47 results about "Aviation engineering" patented technology

The invention provides a health state assessment method and system for a rescue helicopter, and relates to the technical field of aviation engineering, and the method comprises the steps: 1, building an integrated data aggregation platform, synchronously collecting flight operation parameters, maintenance files, abnormal event records and surrounding environment monitoring indexes, and carrying out the maintenance of the maintenance files; comprehensively covering key dimensions of a rotor device, a power transmission unit, a propelling device and sea climate conditions to form a comprehensive original data set; and step 2, based on the comprehensive original data set, applying a self-adaptive parameter adjustment mechanism to perform comprehensive calculation on an accumulated operation time length index, a periodic component residual efficiency value, a core health score, an abnormal repeated occurrence probability, typical problem association strength, an emergency severity degree and a marine climate erosion factor, and generating a comprehensive state index. The method accurately evaluates the health state of the rescue helicopter, reduces the flight risk, and improves the rescue decision and execution efficiency.

The invention provides a geometric and physical fusion method for predicting the assembly deviation of an easy-to-deform structure, and belongs to the technical field of aeronautical engineering manufacturing, and the method comprises the following steps: carrying out the structural decomposition of the easy-to-deform structure, obtaining a space vector point set, and representing and calculating a consistent manufacturing deviation, an assembly positioning deviation and an assembly deformation deviation; constructing an assembly deviation transfer diagram, and mapping the assembly deviation transfer diagram into an assembly state space diagram by using the three multi-source deviations; performing finite element analysis on the assembly precision of the easily-deformed structure by utilizing the assembly state space diagram, and converting the obtained assembly deviation cloud diagram into an assembly deviation contour map; constructing an initial generative adversarial network model, and training and verifying the initial generative adversarial network model by using data in the assembly deviation contour map to obtain a final generative adversarial network model; and inputting the actual deviation image of the easily-deformed structure into the final generative adversarial network model for assembly deviation prediction. The assembly precision and the assembly efficiency of the easy-to-deform structure can be improved.

The invention discloses a wing optimization method and device, equipment, a storage medium and a product, and relates to the technical field of aeronautical engineering.The wing optimization method comprises the steps that a wind tunnel device is started in response to a test instruction for a wing of an unmanned aerial vehicle, so that airflow of the wind tunnel device acts on the wing and drives a coating on the wing to flow; shooting to obtain a coating picture until the coating on the wing is stable; and performing feature extraction on the coating picture, and generating a wing optimization scheme based on the obtained optimization features. According to the method, the wing optimization scheme can be generated based on the obtained optimization characteristics only by observing the characteristics of the coating on the wing, and the problem that the equipment cost and the operation complexity of a traditional eddy current detection technology are remarkably poor can be avoided.

The invention provides a method for detecting global temperature distribution of a nacelle loaded with a turboprop engine, and relates to the technical field of aeronautical engineering.The method comprises the steps that a first nacelle model is generated according to temperature distribution of the surface of the turboprop engine at different stages and the ventilation requirements of air intake and exhaust of the nacelle; extracting a fluid computational domain of the first nacelle model, and performing temperature field and airflow field analysis on the fluid computational domain to obtain a temperature distribution cloud picture and an airflow vector distribution picture of the first nacelle model; if the temperature distribution cloud picture and the airflow vector distribution picture both meet the preset requirements, aerodynamic configuration details and internal solid components of the first nacelle model are added, and a second nacelle model is generated; and performing fluid-structure interaction conjugate heat transfer analysis on the second nacelle model under different working conditions, optimizing the nacelle air intake and exhaust layout according to the result of the conjugate heat transfer analysis, and generating a target nacelle ventilation and heat dissipation scheme. The nacelle has the technical effect that the overall performance and reliability of the nacelle are improved.

The invention provides a variable step route planning method and system based on unmanned aerial vehicle performance, and relates to the technical field of aeronautical engineering.The method comprises the following steps that flight information of an unmanned aerial vehicle is obtained, the position of a starting point of the unmanned aerial vehicle is initialized, and the course of the starting point serves as the initial flight direction of the unmanned aerial vehicle; the starting point position of the unmanned aerial vehicle is added to a preset to-be-explored node storage list, and a node with the minimum search cost is selected from the to-be-explored node storage list to serve as a current node; and moving the current node into a preset explored node storage list, expanding a neighborhood point of the current node by using a waypoint neighborhood search method, and updating the current node in combination with the search cost and the list position of the neighborhood point. According to the invention, route planning can be carried out on the premise of considering the kinetic energy of the fixed-wing unmanned aerial vehicle, and the traceability of the planned route is ensured.

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 present application proposes a geometric and physical fusion method for predicting assembly deviation of deformable structures, belonging to the field of aviation engineering manufacturing technology, and comprising the following steps: decomposing the deformable structure to obtain a spatial vector point set, and characterizing and calculating the consistent manufacturing deviation, assembly positioning deviation, and assembly deformation deviation; constructing an assembly deviation transfer graph, and mapping it into an assembly state space graph by using these three multi-source deviations; performing finite element analysis on the assembly accuracy of the deformable structure by using the assembly state space graph, and converting the obtained assembly deviation nephogram into an assembly deviation contour map; constructing an initial generative adversarial network model, and training and verifying it by using the data in the assembly deviation contour map to obtain a final generative adversarial network model; inputting the actual deviation image of the deformable structure into the final generative adversarial network model for assembly deviation prediction. The present application can improve the assembly accuracy and assembly efficiency of deformable structures.

The invention discloses a novel super-hydrophobic ice suppression performance detection device in the field of aviation engineering technology and materials science, including a base, on which an experimental module, a control module, a test module and a control module are installed; a groove is provided on the base, and the experimental module and the control module are symmetrically distributed on both sides of the groove, and the experimental module and the control module have the same structure, and the test module includes a support plate, and a support assembly is hinged on the support plate, and shearing assemblies are installed on both sides of the support assembly, and environmental simulation components and water discharge components are installed on the shearing assembly; the control module starts and stops the environmental simulation components and the water discharge components, and the control module can also collect and analyze the data of the shearing assembly in real time. The present invention can simulate a variety of climatic conditions, and super-hydrophobic material samples are accurately and repeatably tested, while realizing the automation and intelligence of the test process, to improve test efficiency and data accuracy.

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 present disclosure provides a method, an apparatus, a device, and a storage medium for simulating an arc ground effect, relating to the field of aerospace engineering. The method includes: setting a reference system that moves uniformly in a set direction; respectively determining a relative motion speed of a supersonic aircraft in the reference system, a horizontal motion speed and a sinking motion speed of a flat ground directly below the supersonic aircraft in the reference system; and performing dynamic simulation according to the relative motion speed of the supersonic aircraft in the reference system, the horizontal motion speed and the sinking motion speed of the flat ground in the reference system to determine the aerodynamic characteristics of the supersonic aircraft. The present disclosure can be applicable to simulating supersonic, large-curvature, and long-distance arc ground effects, and improving the calculation efficiency in the numerical simulation process.

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 an aviation floor beam anti-corrosion single-sided adhesive tape reinforced net layer mounting mode, and relates to the technical field of aviation engineering, and the aviation floor beam anti-corrosion single-sided adhesive tape reinforced net layer mounting mode comprises the following steps: providing a base material with a polytetrafluoroethylene (PTFE) film on one side, coating a layer of thin polyurethane same-series adhesive on the surface of the PTFE film, and then coating a layer of polyurethane same-series adhesive on the surface of the base material; then, glass fiber gridding cloth is attached to the PTFE film coated with the polyurethane glue, and a reinforced net layer structure is formed; the coating thickness of the polyurethane same-series adhesive is 0.01 mm to 0.1 mm, so that good adhesion between the gridding cloth and the PTFE film is ensured; before the glass fiber gridding cloth is attached, the surface of the PTFE film is cleaned, so that dirt and impurities on the surface are removed, and the bonding effect is improved; the warp and weft density and strength of the glass fiber mesh cloth are adjusted according to the bearing requirement of the aviation floor beam so as to ensure the structural stability and strength of the reinforced mesh layer, and the effects that wrinkles are avoided, the corrosion resistance is improved, the installation process is simplified, and the application range is wide are achieved through the mode.

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 relates to the technical field of aviation engineering equipment, in particular to a shut-off valve. The shut-off valve includes a housing, a bushing, a cap, an execution valve, a nozzle, and a spring. The lining is arranged in the containing cavity of the shell, the cap abuts against the lining, the execution valve is sleeved with the lining, a control cavity is defined by part of the outer circumferential wall of the execution valve and part of the inner circumferential wall of the lining, a spring cavity is defined by the inner side of the execution valve, part of the inner circumferential wall of the lining and the cap, and the control cavity is suitable for being communicated with the spring cavity through the electromagnetic valve. The spring is arranged in the spring cavity, one end of the spring abuts against the cap, the other end of the spring abuts against the end, away from the cap, of the bush, and the elastic force of the spring or the pressure difference between the control cavity and the spring cavity is suitable for driving the execution valve to move. Therefore, the defect of re-closing caused by instant pressure reduction after the execution valve is opened is overcome, and normal work of the shut-off valve is achieved.

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.