324 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 preparation method of an interface self-reaction strong chemical bonding radiant heat protection multilayer gradient coating, and belongs to the technical field of coating surface modification. The invention aims to solve the problems that the existing coating is high in preparation temperature, long in period, high in cost and not suitable for industrial batch production, the interface bonding between the coating and a nickel-based high-temperature alloy matrix is poor, the coating is easy to crack and even lose efficacy due to mismatching of thermal expansion under high-temperature service, and the thermal protection performance of the traditional coating is insufficient. The interface self-reaction strong-chemical-bonding radiant heat protection multilayer gradient coating and a nickel-based high-temperature substrate are subjected to self-reaction at the interface to form strong chemical bonding, the bonding strength of the coating reaches 15 MPa or above, the emissivity value of the coating reaches up to 0.9, the structure of the coating is kept complete and does not fall off after the coating is oxidized in 1250 DEG C high-temperature static air for 30 h, and the coating has the advantages of being simple in preparation process, low in cost and the like. The method has application potential in a nickel-based high-temperature alloy anti-oxidation radiation thermal protection multifunctional integrated metal thermal protection system for a high-speed 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 the field of aircraft wing design, in particular to a multi-stage telescopic wing for improving the flight performance of an aircraft, the extension length and the area of the wing are changed through multi-stage stretching of the wing, and different requirements of the aircraft for the wing in low-speed and high-speed flight states are met. When the aircraft takes off, lands and flies at a low speed, the telescopic wings stretch out to increase the area of the wings and enable the aircraft to obtain enough lift force, and when the aircraft flies at a high speed, the telescopic wings retract to reduce the extension length and area of the wings and enable the aircraft to obtain enough lift force and reduce the resistance of the wings, so that the extension length and area of the wings are changed through extension of the wings; and the aircraft can simultaneously meet different requirements on the extension length and the area of the wings in low-speed and high-speed flight states.

The invention discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic thermal constraint, which comprises the following steps of: determining a rudder leading edge radius and a rudder leading edge sweepback angle by taking a current state nominal trajectory as a basis and taking a rudder leading edge thermal response temperature as a judgment standard; secondly, the cone angle of one conical surface of the folded rudder is determined through iterative calculation, so that the chordwise size of the high-temperature-resistant refractory metal material and the thickness of a skin and a heat-proof coating at the cone position are designed; and thirdly, through the design of the interference position of the arched shock wave of the heat-proof cup, the gap between the fixed rudder and the movable rudder avoids the interference area of the arched shock wave, so that the chordwise distance between the folding gap and the rudder shaft is determined. By means of the aerodynamic configuration design method, the radius of the front edge of the rudder can be as low as 3 mm, the taper angle is as low as 4 degrees, it can be guaranteed that ablation deformation does not occur when the maximum heat flow of the front edge of the rudder does not exceed 10 MW / m < 2 >, the resistance characteristic of a high-speed aircraft can be effectively reduced, the characteristic of near-zero ablation is achieved, and the good aerodynamic configuration is maintained.

The utility model discloses a high-speed large-maneuvering target drone, which relates to the field of unmanned aerial vehicles and comprises a drone body, wings, a horizontal tail, a vertical tail and an air inlet channel. The fuselage can reduce resistance to a great extent to realize high-speed flight; the wings adopt sweepback upper single wings to reduce the resistance during high-speed flight, the problem of overlarge transverse stability caused by sweepback and high placement of the wings can be avoided through the downward reverse design, and meanwhile, the wings can be integrally disassembled and assembled and have better operability and convenience; the control efficiency of a full-motion horizontal tail is high, and the heading stability under a large-angle-of-attack attitude can be improved through the design of the inverted-V-shaped tail; the vertical fin without the control surface can improve the course stability and ensure the matching property of the transverse course stability, and is light in structural weight and small in flight resistance; the bearing separation channel type air inlet channel can avoid landing damage of the air inlet channel in the parachute landing recovery process, and meanwhile the requirements for high total pressure recovery, low air inlet distortion, radar stealth and the like are met.

The invention discloses a GAN-based high-speed aerodynamic heat flow field reconstruction method and system, and relates to the technical field of analogue simulation, and the method comprises the steps: collecting historical heat flow field data of a high-speed aircraft under different flight conditions, and constructing a real aerodynamic heat flow field data set; establishing a heat flow field generator and a discriminator through feature extraction and simulation discrimination training; performing grid initialization and adaptive adjustment according to the distribution change of the heat flow field, and dividing into N hierarchical reconstruction regions; alternately training the generator and the discriminator in each region by adopting a balance training strategy, and constructing N heat flow field reconstruction networks; and monitoring flight condition update data, optimizing by using each reconstruction network, and generating a simulated heat flow field. Therefore, the technical effects of simplifying parameter configuration, reducing process cost and improving timeliness are achieved.

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 discloses an aerodynamic force / thermal agent model-based rapid calculation method for the thermo-dynamic elasticity of a thermal protection plate of a high-speed aircraft, and the method comprises the steps: building an unsteady aerodynamic force / thermal agent model of a thermal protection wall plate in a high-speed flow based on a Kriging agent model method; a structural dynamics method based on a finite element method is used for establishing an aeroelastic module, a transient heat conduction equation is used for establishing an aerothermal module, and a thermal protection wallboard thermodynamic elasticity analysis frame is built. Finally, the aerodynamic force / thermal agent model is used for carrying out thermal protection wall plate thermodynamic elasticity calculation and analysis. Compared with a traditional high-order method, the calculation efficiency is improved by four orders of magnitude.

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 discloses a gas-phase sweating cooling thermal protection system based on self-sustaining channel supply. The gas-phase sweating cooling thermal protection system comprises a gas source supply device and a gas-phase sweating cooling thermal protection device. The gas source supply device comprises a self-sustaining channel, a gas storage tank, a regeneration cooling channel and a spray cooling assembly; the gas-phase sweating cooling thermal protection device comprises a porous medium layer and a cooling working medium supply channel; the self-sustaining channel comprises a front section part and a rear section part and is used for sucking high-speed incoming flow and conveying the high-speed incoming flow to the gas storage tank; the regeneration cooling channel is arranged on the periphery of the front section part and is used for cooling high-speed incoming flow for the first time; the spray cooling assembly comprises a plurality of spray coolers arranged in the rear section part and is used for cooling the high-speed incoming flow for the second time; the porous medium layer is arranged on the outer surface of the aircraft. The device is applied to the technical field of thermal protection of high-speed aircrafts, the mass of the coolant carried by the aircrafts is greatly reduced, and the cruising ability of the aircrafts is improved while the cooling efficiency is guaranteed.

The invention discloses an intelligent flow control method and system for an air-breathing high-speed aircraft, and relates to the field of flow control of aerodynamic performance of aircrafts, and the intelligent flow control method comprises the following steps: S1, arranging a suction module on a compression surface of an air inlet channel of the aircraft and / or a position, which is easy to cause flow separation, in the aircraft; s2, an air blowing module is arranged in the area, needing drag reduction, of the outer surface of the aircraft; and S3, the air source obtained through control of the driving module from the suction module is conveyed to the air blowing module through a regulating valve through the control system. According to the intelligent flow control method for the air-breathing type high-speed aircraft, air-source-free carrying of aircraft flow control is achieved, suction and blowing flow control can be achieved at the same time, namely, the aerodynamic performance of an air inlet channel is improved through suction, the aerodynamic performance of the outer surface of the aircraft is improved through blowing, and resistance is 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 relates to the technical field of projectile trajectory detection, in particular to a trajectory detection and calculation method for a high-speed flying projectile, and the method comprises the steps: arranging a plurality of binocular vision collectors, a pressure sensor array and a plurality of high-speed cameras at two sides of a flight path of the projectile; three-dimensional point cloud data, shock wave pressure pulsation signals and moving track images of the projectile are collected respectively. Collected data are subjected to filtering, registration, mesh generation, splicing, deblurring and the like and then are aligned and fused according to timestamps, more accurate position and speed information is obtained through a weighted average method, then trajectory points of projectiles are calculated, and noise is eliminated through a smoothing algorithm. Finally, the method can generate the complete space trajectory of the projectile in the flight process, and the analysis result is displayed in a visual mode. According to the technical scheme, through multi-sensor data fusion, the precision and stability of projectile trajectory detection are improved, and the problem that in the prior art, trajectory calculation is low in accuracy is effectively solved.

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 multi-motion-mode wheel-legged robot. The multi-motion-mode wheel-legged robot comprises a robot body, bionic leg walking mechanisms and wheel paddle mechanisms; an electric power storage control system is mounted on the machine body; the number of the bionic leg walking mechanisms is four, and the four bionic leg walking mechanisms are symmetrically installed on the two sides of the robot body through the side swing leg lifting assemblies correspondingly. The side swing leg lifting assembly is used for obstacle avoidance of the bionic leg walking mechanism. The four groups of wheel paddle mechanisms are respectively mounted at four corners of the machine body; the wheel paddle mechanism comprises walking wheel assemblies and propeller assemblies, the walking wheel assemblies are installed at the four corners of the fuselage through the mode switching assemblies, and the propeller assemblies are installed on the walking wheel assemblies. The high-speed moving characteristic of a wheeled robot, the advantage of obstacle crossing ability of a legged robot and the high-speed flying ability of an aerial robot are integrated, and the all-terrain robot has excellent all-terrain adaptability and can efficiently pass through various complex terrains.

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 cross-medium projectile body vertical water entry impact overload / load reduction response characteristic test device and method, belongs to the field of projectile body cross-medium high-speed flight impact tests, and aims to solve the problem of how to give consideration to reduction of impact load borne by a projectile body during cross-medium motion and reduction of self weight. The device comprises an underwater shock wave emitter, a projectile body, a test system and the like. The underwater shock wave generator is composed of a light-gas gun launching tube, a sabot flying plate, a piston, a dynamic pressure sensor and the like. The projectile body is divided into a buffer cap and a non-buffer cap; the test system is composed of a dynamic strain gauge, an axial acceleration sensor, a radial acceleration sensor, a three-dimensional DIC system and the like. According to the invention, different impact wave intensities in water are used for simulating conditions of vertical entry into water at different speeds, impact waves directly act on the front end face of the projectile body to realize impact loading on the projectile body, and the test system is used for realizing research on impact overload / load reduction response characteristics of the projectile body.

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 invention provides a plasma virtual control surface and control mechanism control distribution method for reentry flight of a high-speed aircraft. The method comprises the following steps: firstly, determining that the aircraft enters a reentry flight stage; secondly, the needed torque is calculated according to an attitude controller; then designing control distribution models of a plasma active flow control system, a pneumatic control surface control mechanism and an RCS backstepping control system based on a depth deterministic strategy gradient reinforcement learning algorithm; and the high-speed aircraft completes reentry flight based on the control distribution parameters obtained by the control distribution model. Auxiliary control is achieved through the plasma jet, and a new way is provided for accurate control over the aircraft. According to the method, control means can be reasonably distributed according to the early stage, the middle stage and the later stage of reentry flight, and relevant parameters of all control systems are determined. The reinforcement learning algorithm is used for processing the nonlinear characteristics of the input power and control output of the plasma exciter and the nonlinearity and uncertainty of aerodynamic characteristics in an attitude control model, the stability of a control system can be improved, and the method has a good application prospect.

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 aircraft preset time formation cooperative guidance method with double constraints. The method comprises the following steps: 1, constructing a three-channel decoupling cooperative guidance motion system; 2, constructing a disturbance observer based on preset time; 3, constructing a consistency guidance protocol; 4, constructing a sight angle control protocol based on preset time; and 5, controlling the high-speed aircraft to intercept the maneuvering target by applying the preset time disturbance observer, the preset time consistency protocol and the preset time sight angle control protocol, and outputting a state response diagram. The method has the advantages that three-channel decoupling of the leader following system is achieved through geometric analysis, and then a high-speed aircraft preset time control task suitable for a leader following framework is constructed; 2, the estimation of the residual flight time of the high-speed aircraft is avoided, and the stability of the high-speed aircraft is proved; and 3, designing a lateral angle constraint control algorithm based on preset time, and realizing preset time stability of leader-follower transverse and lateral system states.

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 invention discloses a closed cabin sleeving gap measuring device and method, and belongs to the technical field of industrial automatic production. The problem that in the prior art, the cup joint gap distance is difficult and complex to measure and calculate is solved. According to the method, the cup joint process is simulated through the closed thermal protection model and the closed cabin section model, the characteristic data of the cup joint part are synchronously collected through the multiple gap measuring instruments, the condition of the current cup joint gap is judged by analyzing the characteristic data, and the cup joint gap is adjusted to the target range. The method is mainly used for measuring the sleeving gap of the closed cabin 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.