664 results about "Flight test" patented technology

A system for automated fault analysis and execution of post-flight or pre-flight tests for an aircraft is provided. The system includes a maintenance test client system onboard the aircraft, and a ground-based maintenance test server system that is remote from the aircraft. The maintenance test client system is configured to collect fault data for the aircraft, and to wirelessly transmit the fault data to the ground-based maintenance test server system after the aircraft has landed. The ground-based maintenance test server system is configured to receive the fault data, to generate a test checklist that is influenced by the fault data, and to transmit the test checklist to the maintenance test client system. The maintenance test client system wirelessly receives the test checklist, and initiates execution of at least one test specified in the test checklist.

The invention provides a microminiature rotorcraft experiment platform and an application thereof. The microminiature rotorcraft experiment platform comprises an experiment platform base, a globe joint bearing, a microminiature rotorcraft, a six-dimensional force sensor, onboard circuit hardware and an upper computer, wherein the six-dimensional force sensor is connected between the experiment platform base and a stator of the globe joint bearing; and the microminiature rotorcraft is provided with the onboard circuit hardware. The experiment platform is the microminiature rotorcraft experiment platform with wide application range, the use range of the traditional rotorcraft experiment platform or three degree of freedom experimental device with a single function can be greatly expanded, the experimental requirement of the rotorcraft can be furthest satisfied. The microminiature rotorcraft experiment platform disclosed by the invention can be used for replacing the traditional practical flight test and has the advantages of simple mechanical structure and strong popularity and is easy to realize.

The invention relates to a multi-UAV simultaneous-arrival collaborative control method and belongs to the technical field of flight test, UAVs (unmanned aerial vehicle) and collaborative control. Themethod comprises a decentralized control structure relying only on local information interaction and a decentralized control strategy based on a consistency algorithm. The two are combined according to different features of path planning and speed control, and complementary advantages thereof are utilized to cope with influence of adverse factors of path errors and emergent threats and the like. The method realizes simultaneous arrival of multiple UAVs, thereby improving flexibility, robustness, and reliability of the control strategy, improving multi-UAV cooperative combat efficiency and enabling a multi-UAV combat large-scale system to play a greater role in the war.

The invention relates to a scene matching semi-physical simulation system based on a mechanical arm with six degree of freedom, consisting of hardware systems such as a simulation calculation sub-system, a mechanical arm sub-system with six degree of freedom, a photographing sub-system and a large-screen projection sub-system, and a scene reference map database, a scene matching software, a matching area selecting and a flight path planning software, and a relevant software of a mechanical arm motion model for simulating flight attitude. The actual work process of scene matching is simulated by using the mechanical arm with six degree of freedom, a camera, a large-screen projector and a mechanical arm motion model; a common PC is adopted as a simulation computer; the matching area selecting and the flight path planning software generates matching area data and orbit data; the dynamic characteristics of the system is accomplished by utilizing the mechanical arm motion model and the large-screen projector; the main errors (illumination and smoke) of the system are simulated by controlling the turning-on and turning-off of a fluorescent lamp; and the collection of real-time images, the control of the mechanical arm, the control of the projection speed, and a matching calculation are accomplished in the simulation computer, thus accomplishing the scene matching semi-physical simulation system. The system truly simulates the actual work process and the error of the scene matching, and the experimental result has significant meaning to the research and manufacture and the applications of scene matching engineering samples and can replace the flight-test experimental process in the actual research and development process, thus greatly improving the research and manufacture efficiency, lowering the research and manufacture cost, and having significant economic value and popularization value.

The invention relates to a crosswind information estimation method based on the unmanned aerial vehicle crabbing method. The crosswind information estimation method based on the unmanned aerial vehicle crabbing method comprises the following steps that test data generated during unmanned aerial vehicle crabbing are obtained; the obtained test data are processed; crosswind information estimation is conducted according to the relation between the yaw angle phi and the track angle phi <s> and the relation between the airspeed Va and the ground speed Vd, wherein crosswind information includes the amplitude of crosswind and the crosswind azimuth angle relative to the geomagnetic true north, and a certain included angle exists between the pointing direction of the head of an unmanned aerial vehicle and the crosswind. According to the crosswind information estimation method based on the unmanned aerial vehicle crabbing method, the requirement for real-time crosswind estimation in an unmanned aerial vehicle flight test can be met, the crosswind applied to the unmanned aerial vehicle when the unmanned aerial vehicle flies flat can be measured quite accurately by analyzing the speed relation existing when the unmanned aerial vehicle crabs, the safety coefficient is increased when the unmanned aerial vehicle flies, and the completeness of information is guaranteed.

The invention discloses a load simulation method for a cataplane landing gear, and belongs to the field of strength design of the cataplane landing gear. The load simulation method mainly comprises the following steps: analyzing stress on the whole cataplane, a nose gear and a main landing gear of a cataplane, so as to establish a corresponding mechanic equilibrium equation; establishing a nose gear simulation analysis model, a main landing gear simulation analysis model and a whole cataplane dynamic simulation analysis model according to the nose gear mechanic equilibrium equation; finally, correcting the whole cataplane dynamic simulation analysis model according to testing gravity center data and a front main landing gear load curve during practical catapult launching of the cataplane. The corrected model is the model capable of being used for catapult launching load simulation of the cataplane, problems in load design and analysis of the landing gear during catapult launching of the cataplane are solved, and the flight test verification of a carrier-based aircraft in an aircraft carrier platform and a land-based training ground is guaranteed.

The invention discloses a method for designing a controller of a time-varying aircraft model with uncertainty, which is used for solving the technical problem that the existing robust control theory lacks design steps, so the flight controller is hard to design directly. The technical scheme is as follows: the time-varying system segmentation robust stability and solvability conditions are given, selection of desired closed-loop poles of linear time-varying system state feedback is directly utilized and a constraint condition inequality direct design feedback matrix is given according to the characteristic that all the real parts of all the desired closed-loop poles are negative, so that the engineering technicians in the research field directly design the flight controller for the time-varying aircraft model with uncertainty obtained through wind tunnel or flight tests, thus solving the technical problem that the current researches only give the robust stability inequality but can not directly design the flight controller.

The invention relates to a rotor aircraft test platform, comprising a test platform support, two free rotation supports and a rotor aircraft support seat. The test platform comprises more than three platform supporting legs and a frame body structure which is supported on the more than three supporting legs; the frame body structure comprises a plurality of frames; the frame comprises transverse beam connection members which are connected to the platform support legs; two free rotation supports comprise a PITCH rotation frame and a ROLL turn-over transverse beam; the two ends of the ROLL turn-over transverse beam are supported on the PITCH rotation framework to form hinging; the rotor aircraft support seat is fixed in the middle of the ROLL turn-over transverse beam; and the rotor aircraft is fixed on the rotor aircraft support seat. In the invention, the rotor aircraft test platform limits the rotor aircraft in the framework structure of the platform support, which enhances the safety of the test, and regulate the flight test space through regulating the size of the platform support. The rotor aircraft test platform is reliable and simple in structure, easy in assembling and transformation, and easy to observe.

The invention discloses a semi-physical real-time simulation system of a laser guided missile. The system is used for carrying out simulation test on a missile laser guide head, an electric steering engine and a missile guidance control algorithm. The system comprises a signal simulation part used for generating the laser received by a simulation guide head, a motion simulation part used for simulating the relative motion relation between a missile and a target; a test main control part used for running the missile motion simulation software in real time, driving the whole simulation system, monitoring and recording data in the whole test process, and analyzing the data. The method aims to provide an auxiliary design tool for missile designers, so that designers can know problems possiblyexisting in the missile performance and design. The development period of missiles is shortened, and the number of external field flight tests is reduced. The system has better practicability and higher economic benefits.

The invention discloses a method for designing a controller of a multiple time lag aircraft model, which is used for solving the technical problem that the existing robust control theory lacks the design step, so the flight controller is hard to design directly. The technical scheme is as follows: the multiple lag system segmentation robust stability and solvability conditions are given, selection of desired closed-loop poles of linear system state feedback is directly utilized and a constraint condition inequality direct design feedback matrix is given according to the characteristic that all the real parts of all the desired closed-loop poles are negative, so that the engineering technicians in the research field directly design the flight controller for the aircraft model with multiple time lag uncertainty obtained through wind tunnel or flight tests, thus solving the technical problem that the current research results only give the robust stability inequality but can not directly design the flight controller.

The invention discloses a moonlet satellite-and-ground time synchronization method for a ground whole-satellite test. The purpose of satellite-and-ground time synchronization in a simulated flight test is effectively achieved. The satellite service center computer time, the AOCC time and the ground dynamics simulation time are synchronized with the GPS dynamic simulator orbit simulation time by adjusting orbital parameters of a satellite control subsystem, having control over the orbit simulation starting moment of a GPS dynamic simulator and the satellite-and-rocket separation switch state of a moonlet, and using the satellite time correction function, satellite-and-ground time synchronization is achieved, the real in-orbit situation of the moonlet can be simulated, the basis is provided for satellite power distribution, time sequence analysis calculation and instruction arrangement, and reality and effectiveness of the ground whole-satellite test are ensured. Time users use GPS pulse-per-second signals output by a GPS receiver and corresponding integral second time correction data as the time keeping reference and the time correction reference, the difference between the satellite time and the ground time does not change along with the length of the testing time, and accuracy of a ground test system is improved.

The invention discloses a method for gaining the service life of a helicopter drive system component. Strain test points are determined on the helicopter drive system component to be tested, and estimation data of the service life of the helicopter drive system component to be tested are given; in the process of a ground combined test and a flight test, strain data are obtained from the strain test points; according to the ground combined test and the flight test of a helicopter to be tested under multiple states and multiple load conditions, a criterion for overhaul of the helicopter drive system component to be tested is determined; obtained actual measurement data of the helicopter drive system component are analyzed, and the condition that the strain values of all strain test points of the helicopter to be tested exceed a monitor limit value and a component accumulated fatigue damage value is monitored and analyzed in time; comparison and correction are carried out by combining the estimation data and the actual measurement data, and the service life value of the helicopter drive system component to be tested of the same type is obtained.

The invention belongs to the field of aircraft flight load actual measurement and relates to a method for actually measuring wing torque loads of a high-manoeuvrability aircraft. The method comprises the following steps: 1, determining a structure or a part sensitive to influences of wing aerodynamic loads on wing torque loads by virtue of wing finite element computation and analysis; 2, realizing a torque load test system for each section of wings by virtue of adopting a sensitive torque bridge piece distribution scheme and a torque bridge combination technology; 3, performing a ground calibration test on the torque loads according to the structure or part sensitive to the influences of the aerodynamic loads on the wing torque loads, thereby obtaining conversion relation between external loads and bridge strain gauge signals by virtue of the ground calibration test; 4, performing a flight load actual measurement flight test, thereby obtaining the torque loads of each section of the wings according to the bridge strain gauge signals during flight. The method has the advantages that by virtue of aiming at the implementation difficulty in aircraft wing torque load actual measurement, the implementation problem of a torque load flight actual measurement test is solved, and the accuracy and high precision of the wing torque load test are ensured.

The invention discloses a debugging device for flight test of an unmanned aerial vehicle, relating to a debugging device and comprising an unmanned aerial vehicle part and a ground part, wherein the unmanned aerial vehicle part comprises a DSP (Digital Signal Processor) and a first wireless module, and the ground part comprises a second wireless module, a level switching module and a computer parameter adjusting software module; when changing the control parameter and motor rotating speed in the flight process of the unmanned aerial vehicle, the computer parameter adjusting software module sends a signal instruction which reaches to the second wireless module through the level switching module, and the second wireless module sends a wireless radio frequency signal which is sent to the DSP after being received by the first wireless module of the unmanned aerial vehicle part; the DSP changes the set of the received control parameter in the flight process of the unmanned aerial vehicle and then sends the changed control parameter to the first wireless module, and the first wireless module sends the wireless radio frequency signal, and the second wireless module receives the signal which is then switched to computer parameter adjusting software, thus, changed data can be displayed. The debugging device is simple in structure, high in reliability, easy to operate and capable of shortening the development period of an unmanned aerial vehicle control system.

The present invention discloses a combination missile target imitation flight test system, which includes a simulation development machine used in building a combination missile target flight control model, translating the simulation development machine into an executable code; an off-line simulation operation control model which records a rudder control signal and receives the rudder control signal from a simulation test host, comparing with the rudder control signal of the simulation operation model to judge the accuracy of a flight control software of the control system or the operation of a steering gear; the simulation test host which executes the executable code, sends the relevant track information of executing output to an interface box and sends the rudder control signal of an on-board control system transferred by the interface box to the simulation development machine; the interface box which is used for transferring the signals between the simulation test host and the on-board control system. The combination missile target imitation flight test system can perform a simulation test and a simulation experiment for the flight control system of multi-module and complex logic from model modification to physical simulation once, and is characterized by simple framework, low cost and high efficiency.

The invention discloses a laboratory simulation method for flight attitude disturbance, comprising the following steps: (1) using a high-precision inertia and satellite integrated navigation system to collect and record original data of attitude disturbance in an airborne environment in real time during flight test; (2) post-processing and optimizing the original data to improve the accuracy of measurement; (3) carrying out frequency domain and amplitude evaluation of the post-processing result data, and selecting a six-degrees-of-freedom motion platform of which both the frequency response and amplitude satisfy the requirements of simulation; (4) carrying out attitude disturbance simulation test under laboratory conditions with the post-processing result data as the position input of the motion track of the six-degrees-of-freedom motion platform; and (5) using the high-precision inertia and satellite integrated navigation system to collect and acquire attitude disturbance simulation data and make frequency domain and amplitude analysis, and comparing the amplitude-frequency characteristic of the attitude disturbance simulation data with that of the post-processing result data to confirm the accuracy of simulation. The laboratory simulation method for flight attitude disturbance has the advantages of high accuracy of attitude simulation, good airborne condition reproducibility, good adaptability to models, and simple and convenient operation.

The invention relates to a method for detecting a coverage range of a line-of-sight link during flight path planning of an UAV (unmanned aerial vehicle) before a flight test, and belongs to the technical field of flight tests. According to the method, three-dimensional topographic data is used as a basis, and an action range of the line-of-sight link in the case of terrain masking can be generatedby setting of parameters such as geographic location information of a line-of-sight link station, a radio unshielded distance, and flight height to be inspected. The method combines three-dimensionalterrain with the UAV and a radio transmission path, and completes the calculation and inspection of the radio coverage range of the line-of-sight link of the flight of the UAV, and solves the problemthat the action range of the line-of-sight link cannot be accurately estimated when the UAV flies in complex terrain environment. The method is applicable to the flight path planning check during theflight of the UAV (unmanned aerial vehicle). The method can also be used to calculate an action range of a ground-to-air radar, and is suitable for evaluation of station distribution site selection of radar base stations.

The invention discloses a standard model aerodynamic layout for transition research of a hypersonic boundary layer and a design method. The standard model aerodynamic layout for transition research ofthe hypersonic boundary layer comprises a head ellipsoidal head and a cone body; the head ellipsoidal head and the cone body are in smooth transition with each other; one section, corresponding to the leeward side, on a bottom molded line of the cone body is an upper half molded line; the upper half molded line is linear combined overlapping of an elliptic curve and a CST curve; one section, corresponding to the bottom of an aircraft, on the bottom molded line of the cone body is a lower half molded line; and the lower half molded line is an elliptic curve. The invention further provides thedesign method of the standard model aerodynamic layout for transition research of the hypersonic boundary layer; and the design method of the standard model aerodynamic layout for transition researchof the hypersonic boundary layer can be used for obtaining the standard model aerodynamic layout. The standard model aerodynamic layout for transition research of the hypersonic boundary layer has theadvantages that the standard model aerodynamic layout has typical characteristics of a real aircraft aerodynamic layout, and meanwhile, the whole appearance can be completely analyzed and described by using a mathematical expression to satisfy simplifying demands of flight test and numerical calculation on the aerodynamic layout.

A measuring point arrangement method suitable for hypersonic flight test transition research is disclosed. The measuring point arrangement method is achieved through the following modes: S1, determining whether the natural transition or the forced transition is to be measured according to measurement requirements, and if the natural transition is to be measured, then going to S2; if the forced transition is to be measured, then going to S3; S2, measuring a main flow transition condition and/or a transition condition of a cross flow effect according to measurement requirements, wherein when themain flow transition condition is measured, a measuring point is arranged on a streamline of a hypersonic aircraft in the main flow direction; when the transition condition of the cross flow effect is measured, the measuring point is arranged at a position with a cross flow velocity in the lateral direction; the main flow direction is the central streamline direction of the aircraft and the streamline direction with an included angle of not more than 3 degrees with the central streamline direction; S3, arranging a rough element at a predetermined position on the aircraft, and arranging the measuring point at the downstream of the streamline where the rough element is located; and the measuring point positions realize the measurement of the physical quantity of the surface of the aircraftin the flight test process by installing sensors.

The invention discloses a method for designing a controller of a multi-interval aircraft model, which is used for solving the technical problem that the existing robust control theory lacks design steps, so the flight controller is hard to design directly. The technical scheme is as follows: the multi-interval system robust stability and solvability conditions are given, selection of desired closed-loop poles of linear system state feedback is directly utilized and a constraint condition inequality direct design feedback matrix is given according to the characteristic that all the real parts of all the desired closed-loop poles are negative, so that the engineering technicians in the research field directly design the flight controller for the multi-interval aircraft system model obtained through wind tunnel or flight tests, thus solving the technical problem that the current researches only give the robust stability inequality but can not directly design the flight controller.

The invention discloses an aircraft time lag varying model approximation and controller designing method, and aims to solve the technical problem that the existing robust control theory is lack of the design steps and an aircraft controller cannot be directly designed. The technical scheme is as follows: the robust stability solvable condition of a time varying approximation system is given, a closed loop expecting pole fed back by a linear system state is directly utilized for selection, and a given conditional inequality is given to design a feedback matrix directly according to the characteristic that the real part of all closed loop expecting poles are negative numbers. By utilizing the method, the aircraft model containing time lag varying uncertainty are obtained by engineering technicians in the research field during wind tunnel or flight test, and an aircraft controller can be designed directly, so that the technical problem in the current research that only the robust stability inequality is given and the aircraft controller cannot be designed directly is solved.