36 results about "Flight experiment" patented technology

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 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 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.

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.

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 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 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 discloses a high-integration-level flight control system, relates to the field of wind tunnel dynamic flight tests, and solves the problems that an existing flight control system is limited in interface adaptation type, single in control mode and insufficient in attitude and power cooperative control precision under closed-loop control. Comprising a data receiving interface, a control signal output interface and a control processing module. The data receiving interface is adapted to various communication types such as 485, 232 and the like and is used for acquiring sensing data of multiple components such as inertial navigation, a GPS (Global Positioning System), a remote controller and the like; and the control signal output interface outputs corresponding signals to the control surface control unit, the electronic speed control and electric culvert unit and the like. The control processing module comprises a system self-checking module, a data analysis module and an instruction resolving module, sensing data are analyzed after self-checking is passed, the instruction resolving module can select an open loop mode or a closed loop mode to resolve an instruction and transmit the instruction, and multi-source data collection and multi-execution-unit cooperative control are achieved. The method can be used for flight control of virtual flight, free flight and outfield flight of a dynamic flight model.

The invention discloses an adapter separation performance detection device and method, the detection device comprises an air supply assembly and a test bench, and the air supply assembly is connected with the test bench through an air pipe. The adapter separation performance testing device can adapt to separation performance testing of adapters of multiple models, different structures and sizes under different launching working conditions, and the testing result of the adapter separation performance testing device is verified to be high in accuracy through a flight test. The provided adapter separation performance testing device and method can accurately, effectively, simply, conveniently and rapidly test the separation performance of adapters of different models under different launching working conditions and natural wind speeds, meet the requirements of economy, manufacturability, practicability, operability and the like, and guarantee safe, reliable and error-free launching of aircrafts.

The invention provides a unit workload airworthiness conformity verification method based on a flight test scene, and belongs to the technical field of unit workload evaluation, and the method comprises the steps: S1, building a test scene of simulated flight, sending the test scene to an MOC conformity verification model, and carrying out the parameter calibration of the MOC conformity verification model through combining with a unit workload airworthiness criterion equation; s2, recognizing the face posture of a pilot through a dual-channel input network, and correcting the airworthiness unit workload interval output by the MOC conformity verification model; s3, adjusting influence factors in the workload of the airworthiness unit to determine key indexes, acquiring the physiological index change state of the pilot by an NASA-TLX interpretable framework, and quantifying the influence weight of each key index on the workload of the pilot; and S4, outputting the airworthiness unit workload and the airworthiness integrity value to obtain the optimal unit workload which can be accepted by the pilot.

The invention provides a flapping fatigue test load spectrum compilation method for an elastic ball bowl element, and belongs to the technical field of aircraft component fatigue tests.The method comprises the steps that the total load spectrum of an unmanned aerial vehicle rotor system is determined, and according to the total load spectrum of the unmanned aerial vehicle rotor system and main bearing loads of an upper elastic ball bowl element and a lower elastic ball bowl element, a flapping fatigue test load spectrum of the upper elastic ball bowl element and the lower elastic ball bowl element is compiled; determining flight load spectrums of the upper elastic ball bowl element and the lower elastic ball bowl element; determining a test loading mode according to the flight load spectrums of the elastic upper and lower ball bowl elements; determining a flight test load spectrum of the elastic upper and lower ball bowl elements according to the flight load spectrum in combination with a test loading mode; according to an equal damage principle, determining a tension load in a flight test equal amplitude load spectrum by combining the size of the test piece; and determining the waving angle of the elastic lower ball bowl according to the rigidity of the test piece and the waving load in the flight test load spectrum, and finally obtaining the waving fatigue test constant-amplitude load spectrum of the elastic upper and lower ball bowl elements.

A kind of longitudinal large amplitude adaptive virtual flight test device and method, belong to the field specially for wind tunnel test.This application includes vertical guide rail top through upper end buffer device and spherical hinge support connection, spherical hinge support is equipped on the upper wall surface of wind tunnel through connecting seat, aircraft model is connected with vertical guide rail through pitch yaw sliding simulation device, vertical guide rail bottom is equipped with lower end buffer device, lower end buffer device bottom is equipped with steel wire rope traction seat, steel wire rope traction seat is connected with three groups of front steel wire rope, front steel wire rope is connected with rear steel wire rope through steel wire rope tensioner, rear steel wire rope is connected with the lower wall surface of wind tunnel, so that vertical guide rail lower end is in pre-tightening state.This application releases the degree of freedom of vertical direction completely on the basis of three-axis angle movement, can effectively obtain the real aerodynamic force and damping characteristics of aircraft movement, and reflects the change of height in movement process, more truly simulates actual flight condition, further plays the advantage of virtual flight test.

The invention provides a test flight action automatic identification method and device, and the method comprises the steps: obtaining test flight data which is a data sampling point set with time information; acquiring a to-be-screened action set from the test flight data by adopting a preset sliding window; the duration t of the preset identification sliding window is K times of the single pulse manipulation time t0; the duration of the action to be screened is t; the manipulated variable amplitude absolute value of the to-be-screened action is greater than a first threshold value; the peak-peak value time of the action to be screened is smaller than a second threshold value; the 0-axis crossing frequency of the to-be-screened action is not greater than a third threshold value; and obtaining a cross correlation coefficient of each to-be-screened action in the to-be-screened action set and the standard multiple pulse action, and identifying the multiple pulse excitation action in the to-be-screened action set according to the condition that the maximum value of a cross correlation coefficient vector is not less than a fourth threshold value. Multiple pulse excitation test flight actions can be automatically identified, and the automation degree, execution efficiency and readability of flight test data processing are improved.

The invention discloses a carrier rocket flight test swing nozzle physical swing angle estimation method and a carrier rocket flight test swing nozzle physical swing angle estimation system. The carrier rocket flight test swing nozzle physical swing angle estimation method comprises the following steps: acquiring servo mechanism geometric parameters; acquiring flight test data; determining the absolute length of the two-channel actuator in the flight process; estimating the sinkage of the swing center when the engine works according to the obtained flight test data; according to the estimated sinkage of the swing center when the engine works and the absolute length of the two-channel actuator, the actual physical swing angle of the swing spray pipe is determined; determining an angular displacement length estimation result according to the actual physical swing angle of the swing nozzle and the estimated sinkage of the swing center when the engine works; judging whether the angular displacement length estimation result meets the requirement or not; if the requirement is met, outputting the actual physical swing angle of the swing spray pipe; on the contrary, the sinkage of the swing center is corrected. The accuracy of the calculation result is improved, and a reliable basis can be provided for evaluation of the actual dynamic characteristics of the swing nozzle.

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

The invention relates to a flight test pneumatic identification system and method, and belongs to the technical field of aircraft test data post-processing. A data processing layer module receives flight test data and outputs flight test parameters to a core algorithm layer module; receiving reference data to form an interpolation table; the core algorithm layer module comprises a gravity compensation module and is used for calculating a real-time gravity vector and sending the real-time gravity vector to the kinematics resolving module; the attitude control force compensation unit is used for reading steering engine deflection parameters, outputting control force and torque and sending the control force and the torque to the kinematics resolving module; the kinematics resolving module calculates and outputs aerodynamic force and torque coefficients; the multi-dimensional data interpolation module is used for receiving the flight test parameters sent by the data processing layer module, reading an interpolation table in the data processing layer module and carrying out interpolation based on the interpolation table to obtain aerodynamic force and torque coefficients; and the verification output layer module performs merging according to the time sequence of the flight test measurement data to obtain a comparison curve of the test data and the reference data.

The invention provides a wind tunnel virtual flight test system with linear displacement. The system comprises a heave line motion mechanism, a single-degree-of-freedom angular motion mechanism and a virtual flight wireless rudder control system. The heave line movement mechanism drives a ball screw pair through a motor to drive a support rod to move along a linear guide rail, and uses a pulley block and a counterweight to counteract the dead weights of the model and the support rod, so that the motor only needs to overcome a pneumatic load, and the accurate and rapid control on the heave line displacement of the model is realized; the single-degree-of-freedom angular motion mechanism enables the model to perform pitching motion through a bearing pack and a rotating shaft; the virtual flight wireless rudder control system is integrated in the model, and attitude measurement, wireless data transmission and control surface closed-loop control are realized. The method breaks through the limitation that a traditional wind tunnel test only simulates angular motion, can simulate two-body separation and other complex scenes with relative linear displacement at high precision, and remarkably expands the application range and simulation capability of the wind tunnel virtual flight test.

The utility model provides a mounting structure for thrust paddle test equipment, which belongs to the technical field of flight test and test, and comprises a box body, a cover plate, a coil mounting piece, a crimping assembly and a connector, the cover plate is located above the box body, the coil installation part and the box body are installed in a tight fit mode, the crimping assembly is installed on the cover plate in an embedded mode, the connector is located on the side edge of the box body, and the box body, the cover plate and the coil installation part are symmetrically distributed with the axis of the thrust paddle as the reference center; and the test requirement of the coaxial helicopter on the thrust paddle during the flight test can be met.