106 results about "Flight envelope" patented technology

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 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 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.

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 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 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.

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.

The invention discloses a helicopter demand power model calculation method and system, and belongs to the technical field of helicopter control, and the method comprises the steps: constructing a T-S fuzzy model, and dividing flight modes on a flight envelope based on the constructed T-S fuzzy model, so as to obtain a flight mode set; obtaining time sequence data in the flight process of the helicopter, and determining a current flight mode in the flight mode set based on the obtained time sequence data; and calculating the required power of the helicopter based on the determined current flight mode. A flight mode set is divided on a flight envelope through the constructed T-S fuzzy model, the required power of an engine can be well estimated by recognizing flight modes and calculating the required power in the flight process, power matching with the engine is achieved, and a basis is provided for subsequent high-speed helicopter control.

This invention provides a model-based dynamic anti-windup method and device for an aircraft engine fuel control system. The method comprises establishing a nonlinear component-level model of the aircraft engine and obtaining a state-space model of the aircraft engine using a Taylor series expansion method; establishing a linear dynamic controller based on the aircraft engine state-space model; and designing a dynamic anti-windup compensator for the aircraft engine's fuel supply actuator. The device includes a personal computer terminal, input and output ports for digital-to-analog conversion, an electronic engine control unit designed based on a digital signal processor in the personal computer terminal, and a data bus converter. This invention can improve aircraft engine control performance throughout the entire flight envelope.

The invention provides a height measurement system error budget distribution method suitable for a single machine, and belongs to the technical field of aircraft design, and the method comprises the steps: obtaining a maximum value of an absolute value of an air pressure altitude error and a corresponding test flight state parameter according to the air pressure altitude error obtained through the test flight of an aircraft at each test flight point in an RVSM flight envelope; source identification is carried out on height measurement errors caused by the electronic equipment, the height measurement errors caused by the electronic equipment at all test flying spots are analyzed, and the height measurement error under the worst condition of the electronic equipment is determined; according to the error requirement of a single-machine height measurement system in airworthiness clauses, determining the error margin in the corresponding test flight state when the absolute value of the air pressure height error is maximum; and according to the error margin, analyzing and solving the allowable pressure variation in the test flight state as a basis for evaluating the waviness appearance change of the RVSM key area near the static pressure source during continuous airworthiness. According to the method, the requirement of RVSM operation in AC-21-13 on the error of the height measurement system of the non-grouped aircraft can be met.

The application belongs to the field of unmanned aerial vehicle flight test, and particularly relates to a heading angle fixed three-edge flight unmanned aerial vehicle airspeed system calibration method. The method comprises the following steps: step one, selecting a test point in the flight envelope of the aircraft; step two, performing three-edge flight of the aircraft under the conditions corresponding to the test point, and obtaining flight data of the first edge, the second edge and the third edge in the stable level flight stage; step three, selecting multiple flight data points of the first edge, the second edge and the third edge from the flight data, obtaining the heading angle and the track angle of each flight data point, and judging whether the wind field uniformity in the flight process meets the requirements according to the heading angle and the track angle; if yes, entering step four; if no, returning to step two and re-executing the test point; step four, obtaining the ground speed of each flight data point, and calculating the wind speed according to the ground speed; step five, calculating the true airspeed in the flight process according to the wind speed and the ground speed; and step six, performing test point airspeed calibration according to the true airspeed.

This invention provides a propfan engine speed control method, comprising the following steps: establishing a dual-row propfan aerodynamic mathematical model; aiming at optimal propeller efficiency, performing nonlinear optimization of the pairing relationship between the front row propeller pitch α1 and the rear row propeller pitch α2 at typical operating points within the flight envelope of the dual-row propfan aerodynamic mathematical model, obtaining the optimal control law as a virtual optimal pitch, which is respectively at a certain speed N. s At the flight envelope point, the optimal pitch of the front and rear propeller blades corresponding to the highest propeller efficiency η is determined; the pitch of the dual-row propeller is controlled to be at the optimal pitch based on the front and rear propeller blade pitches corresponding to the highest propeller efficiency. This invention provides an engineering-applicable propfan engine speed control law, which enables free and independent control of the propfan engine core speed and the dual-row propeller speed, meeting the aircraft's requirements for propfan engine use.

An aircraft includes a blended-wing-body airframe that defines a center-body region exhibiting a center-body aerodynamic center forward of a center of mass of the aircraft within a horizontal flight envelope of the aircraft. This airframe also defines an outer-wing region exhibiting an outer-wing aerodynamic center aft of the center of mass of the aircraft within the horizontal flight envelope of the aircraft. Additionally, the aircraft includes a vertical landing subsystem configured to, in a vertical landing configuration, contact a ground surface and structurally support the aircraft, as well as a propulsion subsystem.

This invention belongs to the field of aeronautical acoustics and flight simulation technology, and relates to a method for cabin multi-source sound source evolution and sound field reconstruction under flight state constraints. It solves the problems of traditional blind source separation relying on statistical characteristics, sound field reconstruction relying on measured data, and low sound fidelity in flight simulators. This invention constructs a multi-source mixed acoustic signal-flight parameter synchronous dataset as input, constructs a flight state-sound source feature prior knowledge base as prior constraints, and constructs an objective function including reconstruction error, source independence regularization, and flight state constraint terms. The time-domain signals of each independent sound source are separated using the alternating direction multiplier method, and the sound source feature parameters are extracted as output. Flight state parameters are used as input, and an evolutionary mapping function is constructed based on the mean function. The target flight state is input into the evolutionary mapping function to obtain predicted sound source feature values, driving the equivalent source method to achieve sound field reconstruction, and outputting a three-dimensional sound field distribution. This invention achieves predictive, high-fidelity real-time reconstruction of the sound field within the entire flight envelope.

This invention discloses a method for analyzing the transient characteristics of multi-ducted engine casings under complex loads, belonging to the field of aero-engine technology. This method analyzes the structure and operation of a multi-ducted variable cycle engine, classifying it into three turbofan modes and one turbojet mode. Based on an aerodynamic thermodynamic configuration model, component-level modeling is performed using MATLAB software, generating component-level modeling data. A common set of working equations is established using the modeling data, and the solution is obtained through iterative solving using the Broyden quasi-Newton method. The ultimate aerodynamic loads of the casing system under typical operating conditions are calculated. Combined with the engine flight envelope, 12 typical altitude and Mach number states are selected to solve for the ultimate operating points and analyze the transient load characteristics. This invention can accurately capture the transient response characteristics of multi-ducted engine casing systems under complex load conditions, improving analysis accuracy and efficiency, providing a theoretical basis for the design optimization and reliability assessment of casing systems, and has significant engineering value.

The application discloses a power control method and related device of a compound wing unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles.The method comprises the following steps: after the compound wing unmanned aerial vehicle takes off, the compound wing unmanned aerial vehicle receives a mode control instruction sent by a flight control terminal, wherein the mode control instruction comprises a fixed-wing flight mode and a vertical flight mode; when the mode control instruction received by the compound wing unmanned aerial vehicle is the fixed-wing flight mode, the power condition met by the current battery capacity of the compound wing unmanned aerial vehicle is matched, and the flight envelope and the engine speed of the compound wing unmanned aerial vehicle are determined; when the mode control instruction received by the compound wing unmanned aerial vehicle is the vertical flight mode, the engine speed of the compound wing unmanned aerial vehicle is adjusted to the maximum speed, the countdown of the vertical flight mode is determined, and a transition flight route is planned when the countdown of the vertical flight mode reaches a preset threshold.The power control of the compound wing unmanned aerial vehicle reaches a more ideal effect.

The invention provides a method and a device for determining aircraft flap control parameters. The method comprises the following steps: acquiring flight envelope data, flap geometric data and material data of an aircraft; preprocessing the flight envelope data, the flap geometric data and the material data to obtain preprocessed data; determining a target constraint condition according to the preprocessed data; obtaining an optimal flap parameter according to the constraint condition, a preset optimization target and a preset aircraft aerodynamic model; and determining flap control parameters according to the optimal flap parameters. The shape and the angle of the flap are dynamically adjusted according to the flap control parameters, the purpose of precisely controlling the attitude of the aircraft is achieved, and the aircraft attitude control precision can be improved.