170 results about "Flight envelope" patented technology

The invention relates to the technical field of unmanned aerial vehicle control, and discloses an unmanned aerial vehicle inspection control system and method based on data analysis, and the method comprises the steps: receiving environment feature data of a target region, and constructing a static environment model; obtaining kinetic parameters of the unmanned aerial vehicle, and calculating a feasible flight envelope in combination with the static environment model; after taking off, the unmanned aerial vehicle collects environment data in real time, and a multi-level environment topology model is constructed; acquiring state data in real time, and judging an obstacle avoidance level; if yes, local path planning is carried out, and an optimal flight path is generated; acquiring trajectory deviation and energy consumption data, dynamically calibrating the flight envelope, and optimizing the flight envelope; according to the method, environment perception, trajectory optimization and online learning are combined, the method is adaptive to a complex environment, and the intelligent level of unmanned aerial vehicle inspection is improved.

The invention provides a low-altitude airspace collaborative management method and system based on digital twinning, and the method comprises the steps: analyzing the maneuvering characteristics and flight envelope characteristic parameters of an aircraft based on the obtained real-time state data of the aircraft, and calculating the maneuvering characteristics and flight envelope characteristic parameters of the aircraft by using kinematics and dynamics models, describing motion laws of different aircrafts; real-time flight data of the aircraft are combined, the motion track and attitude change of the aircraft are synchronized in real time in a virtual environment through a digital twin technology, the dynamic behavior of the aircraft is accurately reproduced, and a high-fidelity digital model is provided for flight conflict detection and flight path prediction; a multi-level and multi-granularity digital twin model is constructed according to the difference of the motion characteristics of the aircrafts, and the digital twin model comprises an airspace structure model of a macroscopic level, an aircraft group behavior model of a mesoscopic level and a single aircraft dynamic model of a microscopic level.

The invention provides a reusable rocket landing stage starting height prediction method, a guidance system and a processor, and aims to dynamically and accurately determine the ignition time of an engine by constructing a real-time online starting height prediction model based on an energy conservation principle and combining multiple safety thresholds (including an allowable starting height upper limit and a standby starting height lower limit) so as to predict the starting height of the rocket in a landing stage. And the safety and fuel optimality in the whole flight envelope are ensured, so that the reliability, accuracy and economical efficiency of the reusable rocket landing are remarkably improved. The problems that in the prior art, engine starting height prediction precision and robustness are insufficient, online adaptation to uncertainty of a complex flight environment cannot be achieved, and safety margin consideration is insufficient are solved, and potential safety risks are thoroughly eliminated.

The invention provides a combined dynamic inspiration mode flight envelope design and trajectory optimization method, and the optimization method comprises the following steps: obtaining a flight trajectory design input condition; analyzing the working envelope, flight overload and dynamic pressure flight path design constraints of the air suction power engine based on task requirements; resolving an upper attack angle section alpha up (t) meeting the constraint conditions to obtain an upper boundary of the flight envelope in the height-speed section; resolving a lower attack angle profile alpha down (t) meeting the constraint conditions to obtain a lower boundary of the flight envelope in the height-speed profile; and determining a flight trajectory in the upper and lower boundaries of the height-speed profile, namely, meeting the multi-constraint design condition of the trajectory of the air suction powered aircraft.

The invention discloses a method for rapidly predicting a thermal environment of an aircraft skin structure along a flight envelope, and belongs to the technical field of structural thermal environment prediction. The problems that in the prior art, thermal environment distribution data obtained through a traditional aircraft skin structure thermal environment rapid prediction method is low in precision and low in calculation efficiency are solved. The method comprises the following steps: discretizing a flight envelope to form a flight envelope point set, carrying out aerodynamic heat environment prediction, and carrying out aerodynamic heat data numerical simulation by utilizing a CFD numerical calculation method; according to different wall surface temperatures and heat insulation wall temperatures, a heat flow wall surface temperature correction model is constructed; thermal environment simulation of the aircraft skin structure along the flight envelope is carried out, the extracted temperature of the aircraft skin structure is corrected through a heat flow wall surface temperature correction model, propulsion calculation along the flight envelope time dimension is circularly carried out, and rapid prediction of the thermal environment is achieved. The method effectively improves the accuracy of prediction data, and can be applied to the thermal environment prediction of the hypersonic aircraft structure.

The invention provides a method and device for optimizing the design of a hybrid power system of an aircraft. The method provided by the invention comprises the following steps: aiming at a pre-designed hybrid power system, constructing a simulation model corresponding to the hybrid power system; determining a required power range of a power system according to the flight envelope of the aircraft, and determining load required power of the aircraft under multiple types of preset typical working conditions according to the required power range; determining a working mode of the aircraft under the load demand power according to the load demand power under each type of preset typical working conditions and current and preset working mode determination rules; simulating the working mode by using the simulation model, and obtaining a simulation result; wherein the simulation result comprises operation parameters corresponding to each model; and according to a simulation result, providing an optimization suggestion for a pre-designed aircraft hybrid power system. The method provided by the invention is used for effectively optimizing the design of the hybrid power system of the aircraft and improving the performance and reliability of the hybrid power system.

The invention belongs to the technical field of hypersonic flight vehicle control, and relates to an intelligent control method for a hypersonic flight vehicle driven by coupling information. Firstly, a longitudinal dynamic model of the elastic hypersonic aircraft is established; then, a coupling database based on sampling analysis is constructed, different flight envelope data are simulated, a coupling degree matrix is calculated through a sampling statistical algorithm, and the coupling database is formed to quantify the mutual influence between variables; furthermore, deep neural network training driven by coupling data is carried out, a long-short term memory network is adopted, and a coupling database is used as a data sample to carry out offline training, so that the autonomous fine control capability of the aircraft in a strong coupling environment is improved. And finally, designing an online controller to realize a coupling information intelligent compensation scheme, dividing aircraft maneuvering control into a speed loop and an attitude loop, and designing a corresponding control instruction and a neural network updating law to realize real-time and intelligent flight control.

The invention discloses an attitude-elasticity-pneumatic coupling flutter control method of a flying wing layout aircraft, and relates to the technical fields of mechanics, aeroelasticity, control and the like. The control method comprises the steps that a Lagrangian equation under a general body shafting is synthesized, a full-quantity coupling dynamic model is established, and a disturbance motion dynamic equation used for flight control system design is established through small disturbance linearization; correcting the disturbance motion kinetic equation based on a system identification flight test; designing an attitude-elastic-pneumatic coupling flutter suppression controller based on a traditional or modern control theory, and embedding the attitude-elastic-pneumatic coupling flutter suppression controller into a flight control system; the effectiveness of the flutter suppression controller is verified through a flight test. The problems that existing numerical analysis and wind tunnel experiments cannot accurately reflect the complex coupling effect in flight, flutter cannot be effectively restrained in flight, and the flight envelope cannot be expanded are solved, and the flight safety and stability of the flying wing layout aircraft can be effectively improved.

The invention belongs to the technical field of numerical simulation and analysis of unmanned aerial vehicles, and relates to a rigid-flexible coupled variable-camber-range deformable wing optimization design method, which comprises the following steps of: obtaining working conditions of an aircraft, determining wing types and basic parameters of wings of the aircraft, and establishing a three-dimensional aerodynamic model of the wings; variable-attack-angle pneumatic simulation is carried out on working conditions of the aircraft in take-off and cruise, optimal camber values and corresponding pneumatic loads of wings under multiple working conditions are determined, and the positions of rotating joints of the deformed wing ribs are adjusted to obtain the deformed wing ribs after primary optimization; dividing the optimization range of the deformed wing rib after primary optimization into a fixed area and an optimizable area, loading an aerodynamic load under the optimal camber value in the optimizable area to perform aerodynamic simulation to obtain a deformed wing rib after secondary optimization, and optimizing the structure; and a deformed wing rib sample is made, performance testing is conducted, and the optimizable area is adjusted till the wing performance meets the preset requirement. The aerodynamic efficiency of the full flight envelope can be improved.

The invention relates to the technical field of aircraft control, and discloses a flight program planning method and device for an aircraft, equipment and a medium, and the method comprises the steps: obtaining a nominal flight program and a corresponding flight envelope function which is generated in a preset profile domain based on a process constraint condition and a terminal constraint condition of the aircraft; key nodes are selected in the speed domain of the nominal flight program, random quantities are overlaid on all calculation nodes based on a random walk strategy, and an overlaid result is generated; performing constraint correction on the superposition result through a flight envelope function to generate a corrected attack angle node value; and performing polynomial interpolation fitting on the corrected attack angle node value to generate a random walk flight program. According to the method, the random walk flight program with randomness and constraint conformity can be generated, so that the track unpredictability of the hypersonic gliding aircraft is enhanced, and diversified combat requirements of the aircraft such as high maneuverability and strong penetration are met.

The embodiments of this specification provide an adaptive control method for an aircraft and a computing device, which determine the real aircraft state and the simulated aircraft state corresponding to the current control period. The real aircraft state is determined by an on-board sensor, and the simulated aircraft state is determined by a reference model. An adaptive controller is used to determine the control target compensation amount corresponding to the current control period according to the difference between the real aircraft state and the simulated aircraft state. According to the control target compensation amount and the initial control target corresponding to the current control period, the corrected control target corresponding to the current control period is determined. An inner-loop controller is used to control the aircraft according to the corrected control target and the real aircraft state, and the reference model is used to simulate the aircraft state of the next control period according to the corrected control target and the simulated aircraft state. It can provide accurate control commands under the conditions of large flight envelopes and strong couplings, and improve the anti-interference ability of the aircraft.

The invention provides a supersonic-speed jet adopting supersonic-speed double wings and variant design, and belongs to the field of novel supersonic-speed civil aircrafts, a large-area truncated delta wing with edgings on the inner side is adopted, an engine nacelle is arranged on the upper portion of the outer end of the large-area truncated delta wing, and the supersonic-speed double wings are installed on the outer side of the nacelle; the front edges and the rear edges of the upper wing and the lower wing of the supersonic double-wing adopt intelligent materials to realize a variant function; the empennage can rotate around a mounting shaft in the longitudinal direction of the airplane to change the inclination angle; the front edge and the rear edge of the supersonic double-wing are deformed, and the inclination angle of the empennage is changed to jointly form a supersonic state and a subsonic state. The problem that the conventional double-wing non-design-point performance is poor is solved, high aerodynamic performance can be achieved within the flight envelope range, and it is guaranteed that an aircraft achieves ideal supersonic speed environment-friendly performance such as acoustic explosion and noise and aerodynamic efficiency.

The invention relates to the technical field of data processing, and discloses an aircraft hydraulic pipeline system fault prediction method and system based on digital twinning. The method comprises the steps that flight parameters are acquired through an aircraft data bus for stage identification, and four flight stage identifiers are obtained; carrying out mapping processing on the load spectrum according to the identifier, and constructing a staged digital twin pipeline model; performing weight distribution on the multi-sensor data according to identifiers, and extracting fault feature vectors according to a twin model; and performing prediction processing through the BP neural network in stages to obtain a fault prediction result. The method solves the problem that the existing digital twinborn model cannot adapt to the full flight envelope load difference by adopting fixed parameters, so that the sensor data weight distribution lacks flight stage adaptability, and the fault prediction algorithm cannot accurately identify the differentiated fault features of different flight stages.

The invention relates to the technical field of aircraft trial, and provides a universal test flight platform of a turboshaft engine, which comprises an aircraft, a first test cabin, a first power pack, an independent reduction gearbox and a propeller, two wings are respectively arranged on two sides of the aircraft, the first test cabin is arranged on one wing, the first test cabin and the first power pack are positioned on the same side of the aircraft, and the independent reduction gearbox is arranged on the first test cabin. The first power pack is arranged at the joint of a wing and the aircraft, the first test chamber is provided with a containing cavity, a mounting part is arranged in the containing cavity and used for mounting an engine and enabling the axial direction of the engine to be consistent with that of the containing cavity, the engine is connected with the first power pack through the first test chamber, and the independent reduction gearbox is arranged in the containing cavity. The propeller is arranged outside the containing cavity and located at the end of the first test chamber, and the engine is in transmission connection with the independent reduction gearbox and the propeller in sequence. The turboshaft engine is well matched with the test flight platform, test flight verification of the whole flight envelope of the turboshaft engine is met, and the test flight platform has the characteristic of universality.

Systems, devices, and methods including an aerial vehicle having a global positioning system (GPS) and at least one trace-gas sensor configured to generate gas data; and a processor having addressable memory, the processor configured to: determine a flight envelope based on a received spatial location, a received spatial location of the one or more potential gas sources, a received desired level of confidence, and a received wind data; determine a flight path for the aerial vehicle, where the flight path covers a portion of the determined flight envelope; and determine based on a received gas data whether a gas leak is present in the received spatial location to the received desired level of confidence.

The invention belongs to the field of flight tests of unmanned aerial vehicles, and particularly relates to a calibration method for an airspeed system of a round-trip flight unmanned aerial vehicle with a fixed course angle. Comprising the following steps: step 1, selecting a test point in an aircraft flight envelope; step 2, the aircraft flies back and forth under the condition corresponding to the test point, and flight data of the going path and the returning path in the stable level flight stage are obtained; step 3, respectively selecting a plurality of flight data points of a going path and a returning path from the flight data, acquiring a course angle and a track angle of each flight data point, and according to the course angle and the track angle, judging whether the non-uniformity of the wind field in the flight process meets the requirement; if so, entering step 4; if not, returning to the step 2, and re-executing the test point; 4, acquiring the ground speed of each flight data point, and calculating the wind speed according to the ground speed; 5, calculating the vacuum speed in the flight process according to the wind speed and the ground speed; and 6, calibrating the air speed of the test point according to the vacuum speed.

The present invention relates to the technical field of engine condition monitoring, and particularly relates to an engine performance prediction method using sample adaptive weighting, which comprises the following steps: establishing a first data set and a second data set, wherein the first data set is the already-operated data set of the current engine, and the second data set includes the operation data sets of N other engines of the same model as the current engine; clustering the first data set and the second data set according to the flight envelope distribution to obtain M clustering centers; establishing sub-models such that each of the M clustering centers includes N sub-models; initializing a prediction weight vector W for each of the clustering centers, and performing weighted averaging on the N sub-models included therein to construct M linear weighted prediction models, which are used to predict engine performance; and optimizing the prediction weight vector W of the clustering centers. This method gives full play to data diversity and meets the requirements for accurate performance prediction of aero-engines.

The invention discloses a method for analyzing lubricating oil recovery efficiency in an aero-engine hot-end bearing cavity oil return process, and belongs to the field of aero-engine bearing cavity structure system design. The method comprises the steps that the rotating speed and the pressure change range of a bearing cavity are determined according to the flight envelope of an aircraft; and for each rotating speed and pressure state point, an oil return coefficient, an invalid oil return ratio, an oil return pipe inlet speed parameter and a stability characterization parameter thereof are obtained through simulation analysis. And through dimensionless processing and reasonable weighting, comprehensive efficiency parameters of the oil return efficiency of the bearing cavity under the rotating speed and the pressure are obtained, so that quantitative characterization of the oil return efficiency under a single working condition is realized. Furthermore, in the rotating speed and pressure change range of the bearing cavity, weighted average is carried out on the comprehensive efficiency parameters of the multiple key state points, and comprehensive parameters used for evaluating the oil return efficiency of the bearing cavity in the whole flight envelope are obtained. According to the invention, guidance can be provided for optimal design of a bearing cavity oil return structure, and a bearing system is supported to realize high reliability.

The invention provides a supersonic jet adopting supersonic double wings and belongs to the field of novel supersonic civil aircrafts, an S-shaped bent large slenderness ratio variable cross-section fuselage is adopted as a fuselage, large-area truncated delta wings with edgings on the inner sides are adopted as inner sections of wings, engine nacelles are arranged on the upper portions of the outer ends of the wings, the small-area supersonic double wings are installed on the outer side of the engine nacelle, the tail wing adopts the design of a single vertical tail and an inverted V tail, the acoustic explosion intensity can be reduced, meanwhile, the functions of the wing inner section, the supersonic double wings and an engine nacelle supporting structure are achieved, the structural efficiency is effectively improved, and the structural weight is reduced. By the adoption of the supersonic double-wing design, it can be guaranteed that the aircraft achieves ideal environment-friendly performance and aerodynamic efficiency, the front edge and the rear edge of the supersonic double wings are movable, the problem that the performance of non-design points of conventional double wings is poor is solved, and the aircraft can have high aerodynamic performance within the flight envelope range.

The invention provides an adaptive construction method for a damage probability model fusing knowledge migration, and relates to the technical field of control and regulation, and the method comprises the steps: constructing a feature index of multi-physics field fusion, introducing knowledge migration guided by physical information, and building a prediction-control integrated linkage mechanism, thereby achieving the adaptive construction of the damage probability model. The accuracy and reliability of dynamic instability prediction are remarkably improved, the prediction result is directly applied to active safety control, and a seamless closed loop from intelligent sensing to active intervention is achieved. On the premise that hardware cost is not increased, through upgrading of a software algorithm, the safe flight envelope of the aircraft is effectively widened, and the dynamic robustness of the system in a complex extreme environment is enhanced.

The invention discloses a method for calculating a coupling thermal environment of a hypersonic flight vehicle along a flight envelope, and belongs to the technical field of hypersonic flight vehicle design. The problem of thermal environment calculation of long-time dynamic flight along a flight envelope is solved. The method comprises the following steps: constructing an aerodynamic heat calculation model; constructing a structural thermal environment calculation model for calculating coupling wall surface temperature data; constructing a thermal radiation calculation correction model for calculating the influence quantity of outward thermal radiation on the heat flow; constructing a coupling calculation model for realizing coupling boundary data transmission; on the basis of the models, a coupling boundary data transmission method is built, next-moment propulsion calculation of a flow field and a structure field is carried out, an internal pseudo-time iteration mode is adopted till the iteration convergence requirement is met, and then a hypersonic speed coupling thermal environment data calculation result is obtained. According to the invention, the tight coupling and accurate data transmission process of multiple physical fields can be realized, and the problems of coupling thermal environment, thermal management, thermal protection and accurate prediction of thermodynamic elasticity of the hypersonic aircraft are solved.

The invention discloses a flight safety warning method based on pilot operation intention recognition, which comprises the following steps: S1, collecting EEG data of a subject in each stage of simulated flight through an electroencephalogram signal data collection device; s2, carrying out preprocessing operation on the collected EEG data; s3, extracting and fusing the electroencephalogram signals by constructing a convolutional gradient attention network, and mapping fused features to an operation intention of a pilot; s4, calculating safety risk factors according to the deviation between the flight state and the preset flight envelope, the pilot operation intention and the external environment factors; s5, comparing the risk factor with a preset risk threshold, and dividing the risk into a high level, a middle level and a low level; and S6, the cockpit touch alarm system triggers multi-mode alarm in the cockpit according to the risk level, and adjusts the flight state in real time. The flight accident risk caused by human factors can be reduced, and the aviation safety is improved.

The invention discloses a lubricating oil flow state analysis method for an aero-engine bearing in the oil conveying-oil throwing process, and relates to the technical field of aero-engine bearing lubrication. The method comprises the steps that the rotating speed and lubricating oil flow state points of a bearing cavity are determined according to the flight envelope of an aircraft; aiming at a specific rotating speed and flow state point, obtaining an inlet and outlet pressure difference in the oil transportation-oil throwing process of the state point through simulation analysis, and carrying out quantitative characterization on the lubricating oil flowing state in the oil transportation-oil throwing process of the bearing under a single working condition by using the parameter; and on the basis of the inlet and outlet pressure difference in the oil transportation-oil throwing process of the multiple state points, through reasonable weighted averaging, the characterization parameters of the lubricating oil flow state in the bearing oil transportation-oil throwing process in the whole flight envelope are obtained for comprehensively evaluating the characterization parameters of the lubricating oil flow state in the bearing oil transportation-oil throwing process. The method can be used for guiding the optimal design of a bearing oil conveying-oil throwing structure, so that the cooling and lubricating effects of the bearing are improved, and the reliability of a bearing system is improved.

An aircraft has a first principles takeoff processor (PCE), a predictive flight envelope protection processor (PFEP), and a maximum takeoff weight processor. The PCE is programmed to predict a liftoff location and an energy state of the aircraft at a liftoff on a runway. The PFEP is programmed to assess each of a plurality of potential trajectories for compliance with or violation of a predetermined flight envelope. The maximum weight processor is programmed to: indicate that the aircraft may takeoff at the aircraft weight when any one of the plurality of potential trajectories is in compliance with the predetermined flight envelope; iteratively reduce an input of the aircraft weight to the PCE until the PCE indicates that any one of the plurality of potential trajectories is in compliance with the predetermined flight envelope; and indicate that the input of the aircraft weight as reduced is a maximum allowable takeoff weight.

The invention aims to provide a method for formulating an airworthiness conformity verification test matrix, which comprises the following steps of: continuously updating a test data set by taking the prediction precision of a machine learning model on a test set as an index in a manner of combining the machine learning model and test data set iteration until the prediction precision meets the index; the training condition point set is output to serve as an aero-engine airworthiness conformity verification test matrix, and it is ensured that the airworthiness conformity verification test matrix can cover the vibration characteristics of the blade in the whole flight envelope. And obtaining the dynamic stress response of the engine blade in the whole declared flight envelope.

The invention discloses an inverse design method for a dual-mode internal contraction combined air inlet channel, and relates to a hypersonic air inlet channel of an aircraft. Comprising the following steps: designing a bimodal basic flow field based on a high / low-speed modal cruise Mach number as a design point; the inlet shape design of the air inlet channel is combined with the flight-launch integrated matching requirement; designing an air inlet compression molded surface by a streamline tracking technology; designing the optimal position of the rotating shaft of the splitter plate based on the maximum splitting control efficiency; designing a flow distribution plate rotation angle according to a flow distribution requirement in a flight envelope; the expansion effect caused by rotation of a turbine channel splitter plate and the compression effect generated by area shrinkage in a shared flow channel are comprehensively considered; and a turbine channel and a stamping channel are designed to be separated from each other. The dual-mode conversion adaptability of the internal contraction combined air inlet channel is remarkably improved, and the flow distribution requirements of all channels under different flight working conditions are met.

The present application belongs to the technical field of aviation turbofan engine control design, and specifically relates to a maximum power state control method for the right boundary of the flight envelope of an aviation turbofan engine, comprising: step one, obtaining the design state slip of each operating point at the right boundary of the envelope according to the engine flight envelope; step two, when the engine flight envelope enters the right boundary and reaches the maximum power state, based on the design state slip of the corresponding operating point, the slip is closed-loop controlled by adjusting the nozzle throat area; if one of the engine's high-pressure speed n2, low-pressure speed n1, and exhaust temperature T6 reaches the limit value, it is judged that the engine has reached the maximum power state.

The invention discloses an aero-engine and a power extraction and distribution method thereof, which include a first-level control and a second-level control. The second-level control plans a power extraction and distribution strategy according to flight conditions, power, thrust instructions and electrical load instructions. When the engine is in a steady state, the impact on the engine caused by a sudden change in high-power electrical load and electric energy feedback is balanced. When the engine is operating at the boundary of a flight envelope or in a transition state, the operability of the engine is improved. The output of the second-level control is used as an input instruction of the first-level control to complete closed-loop control of a first generator, a second generator and an energy storage device. The application reduces the impact on the engine dynamics caused by a sudden change in high-power electrical load or the generation of feedback electric energy, and improves the operability of a multi-rotor engine at the boundary of the flight envelope and in a transition state.

The invention discloses an aero-engine direct thrust control method based on migration enhancement near-end strategy optimization, and belongs to the field of control algorithm design. Firstly, a thrust estimator based on a long and short term memory neural network is designed, training is carried out on a data set constructed by random actions, accurate estimation of thrust in a flight envelope is achieved, features extracted by the thrust estimator are optimized and trained on the basis of a near-end strategy, and the thrust in the flight envelope is accurately estimated. And finally, experimental verification is carried out, it is proved that the method can meet the real-time requirement of aero-engine control, and the training stability of the method is superior to that of a common reinforcement learning method.

The invention provides an aero-engine virtual test method and system based on digital twinning, and relates to the technical field of aero-engine virtual tests.The method comprises the steps that flight envelope data of different models of aero-engines are input respectively, and flight envelope data model modeling is conducted on the different models of aero-engines; inputting various engine subject models of different types of aero-engines respectively, and performing aero-engine twin model modeling on the different types of aero-engines respectively; respectively inputting various rack subject models of different types of test racks, and respectively modeling test rack twinborn models for the different types of racks; according to the flight envelope data model, the aero-engine twinborn model and the test bench twinborn model, performing an aero-engine digital virtual test; and performing multi-dimensional analysis and display on the data of the whole process of the digital virtual test. The method has the advantages that virtual-real complementation is achieved, the process is clearer, and the test is more comprehensive and reliable.