172 results about "Flight test" patented technology

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 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 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.

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 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 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 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 invention discloses an envelope extension type flight test state point risk level assessment method, which comprises the steps of designing a test flight state point of an envelope extension test flight subject according to information such as an initial safety value and a limit value of a to-be-extended envelope (speed, overload, angle of attack and the like) involved in the test flight subject; establishing a risk measurement value calculation method of a single envelope extension test flight state point in a flight test; based on the risk measurement value of the single envelope extension test flight state point, establishing a calculation method related to the risk measurement value of a plurality of envelope extension test flight state points in the flight test; designing a risk grading scale, and visually representing the risk grade of the test flight state point through the grade and the color tape; obtaining the risk level of the test flight state point through the position of the test flight state point risk measurement value in the risk analysis scale; and the envelope extension type flight test flight test state point optimization method based on the flight test state point risk level is formed. According to the embodiment of the invention, the problem of evaluation of the risk level of the test flight state point in envelope extension type test flight design is solved, and test flight design personnel can carry out test flight design optimization according to the visual risk level, so that the test flight progress, the test flight cost and the test flight risk are balanced.

The invention discloses a method for constructing a simulation model for a target test flight task. The method comprises the steps of determining a model architecture and a parameter set based on a target test flight task; constructing an initial simulation model for the target test flight task based on the model architecture and the parameter set, wherein the initial simulation model for the target test flight task comprises a conventional sub-model and a key sub-model; performing sensitivity analysis on parameters associated with the key sub-model in the parameter set to screen out one or more sensitive parameters, wherein the influence of the one or more sensitive parameters on the precision of the simulation model facing the target test flight task exceeds a preset threshold value; and optimizing the initial simulation model for the target test flight task by using a parameter optimization algorithm based on the one or more sensitive parameters to obtain an optimized simulation model for the target test flight task. According to the method, model focusing can be performed according to the specific target test flight task, and the parameters are screened for accurate optimization, so that an accurate and reliable simulation model is provided for the specific target test flight task.

The invention belongs to the technical field of aircraft fire extinguishing design, and provides a large amphibious aircraft water throwing efficiency data expansion method, which comprises the following steps: carrying out a pre-flight test of an aircraft to obtain test state points of aircraft water throwing efficiency, including typical test state points and atypical test state points; carrying out water throwing efficiency tests on different test state points, summarizing and analyzing test data, carrying out water throwing efficiency tests on test state points corresponding to invalid data again, summarizing the test data again, and fitting the test data to obtain different water throwing efficiency relation curves; and typical test state point results are extracted to fit coefficients of the water throwing efficiency relation curve, the coefficients are combed, a change relation curve of the water throwing efficiency and different test state points is constructed, and a change curve of the water throwing efficiency along with the water throwing height under each test state point of the airplane is expanded. According to the method, data validity judgment can be improved, the test process is accelerated, and unpredictability of environment variable control caused by an overlong test period is avoided.

The invention provides a flight potential standardization evaluation method based on rapid teaching and flight path assistance, and relates to the technical field of flight simulation evaluation. The method comprises the following steps: firstly, establishing a target capability model including understanding and accepting capability, control capability, attention capability and strain capability, and compiling a simulated flight test according to a cognitive diagnosis theory; constructing a scene library based on the target capability model, configuring track assistance for each flight scene, and marking excitation intensity and discriminability parameters of each flight scene for each capability; short-time rapid pre-teaching is implemented according to a standardized teaching script before evaluation; during evaluation, a multi-capability probe is embedded in a flight scene, evaluation data is collected, a capability estimation value is updated based on a capability-scene matching model, a subsequent flight scene is adaptively selected or scene parameters are adjusted according to the capability estimation value to generate a personalized examination path, a comprehensive score is calculated, and a four-dimensional capability profile map is output. The objectivity of the evaluation result can be improved.

The present application belongs to the technical field of flight test, and particularly relates to a kind of troposphere manned airship fuel system hot climate test flight method. Including: S1: determining test flight condition;S2: make airship climb to the minimum safe height after stable flight first time, gradually increase the power of transmitter to maximum continuous power;S3: after speed stabilization, climb to maximum use height at maximum continuous power;S4: monitor fuel pressure, fuel flow and engine operation during climbing, monitor whether there is gas lock phenomenon caused by fuel overheating, which leads to unstable engine operation;S5: if there is no gas lock phenomenon caused by fuel overheating, which leads to unstable engine operation, airship climbs to maximum use height, returns and lands;S6: if fuel pressure is within normal range, fuel flow is within normal range, and engine does not work abnormally due to fuel interruption during test, then fuel system hot climate test flight is completed.

The present application belongs to the field of flight test technology, and relates to a kind of indication recording system test flight method suitable for troposphere manned airship.The method comprises: hovering 360 in the three time periods of sunrise, strongest sunlight and sunset respectively;In the above hovering process, it is observed whether the display data displayed by the display is easy to see under different angle illumination conditions;If there is an abnormal working condition in the above process, optimization improvement suggestions are proposed for it, and then the above steps are repeated until the qualified criterion is met.

The invention belongs to the technical field of unmanned aerial vehicle design, and particularly relates to an open-loop state large-angle-of-attack range longitudinal aerodynamic coefficient identification calculation method, which comprises the following steps of: comparing simulation data with test data in an actual flight test, calculating an error sum of squares of the simulation data and the test data, and minimizing the error sum of squares; dividing the whole attack angle range into a plurality of sub-intervals, respectively obtaining the optimal aerodynamic parameters of each sub-interval based on the minimum sum of squares of errors, inputting each optimal aerodynamic parameter into the flight simulation model, comparing the optimal aerodynamic parameters with test data for verification, and executing the next step when the difference value between the optimal aerodynamic parameters and the test data is within a set range; and drawing an aerodynamic characteristic curve according to each optimal aerodynamic parameter, verifying the applicability of the flight simulation model, and optimizing the robustness of the flight simulation model. The method is suitable for longitudinal aerodynamic coefficient parameter identification of different types of unmanned aerial vehicles, the parameter identification efficiency is improved, and meanwhile the applicability of the model under the complex flight working condition is enhanced.

The application discloses a piezoelectric guided wave online damage alarm method for an aircraft test flight structure, and belongs to the field of aircraft structure health monitoring, and comprises the following steps: constructing a guided wave sensing network; obtaining a reference feature sample set by acquiring a reference feature through a variable step window sliding method; performing probability statistical modeling on the reference feature sample set, obtaining a probability statistical correlation feature and a reference feature sample set residual error, and taking an absolute value to obtain a reference alarm feature; establishing a monitoring feature sample set, multiplying the monitoring feature sample set and the probability statistical correlation feature to obtain a monitoring feature sample set residual error, and taking an absolute value to obtain a structure damage alarm feature; performing network weighted fusion on the structure damage alarm feature, determining a damage high-probability channel according to a channel screening threshold, calculating a structure damage alarm threshold of the damage high-probability channel, and realizing structure damage alarm. The method is simple and efficient in implementation process, and realizes reliable alarm of structure pit damage under the influence of time-varying factors.

The application provides a five-degree-of-freedom virtual flight test device based on motion compensation, wherein a test model of the device can complete three-axis rotation motion with the help of a three-degree-of-freedom hinge, the test model and an abdominal support strut can move up and down along the axial direction of a three-pronged suspension support to complete the release of the heave degree of freedom, and the test model and the three-pronged suspension support can realize lateral undamped motion under the support of a pneumatic suspension bearing, so as to achieve the purpose of releasing the lateral translation degree of freedom. A parallel six-degree-of-freedom mechanism base at the bottom is responsible for motion compensation of the lateral translation, heave, roll and pitch degrees of freedom, and based on the above structure, the device can realize the five-degree-of-freedom coupled motion test capability except the forward and backward motion. The realization of the above functions has important significance for improving the restoration capability of the virtual flight test of a wind tunnel to the real dynamic characteristics of an aircraft model.

The application provides a neuromuscular coordination ability evaluation and training system with an environment simulation function, which introduces the movements of six degrees of freedom of the subject environment, such as up and down, left and right, and pitching, to simulate flight movements, so that the perception effect of the subject on the environment during the test and training process can be enhanced, and the result is closer to the result in the flight test; four degrees of freedom of the upper and lower limb interaction device and nine test paradigms are designed, the coordination test of two dimensions of coordination and reaction is involved, the test of multiple dimensions and indexes is involved, and the reliability and validity of the test can be improved; five single abilities of coordination ability and four single abilities of reaction ability are designed, the comprehensive ability of two dimensions is obtained through a weight coefficient, and all the results are recorded and managed; according to the daily record of the single abilities, targeted training tasks can be carried out, and the training efficiency is greatly improved; during the test process, the load can be parameterized and adjusted, and the difficulty of the test can be quantitatively controlled.

This application provides a method and system for determining risk sources of distributed electric aircraft based on flight profiles, applied in the field of aircraft simulation technology. It includes: acquiring the overall construction content of the flight profile and topology of the distributed electric aircraft, and obtaining typical operating conditions for each stage in multiple phases corresponding to the flight profile. The overall construction content includes a high-power-density battery pack, a DC-DC buck converter, an electronic speed controller, a DC brushless motor, and rotors / propellers; simulating the mathematical model of the overall construction content under the target typical operating conditions of the target phase, obtaining simulation results for the target typical operating conditions; and accurately determining the flight test risk sources of the distributed electric aircraft under the target typical operating conditions from the high-power-density battery pack, DC-DC buck converter, electronic speed controller, DC brushless motor, and rotors / propellers based on the simulation results. This significantly reduces the flight risk of new aircraft configurations (i.e., distributed electric aircraft) and improves flight test efficiency.

This application provides an airborne integrated data acquisition and recording system and method, applicable to the field of aviation flight testing. The system includes: for each data acquisition module, each data acquisition unit acquires target data under the current flight state; each data processing unit determines the cache priority of the target data based on the importance of the data type attributes of the data acquisition module under the current flight state, as well as the data transmission rate and data change entropy of the target data; wherein the importance of the data type attributes of the data acquisition module differs under different flight states; each data output unit sends the target data and its cache priority to the main control module; the main control module receives the target data and corresponding cache priorities sent by multiple data output units; and caches multiple target data in descending order of cache priority. This enables real-time and reliable acquisition and recording of multi-source heterogeneous data in complex flight test missions.

The utility model relates to the technical field of unmanned aerial vehicles, and discloses a mounting mechanism for a hang-flight test, which comprises a fixing plate, a fixing mechanism is mounted on the outer side of the fixing plate, a balancing weight is fixedly connected to the outer side of the fixing mechanism, a weight bearing mechanism is mounted at the bottom of the fixing plate, and a supporting mechanism is mounted at the bottom of the weight bearing mechanism. The fixing mechanism comprises a connecting assembly, a telescopic assembly, a positioning assembly and an adjusting assembly, the telescopic assembly comprises a supporting plate, the supporting plate is slidably connected to the interior of the fixing plate, a fixing strip is fixedly connected to the outer side of the supporting plate, and a positioning hole is formed in the outer side of the supporting plate. According to the mounting mechanism, the fixing plate, the supporting plate, the two-way threaded rod, the worm gear, the worm and other structures are matched with one another to drive the sliding block and the arc-shaped block to be adjusted according to the size of the unmanned aerial vehicle support, so that the mounting mechanism is conveniently mounted on the unmanned aerial vehicle, and the mounting mechanism can be mounted on the unmanned aerial vehicle without the support under the cooperation of the connecting plate and the fixing bolt structure.

Systems and methods of evaluating aircraft behavior. The systems and methods perform simulated testing using computer-based models to determine predicted behavior of the aircraft. The systems and methods further use measurement data collected during operation of a test aircraft to determine actual behavior of the aircraft. The systems and methods provide for comparing the predicted and actual behaviors to provide for more accurate analysis of the aircraft. In some examples, the systems and methods provide for live flight test optimization which eliminates and / or reduces the need for extraneous flight testing.

The invention belongs to the technical field of aircraft flight tests, and discloses an aircraft control surface jamming fault test method and a test modification system thereof.In a flight test flight test, a control surface jamming function is started, then a fault is injected into an aircraft transmission matching actuator control instruction, and an actuator instruction bias is formed in combination with an original actuator control instruction; and after the jamming bias position is set, the jamming bias position and the initial jamming position are incrementally input into an actuator instruction. On the premise of not changing the entity configuration of airborne equipment, a data transmission link is realized through the flight test control panel, the flight test interface unit and the flight control module; a flight test control panel is used as instruction input, and control plane jamming and jamming parameters are selected; and a flight test software module version of the flight control module is taken as a reference, and a control instruction is issued to realize jamming of different control surfaces.