fixed in a traditional wind tunnel test is solved, and an effective means is provided for verification of a comprehensive attitude control strategy.

The invention discloses a flight test system for realizing online verification and debugging of a fault diagnosis algorithm, and relates to the field of aircraft fault monitoring. The flight test system selects a fixed-wing unmanned aerial vehicle and a secondarily developed flight control algorithm according to a typical fault generation mechanism of an execution mechanism of the fixed-wing unmanned aerial vehicle; the method comprises the following steps of: simulating typical faults of two actuating mechanisms, namely control surface jamming and control surface loosening and floating, and performing online real-time monitoring on multi-dimensional flight state information of the fixed-wing unmanned aerial vehicle by combining an airborne computer carrying an online fault diagnosis algorithm and utilizing flight state data sent by a flight controller in real time; in cooperation with an independently developed fault diagnosis information monitoring and parameter adjusting ground station, online verification and debugging can be carried out on a fault diagnosis algorithm aiming at the fault of the executing mechanism of the fixed-wing unmanned aerial vehicle.

The application discloses a kind of aerodynamic thermal ablation matching coupling simulation design method, S1, for the aircraft shape of un-ablation, select flight orbit typical state point to calculate incoming flow condition, simulate the thermal environment of different types of thermal protection material lap joint part;S2, carry out ablation simulation analysis to determine the ablation amount and ablation profile of the two sides of lap joint part, calculate the initial ablation matching step height h0 of two sides material;S3, estimate the ablation deterioration factor, estimate the ablation matching step height H after ablation deterioration effect based on h0;Ablation step shape modeling is carried out here, obtain the aircraft shape with local ablation step shape;S4, regenerate aircraft surface and outflow field grid, carry out three-dimensional aerodynamic thermal simulation design, give the thermal environment of material lap joint part;S5, carry out ablation simulation analysis again, calculate the secondary ablation matching step height h1;S6, compare h1 with H, confirm the simulation end after estimating the ablation amount in S3 envelope flight test ablation amount.

The invention discloses a high-altitude rocket rarefied flow field reconstruction method and device based on flight test data. The method comprises the following steps: constructing a high-altitude rocket rarefied flow field reconstruction model and a Boltzman-BGK equation non-equilibrium term prediction model based on a Shakhov equilibrium state distribution function; the unbalanced term prediction model comprises a first sub-network and a second sub-network which are arranged in parallel and respectively output an unbalanced term of a density distribution function and an unbalanced term of a temperature distribution function; a loss function used in the training process comprises a first loss function constructed based on a residual term of a Boltzmann-BGK equation, and a second loss function constructed based on a residual term of the Boltzmann-BGK equation. The second loss function is constructed on the basis of a residual term of physical constraint represented by an unbalanced term; the third loss function is constructed on the basis of flight test data and a predicted value of flow field data output by the high-altitude rocket rarefied flow field reconstruction model; and inputting a to-be-predicted flow field coordinate on the high-altitude rocket into the trained high-altitude rocket rarefied flow field reconstruction model to obtain a predicted value of the flow field data. According to the invention, the flow field reconstruction precision can be improved.

The invention discloses a transition measurement method based on plasma probes. The plasma probes are arranged at different positions of a non-interference part on the surface of an aircraft; carrying out flight process aerodynamic heat and temperature field simulation analysis on the probe arrangement position, and selecting a probe material according to a simulation result; a three-dimensional plasma flow field simulation result is obtained according to the typical state in the flight profile transition window interval; extracting electron density data of the three-dimensional plasma flow field, determining the measuring range of a plasma probe according to the electron density data, and carrying out probe circuit design to obtain a plasma probe measuring device; a plasma probe measuring device is used for obtaining a probe actually-measured current-voltage curve and a saturated ion current, electron density data, namely a flight test probe plasma measuring signal, is calculated according to the current-voltage curve and the saturated ion current, and the step change moment of the flight test probe plasma measuring signal serves as the measured transition occurrence moment; and obtaining a transition measurement result according to the measured transition occurrence moment in combination with flight test telemetry data, and confirming the validity of the transition measurement result.

The application relates to the technical field of flight test mission evaluation, in particular to a flight test mission efficiency evaluation method, system, device and medium, which comprises the following steps: constructing a flight test concept meta-model library; generating a computable task model based on the requirements of a target flight test mission; constructing a multi-level index system based on the computable task model; establishing an index mapping relationship table of performance indexes and test data sources; obtaining test data, extracting specific parameters of each performance index from the test data based on the index mapping relationship table, performing standardized assignment, data feature analysis and weight distribution, and obtaining a weight coefficient system of the multi-level index system; and based on the standardized assignment result of the performance index and the weight coefficient system, calculating the evaluation value of the efficiency index through index aggregation to obtain the efficiency evaluation result of the flight test mission. The application can improve the comprehensiveness, accuracy and adaptability of flight test efficiency evaluation.

The invention provides a carrier rocket safety margin design method and system based on probability calculation and a processor. According to the invention, through probability statistical calculation, a forward numerical design method for directly solving the safety margin based on the known depletion shutdown / injection probability is realized. The method does not depend on accumulation of historical design experience, the safety margin can be directly obtained through modeling calculation, and therefore the method is especially suitable for research and development of brand new rockets lacking reference models. According to the method, by adopting the forward design, the selection of the traditional'safety margin cluster 'and a large amount of random targeting iteration are fundamentally avoided, so that the design period is remarkably shortened, and the design efficiency is improved. The method provided by the invention has been successfully applied to the safety margin design of various models of carrier rockets, and multiple simulation results and flight test data prove the effectiveness of the method.

This invention belongs to the field of flight test data processing technology and discloses a method and system for quality assessment of multi-source flight test data based on physical-temporal constraints. The method includes the following steps: data preprocessing and benchmark calibration; physical-temporal joint verification; multi-source spatiotemporal alignment; scenario adaptive evaluation; and comprehensive quality scoring and feedback optimization. This invention solves the problems of insufficient correlation and inconsistent spatiotemporal benchmarks in existing flight test multi-source data, achieving full-dimensional and high-precision quality assessment of multi-source flight test data.

The invention relates to a new-generation information technology, and discloses a method and a system for controlling a flight area of an aircraft, so as to improve the adaptability to a test environment. The method comprises the following steps: calculating a conversion matrix between associated coordinate systems; in the flight process of the aircraft, the coordinates of the current point and the corresponding safe area boundary in the emission coordinate system are solved in real time according to the conversion matrix between the associated coordinate systems, the latitude and longitude information of the current point and the azimuth angle data of the flight test reference trajectory; according to the coordinates of the emission coordinate system corresponding to the safe area boundary, the earth rotation angular velocity of the emission coordinate system and the original point vector diameter of the emission coordinate system, the longitude and latitude corresponding to the safe area boundary are calculated; according to the latitude and longitude of the current point and the latitude and longitude corresponding to the safe area boundary, judging whether the current point is in the safe area range, and if the current point is in the safe area, resolving a normal flight control program; otherwise, resolving the safety control program to enable the aircraft to dive to land.