133 results about "Flight envelope" patented technology

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 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 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 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 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 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.

The invention relates to the technical field of flight control, and discloses a control method for a transition process of a tilt-body vertical take-off and landing aircraft, and the method comprises the following steps: S1, building a transition corridor model used for calculating a full flight profile of the tilt-body vertical take-off and landing aircraft; and S2, carrying out full-speed-domain and full-angle-of-attack simulation calculation on the aircraft by adopting a CFD (Computational Fluid Dynamics) method to obtain high-fidelity aerodynamic parameters. The control method for the transition process of the vertical take-off and landing aircraft with the tilting airframe has the advantages of being flexible in control method, stable in safe state and the like, and solves the problem that in the prior art, the precision of an optimization calculation result is difficult to guarantee, and the control effect is reduced if the robustness of the aircraft is excessively enhanced. The problems that in the transition process of an airplane, generally, attitude and accelerator control are seriously coupled, the control method is not flexible, coordinated control over an accelerator and an angle cannot be achieved, and it is difficult to guarantee that the airplane is always in a safe state, namely in a flight envelope are solved.

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 discloses an airborne electronic equipment cooling device and method. Comprising a heat exchanger (1), a hot side channel of the heat exchanger (1) is communicated with an electronic equipment cabin, a hot side channel outlet of the heat exchanger (1) is connected with a first connector of a three-way valve A (5), a second connector of the three-way valve A (5) is communicated with a rear stage of an evaporation fan (4), a front stage of the evaporation fan (4) is connected with a single-phase medium channel outlet of an evaporator (3), and a single-phase medium channel inlet of the evaporator (3) is communicated with the external environment. The full-flight envelope heat dissipation requirement of airborne electronic equipment can be met.

The application relates to the technical field of aircraft design, and discloses a continuous variable-camber rigid wing driven by multiple rudders and a control method thereof, wherein the continuous variable-camber rigid wing comprises a wing leading edge section, multiple rigid wing sections, a distributed rudder driving mechanism and a transmission mechanism; the multiple rigid wing sections comprise a first wing section, a second wing section and a third wing section which are sequentially hinged through rotating shafts; the distributed rudder driving mechanism is arranged in the wing leading edge section and is used for independently driving the deflection of each rigid wing section; and the transmission mechanism is connected between the distributed rudder driving mechanism and the corresponding rigid wing section and is used for converting the output motion of the distributed rudder driving mechanism into the deflection motion of the corresponding rigid wing section. According to the application, the rigid wing is segmented, multiple rudders are used for independent driving, an advanced aerodynamic analysis model and a closed-loop feedback control are combined, the wing camber is self-adaptively, continuously and accurately regulated within a full flight envelope, and thus the optimal aerodynamic performance can be maintained at all times.

The application discloses a large-size free piston driven high-enthalpy pulse device, and relates to the field of wind tunnel test. The device contains four operation modes, i.e. a high-total-pressure high-enthalpy reflection type shock tunnel, a low-dissociation high-enthalpy straight-through type shock tube, an ultra-high-speed expansion tube and an ultra-high-speed expansion wind tunnel. The four operation modes expand the flight envelope simulated by the wind tunnel, so that the flight environment with extremely high dynamic pressure, extremely high temperature, extremely high speed and extremely high brightness can be simulated for a long time. The application not only can simulate the flight environment in a wide area and extremely high speed, but also can provide high-quality wind tunnel test aerodynamic data for aircraft development, and expand the test capacity of the high-enthalpy pulse device.

The invention belongs to the technical field of helicopter external hanging state monitoring, and discloses a method and system applied to helicopter external hanging state monitoring, and the method comprises the steps: firstly building a helicopter body coordinate system with the gravity center of a helicopter as an original point, and obtaining real-time helicopter positioning data, helicopter attitude angle data and external hanging object positioning data; then calculating the position of the externally hung object in the body coordinate system according to the data; then calculating the position of a hanging point, and correcting the position of the externally hung object according to the position of the hanging point; calculating the actual gravity center position of the externally hung object, and judging whether the actual gravity center position exceeds the flight envelope of the helicopter; finally, a three-dimensional image of the attitude relation between the externally hung object and the helicopter is displayed in real time, and if the attitude of the externally hung object exceeds the flight envelope, the helicopter gives an alarm and generates a corresponding flight control adjustment instruction to assist a pilot in controlling the helicopter. Through accurate position calculation, a dynamic correction mechanism and an effective flight control adjustment strategy, the external suspension stability, safety and transportation efficiency of the helicopter are improved.

This invention provides an online monitoring method and system for airborne flight runaway precursors of civil aircraft based on flight test data. The method includes: constructing a set of input flight parameters for identifying airborne flight runaway precursors; acquiring daily operational data and flight test data of the target aircraft model; and extracting physical feature information of airborne flight runaway by referencing the aerodynamic mechanism model and extreme flight envelope boundary of the target aircraft model. Based on the domain adaptation concept, an offline precursor recognition model is constructed that integrates physical feature information embedding, meta-learning, and multi-instance learning. Based on knowledge distillation technology, the recognition capability of the offline precursor recognition model is transferred to a lightweight network constructed from gated recurrent units to generate an online precursor monitoring model, enabling real-time precursor probability calculation and early warning. Finally, an intelligent agent model based on a dual-delay deep deterministic policy gradient algorithm is constructed to verify the effectiveness of the online precursor warning. This invention overcomes the cross-domain data gap and meets the requirements of online lightweight computation.

The present application belongs to the field of engine design technology, and specifically relates to a method for determining the flight envelope of an aircraft turbofan engine. The method mainly includes: step S1, obtaining the load spectrum of the power extraction of aircraft accessories within the full envelope range and the comprehensive distortion index allowed by the engine inlet; step S2, determining the speed of the compression component of the turbofan engine when the engine stability margin is weakest; step S3, determining the stable working boundary of the engine after power extraction based on the speed of the compression component when the engine stability margin is weakest. The present application obtains the speed of the critical distortion index of the engine under different altitudes and different speed conditions through pressure distortion tests, thereby determining the flight envelope of the engine under the influence of actual power extraction, greatly improving the engine aerodynamic stability evaluation method, thereby giving the power extraction limit within the envelope range, and avoiding problems such as engine surge.

The invention discloses a flight target joint tracking and identification method based on position and flight envelope information, and the method comprises the steps: building a shareable and parsable motion model set based on the motion state of a flight target; receiving measurement data at an initial moment, and initializing flight target track and motion feature information; performing filtering estimation on the motion state of the flight target by using an optimal model expansion algorithm, and extracting motion state features of the flight target; calculating an acceleration-velocity flight envelope likelihood function; identifying a flight target by using a sequential hypothesis testing method; and judging a recognition result, and repeatedly carrying out flight target motion state filtering estimation, feature extraction, likelihood function calculation and flight target recognition until flight target track estimation, recognition and classification are completed. According to the invention, high-precision and stable tracking of the flight target based on the position data and the flight envelope information, and estimation, identification and classification of the position, the speed and the acceleration state are realized.

A system and method for sampling a flight envelope of an air vehicle under test. The method includes obtaining engineering models of the vehicle to identify how the vehicle behaves, obtaining empirical models through test data, and obtaining historical models of similar legacy aircraft. The method further includes ensembling the models into an integrated model and assessing the ensemble uncertainty at any particular test point, for a variety of aircraft conditions, to determine model reliability. The method includes determining the optimal order of test points based on local model uncertainty, testing the air vehicle at a point, revising the integrated model with that data and reassessing the uncertainty, and determining if a stop testing condition has been met, in which case the testing series is closed. The last calibrated integrated model could then be used to generate trusted data for any remaining required test using a certification by analysis method.

The invention relates to a multi-degree-of-freedom deformable wingtip, belongs to the technical field of aircraft design and aeroelasticity design, and solves the problem that a deformable wingtip in single-degree-of-freedom design in the prior art cannot maintain high aerodynamic efficiency in a wider flight envelope, the multi-degree-of-freedom deformable wingtip comprises: a root wing section (100) connected to a wing main body; the middle wing section (200) is rotatably connected to the outer end of the root wing section (100); the winglet (300) is connected to the outer end of the middle wing section (200) in a twisting manner; and the power assembly comprises a bending servo motor (400) for driving the middle wing section (200) to rotate and a torsion servo motor (500) for driving the winglet (300) to twist.