76 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 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 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 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 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.

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 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.

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.

The invention discloses a load calculation method for a complex pneumatic profile nozzle, and the calculation method comprises the steps: rapidly obtaining the large-batch load data of the complex pneumatic profile nozzle in a flight envelope through introducing a concept of a load impact factor of the complex pneumatic profile nozzle, so as to meet the requirement of fatigue life evaluation of the complex pneumatic profile nozzle. The method comprises the specific steps that firstly, basic characteristics of load distribution of the complex aerodynamic surface nozzle are obtained based on a CFD numerical calculation method or a nozzle test, and load influence factors of the complex aerodynamic surface nozzle are obtained; and then the nozzle load variable quantity caused by the flight height, the flight Mach number and the engine state change is calculated by utilizing a nozzle load theoretical formula derived by a one-dimensional isentropic internal flow theory.

The wide-range intelligent variable aircraft design method based on the curved shock wave theory comprises the following steps: 1) according to a flight envelope, the aerodynamic design of the wide-range intelligent variable aircraft is divided into a minimum resistance state, an optimal lift-drag ratio state and a maximum lift state, wherein the minimum resistance state corresponds to a boost climb section and a free flight section, the optimal lift-drag ratio state corresponds to a re-entry glide section, and the maximum lift state corresponds to a terminal down pressure section; 2) the aerodynamic design of the minimum resistance state is as follows: the aerodynamic profile of the wide-range intelligent variable aircraft is inversely designed based on a full three-dimensional curved shock wave surface; 3) the aerodynamic design of the optimal lift-drag ratio state is as follows: the aerodynamic profile of the wide-range intelligent variable aircraft is designed based on a full three-dimensional curved shock wave surface and a wall surface pressure distribution simultaneously controllable inverse method; 4) the aerodynamic design of the maximum lift state is as follows: the aerodynamic profile of the wide-range intelligent variable aircraft is inversely designed based on a three-dimensional wall surface pressure distribution.

The invention relates to the technical field of aviation aircraft design, and discloses a continuous variable camber rigid wing driven by multiple steering engines and a control method thereof, and the continuous variable camber rigid wing comprises a wing leading edge section, a plurality of rigid wing sections, a distributed steering engine driving mechanism and a transmission mechanism; the plurality of rigid wing sections comprise a first wing section, a second wing section and a third wing section which are sequentially hinged through a rotating shaft; the distributed steering engine driving mechanism is arranged in the front edge section of the wing and is used for independently driving each rigid wing section to deflect; the transmission mechanisms are connected with the distributed steering engine driving mechanisms and the corresponding rigid wing sections and used for converting output motion of the distributed steering engine driving mechanisms into deflection motion of the corresponding rigid wing sections. According to the invention, the rigid wing is segmented and is independently driven by a plurality of steering engines, and an advanced pneumatic analysis model and closed-loop feedback control are combined, so that the camber of the wing is adaptively, continuously and accurately adjusted in a full flight envelope, and the optimal pneumatic performance is always maintained.

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 application provides a kind of low airspace collaborative management method and system based on digital twinning, comprising: based on the real-time state data of aircraft obtained, analyzing aircraft maneuverability, flight envelope characteristic parameters, characteristic parameters include maximum speed, maximum acceleration, minimum turning radius, utilize kinematics and dynamics model, depict different aircraft movement law;Combined with real-time flight data of aircraft, through digital twinning technology, real-time synchronization aircraft motion trajectory and attitude change in virtual environment, accurately reproduce aircraft dynamic behavior, provide high-fidelity digital model for flight conflict detection, flight path prediction;For aircraft motion characteristic difference, construct multi-level, multi-granularity digital twinning model, digital twinning model includes airspace structure model at macro level, aircraft group behavior model at mesoscopic level and single aircraft dynamic model at micro level.

The invention belongs to the technical field of helicopter airworthiness conformity verification, and particularly relates to an airworthiness conformity verification method for helicopter near-ground control characteristics. The method comprises the following steps: 1, determining a test flight state of near-ground control characteristics according to an expected flight envelope and performance; 2, determining test flight test parameters; 3, preparation before test flight is carried out; 4, in a test flight airport with the altitude of 0 m, ground-effect near-ground control characteristic test flight is carried out with the combination of the maximum ground-effect hovering weight GIGE, the maximum ground-effect-free hovering weight GOGE and different limit gravity centers; 5, determining a critical center of gravity according to the test flight result of the combination of each weight and center of gravity in the ground effect test flight and the ground effect-free test flight; 6, carrying out ground-effect-free near-ground control characteristic test flight at the critical gravity center in the test flight airport from other heights to the maximum take-off and landing height; and step 7, according to the test flight results of the step 4 and the step 6, determining the maximum side flight and back flight speed and the hovering wind speed envelope of the helicopter.

The invention provides an atmospheric parameter inversion method and system for a high-speed orbiter. The method comprises the following steps: acquiring pressure data of a plurality of measuring points arranged on the surface of the high-speed orbiter; inputting the pressure data into a pre-established mapping relation model to obtain and output corresponding atmospheric parameters; wherein the mapping relation model is a nonlinear model obtained by training by adopting a neural network fused with extended Kalman filtering according to the corresponding relation between the measuring point pressure data of the high-speed orbiter in the flight envelope and the atmospheric parameters. On the basis of the theoretical basis of the one-to-one correspondence single-valued mapping relation between the surface pressure data and the corresponding atmosphere data, the design method for inversing the atmosphere data based on surface pressure reverse modeling is provided, and the technical problem that the surrounding environment parameters of the high-speed orbiter cannot be obtained in real time through a traditional design method is solved; meanwhile, the attack angle and the sideslip angle of the orbiter are solved, and the technical problem of obtaining the high-precision attitude angle of the high-speed aircraft is solved; the applicability is high, and the calculation precision is high.

The invention belongs to the technical field of aircraft heat management, and particularly relates to a multi-heat-sink-based aircraft-engine integrated multistage adjustable heat management system which comprises an aircraft system, a phase change heat storage device and an engine system. By integrating various heat sinks of fuel oil, air, a coolant, a phase change material and liquid hydrogen, a multi-stage adjustable heat management network is constructed. The system dynamically switches heat flow paths according to low, medium and high Mach number working conditions, wherein fuel oil and air are used as main heat sinks when the Mach number is low; when the Mach number is medium, an afterburning fuel path is started to enhance cooling; the liquid hydrogen system is activated when the Mach number is high, fuel oil is deeply cooled through the hydrogen-fired heat exchanger, and heat load balance is achieved in cooperation with the phase change heat storage device. Through cooperative control of multiple valves, on-demand matching of a heat source and a heat sink is achieved, and the heat dissipation capacity and adaptability of the system under wide envelope flight are remarkably improved.