2832results about "Aircraft components testing" patented technology

The embodiment of the invention provides a multi-configuration relay-oriented aircraft electrical box parameter measurement system and method, and belongs to the field of aircraft electrical box test and diagnosis. According to the hardware level, flexible access of a multi-configuration relay electrical box is achieved based on a universal adapter interface, core parameter measurement is completed in combination with an integrated measurement and control circuit, and synchronous acquisition and processing are conducted on multiple paths of signals through a signal acquisition and processing unit; according to the software level, the pull-in / release time, the rebound time and the voltage parameters are accurately calculated through a parameter calculation module on the basis of waveform feature point recognition and dynamic excitation-response theories, and the functions of fault diagnosis, safety protection and data storage are integrated. According to the method, the crossing from on-off judgment to parameterized precise evaluation is realized, the detection efficiency, the measurement precision and the equipment reliability evaluation capability are remarkably improved, and an integrated solution is provided for external field maintenance and state prediction of the aircraft relay.

The invention discloses an output voltage detection method and system based on an aviation power supply, and relates to the technical field of aviation power supply voltage detection. A differential voltage sensor is connected to the output end of an aviation power supply, a high-frequency suppression capacitor is connected in parallel, the sampling rate of 20-50 times of the working frequency is set according to the type of the power supply, and the signals are collected and converted into analog voltage signals; the analog signal is subjected to two-stage filtering, is amplified to 0-5V through an instrument amplifier, and is converted into a digital voltage sequence through a 16-bit ADC (Analog to Digital Converter); calculating AC power supply effective value, peak value and other parameters, DC power supply average value, ripple coefficient and other parameters; comparing the characteristic parameters with a preset threshold value, determining that the characteristic parameters exceed the threshold value and recording related information; and a detection result is transmitted to a monitoring terminal through an RS485, data is locally stored, and a switching value is output to an alarm device when the data is abnormal. The method adapts to aviation power supply load fluctuation and environment temperature change, detection precision is high, data transmission and storage are safe, abnormity is accurately identified and early warning is carried out in advance, and aviation power supply operation and maintenance reliability and efficiency are improved.

The invention discloses an aircraft landing gear shimmy test bed and a test method, the aircraft landing gear shimmy test bed comprises a main body rack, a hanging basket system, a large flywheel, a flywheel driving system, an excitation device and a follow-up displacement measurement system.The aircraft landing gear shimmy test bed can simulate the shimmy of an aircraft after undercarriages of different design configurations are subjected to interference excitation force or interference torque in the process of takeoff, taxiing, landing and the like of the aircraft; the control effect of the shimmy damping system on the stability of the undercarriage has short excitation time so as to simulate real working conditions, the shimmy stability performance of the undercarriage can be accurately detected, excitation is continuously carried out under the condition that the test state is continuously changed, and the working efficiency is greatly improved. The follow-up displacement measuring system designed by the invention can enable the exciting force and the center of the axle to be always kept on the same horizontal plane, can apply torque control and torque excitation control, is suitable for undercarriage shimmy tests with different axle heights, different tire diameters and different excitation modes, and has certain universality.

Disclosed in the present application are an aircraft final assembly and integration testing platform and testing method. On the basis of a rational platform architecture, structured process data preparation, model-based operation guideline document preparation, and dispatch assistance decision-making are realized by means of an application layer, such that uniform data sources in the field of final assembly can be achieved. Data uploading and issuing between an industrial control network and an internal network, and data interaction control between layers are implemented by means of the mutually collaborative operation of the application layer, a storage layer, an edge layer and an execution layer, thereby improving the information acquisition efficiency and the operating efficiency. The edge layer and the execution layer collaboratively operate, so as to realize the interactive collaboration between the assembly guidance work for operators and automatic testing, thereby improving the efficiency of data information acquisition, operation instruction and fault handling during a final assembly and testing process of an aircraft.

The invention provides an inclined variable-angle aircraft interstage separation ground test device. The inclined variable-angle aircraft interstage separation ground test device comprises an aircraft separation device, an angle adjusting device, a temperature control and acquisition device, a vibration control and impact acquisition device, an ignition control device and a device base. The angle adjusting device comprises a device side plate, an arc groove, an angle ruler, a separation mechanism mounting plate, a supporting hinge, a supporting rod, an adjusting sliding rail, an adjusting base, a locking device, a device base gradienter, a rotating bearing and an auxiliary fixing block. The temperature control device comprises a quartz lamp array heating cage, an outer wall temperature control thermocouple and an inner wall temperature acquisition thermocouple, and can simulate a pneumatic heating environment at a separation moment, so that the thermal deformation matching capability of the separation mechanism in a high-temperature environment and the operation reliability of the separation mechanism are obtained. The vibration control and impact acquisition device is composed of a vibration table, a vibration control sensor and an impact sensor.

The invention discloses an unmanned aerial vehicle load engine state identification method based on deep learning, and the method comprises the following steps: S1, collecting the multi-mode operation data of an unmanned aerial vehicle load engine, and completing the standardization preprocessing; s2, inputting the spatial data into a spatial pyramid pooling network, and extracting a multi-scale spatial feature matrix; s3, inputting the spatial features into a masking auto-encoder, performing feature masking, encoding and reconstruction, and generating reconstruction features and error information; s4, fusing the original features, the reconstructed features and the non-spatial data, inputting a discrimination network, and outputting a health label and a risk score; and S5, according to an identification result, a flight control system or a maintenance platform is linked to realize alarm and scheduling. According to the method, intelligent identification and abnormal risk early warning of the state of the load engine of the unmanned aerial vehicle are realized by fusing multi-scale spatial feature extraction and a covering self-encoding reconstruction mechanism.

A system and method for gust load reduction testing of an elastic full-aircraft model in a large low-speed wind tunnel, which belong to the technical field of wind tunnel test apparatuses. The system in the present invention comprises a gust generator and a supporting device that are mounted in a wind tunnel. The supporting device comprises a model sliding guide rail, mounting bases and a trolley, wherein the mounting bases are fixedly mounted at two ends of the model sliding guide rail, the trolley is mounted on the model sliding guide rail, an aircraft model is mounted on the trolley, two parallel sliding rails are mounted on each mounting base, a slider is mounted on each of the two sliding rails, the two sliders are connected by means of a slider connecting base, a lifting ring assembly plate is mounted on the slider connecting base, a plurality of first lifting rings are welded on the bottom face of the lifting ring assembly plate, second lifting rings are welded on the mounting base, springs are hooked between the first lifting rings and the second lifting rings, and a smooth shaft sleeve and a smooth shaft are mounted in the mounting base, one end of the smooth shaft being fixedly mounted in the slider connecting base by means of a bolt, and a limiting block being fixedly mounted at the other end of the smooth shaft. The present invention has a compact structure, involves simple installation and is low cost.

The invention discloses a deicing system real-time state evaluation and fault prediction method and system, and the method comprises the steps: simulating the fault state of a deicing system through a ground state monitoring test of the deicing system, and obtaining sensor data in different fault states; constructing a data set by using sensor data in different fault states to train a state evaluation model and a fault prediction model; the method comprises the following steps: installing a sensor on an aircraft deployed with a deicing system to collect working data of the deicing system; and inputting data acquired by the sensor into the trained state evaluation model and the fault prediction model, evaluating the real-time state of the deicing system, and predicting possible faults. According to the invention, the current working state of the anti-icing and deicing system can be evaluated according to the current state monitoring data, and possible faults of the system can be predicted.

The invention relates to an abnormal unmanned aerial vehicle detection method fusing internal anomaly detection and external environment adaptation, which adopts a three-layer collaborative architecture design and comprises an external environment perception module, an internal anomaly detection module and a virtual-real fusion digital twinning module. And the external environment sensing module is responsible for external environment interference identification, regularly acquiring data of each unmanned aerial vehicle in the unmanned aerial vehicle group, and realizing stability discrimination. Environmental stability judgment and task switching are regarded as conditions of entering a buffer state, and are used for internal anomaly detection of the unmanned aerial vehicle group. In an internal anomaly detection module, real-time operation data of multiple unmanned aerial vehicles are analyzed based on a data flow driving mechanism, and abnormal individuals are identified. And the virtual-real fusion digital twinning updating module is responsible for constructing a high-fidelity digital twinning virtual space and dynamically updating internal and external constraint information to a digital twinning platform to support the next decision. The collaborative fusion of the three modules breaks through the limitation of a single constraint view angle, and realizes the collaborative constraint of internal state monitoring and external environment perception.

The invention provides a helicopter cable fault detection system and method, and belongs to the technical field of helicopter electrical system detection. The system comprises a positioning and matching device and a multi-modal signal analysis device, the positioning and matching device determines the relative position with a to-be-detected aviation plug through a dual-spectrum camera, generates a contact coordinate matrix and a contact and probe mapping table, injects a test signal to identify the contact material so as to obtain a topology result, and detects the contact pressure; and the multi-modal signal analysis device is used for matching parameters based on the topology result and the mapping table, setting a detection signal, positioning the contact, applying a signal in combination with pressure, acquiring a response signal, establishing a mapping table, and analyzing the response signal to generate a contact and fault result mapping table. According to the invention, the whole process automation from aviation plug positioning, signal detection to fault analysis is realized, the system can complete the core links of contact coordinate generation, signal mapping, fault determination and the like without manual intervention, and the detection efficiency is greatly improved.

The invention relates to the technical field of multi-rotor aircrafts, in particular to a blade protection cover detection method, device and equipment and a medium. The method comprises the steps that if it is determined that a multi-rotor aircraft is in a stable hovering state based on the detected moving speed, the detected accelerated speed and the detected angular speed; if yes, acquiring the detected climbing speed and the blade rotating speed of the multi-rotor aircraft; the nominal weight of the multi-rotor aircraft is obtained, and the theoretical climbing acceleration is obtained according to the blade rotating speed and the nominal weight; obtaining a theoretical climbing speed and a disturbance climbing acceleration according to the detected climbing speed and the theoretical climbing acceleration; and determining whether the model of the multi-rotor aircraft is matched with the model of a blade protection cover mounted on the multi-rotor aircraft according to the theoretical climbing speed and the disturbance climbing acceleration. According to the scheme, whether the model of the multi-rotor aircraft is matched with the model of the blade protection cover installed on the multi-rotor aircraft or not can be determined, an operator can be conveniently reminded to replace the unmatched blade protection cover in time, and the user experience is improved.

Disclosed here is a landing gear support fixture for an aircraft under test having a landing gear. The disclosed landing gear support fixture includes: an axle attachment assembly that is compatibly shaped, sized, and configured to mate with and attach to at least one axle of the landing gear of the aircraft under test; a support platform to support the axle attachment assembly, wherein the support platform is in a fixed position relative to a stationary reference foundation; and a force sensor couplable between the axle attachment assembly and the support platform to provide sensor output associated with load imparted by the landing gear of the aircraft under test.

The invention belongs to the technical field of aircraft strength design, and particularly relates to an aircraft structure strain field reconstruction method and device based on a deep learning neural network, and the method comprises the steps: S1, carrying out the dimension reduction mapping of irregularly distributed three-dimensional grid nodes of a finite element model, and obtaining regular two-dimensional grid nodes; s2, obtaining calculation results of each two-dimensional grid node under a plurality of working conditions based on finite element simulation, and constructing a training set; s3, training the U-NET deep convolutional network through the training set to obtain a strain field proxy model; and S4, based on the strain field agent model, processing an input image with a small number of measurement point position measurement results to obtain an output image containing an aircraft structure strain field. According to the invention, the strain field of the whole aircraft can be reconstructed through displacement data of a small number of points measured by the static test of the whole aircraft.

The invention discloses a complex wind field simulation method suitable for a small and medium-sized unmanned aerial vehicle wind resistance test, and relates to the technical field of intelligent wind control, and the method comprises the steps: based on a plasma exciter control instruction set, generating a directional micro vortex group in a wind tunnel through a plasma exciter array, meanwhile, the spatial distribution of the directional micro vortex group is regulated and controlled by adopting an ultrasonic interference method, and a three-dimensional turbulent flow field is generated; an unmanned aerial vehicle anti-risk test is executed, anti-wind performance data is collected and combined with PIV to verify the stability of the nest ring structure, and dynamic adjustment is carried out through a fuzzy PID control algorithm; and constructing a judgment matrix based on the adjusted wind resistance test data, calculating a wind resistance performance index of the unmanned aerial vehicle by adopting a dynamic time warping algorithm, and generating a performance evaluation report. According to the invention, through vortex ring stability feature extraction and fuzzy PID dynamic adjustment based on the deep space-time convolutional network, closed-loop stability control in a complex wind field environment is realized.

The utility model discloses an aircraft sliding door pressure relief test device, which comprises a clamping mechanism and an actuator, the clamping mechanism comprises a vertically arranged rectangular frame, a sliding door is vertically arranged in the rectangular frame, and an upper frame and a lower frame of the rectangular frame are respectively fixed on the upper side and the lower side of a framework of the sliding door. The left frame and the right frame of the rectangular frame do not make contact with the left side and the right side of the framework of the sliding door, the transmission end of the actuator acts on the pressure relief opening cover on one side on the plane in the heading direction of the sliding door to form a loading point, and the actuator loads and applies force to test whether the pressure relief opening cover is separated from the pressure relief lock at the opening value and is normally opened relative to the framework. On the basis of analysis of structural forms and pressure relief loads of the pressure relief opening cover and the pressure relief lock of the aircraft sliding door, the open test device is provided, a confined space pressing test is not needed, the actuator directly acts on the pressure relief opening cover of the sliding door to load and apply force, the test device can be repeatedly used, the test implementation process is easy to operate, and the test efficiency is improved. Test results are accurate, and test cost can be reduced.

An aircraft-based fuel quality monitoring and alerting system is disclosed herein. The aircraft-based fuel quality monitoring and alerting system can include a first sensor configured to output a signal associated with contamination of fuel entering an aircraft through a first fuel inlet. The system can further include a processor configured to determine a first contamination level of the fuel based on the signal and an alerting system configured to deliver an alert if the first contamination level is outside a predetermined range.

The invention relates to the technical field of vibration tests, in particular to a ground resonance test suspension device and a test method. The ground resonance test suspension device is provided with a bearing frame body, a hoisting mechanism, a bearing mechanism and a suspension mechanism, the bearing frame body is composed of a bearing supporting column, a cross beam and a longitudinal beam, and the position of the cross beam can be adjusted; a bracket of the bearing mechanism is provided with a bottom beam, an auxiliary supporting beam and a supporting beam, and a flexible air cushion is arranged on the supporting beam. A flexible supporting unit of the suspension mechanism is provided with a main spring rope, an upper hanging scaffold, a lower hanging scaffold and an auxiliary spring rope, and an auxiliary spring rope connecting unit can control the state of the flexible supporting unit. The invention further provides a ground resonance test method which comprises the steps of suspension pretreatment, aircraft suspension and test, and the equivalent stiffness of the suspension mechanism can be adjusted. The aircraft suspension system can adapt to different aircraft sizes and weight distribution, and the equivalent stiffness of the whole suspension system can be adjusted to meet the test support frequency requirement.

The invention belongs to the field of wind tunnel experiments in a hypersonic flow field environment, and discloses a multi-degree-of-freedom high-frequency motion simulation device for an aircraft dynamic stability test and a calculation method for realizing an actuation angle. According to the technical method, calculation is carried out on the basis of an electromechanical coupling equation, determined by electrical parameters and mechanical parameters, of a piezoelectric material. Modeling can be carried out according to a piezoelectric equation to obtain a calculation method between the voltage applied by the piezoelectric actuator and the pitching and yawing actuation angles of the device, so that the excitation effect of the device in the ground test process can be accurately controlled and predicted. The method is not only suitable for pitching and yawing two-degree-of-freedom actuation calculation, but also can be used for calculating the actuation angle of a three-axis degree-of-freedom experimental device added with rolling drive through coordinate transformation under the condition of adding parameters such as rolling angular velocity and the like in the face of the three-axis degree-of-freedom experimental device added with rolling drive. At present, according to a calculation method based on the device, calculation is carried out only by considering the degree of freedom of the pitching axis and the yawing axis.

The present invention relates to equipment that can realistically reproduce various types of airflow occurring in environments such as an urban area with tall buildings or regions with mountains and valleys, thereby enabling verification of the flight safety of a multicopter-type drone. The present invention comprises: a housing that provides a tunnel-shaped test space in which a plurality of blowers, together with controllers, are installed along the longitudinal direction of both sidewalls of the interior, the ceiling, and the floor, to generate airflows such as crosswinds, updrafts, downdrafts, and vortices; and an integrated control unit that selectively or simultaneously controls some or all of the plurality of controllers installed in the housing so as to form airflows in pattern corresponding to user input or settings.

The invention discloses a test device and a test method for acquiring water-touching slip motion data. The test device comprises a wind tunnel test section, a water tunnel test section, a wave absorbing plate and a free supporting mechanism. The wind tunnel test section and the water tunnel test section are horizontally arranged and communicated, the wind tunnel test section is located above the water tunnel test section, and a wave absorbing plate is arranged between the wind tunnel test section and the water tunnel test section. The wind tunnel test section and the water tunnel test section jointly form a two-phase flow field of air and water. The free supporting mechanism is installed on a wind tunnel test section and is connected with the test model. The mode that the model fixes water-gas medium movement is adopted, waves of the water-gas medium are weakened through the perforated wave absorbing plate, simulation of water-gas two-phase flow of the aircraft moving near the water surface is achieved, high-precision and high-repeatability aircraft water touching slippage movement data can be obtained, and the high engineering application value is achieved.

The invention discloses a rotorcraft flutter model design and test method for a low-speed wind tunnel test, and belongs to the technical field of low-speed wind tunnel tests. The method solves the difficult problem of how to separate key nacelle modes in rotation flutter of the tilt rotorcraft. A fuselage is connected with the side wall of a wind tunnel, wings are of a double-girder straight wing structure, one end of each wing is connected with the fuselage, the other end of each wing is connected with a nacelle mechanism, a rotating paddle is connected with the nacelle mechanism, and a motor, a yawing mechanism and a pitching mechanism are arranged in the nacelle mechanism. The yawing mechanism and the pitching mechanism are connected in a matched mode to achieve decoupling in the yawing direction and the pitching direction of the model, the motor is installed on the pitching mechanism, and the execution end of the motor is connected with a main shaft of the rotating paddle. According to the rotorcraft flutter model design and test method for the low-speed wind tunnel test, decoupling of modes in the pitching direction and the yawing direction of the model can be achieved, the flutter characteristic of the model is shown as coupling of the pitching mode and the yawing mode of a nacelle, and the test requirement of the model wind tunnel test is met.

The invention belongs to the technical field of rotor craft flutter tests, and discloses a rotor craft spiral flutter ground test system and test method. The rotor craft spiral flutter ground test system comprises a hardware part and a software part, and the hardware part comprises a rotor test piece, a support system, a force loading device, a vibration signal sensor, a force sensor, a vibration signal acquisition card, an industrial control computer, a force signal output card, a power amplifier and a force sensor signal acquisition card. The software part of the rotor craft spiral flutter ground test system comprises a rotor force and torque calculation program and a multi-input multi-output force control program which are operated on an industrial control computer. The method can retain all dynamic characteristics of the aircraft structure, is not limited by the size of a wind tunnel test section, does not need to perform scaling similarity on the aircraft structure, is low in cost and short in period, and can quickly deploy and complete a ground test.

The invention belongs to the field of airplane strength fatigue life prediction, and particularly relates to an equivalent stress-based airplane typical structure multi-axial fatigue life prediction method. The method comprises the steps of determining a typical aircraft high-lock connection structure test piece; determining a fatigue load spectrum; constructing a finite element model of the typical high-lock connection structure of the aircraft; determining a dangerous part of the airplane typical high-lock connection structure test piece in finite element analysis; determining the maximum principal stress at the dangerous part of the typical airplane high-lock connection structure test piece in finite element analysis; determining the fatigue life of the test piece; fitting S-N curve data under different stress ratios according to the fatigue life of the test piece and the load spectrum stress ratio, and determining a corrected S-N curve; determining an equivalent stress-life equation of the typical high-lock connection structure of the aircraft, and substituting the maximum principal stress and fitting parameters of the corrected S-N curve into the equivalent stress-life equation to calculate and obtain a fatigue prediction life; and correcting the fatigue prediction life according to a test result.

The invention relates to the technical field of processing, testing or inspection of aircraft parts, and provides a method for detecting surface defects of an aluminum alloy plate with an epoxy coating. The method comprises the following steps: applying an epoxy coating doped with a zinc yellow pigment to an anodic oxide film to construct a coating test piece with a color difference indication function, and placing the coating test piece in a coastal atmospheric environment for exposure sampling. In the detection process, the gloss loss rate, the chromatic aberration, the contact angle, the defect area proportion and the corrosion current density are collected, a dynamic correction coefficient is calculated according to the hydrolysis degree of the zinc yellow pigment, all parameter thresholds are adjusted in a self-adaptive mode, and a threshold set is constructed. And in combination with scanning electron microscope image segmentation and Tafel extrapolation analysis, multi-source indexes are fused to judge whether the coating is invalid. According to the method, the hydrolysis behavior of the zinc yellow pigment can be tracked in real time, the judgment standards of optical, morphology and electrochemical indexes are automatically corrected, and the accuracy and reliability of failure detection of the epoxy coating aluminum alloy plate under the coastal service condition are improved.

The invention belongs to the field of electromechanical comprehensive thermal management, and particularly relates to a real-time analogue simulation method and system for electromechanical comprehensive thermal management of an airplane. The method comprises the following steps: constructing a comprehensive thermal management environment model of an aircraft operation environment; constructing a comprehensive thermal management system model; performing real-time coupling on the comprehensive thermal management environment model and the comprehensive thermal management system model, and simulating a real-time comprehensive thermal management process of the airplane based on a comprehensive thermal management strategy to obtain a real-time simulation result; and performing optimization iteration on the comprehensive thermal management system model according to the real-time simulation result. According to the method, through accurate simulation of the electromechanical comprehensive thermal management process of the airplane, powerful support can be provided for performance optimization, safety improvement and cost control of the airplane.

The invention discloses a distributed jet active flow control-based rotor wind tunnel test method, which belongs to the technical field of aviation aerodynamic wind tunnel tests, and comprises the steps of rotor test bed dynamic characteristic test, manipulation matrix calibration, preposed preparation of rotor model dynamic balance and common taper adjustment, and dynamic balance and common taper adjustment. And then hovering and forward flight wind tunnel tests of distributed jet active flow control are carried out in sequence, test data are collected and processed, and the aerodynamic performance of the rotor wing under the condition that jet control exists or not is compared and analyzed. According to the invention, the industrial blank of the lack of a distributed jet active flow control rotor wind tunnel test method is filled, and key technical support is provided for theoretical research, technical optimization and engineering application of the rotor distributed jet active flow control method.

The invention discloses a typical mission profile-based water surface aircraft water surface load spectrum compilation method. The method comprises the following steps of: (1) determining a typical mission profile and a take-off and landing sea area of a water surface aircraft; (2) determining a wave parameter distribution condition of the take-off and landing sea area, and further determining a wave parameter distribution diagram; (3) calculating to obtain a gravity center overload time domain curve of the water surface aircraft under different wave parameters of each mission profile; carrying out filtering and statistical analysis on the gravity center overload time domain curve of the water surface aircraft to obtain a gravity center overload exceeding number of each task profile in each state; (4) weighting according to the proportion time of each working condition of the typical mission profile to obtain a gravity center overload override curve of each stage of the water surface aircraft; and (5) compiling a water surface load spectrum table. Technical reference can be provided for evaluation of the take-off and landing fatigue life of the water surface aircraft.

The invention belongs to the technical field of unmanned aerial vehicle testing, and discloses an air-ground unmanned aerial vehicle wind resistance performance testing platform based on a physical information neural network. The air-ground unmanned aerial vehicle wind resistance performance test platform is composed of a wind field simulation system, a physical information neural network control system and an unmanned platform indoor positioning system. The wind field generated by the wind field simulation system is accurately controlled based on the physical information neural network control system, the complex wind field required by the air-ground unmanned aerial vehicle in the wind resistance test process of the air and ground subsystems is simulated, and the uniformity of each index of the wind field is ensured.

The aircraft underwater launching simulation experiment system comprises a water tank module, a launching module, a bullet changing module, a film changing module, a recycling module, a pneumatic knife gate valve module and a control cabinet. The launching module is arranged below the water tank module, the bullet changing module is arranged between the launching module and the water tank module, the film changing module is arranged in the water tank module and located above a bullet inlet hole of the water tank module, the recycling module is arranged on the upper portion in the water tank module, and the pneumatic knife gate valve module is arranged at the bottom of the water tank module. Through cooperation of the water tank module, the launching module, the bomb replacing module, the membrane replacing module, the recycling module and the pneumatic knife gate valve module, continuous launching of experimental bombs in an aircraft underwater launching simulation experiment is achieved, through cooperation of the control cabinet, tedious operation of bomb replacing is relieved, and the test efficiency is improved; the influence of water in an incidence pipeline on the launching state of the projectile body caused by the fact that the diaphragm is placed on the lower portion of the water tank is reduced.