92 results about "Flight test data" patented technology

The invention discloses a large-angle oblique imaging aero-camera targeted target positioning method. In accordance with the problems that the shooting distance of a large-angle oblique imaging aero-camera is long, and the operating distance of laser distance measuring equipment is limited, and the invention provides a direct targeted target positioning algorithm which is independent of the distance measuring equipment. According to the position and the attitude information of a carrying aerial vehicle, measured by the carrying aerial vehicle POS ( position and orientation system) and frame angular position information measured by a position encoder in the aero-camera; a homogeneous coordinate conversion method is used for solving the direction of a target under a geodetic coordinate system; and an earth ellipsoid model and a digital elevation model are used for determining latitude and longitude information of the target point. According to the large-angle oblique imaging aero-camera targeted target positioning method disclosed by the invention, flight test data is used for verifying the validity of the target positioning algorithm; and when the flight height is 18000m, and the rolling angle of a shooting frame is smaller than 63 degrees, the target positioning circular probable error is smaller than 70m, so that engineering actual requirements can be met.

The invention belongs to the technical field of embedded air data systems for aerospace vehicles, provides a monitoring and processing method for pressure faults of an embedded air data system, and aims to solve the problem of system final fault caused in case of pressure hole blockage, pipeline leakage, pressure sensor faults, acquisition and processing circuit faults, signal transmission faults and the like occurred in an FADS, or the problem of parameter operation divergence which finally causes impacts on flight control. A correlation matrix among pressure spots is determined by using a wind tunnel test and flight test data according to the distribution of pressure holes, and the correlation matrix is pre-assembled in the FADS; fault location is performed through a finally obtained pressure value, the pressure value of a fault point is estimated by using a least squares curve fitting method, and the pressured value obtained through evaluation is used for participating follow-up operation. The embedded air data system pressure fault monitoring and processing method provided by the invention can effectively process pressure hole blockage, pipeline leakage, pressure sensor faults, processing circuit faults, signal transmission faults, pressure value fault and the like, can provide reasonable atmospheric parameters to a mobile system, and can effectively improve the practicability of an embedded air data system.

The invention discloses a flight test and ground simulation aerodynamic force data comprehensive modeling method, which is used for establishing a full-airspace and full-speed-domain aerodynamic forcemathematical model of an aircraft according to more than 10 flight test data samples and aerodynamic force ground simulation data in a flight envelope range of the aircraft. The method comprises thefollowing steps: firstly, performing aerodynamic force calculation ground simulation within a flight envelope range of an aircraft, obtaining aerodynamic force simulation data in a full airspace and afull speed domain, and converting the aerodynamic force simulation data into an aerodynamic force simulation sample; And performing a flight test, obtaining more than 10 times of flight test state measurement data, and obtaining an aerodynamic force identification data sample of the flight test according to an aerodynamic force identification method. And finally, taking the aerodynamic force identification data sample and the aerodynamic force simulation sample together as a total sample, and establishing an aircraft full-airspace and full-speed-domain aerodynamic mathematical model integrating flight test and ground simulation data through machine learning by utilizing an artificial intelligence-based BP neural network method.

The invention belongs to the field of unmanned aerial vehicle flight control, and relates to an unmanned aerial vehicle shallow grey model correction method, particularly to a method for correcting atilting rotor wing unmanned aerial vehicle nonlinear shallow grey model by combining mechanism modeling and identification modeling. The method comprises the steps: providing a nonlinear light grey model of a tilting rotor wing unmanned aerial vehicle through mechanism analysis, and after linearization, determining an identification model structure according to identification requirements; designing an identification scheme including designing an input signal and selecting a data sampling frequency and a data length, and preprocessing acquired flight test data; based on the flight test data, utilizing CIFER software to obtain a linear model of the tilting rotor wing unmanned aerial vehicle in each mode; and finally correcting the non-linear model by utilizing an identification result, andif necessary, continuously iterating the process until the linearized parameters of the non-linear model are consistent with the structure and the identification model. By means of the method, the accurate nonlinear model of the tilting rotor wing unmanned aerial vehicle can be obtained.

The invention discloses a flight test data recording device and method. The device comprises a master control circuit and a storage circuit, wherein the master control circuit comprises an external electrical interface circuit, a master control end power supply and a clock management circuit, a gigabit Ethernet interface circuit, an RS-485 interface circuit, a master control end central processing circuit, a configuration storage circuit and a storage stack interface circuit; the external electrical interface circuit is respectively connected with the master control end power supply and clock management circuit, the gigabit Ethernet interface circuit and the RS-485 interface circuit; the gigabit Ethernet interface circuit, the RS-485 interface circuit, the configuration storage circuit and the storage stack interface circuit are respectively connected with the master control end central processing circuit; and the storage circuit comprises at least two storage stacks with the same structures, and the two storage stacks reserve each other. By the flight test data recording device and method, the requirement of simultaneously recording a plurality of load transmitting data can be met, and high efficiency of data recording is guaranteed.

The invention discloses a comprehensive processing system for experimental flight-test data. The system can solve the problems of large quantity, various sorts and irregular formats of the experimental flight-test data to be processed, and can meet various kinds of processing requirements. A client platform provides a man-machine interaction interface, the man-machine interaction interface supplies a fixed function, and a selective function is stored in a client algorithm UI unit in an algorithm form and allows the client platform to call; a server-side algorithm service platform integrates a system service module for achieving system service and a background calculation module for achieving various algorithms required by the experimental flight-test data, calls corresponding modules according to instructions of the client platform, and operates the called corresponding module by combining the storage content of a server-side disc array. The client platform cooperates with the server-side algorithm service platform to achieve system service and algorithm achievement.

The invention relates to a transition experiment data expansion technology, and discloses a cross flow transition experimental data expansion technology based on a stability method. The method comprises the steps: employing stability analysis means with reliable theoretical support to analyze existing wind tunnel experiment and flight test data; according to incoming flow parameters or physical property parameters such as incoming flow Reynolds number, attack angle and surface roughness which have significant influence on the cross flow transition position, changing the parameters in a range suitable for a stability theory, and obtaining a theoretical solution of the cross flow transition position, so continuation of experimental data is realized. Compared with a flight test and a wind tunnel test, the method has the characteristics of low cost, easiness in implementation and high reliability, is used for supplementing and expanding experimental data, and makes up the defects of high cost, long period and high difficulty of direct data supplementation at present.

Systems and methods for navigating an unmanned aerial vehicle are provided. One example aspect of the present disclosure is directed to a method for modeling engine health. The method includes receiving, by one or more processors, engine data. The method includes receiving, by the one or more processors, flight test data. The method includes generating, by the one or more processors, one or more coefficients for a power assistance check (PAC) based on the engine ATP data and the received flight test data using a machine learning technique. The method includes transmitting, by the one or more processors, the one or more coefficients for the PAC to a vehicle, wherein the vehicle uses the one or more coefficients in the PAC to predict engine health.

The invention provides a method for measuring the transition process characteristics of an aircraft boundary layer in a hypersonic flight test. The method comprises the following steps of 1), carryingout the simulation calculation of a flow field around an aircraft, and obtaining the surface flow characteristics of the aircraft, 2) determining an area suitable for boundary layer transition process measurement on the surface of the aircraft,3) carrying out simulation calculation on the surface heat flow and the structural thermal response of the aircraft, obtaining the calculation results of the surface heat flow and the structural temperature of the aircraft along the whole flight profile, and carrying out model selection on a heat flow sensor and a temperature sensor,4) evaluating feasibility of sensor installation in the transition measurement area, 5) determining the positions of sensors, the number of sensor measuring points and the distance between the sensor measuring points according to the analysis result of the surface flow and transition characteristics of the aircraft,6) analyzing flight test data obtained by the transition process measurement schemes determined in thesteps 1)-5), and drawing a curve of surface heat flow or temperature changing along the flow direction, and 7) analyzing the change rule of the curve of heat flow or temperature changing along the flow direction, and determining the transition process characteristics of the boundary layer.

The invention discloses an all-intelligent aircraft flight test error analysis system. The system comprises a test data management module, a flight test data analysis module, an intelligent fault diagnosis module and a human-computer interaction module. The test data management module supports batch import and unified management of test data. The flight test data analysis module performs data processing and feature extraction on the test data. The intelligent fault diagnosis module conducts intelligent reasoning and rule matching and diagnoses system faults according to symptom characteristicsobtained by the data analysis module and a knowledge rule of a knowledge base. The human-computer interaction module can control the whole process, and dynamically display the data analysis and faultpositioning processes and results of the system in real time. The system has the advantages that the data analysis results can be quickly given, system faults are accurately positioned, the test datais imported in batches, the test data is managed in a unified mode, the diagnosis process is automatically carried out, the diagnosis results are dynamically displayed in real time, the diagnosis process is autonomous and controllable, and the openness is good.

The invention discloses an association analysis method for high speed aircraft pneumatic heat flight test data, which comprises steps of analyzing and fitting the three-dimensional streamline relation based on the three-dimensional streamline relation of the heat flow between the first typical part and the second typical part of the aircraft to obtain an association simple form of he heat flow between the first typical part and the second typical part. According to the association simple form, the pneumatic heat data of different typical parts are subjected to association analysis. With the method, the utilization efficiency of the pneumatic heat data of the typical parts is improved, the testing cost is reduced, and the testing efficiency is improved.

A method and system for dynamically selecting and retrieving real-time data from a time / space partitioned flight computer system located on board an aircraft. The system used to implement this method consists of an on-board memory read-out device attached to a serial data bus capable of communicating with the flight computer and / or a ground-based memory read-out device capable of communicating with the flight computer via radio telemetry. This data consists of any valid memory location within the flight computer, including but not limited to flight data, aircraft status, and flight computer state.

The invention discloses on-line detection equipment and an on-line detection method for a fly-by-wire flight control system, and belongs to the field of aeronautical engineering testing. The on-line detection equipment is characterized by consisting of a signal transmission and disconnection device and flight testing interface equipment, wherein the signal transmission disconnection device is used for completing transmission and disconnection of related crosslinking signals of a fly-by-wire flight control computer; the signal transmission disconnection device is of a modular combined structure and consists of a plurality of unit modules which are relatively independent in structure and relatively complete in function; each independent unit module can complete transmission and disconnection of similar or approximate signals; the flight testing interface equipment consists of an operation controller, a flight testing data bus interface plate, an onboard data bus interface plate, an analog quantity signal interface plate, a switching value signal interface plate, a digital value signal interface plate, a tester bus or a computer bus and detection application software, and is used for extracting, analyzing and displaying and output parameter information of the fly-by-wire flight control computer in various tests on the fly-by-wire flight control computer or the fly-by-wire flight control system.

The invention discloses a quasi-static undercarriage dynamics model construction method, and belongs to the technical field of aircraft simulation. The method includes: step 1, inputting aircraft state parameters, and internally constructing main undercarriage-system parameters; step 2, constructing a tire coordinate system, a half-wheel-axis coordinate system, a shock-absorbing pillar coordinate system and an aircraft body coordinate system, and establishing conversion matrices of mutual conversion of all the coordinate system; step 3, calculating whether aircraft wheels are in contact with ground and whether shock-absorbing pillars are compressed, and calculating corresponding tire pressure; step 4, using the iterative initial values to calculate stress of each shock-absorbing pillar, stress of each wheel axis and stress of each tire until difference values between two frames thereof meet a threshold value requirement; step 5, obtaining received resultant force and a resultant moment of a body axis system at a gravity center of the aircraft by solving through coordinate conversion; and step 6, outputting calculation process parameters according to needs. According to the method, a solving process of undercarriage system parameters is simplified while aircraft ground-motion simulation precision is guaranteed, running efficiency of a simulation program is improved, and dependence of the model parameters on flight test data is reduced.

The invention relates to a flight and ground correlation method of natural transition test data of a boundary layer. A test similarity criterion and a data fitting formula are adopted as design principles of the method. According to an implementation mode of the method, firstly, analysis of the differences between a flight and the ground is carried out, and main influence factors influencing transition are concluded; secondly, through fitting of existing ground tests and flight test data, modification formulas of the influence factors is constructed; finally, according to the modification formulas, extrapolation is conducted on ground test data to obtain a transition Reynolds number under a flight condition or correlation between the ground tests and the flight test data is achieved. The method has the advantages that since an emphasis is placed on the test similarity criterion, the influence quantities to be modified are reduced; through classification of the differences between the flight and the ground, the modification formula of each influence quantity in a decoupling form is constructed. Since the modification formula of each influence factor is obtained by fitting a great amount of test data, the reliability is high.

The invention discloses a small unmanned aerial vehicle control object modeling method based on test data. According to the method, through transformation of control signals, an established control object model can be applied to the process with small variation and can also be applied to the process with great variation, and an established model system is high in identification precision and good in model prediction capacity. According to the method, by the utilization of test flight data, high-cost wind tunnel tests can be avoided, and therefore early-stage development cost is lowered. According to the method, through changing of mathematical features of input signals (namely the control signals) in the test flight test data, the difference between the different input signals is amplified, therefore, the error range of an estimation model and a true model is narrowed down, and the accuracy and the prediction capacity of the estimation model are improved.

The present disclosure provides systems and methods that enable flexible access of internal data of an avionics system (e.g., a Flight Management System). Aspects of the present disclosure enable enhanced, flexible and robust recording of flight test data, testing, debugging, and analyzing in-service problems. One example system includes a Configuration Tool that generates a Loadable Configuration File and a Configuration File based on user selections, structure of data in Data Stores, platform information, processor information, and Receiving Component requirements; a Data Access Component that subscribes to data in the Data Stores based on the Loadable Configuration File and outputs the data; and a Receiving Component receives the data, and decodes the data based at least in part on the Configuration File. Aspects of the present disclosure provide the ability to change which data gets output by the avionics system without the need to recompile and recertify the avionics system software.

The invention discloses a multi-source heterogeneous flight test data processing method and a multi-source heterogeneous flight test data processing system, and relates to the technical field of aeronautic testing. According to the multi-source heterogeneous flight test data processing method, data source information, database server information and engineering data sending are defined by utilizing a preprocessing module, and defined data is sent to a real-time fusion processing module; and the real-time fusion processing module receives data packets downloaded by multiple paths of standard measurement and control communication links of an unmanned aerial vehicle and PCM data acquired, remotely tested and downloaded by an airborne testing system, forms uniform fused data after carrying out fusion processing on the received data packet and PCM data, and sends the uniform fused data to a monitoring terminal. The multi-source heterogeneous flight test data processing system is used for implementing the multi-source heterogeneous flight test data processing method. The multi-source heterogeneous flight test data processing method and the multi-source heterogeneous flight test data processing system which are disclosed by the invention have the advantages that various communication link formats of data and PCM data can be simultaneously processed; flexibility and real-time performance of the system are reinforced; and the data processing part is completely universalized, so that equipment implementation cost is reduced.

The invention belongs to the technical field of aviation test flight, and particularly relates to a flutter data processing framework based on a flight test. The flutter data processing framework based on the flight test comprises a platform management system, a basic function module, an auxiliary function module and a bottom layer supporting module; the platform management system is used for resource dispatch and operational process management of a platform, the basic function module is used for processing data of a general flutter flight test, the auxiliary function module is used for analyzing flutter test flight data which are specific or are not provided with a currently established clear data processing technology system, the bottom layer supporting module is used for providing a basic environment supporting normal operation of the platform and basic software shared by each function module of the platform. The flutter flight test data processing platform concept and a top layer design framework are primarily raised, a plurality of layers such as a flight test mode, an incentive mode, the observation signal property, a processing mode, a processing task, a processing method and the like are projected in one unified framework, and therefore tradition, modernization, existence and development are coordinated in a higher level.

The invention provides an aerodynamic force multi-source data fusion and identification method and system and a medium. The method comprises the following steps: 1, building a globally related aerodynamic force mathematical model; 2, according to the existing calculation data, the wind tunnel test data and the flight test data, calculating standard deviations of multiple source data; 3, establishing an overdetermined equation set by taking existing calculation data, wind tunnel test data and flight test data as data samples of the aerodynamic mathematical model; and step 4, solving the overdetermined equation set by adopting a weighted least square method, obtaining the weight by calculating the standard deviation, solving the coefficient in the aerodynamic mathematical model, and completing the establishment of the multi-source data fusion aerodynamic mathematical model. According to the invention, the problems of different reliability of data from different sources and error decomposition to sample points in aerodynamic mathematical model identification and correction are solved, an overdetermined equation set can be comprehensively solved from data from three sources of calculation, a wind tunnel test and a flight test, and identification and model correction are completed in one step.

The invention provides an aerodynamic modeling and full-aircraft flutter analysis method based on wind tunnel test response, and belongs to the technical field of aircraft design. According to the method, an unsteady aerodynamic model is identified through wind tunnel test data, the unsteady aerodynamic model is obtained, an unsteady aerodynamic coefficient of a scaling wind tunnel model is converted to the full-aircraft model through the similarity law, and flutter analysis is made on the full-aircraft model by using a converted aerodynamic model so as to obtain a flutter boundary. Compared with flutter analysis by using flight test data in the prior art, aerodynamic modeling is performed by using the response data of the wind tunnel test model, and since sensors are relatively easy to embed in the wind tunnel test, more effective data can be obtained based on the test data of the wind tunnel, the aerodynamic model is more accurate, and the reliability of the aerodynamic model is improved. Thus, the flutter analysis result is more credible, especially aerodynamic modeling of a transonic interval.

The invention discloses an aircraft modeling method based on a variable measurement number maximum information criterion. The aircraft modeling method is used for solving the technical problems of poor proving correctness of pneumatic models and parameters provided by flight tests due to the maximum information criterion when observation vectors of existing different flight tests are not the same. The technical scheme includes that during aircraft model modeling and model validation, different flight test factor amendment modeling and model validation criteria of an aircraft are acquired by analyzing the influence of different observation vectors to the maximum information content criterion, and scalar model selection and testing discriminant is acquired through U-D decomposition of measurement variance estimation Rj and Rj+1. By the method, aerodynamic force models and moment models of the aircraft can be established directly conveniently according to flight test data, the technical problem that the aerodynamic force models are established and tested incorrectly according to different flight test data due to the fact that influence of different observations to the maximum information criterion is not considered directly.

The invention discloses a hybrid laminar flow wing air suction energy loss engineering calculation method, which comprises the following steps: according to CFD calculation data, wind tunnel test data and flight test data analysis, respectively fitting a wing air suction device energy loss factor calculation function taking a Rayleigh number REN, a lift coefficient CL, a wing front edge sweepback angle lambda w and a cruise Mach number as independent variables; calculating an energy loss factor generated by airfoil mass by using an eN airfoil boundary layer-by-layer transition position calculation method considering air suction by taking the upper and lower airfoil layer-by-layer transition positions obtained by the X21 flight test as a reference, and finally, calculating the total area of the energy loss factor generated by the airfoil mass by summing the area integral of the energy loss factor of each airfoil section in the airfoil span direction, and obtaining the energy loss factor of the wing air suction device of the whole aircraft. The method provided by the invention is suitable for demonstrating the power demand of the air suction device at the front edge of the hybrid laminar flow wing with complex configuration, and provides technical support for the thrust-weight ratio design of a hybrid laminar flow control technology aircraft and the optimal design of the wing configuration based on the air suction system pipeline design of power distribution optimization.

The invention discloses an aircraft modeling method based on a maximum information-reliability online identification criterion. The technical problems that the proving correctness of aerodynamic models and parameters verified by a flight test is poor caused by a simplified model identification criterion are solved. The technical scheme includes that data lengths are considered and reliable parameters are introduced in a maximum information criterion, an online modeling criterion is corrected, the division of U-D of Rj and Rj+1 is evaluated for the measurement variance, and a scalar online model selection and verified discriminant is obtained. The method has the advantages that aerodynamic models and moment models of an aircraft can be established on line directly in accordance with flight test data, and the technical problems of aircraft high attack angle model establishment and verification according to flight tests caused by a simplified model determination criterion are solved.

The invention discloses an aircraft aerodynamic identification correction method and a medium, and belongs to the technical field of aerodynamic design. The method aims at a nonlinear aerodynamic mathematical model of an aircraft with complex aerodynamic characteristics. Based on the fact of flight test state diversity, flight test data and wind tunnel test data are jointly used as modeling samplepoints, a corrected aerodynamic mathematical model is established by solving an overdetermined equation set, identification can be carried out along with the development stage, the flight test samplepoints are continuously increased, and the precision of the aerodynamic mathematical model is improved.

The invention provides a flight test method for measuring the stability margin of a multi-input multi-output flight control system. The steps include: (1) inputting a frequency sweep excitation signal to the flight control system to be tested; (2) obtaining the open-loop transfer function of the flight control system Matrix (3) Calculate the minimum singular value σ(P) and the minimum eigenvalue λ(P) of the system hysteresis matrix P=I+G; (4) Calculate the guaranteed amplitude margin and guaranteed phase according to the MIMO system stability margin formula Margin; (5) In order to reduce conservatism, replace σ(P) with λ(P) and substitute it into the calculation formula of MIMO system stability margin to obtain amplitude margin and phase margin. The invention proposes a new index for evaluating the MIMO system stability margin, a new index for evaluating the MIMO system, and proposes a set of flight test methods for obtaining the MIMO system stability margin, including test principles, test actions and flight test data processing procedures. Also proposed two conservative methods to reduce the stability margin of the MIMO system, improved the stability margin test flight technology of the single-input single-output (SISO) flight control system, and provided a basis for the flight test research of the multiple-input multiple-output (MIMO) flight control system and the new generation of fighters. The development provided technical support.

The invention discloses a method for correcting cruise fuel consumption characteristics based on circling test flight data. The method comprises the following steps of: a, selecting a height point, aspeed point and a flight state, and calculating the hourly fuel consumption during stable circling at different roll angles; b, carrying out oil consumption characteristic correction; c, calculating the hourly oil consumption of the engine through the function relationship; e, correcting to the target weight, which comprises circling oil consumption calculation and level flight oil consumption calculation; correcting the oil consumption per hour to a target weight level flight state through the performance data; the method comprises the following steps: calculating the hourly oil consumption of an engine, the hourly oil consumption of the engine and the oil consumption in a level flight state during different rolling stable circling; the flight test data can be analyzed and calculated at different speeds, cost-reduced oil consumption data can be reliably and accurately measured, Meanwhile, multiple groups of effective flight test data can be obtained by selecting multiple groups of stable circling with different roll angles and changing the angle of the stable circling.

The invention belongs to the field of helicopter structure fatigue design, and particularly relates to a static load and dynamic load separation method for a rotating component flight test. The methodcomprises the following steps: synchronously acquiring a helicopter blade azimuth angle Pj and a rotating part load; standardizing the helicopter blade azimuth angle Pj to obtain a standard blade azimuth angle signal; calculating according to the standard blade azimuth angle signal to obtain a fundamental frequency period of the standard blade azimuth angle signal and a data point number of the rotating part load in a single period; determining a static load in the load of the rotating part in the ith fundamental frequency period; determining a dynamic load in the load of the rotating part inthe ith fundamental frequency period; and determining a rotating component flight test static load and a rotating component flight test dynamic load according to the flight test state of the helicopter. According to the method, helicopter flight test data can be separated based on helicopter blade azimuth signals, and static and dynamic loads of part flight test are obtained.

The invention belongs to the technical field of aircraft flight test power device testing and particularly relates to a method for directly measuring the flight tension of a propeller. The method, byperforming strain gauge modification and load measurement on a mounting joint / pull rod system, directly obtains the tension of the air propeller according to actually measured flight test data. The method comprises modifying a strain gauge on the mounting joint / pull rod system, performing a tension calibration test on the ground, obtaining a propeller tension-mounting joint / pull rod system straincalibration equation, and finally obtaining the tension of the propeller based on the actually measured flight test data under a flight condition. Compared with a model-based indirect calculation method, the tension direct measurement method does not require a large number of calculation models, and is low in sensor placement, production, installation and calibration cost. Thus, the method is simpler and less expensive. In addition, due to the high dynamic response capability of the strain gauge, the direct method is suitable for dynamic tension measurement and is also suitable for real-timemonitoring.