287 results about "Flight test" patented technology

The invention belongs to the technical field of flight testing, and discloses an aircraft testing method and system for simulating a flight environment, and the method comprises the steps: obtaining design parameters of an aircraft and environment data of a real scene, and constructing a simulation test model; in the simulation test model, generating a simulation scene based on a preset test task; collecting multi-source simulation data of a simulation scene where the virtual object is located; inputting the multi-source simulation data into a pre-constructed state perception prediction model, and predicting the state perception of the virtual object to obtain a state prediction result; the state prediction result is input into a pre-constructed test control model, a navigation control instruction is output, and the virtual object sails along a test track when responding to the navigation control instruction; the test trajectory is generated based on a preset trajectory planning algorithm; and recording test data of the virtual object in the process of sailing along the test track to obtain a simulation test result. The method has the advantages of low test cost, few limited conditions, high controllability and the like.

The invention provides an unmanned aerial vehicle laboratory testing method and device, and belongs to the technical field of unmanned aerial vehicle testing. The scheme comprises the following steps: determining a global reference point based on the position of a laboratory, and further determining a conversion relation between a local northeast high coordinate system and a global coordinate system and a plurality of fixed reference points; according to the fixed reference point, sequentially executing stand setting and starting of an automatic measurement tracker in a laboratory, and further obtaining real-time coordinates of the to-be-tested unmanned aerial vehicle by using the automatic measurement tracker; the GNSS signal generator generates a simulation navigation signal according to the real-time coordinates of the to-be-tested unmanned aerial vehicle and records a simulation track; the to-be-tested unmanned aerial vehicle performs indoor flight test based on the simulation navigation signal; and comparing the simulation track with a flight track recorded by a flight control system of the to-be-tested unmanned aerial vehicle so as to output a laboratory test result. The scheme does not depend on hardware modification of the unmanned aerial vehicle, the implementation difficulty is reduced, the implementation cost is saved, and controllable testing of medium and large unmanned aerial vehicles in the indoor environment is facilitated.

The invention discloses a rocket aircraft aerodynamic performance prediction method based on heaven and earth test and simulation data, and relates to the technical field of aerodynamic design. The method mainly comprises the following steps: obtaining a large number of aerodynamic performance parameters by using a non-viscous rapid design theory to construct a low-precision flow field training data set; performing heaven-earth correlation error correction on ground wind tunnel test data through a flight test; performing simulation error correction on a large number of aerodynamic performance parameters obtained by CFD numerical simulation through test data to obtain a high-precision training data set; and performing data fusion on the high-precision data set and the low-precision data set, and constructing a pneumatic prediction model. The invention provides the method for rapidly predicting the aerodynamic performance of the rocket aircraft, and an efficient and feasible solution is provided for rapid prediction of the aerodynamic performance.

The embodiment of the invention provides an equivalent model design method and system for an unmanned aerial vehicle carrying mission load flight test, and the method comprises the steps: obtaining mission load demand parameters and test condition parameters, the demand parameters comprise a long side A1 and a short side A2 of a rectangular airspace, A1 > = A2, and a demand speed v0; the test parameters comprise a test airspace long side B1 and a test airspace short side B2, B1 is greater than or equal to B2, the test machine minimum speed v1 and the test machine maximum speed v2, and v2 is greater than or equal to v1; when A1 < = B1, A2 < = B2 and v1 < = v0 < = v2 are not satisfied, calculating an equivalent coefficient p, and if the airspace condition A1 < = B1 and A2 < = B2 are not satisfied, calculating and determining the equivalent coefficient p; if the airspace condition A1 < = B1 and A2 < = B2 are satisfied, but the speed v1 < = v0 < = v2 is not satisfied, determining an equivalent coefficient p; based on the reference point coordinates and the equivalent coefficient p, longitude and latitude are achieved through geocentric coordinate system conversion, and horizontal navigational speed mapping is carried out. According to the technical scheme, by simulating the flight profile required by the mission load, the problem that the unmanned aerial vehicle cannot fly out of the flight profile required by the mission load in a flight test carrying the mission load can be solved.

The invention relates to a high-speed aircraft transition identification method based on a flight test high-frequency vibration signal. The method comprises the following steps: selecting high-quality measurement point data suitable for analysis from flight test high-frequency vibration data as a vibration signal to be analyzed; performing time-frequency analysis on the to-be-analyzed vibration signal, outputting a time-frequency diagram, and representing the change condition of energy in a frequency domain by using a centroid frequency and an envelope difference of the centroid frequency; carrying out preliminary identification on the transition occurrence moment; decomposing a to-be-analyzed vibration signal into a plurality of intrinsic mode functions with different center frequencies through variational mode decomposition; quantizing the change condition of the energy of each modal component along with time; and aiming at the change condition of the energy of each modal component along with the time, the transition is identified according to the energy of the mode excited by the transition, and the energy sudden increase point is the moment when the transition begins to occur. According to the method, the transition occurrence time can be effectively identified under the condition that no hole is formed in the surface of the aircraft.

The invention relates to a tilt rotorcraft semi-physical simulation system and method, and relates to the technical field of wind tunnel flight test semi-physical simulation. Based on the initialized system with the test parameters set, six-component balance load, wind speed and attitude data are collected in real time, and the collected key physical data are filtered and then sent to the flight simulation subsystem; the flight simulation subsystem carries out calculation based on a flight dynamics model, a rotor wing power model, a control instruction preprocessing model and the received data and outputs an actuation instruction; the actuator receives an actuation instruction and drives the aircraft model, and the airborne sensor feeds back new attitude data to form a control closed loop; the flight simulation subsystem sends state data to the 3D visual simulation module in real time to realize synchronization of control and visual; in this way, the stable state under the combination of different tilting angles and flight speeds is recorded, and a tilting transition corridor is obtained through fitting. Therefore, simulation precision, test safety and iteration efficiency are taken into consideration, and fused semi-physical simulation is formed.

The invention belongs to the technical field of flight test, and discloses a cockpit vision field evaluation test flight method suitable for a troposphere manned airship, which comprises the following steps: selecting an airship test condition; the airship climbs to a preset field height and keeps level flight; a preset flight test action is completed; and commenting by a flight tester, and filling in a cockpit vision field commenting card. According to the method, the blank of the flight test method for evaluating the visual field of the cockpit of the troposphere manned airship with the vector propulsion system is filled, and the flight test efficiency and the flight test quality of the manned airship are improved.

The invention discloses a multi-modal unmanned aerial vehicle flight control system and a switching method, and relates to the technical field of flight control systems, the multi-modal unmanned aerial vehicle flight control system comprises a parameter generation module for performing real flight test on an unmanned aerial vehicle to form a parameter combination and a modal mapping table; the mode switching module is used for collecting current multi-dimensional data of the unmanned aerial vehicle for preprocessing, generating an undetermined control mode in combination with the mode mapping table for verification, obtaining a target mode and executing mode switching; the instruction processing module is used for executing smooth operation to obtain an initial control quantity during mode switching, calculating a final control law and an inverse matrix, executing matrix multiplication based on the final control law and the inverse matrix, generating a torque instruction and taking the torque instruction as an execution instruction of motor thrust distribution or steering engine deflection in the unmanned aerial vehicle; according to the invention, the synchronism and consistency of mode switching and control law updating are realized, and the real-time performance and accuracy of the unmanned aerial vehicle during flight task execution are improved.

The invention discloses a fixed-wing aircraft non-airport take-off and landing adaptability evaluation method based on test flight performance data, and belongs to the field of aviation aircraft design. Comprising typical difference analysis of a non-airport environment and an airport environment, takeoff and landing sliding tolerance analysis of the unmanned aerial vehicle in the non-airport environment and takeoff and landing lateral deviation tolerance analysis of the unmanned aerial vehicle in the non-airport environment. The non-airport take-off and landing adaptive capacity of the airplane is analyzed through take-off and landing data of previous test flights. Key differences and possible problems faced by taking off and landing in the non-airport environment are found out by comparing the test flight airport environment with the non-airport environment, and key constraints of taking off and landing in the non-airport environment are formed; taking-off and landing taxiing section flight parameters in the test flight data are extracted, conformity analysis is carried out from all constraint dimensions, and the conformity analysis comprises analysis of taxiing tolerance based on grounding point deviation and analysis of taxiing feasibility of all levels of roads based on taxiing side offset distance probability distribution; whether field cleaning work needs to be carried out or not during taking-off and landing is evaluated by combining airplane height and highway guardrail design standard comparison.

The invention relates to a closed-loop virtual test flight pre-evaluation method and a medium, and relates to the field of flight tests. The method comprises the steps of constructing a digital twin module which comprises a pneumatic sub-model, a flight control sub-model, a power sub-model and a structure sub-model, defining a tilt transition specific high-risk subject and generating a test case, executing parallel Monte Carlo simulation and hardware-in-the-loop test, performing flight safety risk identification and envelope boundary quantification, and determining a tilt transition specific high-risk subject according to the tilt transition specific high-risk subject. And generating a risk assessment report and optimizing a real aircraft test flight outline. Wherein the aerodynamic sub-model adopts a dynamic vortex grid correction method based on CFD offline calibration and online query; the flight control sub-model comprises a configurable fault injection module based on a script engine, and the configurable fault injection module is used for defining a coupling fault sequence of sequential logic and a condition triggering relation. Compared with the prior art, the accuracy and the applicability of the pre-evaluation of the virtual test flight are improved.

The present invention discloses a variable-pitch propeller design verification method and a dedicated simulation test device. The method adopts a variable-pitch propeller design verification method based on blade engineering multi-segment three-dimensional surface elements and PSP surface pressure distribution measurement technology. There is no need to use a large-scale wind tunnel to test the propeller, nor does it require a high-power engine / motor to drive the propeller to rotate. Only the surface pressure distribution of a series of three-dimensional surface elements needs to be measured in the test device proposed by the present invention. The method has the advantages of being applicable to a wide range of propellers, a relatively simple test system, low cost and high efficiency. It has more prominent advantages than full-scale wind tunnel tests and flight tests, is easier to implement, and has higher safety.

The invention discloses a coaxial double-propeller multi-rotor unmanned aerial vehicle stability control method based on aerodynamic disturbance optimization, and the method comprises the following steps: S1, collecting flight data, preprocessing the data, and generating an aerodynamic disturbance training data set; s2, based on the aerodynamic disturbance training data set, constructing an improved Fourier neural operator network model; s3, optimizing parameters of the model by adopting an improved chaos grey wolf optimization algorithm; s4, embedding the optimized model into an unmanned aerial vehicle flight control system; s5, according to the predicted aerodynamic disturbance data, a model prediction control algorithm is adopted to adjust the thrust and the attitude; and S6, verifying the stability of the control method through a simulation test and an actual flight test. The optimization method for improving the flight stability of the unmanned aerial vehicle in the complex aerodynamic environment is described in detail, and the method has the advantages of being high in calculation efficiency, accurate in aerodynamic disturbance prediction and high in control system robustness.

The invention provides a ground GNSS differential data correction method and device, and the method comprises the steps: obtaining a pseudo range and a carrier phase, calculating a whole cycle jump value of the carrier phase through employing an M-W combination function and an ionosphere delay change rate, and correcting a carrier phase value through employing the calculated whole cycle jump value of the carrier phase. The influence of station-satellite geometric distance is eliminated, the influence of ionosphere change is considered, and the method is suitable for cycle slip detection in a dynamic environment; according to the method, a self-adaptive threshold model is constructed on the basis of Turbo Edit cycle slip detection in a flight dynamic observation environment, cycle slip misjudgment at a low satellite elevation angle is reduced, and the influence of too low sampling rate on GF combined cycle slip detection performance is weakened to a certain extent; and a technical guarantee is provided for applying the GNSS differential data to the flight test.

The invention provides a vortex generator modified part mounting method for a large civil aircraft, and belongs to the field of test flight verification of a flow control technology of the large civil aircraft. The method comprises the steps that 1, the specific structure and the installation position of the vortex generator are determined; 2, determining the structure of the modified part of the vortex generator; 3, a prefabricated body of the vortex generator modified part is manufactured on the ground; step 4, ensuring that the installation area is prepared; 5, marking the mounting position of the vortex generator modified part; step 6, fixing the vortex generator at a preset position by using a high-strength structural adhesive; step 7, cleaning the surface, and sealing the edge after the impregnating adhesive is completely cured; 8, after installation is completed, a flight test is conducted, and the improvement effect of the vortex generator on the lift coefficient and stall delay is evaluated. The structure and the mounting position of the vortex generator are optimally designed, so that the lift coefficient during aircraft landing is effectively improved, and stall is delayed.

This invention discloses a method and system for real-time monitoring of aircraft flight tests based on telemetry technology. By utilizing telemetry technology and existing aircraft telemetry data communication networks and links to acquire, monitor, and respond to telemetry data in real time, it enables real-time monitoring of the flight test aircraft by the ground system, even when the aircraft does not have ACARS functionality in the early stages of flight testing. It also facilitates monitoring of multiple aircraft and multiple locations for flight tests, as well as remote monitoring. The method and system for real-time monitoring of aircraft flight tests of this invention also have better reliability and can save significant manpower and other costs.

The application provides a fixed-wing aircraft airspeed calibration flight test method based on a standard machine. The method comprises the following steps: determining the weight of the fixed-wing aircraft as a first weight, setting the aircraft to a specified take-off configuration, and taking off according to a take-off procedure; the standard machine flies to a specified airspace according to a plan; at an initial specified flight height, a tow cone is released to a specified length, and the tow cone can obtain an accurate static pressure value in the current area without being affected by a spoiler; the flap and slot wing configuration of the fixed-wing aircraft is set to a first flap and slot wing configuration; the aircraft is trimmed at a given speed; the speed in the above step is set to other speeds within the speed envelope of the first flap and slot wing configuration, and the steps are repeated; the standard machine obtains airspeed errors under the first flap and slot wing configuration and different speeds; and the airspeed errors under the first flap and slot wing configuration and different speeds are fitted to obtain an airspeed correction curve under the first weight, the first flap and slot wing configuration. More accurate airspeed errors can be obtained.

The invention relates to the technical field of signal regulation, and discloses an anti-interference signal dynamic regulation method for an airborne comprehensive regulator, which comprises the following steps: acquiring interference signal data of the airborne comprehensive regulator; classifying the interference signal data based on a preset interference classification model, determining a corresponding dominant interference data type, and calling a corresponding adjustment algorithm according to the dominant interference data type to determine a dynamic adjustment parameter; adjusting the signal data of the airborne comprehensive regulator according to the dynamic adjustment parameter, determining the signal quality after adjustment, and carrying out iterative optimization on the dynamic adjustment parameter by adopting a gradient descent algorithm until the signal quality meets a preset index; and acquiring a flight test stage, presetting an adjustment mode according to the interference characteristics of the flight test stage, and automatically switching the adjustment mode according to the matching degree of the interference signal data and the adjustment mode. According to the anti-interference signal dynamic regulation method, the stability and reliability of the airborne comprehensive regulator are remarkably improved.

The application discloses a route editing method, an electronic device, a system and a storage medium. The method comprises the following steps: displaying a three-dimensional space map of a selected position based on a display of a head-mounted device; and in response to an interactive operation of a remote controller, performing route editing in the three-dimensional space map based on an interactive interface to obtain a flight route of a movable platform. In this way, route editing can be performed in the three-dimensional space map based on the interactive operation of the remote controller, immersive route editing is realized, the intuitiveness of route editing is improved, and compared with a flight test in a real world, the editing efficiency can be effectively improved. In addition, due to a high degree of visualization, the relationship between the route and buildings or terrains in the three-dimensional space map can be effectively checked, and the reliability of route editing is improved.

The invention belongs to the field of helicopter flight mechanics design, and particularly relates to a helicopter floating system air opening safety analysis and verification method. Comprising the steps of 1, determining a security analysis framework; based on the unfolding process of the helicopter emergency floating system, a multi-dimensional safety analysis framework for flight state change analysis, packaging covering cap trajectory analysis and buoy deformation analysis is constructed; 2, determining a test verification scheme; a theoretical analysis result is verified through a wind tunnel test, collision damage verification and a ground simulation test; 3, carrying out a flight test; and the unfolding process of the emergency floating system is tested in actual flight.

The invention discloses a longitudinal large-amplitude self-adaptive virtual flight test device and method, and belongs to the field of special tests for wind tunnels. The top of a vertical guide rail is connected with a spherical hinge support through an upper end buffer device, the spherical hinge support is installed on the upper wall face of a wind tunnel through a connecting base, an aircraft model is connected with the vertical guide rail through a pitching and yawing sliding simulation device, a lower end buffer device is installed at the bottom of the vertical guide rail, and a steel wire rope traction base is installed at the bottom of the lower end buffer device. The steel wire rope traction seat is connected with three groups of front-section steel wire ropes, the front-section steel wire ropes are connected with rear-section steel wire ropes through steel wire rope tensioners, and the rear-section steel wire ropes are connected with the lower wall surface of the wind tunnel, so that the lower ends of the vertical guide rails are in a pre-tightening state. On the basis of three-axis angular motion, the degree of freedom in the vertical direction is completely released, the real aerodynamic force and damping characteristics of aircraft motion can be effectively obtained, the height change in the motion process is reflected, the real flight condition is simulated more truly, and the virtual flight test advantages are further played.

The present application belongs to the water surface take-off and landing control technology field of water surface aircraft, and particularly relates to a kind of amphibious aircraft water surface take-off and landing longitudinal control method.The present application determines the corresponding elevator deflection angle of the aircraft when sliding at the median pitch angle at different speeds through flight test, and when taking off, the corresponding relationship of the flight test speed, median pitch angle and elevator deflection angle is used to draw the elevator control rate curve of the aircraft during the take-off process respectively, so as to obtain the elevator control rate of the aircraft during the take-off process on the static water surface and wave water surface;when landing, the pilot gradually brings the rod after the aircraft touches the water, so that the elevator deflection angle is in the minimum state of the maximum available deflection angle combination until the aircraft completes the deceleration on the water surface.The present application effectively solves the unstable influence caused by the interference of the wave water surface on the take-off and landing of the amphibious aircraft, and improves the take-off and landing safety.

This application discloses a simulation method and system for a launch vehicle attitude control system. The simulation method specifically includes the following steps: a host computer compiles a simulation model and generates an executable program, which is then downloaded to a simulation computer; the host computer runs the simulation model and selects a simulation mode; an industrial control computer simulates the flight test process, uploads the flight control software to the flight control assembly, and starts the process; the host computer controls the simulation computer to send ignition and takeoff command signals to the flight control assembly; the simulation computer calculates and obtains attitude angle and attitude angular velocity measurement information and sends it to the flight control assembly; the flight control assembly operates after receiving the ignition and takeoff command signals, issuing actuator control commands based on the received measurement information. This application achieves step-by-step, layered verification of the theoretical controller, flight control software algorithm, flight control assembly, and servo mechanism, facilitating problem localization and analysis, and ensuring the safety and reliability of the experimental process.

The invention belongs to the technical field of network testing, and particularly relates to a TTE network systematic testing method for an airborne system of an air-space platform. Comprising the following steps: checking network appearance; checking the effectiveness of the electrical characteristics of the network; checking the validity of the communication protocol; checking the effectiveness of switch communication; performing fault injection; and checking the flight state. According to the air-space platform airborne system TTE network systematic test method, TTE network conventional inspection items are covered, test flight requirements are met, meanwhile, the tail end of a development process to which test flight belongs is considered, and customized design is carried out on a communication protocol, fault injection and the like in combination with system characteristics, so that correctness and integrity of the TTE network before test flight are ensured.

The application belongs to the technical field of flight test, and particularly relates to a high-temperature strain gauge measurement error correction method and system. The method comprises the following steps: S1, obtaining a first strain variable of a specified component in aircraft test through a high-temperature strain gauge; S2, obtaining a measuring point temperature and a temperature rise rate of the specified component based on a thermocouple arranged on the specified component; S3, determining a second strain variable caused by temperature according to the measuring point temperature and the temperature rise rate in a preset interpolation table, wherein the interpolation table gives the relationship between the strain variable and the temperature under different temperature rise rates; and S4, subtracting the second strain variable from the first strain variable to obtain a third strain variable of the specified component caused by load. The application can accurately calculate the strain of the thermal structure of the aircraft under high-temperature conditions, and provide benchmark data for heat protection and heat transfer calculation of key components on the full flight envelope.

The invention relates to a test point cooperation system and method applied to civil aircraft testing, and the method comprises the steps: building a mapping relation between the name of each civil aircraft part and the position of the civil aircraft part in a civil aircraft model, and setting a test refitting model library module comprising alternative test refitting models; therefore, when a user inputs the test point selection information, a corresponding model matched with the to-be-tested civil aircraft part can be generated, and a model formed by loading the corresponding test modification model on the to-be-tested civil aircraft part can be generated according to the model loading instruction, so that data transmitted by a system can be transmitted on different devices in a model format. Therefore, more visual information data can be obtained among departments through the received model data. The testing point cooperation system and method applied to civil aircraft testing can effectively improve intuition, are convenient to operate, are suitable for cross-department cooperative work, guarantee the convenience and accuracy of test flight test refitting design, and effectively improve the efficiency.

The present application relates to the technical field of flight test, in particular to a high-speed target acquisition system equipment for high-speed aircraft, which comprises a high-speed cavity and a high-speed cover, the high-speed cavity is provided with a high integration mechanism on the surface of a high-speed connector panel, and a sealing and waterproof mechanism is arranged between the high-speed cavity and the high-speed cover and between the high-speed cavity and the high-speed connector panel.The present application not only enables the equipment to be soaked in water during use, ensures that data will not be damaged due to water ingress, and can resist the influence of corrosive liquids such as seawater and harsh environments in space missions, but also can simplify complex process flow, improve production efficiency, reduce cost, improve product quality, better meet the increasingly complex and diversified demands of modern industrial field, realize the integration of acquisition, transmission and communication in limited space, and meet the space limited requirements of aircraft.

The invention discloses a rigid-elastic coupling aeroelasticity control flight test platform, and relates to the technical fields of unmanned aerial vehicles, mechanics, aeroelasticity, control and the like. The rigid-elastic coupling aeroelasticity control flight test platform comprises a double-triangle high-aspect-ratio sweepback tailless flying wing pneumatic structure with a rigid-elastic coupling effect, a lightweight aircraft structure design and aeroelasticity control airborne equipment. The lightweight aircraft structure design comprises fuselage structure design, wing structure design and fuselage-wing mounting mechanism design; the aeroelastic control airborne equipment comprises control equipment, communication equipment and power equipment. The rigid-elastic coupling aeroelasticity phenomenon can be obviously captured and effectively inhibited in a flight test, a reliable test platform is provided for rigid-elastic coupling aeroelasticity mechanism analysis and inhibition technology verification, and a foundation is laid for deep research of an aeroelasticity control technology.

The invention belongs to the technical field of helicopter safety design, and discloses a helicopter emergency floating system buoy cabin safety analysis method and system, and the method comprises the following steps: precisely determining a severe working condition and an aerodynamic state at a buoy cabin according to the flight working condition of a helicopter; according to the severe working condition and the aerodynamic state of the buoy cabin, the buoy cabin cover and the connecting strength of the buoy cabin cover are subjected to stress analysis and checking; carrying out buoy cabin deformation simulation calculation analysis to obtain buoy cabin deformation; carrying out air blowing and buoy cabin strength test examination under a harsh flight working condition; a buoy inflation expansion test is carried out, and the buoy cabin deformation condition is actually measured; installing a buoy cabin on the helicopter, and carrying out flight test system verification. The problem that deformation characteristics and pneumatic safety of the hard buoy cabin are not systematically analyzed is solved.

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