36 results about "Aviation engineering" patented technology

The invention belongs to the technical field of aviation engineering tests, and discloses a flap control surface loading method and system for an amphibious aircraft ground integration test, and the method comprises the steps: designing a parallel loading system composed of a rocker arm, a pull rod and a loading cylinder, and carrying out the loading of a flap control surface according to a preset flap load spectrum in a flight control ground test, the loading cylinder is used for driving the parallelogram mechanism to simulate the action of the pneumatic hinge moment on the control surface in real flight. The problem that a traditional loading mode is difficult to accurately simulate the distributed load borne by the flap in real flight is solved, and the verification accuracy and reliability are improved.

The invention relates to:-a laminated glass sheet comprising a first glass sheet (1) forming the outer face of the glass sheet, the first glass sheet (1) being connected to a second glass sheet (3) by means of a first interlayer adhesive layer (2), the edge of the first glass sheet (1) being indented relative to the edge of the second glass sheet (3), the peripheral portion of the free surface of the first glass sheet (1), the edge face of the first glass sheet (1), the edge face of the first interlayer adhesive layer (2) and the surface portion of the second glass sheet (3) extending beyond the first glass sheet (1) describe a continuous stepped profile covered with stepped elements (7) in which an adhesive (6) is inserted, said stepped element (7) being made of a polymeric material, which may contain a reinforcing filler, exhibiting a water vapor permeability of at most equal to 5 g / m2 / day; a method for the production thereof; -its use (aeronautical engineering and the like).

The application discloses a simulation analysis method for an aircraft landing gear tire under heavy landing impact load, and relates to the technical field of aeronautical engineering and computational mechanics. First, in the tread and sidewall rubber area, the micro-heterogeneity of the rubber material in the tire is equivalent to the super-elastic mechanical behavior of the macro-continuous medium based on the homogenization theory. This process avoids the extremely dense mesh and complex constitutive relationship required for depicting the micro-features such as filler distribution and interface effect in the traditional model. Then, the rebar element representing force transmission under impact load is created in the pre-constructed rubber matrix grid, and the rebar element is coupled with the rubber matrix. The simplified coupled model can accurately capture the anisotropic stiffness, main force transmission path and reinforcement effect of the cord-rubber composite structure, so as to finally stably and accurately simulate the transient impact process of the tire under heavy landing impact load, which includes the triple nonlinearity of geometry, material and contact.

The invention relates to the technical field of aeronautical engineering, and discloses an aero-engine performance decline trend prediction and repair method, which comprises the following steps: step 1, establishing an intelligent prediction model, monitoring key data of an engine in real time through a sensor, predicting decline trends of all parts in combination with a machine learning algorithm, and providing early warning of repairing opportunities in advance for maintenance personnel; 2, applying a local repairing technology, identifying a part needing to be repaired based on the intelligent prediction model, and accurately repairing a damaged part by adopting laser repairing and metal spraying technologies, so that the maintenance cost is reduced, and the service life of the part is prolonged; and step 3, accurate repair opportunity selection: the system monitors the state of each part of the engine in real time, and automatically sends out repair early warning according to the decline trend. Key data of the aero-engine are monitored in real time by constructing an intelligent prediction model, the performance decline trend of the parts is accurately predicted, and the repair opportunity is early warned in advance. The problems that in traditional maintenance, the repairing scheme is single, and the opportunity is inaccurate are solved.

The invention belongs to the technical field of artificial intelligence and aviation engineering crossing, and provides an aero-engine test document automatic generation method and system based on demand analysis. Segmenting the test technical requirement document by adopting a hybrid text segmentation algorithm based on semantic vector similarity, and constructing a bimodal storage structure; high-precision fusion of multiple paths of retrieval results is realized by utilizing a reciprocal ranking fusion algorithm and a weighted reciprocal ranking fusion algorithm, and a high-correlation context is extracted; structured test points are automatically analyzed and generated based on a fine-tuning large model, and multi-target task grouping is carried out in combination with test rules such as a safety envelope, positive and negative temperature isolation and single-day duration; and checking and outputting compliance test task groups through a rule-driven post-processing mechanism. According to the method, the production scheduling efficiency and the resource utilization rate are remarkably improved, dynamic adjustment is supported, and the method is suitable for multi-equipment and high-complexity aero-engine test scenes.

The invention relates to a self-driven throttle lever damping force device and a driving method, and belongs to the technical field of aviation engineering. Comprising a turbine (1), an operating rod (2), a deep groove ball bearing a (3), a damping force adjusting nut (4), a thrust bearing (5), a disc spring (6), a main friction disc (7), a worm (8), an auxiliary friction disc (9) and a deep groove ball bearing b (10). The auxiliary friction disc (9), the main friction disc (7), the disc spring (6), the thrust bearing (5) and the operating rod (2) are sequentially installed on a shaft of the turbine (1) and fastened through the damping force adjusting nut (4) to form the human-sensing damping assembly, and then the whole human-sensing damping assembly is installed on the deep groove ball bearing a (3) and the deep groove ball bearing b (10). The damping device can be used for providing a stable damping force feeling in the stroke of the rotary throttle lever of the aircraft, meanwhile, the motor can be additionally arranged for self-driving, and the portability and safety of pilot operation are improved.

The invention discloses an aircraft hangar maintenance and warehousing transfer cooperative scheduling method and device, and relates to the technical field of scheduling optimization technology and aviation engineering management, and the method comprises the steps: building a cooperative scheduling problem model; in a transfer scheduling stage, minimizing transfer time and a maintenance urgency proxy target are taken as optimization targets, and an intelligent optimization algorithm is utilized to solve an aircraft warehousing transfer sequence and actual entrance time; in a maintenance scheduling stage, taking actual entrance time as a constraint, taking maximization of machine group wave availability as an optimization target, and solving a hangar maintenance scheme and maintenance completion time by utilizing an intelligent optimization algorithm; the maintenance completion time is fed back to a transfer scheduling stage, iteration is repeated to convergence, and a converged aircraft in-storage transfer and hangar maintenance scheduling scheme is output. Through repeated execution of a transfer-maintenance-feedback cycle, tight coupling between two sub-problems of transfer and maintenance is effectively solved; and the global optimality of the scheduling scheme is obviously improved.

The invention relates to the technical field of aircraft landing gear hydraulic systems, in particular to an aircraft landing gear hydraulic system coordinated extension and retraction function reliability evaluation method, which comprises the following steps: constructing an AMESim-MATLAB-based landing gear hydraulic system joint performance simulation model; determining uncertain parameters and interval boundaries of the uncertain parameters which influence the coordinated folding and unfolding function; on the basis of the interval boundary, constructing a performance function representing the coordinated receiving and releasing function of the system, calculating an output interval of the performance function through joint simulation, and further calculating a non-probability reliability index to evaluate the reliability of the system; and based on the output interval, identifying a key variable influencing the reliability through global sensitivity analysis. According to the method, the non-probability interval theory is adopted, evaluation can be completed only through parameter boundary information, the defect that a traditional probability method depends on mass data and expensive computing resources is overcome, the method is particularly suitable for aviation engineering scenes with scarce experimental data, weak links can be accurately recognized, and a quantitative basis is provided for system optimization and maintenance.

The invention relates to the field of temperature field numerical calculation, in particular to a temperature field calculation method and system based on a continuous submerged space induction point strategy. The method comprises the following steps: firstly, acquiring material attributes of a to-be-measured thermal component and unstructured grids and temperature data in a specified application scene, mapping the unstructured grids and temperature data to a continuous submerged space through an encoder, and generating a structured feature map in combination with induction points; performing dynamic feature fitting and time marching on the feature map by using a Unet module to obtain a feature map at the next moment; and finally, inversely converting the temperature vector into an unstructured temperature vector through a decoder, and optimizing the model through a loss function containing a heat conduction equation and a boundary condition. The problem that a traditional CNN cannot directly process unstructured data is effectively solved, efficient and high-precision solving of a transient temperature field in a complex geometric model is achieved, the efficiency and reliability of engineering thermal simulation are remarkably improved, and the method is suitable for aviation engineering thermal component design and a digital twin system.

The application belongs to the technical field of aeronautical engineering remanufacturing, and discloses a kind of aircraft remanufacturing data enhancement methods based on physical failure-bayesian deep learning fusion, which comprises the following steps: collecting multi-source heterogeneous data and preprocessing;Establish a multi-physics field coupled physical failure model, and construct a physical prior knowledge base;Build PoF-BGAN, and add the output of the physical failure model as a regular term to the loss function;Adaptive clustering and cross-modal mapping are used to realize multi-modal data alignment;Based on physical failure sensitivity analysis, edge fault samples are generated;From the statistical, physical and task dimensions, the generated data quality is evaluated and a feedback loop is established;The remanufacturing quality evaluation model is trained using the enhanced data set.The application solves the problems of fault sample scarcity, data heterogeneity, poor physical consistency of generated samples and missing extreme working condition samples in the field of aviation remanufacturing, and provides high-quality data support for remanufacturing quality evaluation and process optimization.

The invention provides a multifunctional wind tunnel test device for the surface of a turbulence drag reduction micro-rib, and belongs to the technical field of aviation engineering. The objective of the invention is to solve the problem that an existing test device is unfolded around a low-speed wind tunnel and is difficult to adapt to research requirements of small high-speed wind tunnel size and high flow speed in the aviation field. Optical windows are arranged on four walls of a test area of a wind tunnel test section hole body, a first flat plate is mounted on the wind tunnel test section hole body and connected with a third flat plate, and a second flat plate is mounted at the bottom of the first flat plate; the test area bottom plate is mounted in the first flat plate groove through the test area flexible interlayer, and the micro rib surface test sample piece is paved on the test area bottom plate; a PIV test particle release groove is formed in the first flat plate and is connected with a PIV test particle release pressurizing cabin box arranged at the corresponding position of the bottom of the second flat plate; and a turbulent flow array is arranged on the third flat plate. The problem that an existing test device is unfolded around a low-speed wind tunnel and is difficult to adapt to research requirements of small size and high flow speed of a high-speed wind tunnel in the aviation field is solved.

The application discloses an aircraft hangar maintenance and warehouse entry transfer collaborative scheduling method and device, relates to the technical field of scheduling optimization and aviation engineering management, and comprises the following steps: establishing a collaborative scheduling problem model; in the transfer scheduling stage, taking the minimization of the transfer time and the maintenance urgency proxy target as the optimization target, solving the aircraft warehouse entry transfer sequence and the actual entry time by using an intelligent optimization algorithm; in the maintenance scheduling stage, taking the actual entry time as the constraint and taking the maximization of the fleet wave availability as the optimization target, solving the hangar maintenance scheme and the maintenance completion time by using the intelligent optimization algorithm; feeding back the maintenance completion time to the transfer scheduling stage, repeatedly iterating until convergence, and outputting the converged aircraft warehouse entry transfer and hangar maintenance scheduling scheme. Through repeated execution of the cycle of "transfer-maintenance-feedback", the close coupling between the two sub-problems of transfer and maintenance is effectively solved, and the global optimality of the scheduling scheme is significantly improved.

The invention discloses a small stackable plasma propulsion engine for an unmanned aerial vehicle, and belongs to the technical field of aviation propulsion. Comprising a plurality of engine modules, each engine module comprises a cylindrical shell, the diameter of the shell is 10 cm, and the height of the shell is 8 cm; a plurality of emission electrode groups and a plurality of collection electrode groups are arranged in the shell in a surrounding manner; a connecting structure is arranged between every two adjacent engine modules. The connecting structure comprises an internal connecting mechanism and an external connecting mechanism, the internal connecting mechanism is a threaded connecting mechanism, and the interiors of the adjacent engine modules are connected in a screwing-in mode; the external connecting mechanisms are bolt connecting assemblies, and the exteriors of the adjacent engine modules are connected and fixed through the bolt connecting assemblies. The invention has the advantages of compact structure, adjustable thrust, high energy conversion efficiency, large thrust-weight ratio and the like, can be widely applied to tasks such as deep space exploration, orbit maintenance, attitude adjustment and the like, and has important aviation engineering application value.

The utility model relates to the technical field of aeronautical engineering, and discloses a quick-connection detachable blowdown vehicle for airplane blowdown, the top of a connecting piece is provided with a drain pipe, the surface of the drain pipe is provided with three grooves, the bottom end of a connecting pipe is fixedly connected with a corrugated pipe, the bottom end of the corrugated pipe is fixedly connected with a water outlet pipe, and the bottom end of the water outlet pipe is fixedly connected with a water inlet pipe. A valve is fixedly connected to the bottom end of the water outlet pipe, a water bucket is fixedly connected to the top of the bottom plate, a fixing block is fixedly connected to the top of the water bucket, a plurality of hasps are rotationally connected to the tops of the fixing block and the connecting block, and a plurality of locking blocks are fixedly connected to the surfaces of the connecting pipe and the water outlet pipe. According to the portable sewage discharging device, the independent portable sewage discharging device formed by connecting the connecting pipe, the corrugated pipe and the water outlet pipe can improve the sewage discharging efficiency, increase the portability, is small in occupied area, can be carried and used along with an airplane, can be adsorbed on an airplane body through the suction cups to collect water discharged from a discharging opening, and avoids the risk that the water escapes to pollute an airplane apron.

The application provides a health state evaluation method and system of a rescue helicopter, and relates to the technical field of aeronautical engineering.The method comprises the following steps: step 1, an integrated data aggregation platform is established, flight operation parameters, maintenance archives, abnormal event records and surrounding environment monitoring indexes are synchronously collected, key dimensions of rotor devices, power transmission units, propulsion devices and sea climate conditions are comprehensively covered, and a comprehensive original data set is formed; and step 2, based on the comprehensive original data set, an adaptive parameter adjustment mechanism is used to comprehensively calculate cumulative operation time length indexes, periodic component residual efficiency values, core health scores, abnormal repeated occurrence probabilities, typical problem correlation strengths, sudden event severity and marine climate erosion elements, and generate comprehensive state indexes.The application can precisely evaluate the health state of a rescue helicopter, reduce flight risks and improve rescue decision-making and execution efficiency.

The application relates to a kind of comprehensive performance ground test device and method of turbulent flow drag reduction micro-rib surface, belong to aeronautical engineering technical field.It includes a set of replaceable test assembly installed on ground station main body;Ground station main body is connected with high-pressure gas source interface;Particle release groove is installed on ground station main body, and is connected with dust release device and water droplet release device respectively;It also includes data acquisition instrument and test data comprehensive processing computer connection;Test assembly is the first component for micro-rib surface drag reduction performance measurement, or is the second component for micro-rib surface wear resistance and stain resistance performance measurement.Can effectively simulate high-speed turbulent boundary layer flow, high-friction wall turbulent flow, water droplet / dust and its combination high-speed wall turbulent flow, to support the effective development of micro-rib surface test sample drag reduction performance, wear resistance, stain resistance and other test tests, support micro-rib surface test sample comprehensive performance evaluation, and then promote the technical iteration of micro-rib surface design and preparation.

The invention discloses a tail skid system mounting device, which belongs to the technical field of civil aviation engineering, and comprises a front reinforcing plate, a rear reinforcing plate, an upper reinforcing plate, a lower reinforcing plate, a first mounting piece, a second mounting piece and a supporting piece, and the rear reinforcing plate and the front reinforcing plate are arranged in a mounting groove at an interval and are both connected with a fuselage skin; the upper reinforcing plate is located between the front reinforcing plate and the rear reinforcing plate, connected with the tops of the two and connected to the fuselage skin. The lower reinforcing plate is attached to and connected with the fuselage skin, the bottoms of the front reinforcing plate and the rear reinforcing plate are connected to the lower reinforcing plate, and a box type frame is defined by the front reinforcing plate, the rear reinforcing plate and the upper reinforcing plate. The first mounting piece is arranged on the lower reinforcing plate and used for rotationally mounting the friction block assembly; the second mounting piece is fixed on the rear reinforcing plate and is used for rotationally mounting the actuator cylinder assembly; the supporting piece is located in the box type frame and fixedly connected with the front reinforcing plate, the rear reinforcing plate, the upper reinforcing plate and the lower reinforcing plate. The problem of fuselage structure damage caused by load concentration can be solved, and therefore test flight safety is guaranteed.

The invention discloses an aviation plug detection method and system, a medium and a program product, and relates to the field of aviation engineering. The aviation plug detection method comprises the following steps: determining a template picture pinhead coordinate according to an aviation plug template image, and obtaining a to-be-detected aviation plug image after a to-be-detected aviation plug is fixed and a pinhead is colored; according to the to-be-detected aviation plug image and the needle coordinates of the template picture, identifying the current pin state; and when the current pin state is a pin shrinking state, rechecking the current pin state to obtain a current pin state result. According to the technical scheme of the embodiment of the invention, the complexity of aviation plug detection is not high, and the plug state can be accurately identified.

The utility model discloses a fixed wing aircraft assembly fixture structure, which relates to the technical field of aviation engineering, and comprises an assembly base, the assembly base comprises a wing fixing mechanism, an empennage fixing mechanism and a fixing seat, the wing fixing mechanism and the empennage fixing mechanism comprise a vertical seat and a fixing column, and the vertical seat and the fixing column are connected through a connecting rod. A motor is arranged on one side of the fixing column, the output end of the motor is connected with a first bevel gear, the first bevel gear is in meshed connection with a second bevel gear, the inner ring of the second bevel gear is connected with a lead screw, the outer ring of the lead screw is sleeved with a first sliding seat, and one side of the first sliding seat is connected with an upper cover plate. A first sliding groove is formed in one side of the fixing column. The fixing mechanism has the effect of being high in position clamping stability, can ensure that the airplane structural part is kept at the correct position and posture in the assembling process, and reduces assembling errors caused by displacement or deformation.

The invention belongs to the technical field of aviation engineering design, and discloses an aircraft hidden fault indication recording system and an alarm database design method.The aircraft hidden fault indication recording system comprises a plurality of sensors used for feeding back state parameters of an aircraft in real time, and the state parameters comprise voltage, current, vibration, frequency, Hall, displacement, temperature and humidity; the at least one processing unit is used for receiving the state parameters of the sensor and external system data so as to realize a system design expected function; the at least one resident application is used for synthesizing the abnormal state of the whole aircraft based on the feedback data of each processing unit and the sensor data, then giving out comprehensive judgment and feeding back possible hidden faults; wherein the system is composed of the sensor, the processing unit and the resident application according to the complexity of cross-linking of the aircraft system. According to the invention, the problem that the existing hidden fault of the airplane is difficult to position and troubleshoot is solved, so that the hidden fault of the airplane is quickly positioned and troubleshot.

The utility model relates to the technical field of aviation engineering, in particular to a sealing ring with a good sealing effect, which comprises a sealing ring body, a shell is arranged on the surface of the sealing ring body, a connecting plate is welded on the surface of the shell, a bolt is arranged on the surface of the connecting plate in a penetrating manner, and an adjusting mechanism is arranged on the inner wall of the sealing ring body. The adjusting mechanism comprises a cavity, the cavity is formed in the inner wall of the sealing ring body, a supporting plate is movably connected to the inner wall of the cavity, and a supporting rod and a reset spring are fixedly installed on one side of the supporting plate. Through the cooperation of the cavity, the supporting plate, the supporting rod, the reset spring and the sealing inner ring, the sealing ring is made of high-elasticity materials, can generate elastic deformation under the action of system pressure, is tightly attached to a sealing face, can adapt to tiny deformation such as vibration and impact under different working conditions, and ensures the stability and durability of the sealing effect.

The invention belongs to the technical field of aviation engineering design, and discloses a connecting device and a connecting method for a front section and a rear section of a rear fuselage of an aircraft. The connecting device for the front section and the rear section of the rear fuselage of the aircraft comprises a first connecting piece, a second connecting piece, a first adjusting assembly and a second adjusting assembly, wherein the first connecting piece is connected to the second connecting piece to fix the front section and the rear section of the rear fuselage; the first adjusting assembly comprises a first adjusting piece sliding in the second direction and a plug connector fixed to the first adjusting piece. The second adjusting assembly comprises a second adjusting piece sliding in the third direction, and the second adjusting piece is provided with an inserting hole connected with the inserting piece. According to the connecting device for the front section and the rear section of the rear fuselage of the aircraft, aligned installation of the front section and the rear section of the rear fuselage of the aircraft is guaranteed, the installation strength and the installation efficiency are improved, the connecting method for the front section and the rear section of the rear fuselage of the aircraft guarantees the machining precision of the front section and the rear section of the rear fuselage of the aircraft, and the connecting efficiency is improved; and the overall structural strength and the flight safety of the rear fuselage are ensured.

The invention provides a power battery temperature prediction method and device based on eVTOL typical flight characteristics, and belongs to the technical field of aviation engineering such as an aircraft technology, a power technology, an avionics technology and an electromechanical system technology. The method comprises the following steps: collecting time sequence data of current and temperature in real time after the eVTOL takes off, calculating a time-current curve curvature through a high-precision non-equidistant five-point curvature method, and determining a flight state conversion time point; according to typical characteristics of different flight states, fitting a temperature-time change relationship between the take-off stage and the initial stage of the cruise stage through partial data of the take-off stage and the cruise stage; and predicting the temperature change of the remaining voyage in the flight process by using the obtained fitting parameters in combination with the remaining cruise duration and the landing duration. The method has the characteristics of simplicity, high efficiency and high practicability, can predict the temperature change in the later flight period in the initial flight period, provides effective data support for eVTOL power battery thermal management, and lays a foundation for safe flight.

The invention belongs to the technical field of aeronautical engineering structure tests, and provides a movable airfoil aerodynamic load processing method in a full-size aircraft fatigue test, and the method comprises the steps: converting the original coordinates and loads of a plurality of discrete nodes on a movable airfoil under a full-aircraft coordinate system into a local coordinate system with an airfoil rotating shaft as a reference; calculating the loaded pressure center of each working condition based on the load distribution in the local coordinate system, and determining the position of a loading point for applying an equivalent load according to the loaded pressure center of multiple working conditions; on the basis of the node coordinates and the load under the local coordinate system, original rotating shaft torque borne by the movable airfoil under each fatigue test working condition is calculated; and according to the position of the loading point and the original rotating shaft moment, inversely solving an equivalent load needing to be applied at the loading point, and enabling the total rotating shaft moment generated after loading to be equal to the original rotating shaft moment. According to the method, mechanical distortion caused by coordinate simplification in a traditional method can be avoided, and the accuracy of fatigue test load loading and the engineering reliability are remarkably improved.

The invention belongs to the field of aviation engineering, and relates to an amphibious aircraft landing configuration warning system and method. The amphibious aircraft landing configuration alarm system comprises an undercarriage system, an alarm system, a navigation system, an airborne system, a selection switch and a plurality of sensors. The sensor is mounted on the undercarriage system to feed back the current real folding and unfolding state of the undercarriage; the alarm system is used for collecting selection of the selection switch, collecting the current real folding and unfolding state of an undercarriage, collecting state parameters related to an aircraft landing scene from the navigation system and the airborne system, comprehensively judging whether the landing scene is matched with the state of the aircraft undercarriage or not according to the collected information and based on the identification logic of landing configuration alarm, and giving an alarm according to the judgment result. The method is used for recognizing and generating landing configuration alarms in advance.

The present invention belongs to the technical field of test inspection, and specifically relates to a test apparatus for the containment capability of an aero-engine casing, comprising a horizontal base plate, an air cannon device, a support device, a measurement device and a safety protection device, wherein the air cannon device comprises an air cannon base, an air cannon chamber, an electrical cabinet, a chamber controller, a flange, a cannon barrel, and a component detachment device; the support device comprises a cannon barrel bracket, a pair of infrared dot sensor brackets, and a dual-port fixture; the measurement device comprises infrared dot sensors, a signal transceiver, a power supply, a data acquisition instrument, a high-speed camera, and a data processor; and the safety protection device comprises explosion-proof glass, a protective panel and a pair of protective walls. The present invention having a rationally designed structure allows accurate simulation of the impact between screws or nuts after failure and the casing and adjacent components, thereby improving the safety and reliability of the casing, providing significant application prospects in aviation engineering.

The invention relates to the technical field of aviation engineering, and provides a stealth coating in-situ radiation heating curing device and method.The device comprises a radiation heating module, an acquisition module, a processing module and a control module; the radiation heating module comprises a mounting main body and a plurality of radiation heating pieces arranged on the mounting main body; the processing module is connected to the acquisition module; the control module is connected to the processing module; the structural performance parameters of the stealth coating are obtained through the collection module, the design strategy of the radiation heating module is determined based on the structural performance parameters of the stealth coating, and the design strategy comprises at least one of the arrangement position of the radiation heating pieces on the installation body, the number of the radiation heating pieces and the design parameters of the radiation heating pieces; the radiation heating piece is regulated and controlled based on the design strategy, so that the whole stealth coating is rapidly heated and cured through the regulated and controlled radiation heating module according to the curing requirement of the stealth coating, and the maintenance efficiency of the stealth coating can be improved on the premise that the maintenance quality of the stealth coating is guaranteed.

The invention relates to the field of aeronautical engineering and aerodynamic noise control, and provides a noise reduction structure for a flap side edge. The noise reduction structure for the flap side edge comprises a three-dimensional gradient front edge flow guide structure which is arranged at a front end inlet of the flap side edge; the side edge vortex blocking baffle is of a sheet-shaped structure and is arranged on the downstream of the three-dimensional gradual change front edge flow guide structure; the sawtooth-shaped upper edge vortex control structure is arranged on the upper edge, away from the flap airfoil, of the side edge vortex blocking baffle. According to the noise reduction structure for the flap side edge, the speed difference between the noise reduction structure and a root low-energy shearing layer can be reduced, and the forming condition of a primary large vortex is destroyed; the core source of the broadband noise is solved, the high-frequency noise peak is avoided, and the broadband noise intensity is greatly reduced.

The utility model relates to the technical field of aviation engineering, in particular to an integral drive generator dismounting auxiliary device, which is characterized in that an adjusting seat comprises a first rotating shaft connected to a lifting frame, a first seat body connected to the rotating shaft, a second rotating shaft connected to the first seat body, a second seat body connected to the second rotating shaft, a first screw rod and a second screw rod; two ends of the first screw rod are respectively connected with the lifting frame and the first seat body; the second rotating shaft is connected to the first seat body, an included angle is formed between the second rotating shaft and the first rotating shaft, the second seat body is connected to the second rotating shaft, and the two ends of the second screw are connected with the first seat body and the second seat body respectively; the clamping piece is connected to the second base body, the first base body can be driven to rotate around the axis of the first rotating shaft by rotating the first screw rod, and the second base body can be driven to rotate around the axis of the second rotating shaft by rotating the second screw rod, so that the clamping part is more attached to the overall driving generator.

The invention provides a runway feature detection method for large aircraft landing, and belongs to the field of aviation engineering tests. The method comprises the steps of performing ROI region detection and segmentation on a runway feature image acquired by a visual sensor based on a deep neural network, introducing a slope and slope distance clustering algorithm and a scoring mechanism to eliminate most interference linear features on a linear detection result, and completing accurate extraction of runway left and right sideline features. The method automatically realizes measurement data feature point extraction, and has the characteristics of large data sample size, objectiveness, accuracy, small error, high practicability and the like.