231 results about "High-speed flight" patented technology

According to the transition prediction method suitable for the wide Mach domain of the high-speed aircraft provided by the invention, temperature and Mach number coupling correction is performed on key parameters in time scale modeling of a second mode of a transition mode, and simplified processing of a flow compression effect and a heat transfer process by traditional single parameter association is broken through; the influence of an incoming flow disturbance effect is embedded in the mode, and dynamic association between test environment parameters and transition triggering criteria is achieved; according to the improved mode, accurate prediction results are obtained on a plurality of examples with the Mach number range of 4.7-14, the research results provide a more reliable theoretical tool for aerodynamic configuration optimization of a high-speed aircraft with a wider speed range, and the method has important engineering guiding significance for thermal protection system design of a new-generation reusable aircraft.

The invention discloses a hypersonic flow instability multi-field coupling prediction method and system, and relates to the technical field of modern high-speed aircraft design, and the method comprises the steps: obtaining the multi-physical field parameter distribution of a hypersonic flow field; based on a shock wave-boundary layer interaction mechanism, constructing an instability sensitive graph data structure by taking multi-physical field parameters as input, setting graph nodes in a shock wave influence region, a boundary layer region and an intersection region thereof, and calculating connection weights among the nodes; performing multi-scale decomposition on the instability sensitive graph data structure, and extracting instability modes of a shock wave scale, a vortex scale and a molecular scale; screening the instability modes by adopting a hypersonic velocity stability criterion, and determining a dominant instability mode; and constructing an instability evolution prediction model based on the dominant instability mode, and outputting a spatio-temporal evolution rule of hypersonic flow instability. According to the method, the problem that the flow instability phenomenon in the hypersonic velocity complex multi-physical field environment is difficult to accurately predict can be solved.

The invention relates to a high-speed aircraft transition identification method based on a flight test high-frequency vibration signal. The method comprises the following steps: selecting high-quality measurement point data suitable for analysis from flight test high-frequency vibration data as a vibration signal to be analyzed; performing time-frequency analysis on the to-be-analyzed vibration signal, outputting a time-frequency diagram, and representing the change condition of energy in a frequency domain by using a centroid frequency and an envelope difference of the centroid frequency; carrying out preliminary identification on the transition occurrence moment; decomposing a to-be-analyzed vibration signal into a plurality of intrinsic mode functions with different center frequencies through variational mode decomposition; quantizing the change condition of the energy of each modal component along with time; and aiming at the change condition of the energy of each modal component along with the time, the transition is identified according to the energy of the mode excited by the transition, and the energy sudden increase point is the moment when the transition begins to occur. According to the method, the transition occurrence time can be effectively identified under the condition that no hole is formed in the surface of the aircraft.

The embodiment of the invention relates to the technical field of aerospace safety, and provides an aircraft trajectory prediction method and system based on space gridding and heuristic search. According to the method, adaptive grid resolution adjustment is performed on high-speed aircraft trajectory prediction, and a possible high-probability trajectory is quickly calculated according to the type of the aircraft, so that the problems of inflexible calculation resource coordination, insensitive aircraft type perception and low calculation binding physical constraint in the current high-speed aircraft trajectory prediction are solved.

The invention provides a wave rider capable of vertically ascending and descending, and belongs to the technical field of wide-speed-range aircrafts, the wave rider specifically comprises a front edge shock wave generating surface, a central fuselage and sweepback wings located on the two sides of the central fuselage, the two sweepback wings are provided with mounting holes penetrating through the upper surface and the lower surface of the wave rider, and lift fans are mounted in the mounting holes; the lift fans are embedded into the mounting holes and used for providing vertical lift force for the wave rider, the axes of the lift fans are parallel to the vertical axis of the wave rider, and the lift fans on the two sweepback wings are symmetrically distributed on the two sides of the central fuselage. According to the treatment scheme, the vertical lifting capacity is provided for the wave rider, and the high-speed flight aerodynamic performance is considered.

The invention relates to the technical field of undercarriages, in particular to a folding undercarriage device of a low-altitude aircraft, which comprises an aircraft body, a pesticide spraying mechanism for crop fertilization is arranged on the aircraft body, a height-adjustable undercarriage mechanism is arranged at the lower end of the aircraft body, and a landing buffer mechanism is arranged on the undercarriage mechanism. The landing buffering mechanism is used for reducing impact on the undercarriage mechanism at the moment of landing, a linkage auxiliary mechanism is arranged on the aircraft body, the undercarriage mechanism comprises a fixing plate body, and the fixing plate body is detachably installed on the aircraft body. By arranging the undercarriage mechanism and adopting a dynamic folding design, after taking off, the second-stage frame body retracts to drive the telescopic joint to be shortened, so that the overall height is reduced, air turbulence during flight can be effectively reduced, meanwhile, the first-stage frame body rotates to be parallel to the fuselage, the cross sectional area of the undercarriage mechanism is greatly reduced, and the overall height is reduced. And the wind resistance loss during high-speed flight can be obviously reduced.

The invention relates to a method for intercepting a flying object through aerial high-speed net casting, which is used for solving the technical problem of low interception success rate caused by insufficient maneuverability of an interception means and lack of enough large and reliable interception area when approaching a target in the prior art, and comprises the following steps: controlling an unmanned aerial vehicle to fly to approach the target; an unlocking signal is sent to a locking mechanism arranged on the machine main body, so that the locking mechanism synchronously releases all the power systems; the released power system flies away by means of self-rotation inertia, and the catching net connected with the power system is pulled by the pulling rope, so that the catching net is unfolded to form an intercepting area in the flying-away process. The unmanned aerial vehicle keeps flying at a high speed in the pursuit stage, the unmanned aerial vehicle flies away through inertia of a plurality of power systems in the interception stage, the capture net is unfolded to form an interception area larger than the projection area of the vehicle body, and compared with a traditional interception method, the interception method has the higher interception speed and the higher interception success rate.

The invention belongs to the technical field of high-temperature refractory composite materials, and discloses an ablation-resistant connecting piece and a preparation method thereof. The ablation-resistant connecting piece comprises the main body connecting part made of the high-temperature alloy material and the ablation-resistant fixing part, the internal main body connecting part can be effectively protected by the fixing part through ablation resistance and heat resistance, the main body connecting part is made of the high-temperature alloy, the high-temperature strength is high, and the service life of the main body connecting part is prolonged. The fixing part and the high-temperature-resistant binder are co-cured and formed on the main body connecting part, and the bonding strength of the fixing part and the main body connecting part is further improved through the embedding structure, so that the ablation-resistant connecting piece has the heat-proof performance of an ablation material and the high-temperature strength performance of high-temperature metal, is suitable for connection of short-time high-speed aircrafts, and is suitable for being used in a large-scale production line. The preparation method of the ablation-resistant connecting piece is relatively simple in required equipment, convenient to operate, stable in process, high in production efficiency and suitable for large-scale industrial production.

The application discloses a method for preparing long-time aging antioxidant high-emissivity alumina-silica double-layer structure coating by a pre-oxidation method, and belongs to the technical field of coating surface modification. The application aims to solve the problems of uncontrollable coating structure and insufficient high-temperature oxidation resistance of existing coatings during long-time high-temperature service. The long-time aging antioxidant high-emissivity double-layer structure coating prepared by the application comprises an aluminum inner layer and a silicon outer layer, and the double-layer structure mainly composed of the inner layer of Al2O3 and the outer layer of SiO2 is formed in situ during the pre-oxidation process. The double-layer structure coating has good bonding performance, the bonding strength reaches above 13 MPa, the emissivity reaches 0.9, the double-layer structure of the inner layer of Al2O3 and the outer layer of SiO2 is maintained after oxidation in static air at 1150 DEG C for 50 h, the coating remains complete without falling-off phenomenon, and the coating has application potential in a long-time aging antioxidant radiation heat protection system of a high-speed aircraft.

The invention provides a high-speed aircraft online trajectory planning method based on improved polynomial guidance, which solves the problems of calculation delay and poor trajectory smoothness in the existing autonomous path planning, and comprises the following steps: establishing a plane-coordinate system and a motion equation set of a high-speed aircraft, and determining constraint conditions in the flight process of the high-speed aircraft; constructing an optimization model containing an adaptive penalty function based on an improved sequential quadratic programming algorithm, and generating an obstacle avoidance path point set; performing trajectory optimization by adopting an improved polynomial, generating a trajectory meeting aircraft dynamics constraints among the path points generated in the step 2, and selecting a key path point set; based on the obstacle avoidance path point set and the key path point set, adopting a centripetal parameterized Catmull-Rom spline curve to process the generated path to obtain a smooth trajectory; a guidance instruction is obtained based on smooth trajectory analysis and calculation, and the high-speed aircraft is guided to sail according to the planned trajectory. And the real-time calculation efficiency is improved on the basis of realizing the smooth trajectory.

The invention discloses a testing device of an active and passive combined thermal protection system, and relates to the technical field of thermal protection of hypersonic aircrafts. The micro-channel cold plate comprises a micro-channel cold plate assembly, the top of the micro-channel cold plate assembly is fixedly connected with a passive thermal insulation layer, and a plurality of thermocouples are installed between the passive thermal insulation layer and the micro-channel cold plate assembly; a gear pump is installed at an inlet of the micro-channel cold plate assembly through a fuel oil pipeline, a fuel oil tank is installed at the inlet end of the gear pump, and an adjusting valve is installed at the outlet end of the gear pump. By adjusting the flow of the micro-channel cold plate in real time, the change of an external pneumatic thermal environment is well matched, so that efficient thermal protection of a large-area fuselage of a high-speed aircraft under long-time and high-heat-flow conditions is realized, and the consumption of a coolant is effectively reduced; and when the control module adjusts the flow, constant-pressure supply of the cooling working medium is achieved by changing the opening degree of the pressure adjusting valve, the working stability is improved, and the thermal failure problem of the active and passive thermal protection system is effectively avoided.

The invention discloses a self-adaptive optical distortion compensation device and compensation method based on high-speed closed-loop control, and the device comprises an optical path front-end interface which is used for connecting a coupling wind tunnel or a laser emission system; the wavefront sensor is used for measuring a local wavefront slope by adopting a distributed array; the FPGA controller comprises a wavefront reconstruction processor and a closed-loop logic controller and is used for reconstructing wavefront and generating a control instruction; the MEMS deformable mirror array comprises a plurality of MEMS drivers and is used for correcting a wavefront phase; the feedback monitoring module is used for performing dual verification through deformation capacitance measurement and CCD image feedback; and the imaging / laser output end is used for converting the corrected optical signal into an image and outputting the image. The method can be widely applied to the fields of high-speed aircraft wind tunnel testing, laser communication, infrared imaging guide head calibration, high-energy laser beam shaping and the like, a programmable hardware platform, a high-speed parallel operation structure and a predictive control algorithm are introduced, and unprecedented sub-microsecond closed-loop compensation is achieved.

The invention discloses a high-speed aircraft aerodynamic configuration agile optimization method based on an artificial intelligence model. The method comprises the following steps: A1, establishing an aerodynamic force agent model; a2, performing a global optimization algorithm; a3, establishing an intelligent autonomous decision-making model; by constructing a multi-target adaptive weight distribution model, weight dynamic collaborative optimization of proxy model confidence, a constraint penalty function and target function key elements is achieved, the mathematical modeling problem of dynamic matching of weight parameters and optimization targets in the iteration process is solved, a closed-loop optimization system with the autonomous perception-decision-execution capacity is developed, and the optimization efficiency is improved. Therefore, the convergence speed is improved in a complex engineering optimization scene, local optimum is avoided, and the robustness of a global optimal solution is ensured; and A4 pneumatic optimization iteration: based on the proxy model for predicting indexes such as the lift-drag ratio, performing global optimization of design parameters by using an artificial intelligence optimization algorithm, performing verification through CFD, and after verification, putting a CFD calculation result into a training set of the model to re-train the model and continuing optimization until optimization design is converged.

The application discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, determines a rudder leading edge radius and a rudder leading edge sweepback angle based on a current state nominal trajectory and a rudder leading edge thermal response temperature as a judgment standard; secondly, a conical angle of a folding rudder-cone surface is determined through iterative calculation, so that a chordwise dimension of a high-temperature-resistant refractory metal material, and a thickness of a skin and a heatproof coating at the cone are designed; thirdly, through interference position design of a heatproof cup bow shock wave, a gap between a fixed rudder and a movable rudder avoids the bow shock wave interference zone, so that a chordwise distance between a folding gap and a rudder shaft is determined. Through the aerodynamic configuration design method, the rudder leading edge radius can be as low as 3mm, and the cone angle can be as low as 4°, so that ablation deformation does not occur when the maximum heat flow of the rudder leading edge is less than 10MW / m 2 2, the resistance characteristics of the high-speed aircraft can be effectively reduced, the aircraft has a near-zero ablation feature, and a good aerodynamic shape is maintained.

The invention discloses a high-speed aerodynamic derivative analysis modeling method and system based on symbolic regression, and belongs to the technical field of aircraft aerodynamic modeling and artificial intelligence cross, and the method comprises the steps: obtaining an aerodynamic data set of a high-speed aircraft, and carrying out the preprocessing; defining grammar for generating candidate symbol expressions and constructing an expression tree; modeling based on the training set by adopting a layered symbol regression framework to obtain an analytical model; the prediction precision of the analytical model is verified on the test set, and physical consistency analysis is carried out; carrying out uncertainty quantification on the analytical model to obtain predicted uncertainty estimation of the analytical model; and finally, outputting an analytic model in a mathematical expression form for describing the relationship between the aerodynamic derivative and the flight state variable and uncertainty estimation of the analytic model. The method overcomes the defects of a traditional method in precision, generalization and interpretability, realizes full-automatic efficient modeling from data to the analytic model, and improves the modeling efficiency. The method is suitable for control design, stability analysis and real-time simulation of the high-speed aircraft.

The utility model discloses a bionic aircraft side wing, and belongs to the technical field of bionic equipment. The aircraft mainly comprises an aircraft framework and side wing parts arranged on the two sides of the aircraft framework, a fixing frame and a driving component installed on the fixing frame and used for driving the side wing parts to move are arranged at the front end of the aircraft framework, a battery pack and a microcontroller are arranged on the aircraft framework, and each side wing part comprises a front wing, a rear wing and a connecting block. The bionic aircraft side wing is composed of the front wing body, the rear wing body and the connecting block, the front wing body and the rear wing body are connected through the connecting block and provided with the outer contour frames and the wing faces, and the cross structures are arranged in the outer contour frames, so that the rigidity, stability and overall strength of the side wing can be effectively enhanced, and the excellent performance of an aircraft in high-speed flight or complex airflow is guaranteed; meanwhile, the airfoil is made of a thin film material with high tensile strength and elasticity, so that the structural weight is reduced, the aerodynamic performance is optimized, and the efficiency and reliability of the aircraft are improved.

The invention relates to the technical field of high-speed aircraft ground simulation and electromagnetic property calculation, and provides a high-speed aircraft electromagnetic calculation method and system based on a quantum classification algorithm. A variable component sub-classifier is used as an intelligent scheduling core to perform feature recognition and region division on a whole flow field, and a transmission matrix method and a WKB approximation method are adaptively called according to a division result to perform hybrid calculation, so that the technical problem that a traditional single calculation strategy is difficult to collaboratively optimize in precision and efficiency is solved; the method has the advantages that global computing resources can be distributed according to needs, and the overall solving speed is greatly increased while high simulation precision of the key area is guaranteed.

The invention discloses a fixed wing patrol aircraft flight attitude control method, and belongs to the technical field of unmanned aerial vehicles. Aiming at the technical contradiction that high-speed flight maneuverability and weapon shooting stability cannot be compatible, the method achieves breakthrough through the following innovative schemes that a four-wing propeller collaborative framework is designed, and each propeller is provided with a horizontal linear driver to achieve front-back telescopic displacement in the axial direction of a fuselage; in the horizontal flight stage, based on the feedback of a gyro stabilizer, pitching / rolling double-axis attitude accurate control is realized through the cooperative adjustment of propeller position stretching and rotation speed difference; triggering a layered compensation strategy by an ignition event, and synchronously adjusting the horizontal position and the rotating speed of a propeller according to the recoil force direction; and the tail vane intervenes in fine adjustment only when the yaw is out of tolerance. By combining the seamless conversion design of the vertical take-off and landing stage, the three technical problems of height loss without initial-speed deployment, low aerodynamic efficiency of high-speed flight and insufficient shooting precision are solved.

The invention relates to the technical field of high-speed aircraft health management, and particularly discloses an interpretable aircraft state sensing and predicting system and method. The method comprises the following steps: constructing and training a lightweight anomaly diagnosis network with an expansion perception layer; aircraft dynamics, kinematics and fault mechanism formulas are used as soft constraints to be embedded into the future state estimation network; an SHAP interpretation framework is used for sorting the importance of the input features, and a network structure is optimized accordingly; and finally, accurate diagnosis of abnormal types and positions and rapid prediction of aircraft states at a plurality of moments in the future are realized. According to the method, by introducing three innovation points of an expansion perception layer, physical information embedding and feature importance sorting, the diagnosis accuracy and the model interpretability are remarkably improved, and the high-performance calculation requirement in a scene with limited calculation power is met.

The invention discloses a bearing and wave-transmitting integrated design method and system for a variable-density honeycomb structure radome, and relates to the technical field of high-speed aircraft radome design. The method comprises the following steps: carrying out force / heat / electric performance characterization on a honeycomb cell element model; sensitivity analysis is conducted on the relation between the cell element structure parameters and the electromagnetic performance, and key design parameters are obtained; establishing an electrical performance calculation model of the two-dimensional honeycomb structure planar radome; determining an antenna-radome system, dispersing the three-dimensional radome in the axial direction to form a plurality of design points, and calculating a weighted average incident angle and force thermal load distribution of each design point; calculating an initial honeycomb relative density recommendation value at each design point; calculating the wave transmission rate and the aiming error of the formed variable-density radome; calculating the force thermal deformation and stress of the variable-density radome; and judging whether the scheme performance meets constraint conditions or not. By optimizing the distribution rule of the variable-density honeycomb structure, comprehensive optimization of the bearing performance and the wave-transparent performance of the radome is achieved.

The application designs a high-speed aircraft cluster cooperative guidance method based on space-time synchronous attraction domain. An artificial potential field is designed to keep a proper distance between aircrafts, avoid collision while gathering; a speed matching item is designed based on consistency theory to realize dynamic consistency of aircraft speed; a guidance law based on resistance acceleration-speed profile is designed to meet the cooperative requirements of terminal range, height, speed and other aircrafts, and the combination of the three realizes efficient autonomous clustering and guidance of the aircraft. The application can make the aircraft cluster reach the target point with high precision under the conditions of meeting complex constraints such as heat flow, overload, dynamic pressure, collision avoidance, clustering, dynamic consistency of speed and the like through the design of attraction domain, angle of attack and roll angle control.

The application discloses a hypersonic flow instability multi-field coupling prediction method and system, relates to the technical field of modern high-speed aircraft design, and comprises the following steps: acquiring the multi-physical field parameter distribution of a hypersonic flow field; based on a shock wave-boundary layer interaction mechanism, constructing an instability sensitive graph data structure by taking the multi-physical field parameters as input, setting graph nodes in the shock wave influence area, the boundary layer area and the intersection area thereof, and calculating the connection weight between the nodes; performing multi-scale decomposition on the instability sensitive graph data structure, and extracting instability modes of the shock wave scale, the vortex scale and the molecular scale; screening the instability modes by adopting a hypersonic stability criterion, and determining a dominant instability mode; and constructing an instability evolution prediction model based on the dominant instability mode, and outputting the spatiotemporal evolution law of the hypersonic flow instability. The application can solve the problem that the flow instability phenomenon under the complex multi-physical field environment of hypersonic speed is difficult to accurately predict.

The invention belongs to the technical field of high-speed wind tunnel tests, and particularly relates to a high-speed aircraft forward jet flow wind tunnel test device and a design method thereof. A thrust balance, a jet flow device and a ring type balance device are arranged in a jet flow test model of the test device, and the jet flow test model is externally connected with an air supply pipeline and a supporting device. The design method comprises the steps of selecting simulation parameters; carrying out similar simulation on the jet flow momentum; designing a jet flow test model; determining a jet flow test model supporting mode; designing a jet flow device; designing a jet thrust balance; carrying out ground jet flow debugging; and performing wind tunnel test verification. According to the test device and the design method thereof, direct measurement of jet flow thrust and direct measurement of aerodynamic characteristics of a jet flow test model are achieved, direct measurement of the jet flow thrust is achieved under the condition that the internal space of the model is not increased, and the jet flow thrust can be accurately obtained while the axial force of the model is obtained; therefore, the method has an engineering practical value.

The utility model discloses a zero-response temperature measurement assembly resistant to high-speed thermal shock, and relates to the field of temperature measurement. According to the utility model, the problem that the fitting degree between the thermocouple wire temperature measuring point and the housing cannot be accurately ensured and the real-time accurate measurement is influenced because the temperature measuring point of the thermocouple wire is separated from the armored housing by a certain distance in the existing temperature measuring assembly is solved. A temperature sensitive assembly (2) is inserted in an armored protective tube (1) and is installed in the armored protective tube (1) through an installation assembly. The tungsten-rhenium thermocouple (2-3) is installed in the silicon nitride porcelain tube (2-2), the silicon nitride porcelain tube (2-2) is installed in the framework (2-1), the front portion of the tungsten-rhenium thermocouple (2-3) is welded into a sphere and then fixedly connected with the framework (2-1) into a whole in a welding mode, and the framework (2-1) is screwed at the bottom of the front end of the armored protection tube (1). The device is used for measuring the tail flame temperature of the high-speed aircraft.

Disclosed herein are a vertical take-off and landing aircraft, a maneuvering method of the same, and a control apparatus for the same. The vertical take-off and landing aircraft includes: a fuselage; a lift unit including a first propeller disposed at a periphery of the fuselage to provide lift to the fuselage and a duct protecting the first propeller; and a thrust unit including a second propeller disposed at a rear end of the fuselage to provide thrust to the fuselage. Here, the duct includes: a propeller guard portion; a leading airfoil portion; a trailing airfoil portion; and a flow guide portion. The propeller guard portion is convexly formed toward the first propeller to have a central diameter smaller than a diameter of an inlet for air to flow in therethrough and a diameter of an outlet for air to flow out therethrough. The leading airfoil portion is disposed in a front region of the duct colliding with air during forward flight and has a shape corresponding to a shape of a front portion of an airfoil including a leading edge. The trailing airfoil portion is disposed in a rear region of the duct designed to allow air to flow thereover during forward flight and has a shape corresponding to a shape of a rear portion of the airfoil including a trailing edge. The flow guide portion protrudes from the propeller guard portion toward the first propeller and suppresses vortex generation at a tip of the first propeller upon rotation of the first propeller. The vertical take-off and landing aircraft according to the present invention can achieve high performance in terms of high-speed flight and stable flight while having excellent maneuverability and efficiency.

An ultra-high-speed helicopter with stowed single-blade coaxial rotors is provided and relates to the technical field of aircrafts. The ultra-high-speed helicopter includes a fuselage. Wings and tails can be arranged on the fuselage, respectively. Tail blades can be arranged at a tail end of the fuselage. A rotor mechanism can be arranged on a top of the fuselage. The rotor mechanism includes rotors. The rotor mechanism can drive the rotors to be stored in a rotor storage cabin of the fuselage. When the helicopter flies at high speed, the rotors stop working, and tail blades provide propulsive force and the wings provide lift force, so that the defects in the aerodynamic characteristics of the rotors are avoided during high-speed flight. In addition, the rotors can be retracted and stowed in a fuselage to further reduce drag and increase the maximum flight speed.

The invention discloses a low-altitude high-speed flight manned flying saucer, and relates to the technical field of aviation aircrafts, the low-altitude high-speed flight manned flying saucer comprises a saucer-shaped body, a power device and an undercarriage, the saucer-shaped body is composed of a central bearing ball cabin and a peripheral reinforcing ring coaxially sleeved outside the central bearing ball cabin; the central bearing ball cabin is divided into a celestial navigation and communication cabin, a panoramic cockpit and an environment control and life support system cabin from top to bottom; the inner cabin section of the peripheral reinforcing ring forms a passenger cabin; the bearing strength and the space utilization rate of the machine body are effectively improved through the composite structure of the central spherical cabin and the peripheral reinforcing ring; due to the integrated design of the annular duct and the vector engine, the aerodynamic efficiency and the control flexibility in the vertical take-off and landing and high-speed flight states are remarkably improved; an independent environment control system ensures the safety and comfort of passengers in high-altitude flight; the overall layout gives consideration to functionality and maintenance convenience, and has high practical value.

The invention relates to the technical field of data link communication abnormal flow detection, in particular to a high-speed aircraft data link communication abnormal flow detection method and system, and the method comprises the steps: judging a communication message which accords with a preset data link communication protocol specification as a protocol layer compliance message, and carrying out the analysis to obtain a formation cooperation instruction; collecting multi-source task context information of a local high-speed aircraft to generate a task execution field snapshot; projecting the formation cooperation instruction to the task execution field snapshot for comparison and analysis to obtain a multi-dimensional semantic comparison result; if all the multi-dimensional semantic comparison results are passed, the corresponding communication flow is judged to be task-driven normal collaborative flow; and if the judgment result of the multi-dimensional semantic comparison result is that the flow does not pass, judging the corresponding communication flow as the intention camouflage type spoofing abnormal flow. The deception traffic of which the protocol layer is compliant but the tactical intention is abnormal is effectively identified, so that the accuracy of high-speed aircraft data link communication abnormal traffic detection is improved.

This invention relates to the field of target recognition technology, specifically providing a lightweight target recognition method and system for high-speed aircraft terminals. The method includes: extracting a first infrared image sequence when target occlusion is detected; calculating aerosol static distribution index and aerosol dynamic distribution index by analyzing the first infrared image sequence; fusing the two into an aerosol spatial information index; selecting a first target image sequence using a genetic algorithm; calculating optimized parameters for an infrared imaging system by analyzing the aerosol static distribution index and aerosol dynamic distribution index of the first target image sequence; obtaining a second target image sequence using the optimized parameters; and tracking and identifying the target based on the second target image sequence. This invention reduces computational load while maintaining target recognition accuracy, enabling efficient target recognition by a high-speed aircraft terminal in dusty aerosol environments.

The present invention discloses a flight target recognition system based on a photonic neural network and a construction method. The ultrafast flight target recognition system based on a photonic neural network is designed and constructed for high-speed moving aircraft, including but not limited to airplanes, drones, helicopters, gliders, delta wings and other aircraft, as well as missiles, to achieve rapid recognition of high-speed moving targets. The present invention provides a flight target recognition system based on a photonic neural network, including an imaging and filtering module, a light field scaling and coupling module, a photonic neural network module, a mode control module, and an output light field coupling module; a construction method for a flight target recognition system, including artificial neural network design and photonic implementation of the artificial neural network; the present invention designs a photonic neural network transmission and control model based on all-optical passive components, trains for small sample learning of high-speed flying targets, builds static and dynamic target recognition systems, and realizes recognition of high-speed flying moving targets.