82 results about "Flight experiment" patented technology

The invention provides an attitude estimation method of a maneuvering acceleration-assisted extended Kalman filter (EKF) attitude and heading reference system (AHRS). The state quantity of the EKF contains an error of three attitude angles, a null bias error of a three-axis gyroscope and a three-axis carrier maneuvering acceleration error of a carrier system, and the observed quantity of the EKF contains a three-axis acceleration error and a three-axis earth magnetic field error. A strapdown attitude algorithm and the nine-state EKF algorithm are subjected to data fusion to acquire attitude estimation of the AHRS. A single-axis turntable experiment, a vehicle-mounted dynamic experiment and a flight experiment prove that the maneuvering acceleration-assisted nine-state EKF data fusion algorithm has stable attitude angle accuracy under different maneuvering conditions, and the mean square deviation of the attitude angle can be limited within 2 degrees.

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 provides a flight controller for a civil small UAV (Unmanned Aerial Vehicle). The flight controller comprises a flight control microprocessor, a GPS module, a steering engine driving module, a sensor, a data storing module, a power supply and a ground measuring and control module which form a digital control system in which the resource is unified to be dispatched and distributed and all parts coordinately work. According to the flight controller for the civil small UAV, a hardware system design scheme of the small UAV using an STM32F103RE microcontroller as a core is proposed according to the requirements on small size, light weight, simple structure and high reliability of the civil UAV; the overall scheme design of the system and the main functional modules are described in details; the feasibility and reliability of the design scheme are verified through the flight experiment on a small fixed-wing UAV; in addition, the flight controller is fine in design, low in cost, high in reliability, and suitable for mass production and use.

The invention discloses a retractable captive flight support for external hanging equipment, which is arranged at a central cap of a base plate of an aircraft cabin and comprises a fixed guide framework, a movable frame, a movable pull rod, movable pulleys, a rocker arm mechanism, a cover plate and a butt-joint frame plate, wherein the fixed guide framework is connected with the base plate of the aircraft cabin and connected with the movable frame through a concave guide rail, the lower end of the movable pull rod is mounted on the movable frame, the upper end thereof is connected with a pull rod fixing seat arranged at the front end of the cabin cap through a positioning pin, the pull rod fixing seat is fixed on the base plate of the aircraft cabin, the lower frame of the movable frame is welded with the butt-joint frame plate, the movable pulleys are respectively arranged at the right and left ends of the movable frame, and the rocker arm mechanism is arranged at one side of the fixed guide framework through a fixing seat. The retractable captive flight support of the invention has the advantages of good generality, secure folding and unfolding, ability of independently completing captive flight experiments of various guiding heads without developing special equipment, safety and convenience, short and simple experiment organization cycle and greatly-reduced experiment cost.

The invention discloses a micro air vehicle experimental device and relates to experiment and measurement technology. The device can be used in an air vehicle experiment and parameter measurement and consists of a support, a pallet capable of rotating in three dimensions, a movable balance weight and a safety limit element. The support consists of a cross beam and a support rod, wherein the cross beam and the support both have fixing holes and fixing bolts so as to be fixed on a horizontal plane to ensure the stability of the experimental device. The rotating pallet is used to carry the air vehicle receiving experiment, the air vehicle can be fixed on the pallet through a bolt, and the pallet and the air vehicle can rotate on three vertical axes in space under the drive of a spherical joint. The experimental device of the invention is designed with the attitude motion and the degrees of the freedoms of motion and position in the three dimensions of the body of the air vehicle according to the spaceflight characteristics of the air vehicle and realizes flight experiment and the measurement of static and dynamic parameters of the air vehicle under the condition of ensuring the safety of the air vehicle.

The invention discloses a flight test determination method of the multi-input and multi-output equivalent pneumatic servo elastic robust stability, used for solving the technical problems of the conservation and complex calculation of the traditional robust analysis method. Based on the technical scheme, the flight test determination method comprises the following steps of obtaining an open-loop transfer function frequency characteristic matrix of a multi-input and multi-output system through a frequency sweeping flight test and the like; lagging a phase angle through serial gaining at each loop; obtaining a scalar judgment formula of phase and amplitude margin between two adjacent flutter frequency regions according to direct equivalence of the closed-loop frequency characteristics by carrying out characteristic decomposition on the system open-loop transfer function frequency characteristic matrix; and calculating the phase margin and the ASE stability at the flutter frequency in a mode similar to a single-input and single-output system. The invention simplifies the problems through characteristic linear conversion, obtains the scalar judgment formula, gives a calculation method of the flutter boundary stability, the margin and the safety during elastic flight of a canard wing aircraft and reduces the conservation of the traditional method.

An embodiment of the invention discloses an analog simulation system, and relates to the technical field of aviation. By adopting the system, a flight experiment on a near-earth warning system can be conducted at low cost and at low risk. A method adopted by the system is as follows: an aircraft dynamic simulation system is connected with an operating mechanism, and the operating mechanism at least includes a steering wheel, a pedal and a throttle lever; a three-dimension visual display system is connected with the aircraft dynamic simulation system; a warning computer is connected with the aircraft dynamic simulation system and the three-dimension visual display system, and is used to generate warning information according to the aircraft status information and the external flight environment information generated by the three-dimension visual display system; and an instrument display module is connected with the warning computer, the aircraft dynamic simulation system and the three-dimension visual display system. The analog simulation system of the invention is applicable to analog simulation of a warning system.

The invention discloses a lift model-assisted four-rotor aircraft height fault-tolerant estimation method. The method comprises 1, acquiring a lift coefficient of a four-rotor aircraft lift model through a flight experiment, 2, carrying out navigation filtering through two federated Kalman filters in parallel, 3, detecting faults of a four-rotor aircraft height sensor in a navigation filtering process through a fault detection strategy and 4, if the fault of the four-rotor aircraft height sensor is detected, carrying out fault isolation and recovery of the sensor with fault and eliminating thenavigation information of the fault sensor. Through combination of the four-rotor aircraft lift model and airborne sensor, the redundancy of the four-rotor aircraft height sensor is formed, the faultdiagnosis and fault tolerance of the height sensor are realized and the robustness of the navigation system is improved.

The invention discloses an unmanned plane identification experiment ground station capable of assisting in operation and belongs to the technical field of unmanned plane modeling and control. The unmanned plane identification experiment ground station capable of assisting in operation is characterized by having four main functions of data reading, data recording, experiment monitoring and curve display, wherein the data reading function is used for receiving and reading a remote control instruction of a pilot and dynamic response initial data of an unmanned plane; the data recording function is used for recording the initial data for the system identification after an identification experiment; the experiment monitoring function is used for continuously computing the frequency and amplitude of the remote control instruction and the dynamical response data and giving an alarm when the frequency and the amplitude of the data exceed a safety range; and the curve display function is used for displaying the experiment data in a form of curve in real time to assist ground crew in monitoring the identification experiment and providing feedback information to the pilot. The unmanned plane identification experiment ground station can effectively assist the pilot and the ground crew in completing the unmanned plane identification flight experiment and has the advantages of safety, high efficiency, convenience and reliability.

The invention discloses a two-degree-of-freedom wind tunnel virtual flight test method. The two-degree-of-freedom wind tunnel virtual flight test method comprises the following steps: step 1, preparing a wind tunnel virtual flight test; 2, releasing the pitching degree of freedom and the rolling degree of freedom of the test model; 3, executing a pitching and rolling maneuvering flight control law; 4, ending single blowing of the wind tunnel virtual flight test; and 5, executing a wind tunnel virtual flight test in a new parameter state. According to the two-degree-of-freedom wind tunnel virtual flight test method, the advantages of real simulation of a flight physical environment, high efficiency, accuracy, low cost and repeatable simulation of a large wind tunnel are fully utilized, and the wind tunnel virtual flight test method suitable for integrated evaluation and verification of pneumatic / motion / control performance of an aircraft is established; the method is an innovative method means besides existing digital modeling simulation, semi-physical simulation and flight test methods, and has important technical support significance for flight control optimization design and performance evaluation of the aircraft.

The present invention discloses an aircraft support device of an aircraft formation flight experiment and a transonic speed wind tunnel experiment device. The support device comprises: a pedestal fixedly arranged at the bottom portion of a wind tunnel experimental section; a back stay support comprising a first connection arm and a second connection arm, wherein the first connection arm and the second connection arm are connected at an angle, the first connection arm is fixedly connected with the pedestal, and the second connection arm is connected with a lead aircraft through an angle-variable block; the angle-variable block comprising an installation portion and an angle-variable plate connected with the lead aircraft, wherein the angle-variable plate is connected with the second connection arm; and a bent-tail connection rod having one end connected with a grid force measurement mechanism in the wind tunnel experimental section and having the other end connected with the six-degree-of-freedom scale through a scale pole, wherein the six-degree-of-freedom scale is connected with a wing plane. The aircraft support device of the aircraft formation flight experiment and the transonic speed wind tunnel experiment device realize firm support of the lead aircraft and control of the wing plane flight state in the wind tunnel experiment of the aircraft formation flight so as to improve accuracy of a wind tunnel experiment at the aspects of large-scale aircraft formation flight efficiency assessment and formation parameter optimization.

The invention discloses a method for identifying a fixed-order parameter model of an aircraft based on modal segmentation and genetic algorithm, which belongs to the field of aircraft identification modeling. The invention is characterized in that the method comprises four stages, namely, model structure determination, flight data acquisition, model identification and model validation, wherein the stage of model structure determination is used for establishing a modal segmentation model of the aircraft, and comprises three steps, namely, kinetic analysis, model order determination and modal segmentation model determination; the stage of flight data acquisition is used for obtaining the frequency-domain response data of the aircraft, and comprises three steps, namely, sweep-frequency flight experiment, frequency domain transformation and data response of frequency domain; the stage of model identification is used for identifying the obtained dynamic model of the aircraft, and comprises a step of identifying model by using the genetic algorithm; and the stage of model validation is used for verifying the obtained dynamic model, and comprises a step of model validation, wherein, in the stage of model structure determination, the complicated high-order dynamic models are simplified by using the modal segmentation model, so that the precision of the model is not restricted by the model order in the process of identification while all unknown parameters are kept; in the stage of flight data acquisition, the true dynamic frequency response of the aircraft is obtained; in the stage of model identification, the modal segmentation model is upon the true dynamic frequency response to the greatest extent by using the genetic algorithm; and in the stage of model validation, the obtained dynamic model is inspected, in case of meeting the requirements, all identification models meet the requirements, otherwise, an experiment is carried out again.

The invention relates to a flutter test safety protection method, belongs to the technical field of aviation aerodynamic wind tunnel tests, and aims to solve the problem of damage to a model or internal parts of a wind tunnel caused by excessive vibration of a test model in the wind tunnel flutter test process. According to the method, the flow field stability condition is monitored through a dataacquisition and analysis module when a wind tunnel control module starts a test, the value of an acceleration sensor on the test model is monitored, calculation and analysis are carried out, a safetyprotection device is used to carry out safety protection on the model, the control difficulty of the aircraft in the pneumatic simulation flight test is reduced, the control efficiency is improved, the problems of large size of a development board and occupation of model space are solved, the weight of the model is effectively reduced, and the development difficulty is greatly reduced.

The invention discloses an embedded wind tunnel free flight test model pose acquisition method, and relates to the technical field of flight tests. The embedded wind tunnel free flight test model poseacquisition method comprises the steps of: carrying out the interception of inertial sensor data according to a preset rule to obtain test effective data which comprises triaxial accelerometer data and triaxial gyroscope output data; performing low-pass filtering on the triaxial accelerometer data to obtain first data, and removing high-frequency vibration interference; performing band elimination filtering on the first data and the output data of the three-axis gyroscope to obtain second data, and removing pulsation interference of a pneumatic airflow; and carrying out recursive pose calculation on the basis of the second data to obtain six-degree-of-freedom pose information of the aircraft model within effective time in the test process. The embedded wind tunnel free flight test model pose acquisition method can improve the test precision for the application scene of the wind tunnel free flight test.

Provided is an incoming flow parameter determining method applicable to an appearance of a body of revolution. The method comprises the steps of 1, building a surface pressure approximation model applicable for the appearance of the body of revolution, wherein firstly a surface pressure computational formula is determined, and surface pressure is expressed as a sum of a product of incoming flow dynamic pressure qc and a pressure coefficient Cpi and incoming flow static pressure p infinity; then a polynomial form is adopted to express the pressure coefficient Cpi, and polynomial factors are a flight attack angle alpha, a sideslip angle beta and an incoming flow pressure ratio R; finally the factors in the polynomial are obtained through numerical fitting, regression or a least square method to surface measuring point pressure of the body of revolution in multiple sets of states; 2, obtaining the surface measuring point pressure of the body of revolution in a flight experiment, conducting back calculation according to the surface measuring point pressure combining the approximation model, and accordingly obtaining an incoming flow parameter. According to the incoming flow parameter determining method applicable for the appearance of the body of revolution, the model can be used for an embedded air data system, and the prediction accuracy of the system can be effectively improved.

The invention discloses a metal additive manufacturing device oriented to weightless flight and vacuum working conditions. The metal additive manufacturing device comprises a vacuum system, a metal fusion system, a movement system and a monitoring system, metal wires are used as a raw preparation material, a high-energy-beam heat source is adopted as an energy source, and the device is used for metal additive precise manufacturing under weightless flight microgravity and vacuum working conditions. The design of a lightweight cavity and arrangement of the compact-type motion system are adopted,the ratio of an effective space of a molding area to the system occupation space is increased, weightless flight experiment platform mechanical and electrical constraint conditions are satisfied, andmetal additive manufacturing experiments in a weightless flight periodicity microgravity environment are carried out; an adjustable vacuum system is arranged to guarantee that the whole metal wire additive manufacturing process is carried out in a vacuum condition; the manufacturing process and state data are entirely recorded to provide a technical guarantee for the observation and traceback ofthe manufacturing process; the ring-column-type high-energy-beam heat source is adopted for focusing small-diameter light spots in cooperation with a precise wire feeding system, then it is guaranteedthat the heat source and the raw material are symmetric and concentric, and precise manufacturing is achieved.

The invention discloses a composite recognition method for a flight dynamics model of an unmanned helicopter, which belongs to the field of unmanned aerial vehicle dynamics modeling. The composite recognition method is characterized in that: the unmanned helicopter, a flight control computer, a data station, a ground station, a remote control transmitter and a remote control receiver are involved, wherein the flight control computer is used for assisted control for a recognition experiment to keep the relatively more stable flight conditions such as speed and height of the helicopter at the same time of ensuring the flight safety of the helicopter; a ground operator triggers the dynamic response of the helicopter by a remote control instruction; and the flight control computer adds the remote control instruction into the control input of the unmanned helicopter and records the remote control instruction together with an automatic control instruction for dynamic recognition after a flight experiment. In the method, manual remote control and assisted automatic control are simultaneously introduced, so that the flight dynamics model of the unmanned helicopter can be precisely and safely recognized.

The invention discloses a shipboard aircraft cluster motion modeling simulation method. The method comprises the following steps of: motion solution space construction: establishing a motion solutionspace of a shipboard aircraft cluster; motion modeling: establishing a motion control model of the shipboard aircraft cluster; and motion simulation: inputting parameters required for simulating motion of the shipboard aircraft cluster on the basis of the motion solution space and the motion control model, and carrying out shipboard aircraft clyster motion simulation. Through the shipboard aircraft cluster motion modeling simulation method, effective motion solution spaces and motion control models can be provided for motion simulation of shipboard aircraft clusters, so that motion laws of theshipboard aircraft clusters can be researched under various simulation conditions. The method is capable of better simulating rising / landing and cruising behaviors of the shipboard aircraft clustersunder various environments, so as to observe operation data of shipboard aircrafts in real time and provide comprehensive simulation data for real large-scale shipboard aircraft cluster flight experiments.

The invention discloses a ground full-size equivalent test method for a low-temperature pressurized delivery system of a carrier rocket, which comprises the steps of providing a ground full-size test module of the carrier rocket, wherein the ground full-size test module comprises a full-size storage tank and a low-temperature pressurization conveying system; pressurizing a storage tank by adopting normal-temperature helium, and setting a pressurization control band, a filling liquid level height, an initial air pillow and a propellant temperature according to the pressure of the storage tank of the carrier rocket in a real flight state; when the initial parameter of the propellant filling liquid level height meets the requirement, opening a throttling orifice plate of a corresponding discharge pipe based on the output quantity of the carrier rocket propellant in the real flight state, and recording working condition parameters; controlling working condition characteristic parameters, carrying out data analysis according to the consistency comparison condition with the flight state characteristic parameters, and optimizing the flight state parameters. According to the test method, the number of whole-system power test run and flight test times of the carrier rocket can be reduced.

The invention discloses a steel cable type flight experiment table and belongs to the technical field of plant protection unmanned aerial vehicles. The table comprises a frame body, a flying trolley,a clamp and a ground lifting platform. A friction winch is adopted for closed-loop traction, so through the flying trolley, rapid acceleration and deceleration can be achieved, the track length is shortened and the test section distance is ensured to the greatest degree. Meanwhile, a steel cable replaces the track, so that the structure is simplified. In addition, a ground lifting platform adoptsa hydraulic shear type lifting vehicle to be raised and descend, so the height of a to-be-tested target is adjustable through the lifting vehicle, the complexity of a connection clamp of the flying trolley can be greatly reduced, the structure of the flying trolley is further simplified, and the load of the flying trolley is effectively improved. The steel cable type flight experiment table provided by the invention is more suitable for the actual field operation condition, can simulate the actual operation environment, and can simulate the actual spraying effect under the conditions of simulating different flight heights, speeds and angles of a plant protection unmanned aerial vehicle outdoors.

The invention belongs to the technical field of pressure distribution measurement in a flight test, and particularly relates to an intelligent flexible pressure measuring belt for pressure distribution measurement in a flight test. The pressure measuring belt comprises a plurality of pressure measuring units which are arranged on a flexible pressure measuring belt substrate and are in data connection in sequence; each pressure measuring unit comprises a plurality of channel structures, each channel structure comprises a pressure sensing unit, a signal acquisition unit, a data processing unit, a data communication unit and a data transmission unit, and the pressure sensing units, the signal acquisition units, the data processing units, the data communication units and the data transmission units are in data connection in sequence. The pressure measuring belt can be tightly attached to the shapes of parts such as wings, and has the characteristics of high precision, light weight, easiness in modification, convenience in maintenance and the like.

The invention belongs to the field of helicopter model identification, and provides a flight experiment data rapid frequency domain identification method suitable for a telex flight control helicopter, which comprises the following steps of: 1, eliminating and correcting a data outlier, performing low-pass filtering, correcting a sensor position, checking data compatibility and reconstructing data; 2, converting the flight experiment data into a frequency domain, constructing an identification model, and adopting a completely linearized state space model; 3, fast initial value estimation is carried out, and initial value estimation theta of the to-be-identified parameters is obtained; 4, model structure identification is carried out, and before the next step of identification is carried out, the parameters are selected and used; And 5, outputting an error in a fast frequency domain. According to the method, the accelerated optimization algorithm is designed according to the characteristic that the convergence rates of the parameters to be identified are different, so that the calculation efficiency is further improved, the relatively detailed description of the frequency spectrum is realized, and meanwhile, the calculation amount is ensured.

The invention discloses an aircraft aerodynamic parameter identification method based on a recurrent neural network. The method comprises the following steps: 1) performing a simulated flight test by using a training-level simulator in combination with flight simulation software to obtain flight data; 2) taking the six-degree-of-freedom dynamic equation set of the aircraft rigid body as a state equation of the system, and calculating to obtain corresponding aerodynamic force and aerodynamic torque according to data obtained by the test; 3) taking flight data such as an attack angle and a sideslip angle as input, taking the aerodynamic parameters calculated in the step 2 as reference data, and training by utilizing a recurrent neural network combined with a real-time recurrent learning algorithm to obtain an aerodynamic parameter identification model; 4, selecting and loading flight data which do not participate in model training into the aerodynamic parameter identification model of the recurrent neural network obtained in the third step for parameter identification, and corresponding aerodynamic force and aerodynamic moment. The parameter identification model established through the method has good applicability, accurate modeling can be completed for aerodynamic force, and application and popularization can be achieved.

The invention provides a flight test route planning method for measuring and controlling equipment precision identification. The method comprises the following specific steps: 1) determining assessment equipment and equipment coordinates, wherein one route can simultaneously assess one or two pieces of equipment; 2) determining a projection distance of a target endurance point; 3) determining a flight height range of an aircraft trial route segment according to the shortest acting distance of the equipment and the precision-guaranteed tracking pitch angle limit; 4) determining a longitude value and a latitude value span corresponding to the unit surface length according to the local latitude; and 5) determining the route direction and calculating the waypoint coordinates in combination with external conditions such as terrain, weather, low elevation, target RCS change and sun illumination included angle, and completing the flight route planning. The method provided by the invention isclosely combined with the longitude and latitude coordinates in a real map to set the flight route, thereby effectively improving the test efficiency and precision of the external field test equipment.

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.

The invention discloses a synchronous detonation device for a hypersonic aircraft component separation wind tunnel experiment. The synchronous detonation device comprises a synchronous trigger unit (1), a detonation control unit (2) and a constant current unit (3). The synchronous detonation device solves the problems of time synchronization of a detonation process and a wind tunnel effective experiment, monitoring of the detonation process and the like, avoids the problem of erroneous detonation induced by strong electromagnetic interference of wind tunnel aided high-power equipment, provides an experimental foundation for verifying the feasibility of an aircraft component separation scheme with a wind tunnel, allows the component separation verification scheme not to be limited to actual flight experiments, and reduces the aircraft research and manufacturing cost and cycle.

The invention relates to a flight control system and method for a wind tunnel virtual flight test, belongs to the field of wind tunnel virtual flight tests, and aims to solve problems of model yaw angle recognition, single function, incapability of meeting complex wind tunnel test requirements and the like. The system is composed of a programmable autopilot, a remote controller receiver, a remotecontroller, a power supply, an adaptive circuit, a radio station, a gyroscope, an encoder, a flight control program and the like. The three-degree-of-freedom supporting device is connected with a model supporting rod, and the flight control device controls a model control surface according to the control law and the control program; measurement of model attitude and attitude angular velocity is realized, meanwhile, wireless communication between equipment and an external computer is achieved, field manual operation of operators is omitted, control difficulty of the aircraft in the wind tunnelsimulation flight test is lowered, control efficiency is improved, problems that a development board is large in size and occupies model space are solved, the weight of the model is effectively reduced, and development difficulty is greatly lowered.

The invention provides an average temperature calculation method of a level flight process of a stratosphere airship with a solar cell. According to the flight parameters of an airship, the design parameters of the airship, the characteristic parameters of airship body materials, the characteristic parameters of the solar cell and the characteristic parameters of cell thermal insulation materials, an atmospheric environment parameter and an airship thermal environment parameter are calculated, the airship is divided into a plurality of nodes on the basis of the geometrical characteristic and the heat transfer mode of the airship, an energy differential equation of each node is established, and the average temperature data of each node of the flight process of the airship is calculated through the solving of an energy differential equation set of multiple nodes of the airship. The average temperature calculation method has a guiding meaning on aspects including the design, the material selection, the flight experiment planning, the circumvention of potential dangers and the like of the stratosphere airship with the solar cell, can improve the one-time success rate of the design of the stratosphere airship with the solar cell, shortens the design period of the stratosphere airship with the solar cell and lowers the design cost of the stratosphere airship with the solar cell.

The invention relates to the technical field of radars and discloses a double-base forward-looking SAR semi-physical simulation device and method. The method is characterized in that firstly, an echoand state signal simulator based on ZYNQ and a digital-to-analog converter is designed, and echo and state data input is provided for a double-base forward-looking SAR imaging system; secondly, a logic module for synchronously receiving and transmitting echo and state data is designed in the ZYNQ to ensure that the echo and state data can be transmitted to an imaging system in a one-to-one correspondence manner, so the data flow of the imaging system is simplified; thirdly, an FPGA-based echo and state data receiving and packaging module is designed; and lastly, a data cycle sending module isdesigned based on the FPGA to ensure that the DSP can perform signal processing in an assembly line mode. The method is advantaged in that the aim of carrying out hang flight equivalent verification on the hardware double-base forward-looking SAR imaging system in a laboratory is fulfilled, large-scale outfield hang flight experiment expenditure is avoided, moreover, debugging efficiency of the double-base forward-looking SAR imaging system is further improved by operating in a laboratory.

In order to overcome the problem in the prior art that a complex flutter model under aerodynamic force and strength change influences can not be effectively expressed, the invention provides an aircraft flutter analysis grid model Walsh modeling method. The method is characterized in that a plurality of grid points are selected in an aircraft engine body shaft system; under aerodynamic force and strength change influences including different flight speeds, atmospheric densities, airflow environments, different temperatures and the like, according to an engine body shaft system decomposition method, a complex flutter grid model is expressed; according to a requirement for establishing the model, the requirements of installing sensors and data and image records are put forward; through an effective flutter flight experiment, data is obtained, and an excitation function is obtained through a measured value of an airflow sensor; a Walsh function is adopted to carry out approximation and equivalent description on a vibration variable; according to a distinguishing method, the solving of three axial vibration equations on an engine body axial system coordinate grid point can be simultaneously determined, and the technical problem in the prior art that the complex flutter model under aerodynamic force and strength change influences can not be effectively expressed is solved.