316results about "Aircraft components testing" patented technology

Apparatus and methods for ultrasonic inspection of elongated composite members in a single scan pass using pulse echo phased arrays (97, 98, 99) operating in a bubbler method. The system concept is fully automated by integrating an inspection NDI probe assembly (32, 34, 36) to a robot (100) and using the robot to move the inspection NDI probe assembly along the part (i.e., outside of an inspection tank); and by integrating tooling fixtures that move out of the way as the inspection NDI probe assembly travels along the length of the part during the inspection. In addition, the system allows for generally elongated composite members having lengthwise variation in shape, curvature and dimensions.

The application provides a kind of aviation tire dynamic simulation test machine dynamic load calibration device and method, it is related to aviation tire test equipment field, including: fixed frame, it is arranged between the flywheel and loading frame of test machine, and is fixed on the base of test machine;Mobile frame, it is arranged on fixed frame, and can be moved and locked along the loading direction of test machine;Loading plate, it is installed on mobile frame, for the load of aviation tire;Calibration load sensor, for measuring real-time load value in calibration process;Arc tooling, it is installed on fixed frame, and arc tooling is adapted and installed on the arc contact surface of flywheel, arc tooling is connected with one end of calibration load sensor;The other end of calibration load sensor is connected with loading plate.The application realizes load calibration on test machine, and the device can be used for calibrating the dynamic load of tire in particular, so that it can simulate the real working condition of tire, and improve the measurement accuracy of load of aviation tire dynamic simulation test machine.

The application relates to the technical field of aircraft air conditioning system detection, in particular to an airplane trimming air valve leakage testing device and method. The device comprises a support structure, an input shaft, an output shaft, a valve and a position adjusting structure. One side of the input shaft is provided with a first clamping surface, and the input shaft is provided with a first channel for inputting test gas. The first channel is communicated with the first clamping surface. One side of the output shaft is provided with a second clamping surface, and the second clamping surface and the first clamping surface are oppositely arranged for clamping the airplane trimming air valve to be tested. The output shaft is provided with a second channel, and the second channel is communicated with the second clamping surface. The valve is used for controlling the opening or closing of the second channel during the test. The position adjusting structure is used for adjusting and locking the distance between the input shaft and the output shaft. The device solves the problem that, in the existing maintenance operation, due to the lack of standardized tooling, the air valve gas path interface cannot be effectively connected to implement the quantitative indexes of the manufacturer's maintenance manual, and fills the gap of the leakage testing standardized detection tool.

The application provides a high-hypersonic wind tunnel piece water-cooled hinge torque balance measurement system and belongs to the technical field of wind tunnel test measurement.The application aims to solve the problem that the rudder surface is difficult to be externally cooled due to thin structure and the hinge torque is difficult to be measured in the prior art.The aircraft fuselage is connected with the aircraft wing and the piece water-cooled hinge torque balance in the application.The top and bottom of the piece water-cooled hinge torque balance are respectively provided with a balance upper cover plate and a balance lower cover plate.The upper end surface of the balance lower cover plate is flush with the aircraft fuselage.The balance free end, the balance measurement element and the fuselage mounting base of the piece water-cooled hinge torque balance are sequentially connected.The balance free end is connected with the measured rudder surface through an angle ear.The balance measurement element is pasted with a strain gauge, and the balance measurement element is internally provided with a water-cooled groove circuit.The application solves the problem that the measurement precision is greatly reduced due to high temperature in the high-hypersonic wind tunnel test, and the rudder surface is difficult to be externally cooled due to thin structure and the hinge torque is difficult to be measured.

The application discloses a hydrogen fuel cell power supply adaptability ground test method, system, equipment, medium and product of a UAV, relates to the UAV test technical field, obtains the ground test data of the to-be-tested UAV, carries out power demand analysis by using the ground test data, obtains a plurality of motor input total power, a plurality of motor output shaft power and weighted average propeller force efficiency, carries out power supply test on the preset hydrogen fuel cell by using each motor input total power, obtains a safety detection result and a plurality of variable load power response times, when the safety detection result is qualified, then carries out power supply adaptability detection according to each variable load power response time, each motor input total power, each motor output shaft power and weighted average propeller force efficiency, and obtains a power supply adaptability detection result. The application solves the problem that the existing hydrogen fuel cell lacks the safety buffer and power regulation capability of lithium batteries, which easily leads to the risk of stack overheating and reduces the reliability of the hydrogen fuel cell in the application of the UAV.

In some embodiments, techniques are provided for analyzing time series data to detect anomalies. In some embodiments, a machine learning model is used to process time series data. In some embodiments, the machine learning model is trained in an unsupervised manner on a large amount of prior time series data, allowing a highly accurate model to be created from new data. In some embodiments, the training of the machine learning model alternates between a fit optimization and a prune optimization to allow a large amount of training data to be processed that includes unlabeled anomaly records. Because a machine learning model is used, anomalies within complex systems can be detected, including but not limited to autonomous vehicles such as unmanned aerial vehicles. When an anomaly is detected, a command can be sent to the monitored system (e.g., autonomous vehicle) to address the anomaly.

A fibre-reinforced polymer composite component 2 formed from a polymer matrix material containing fibre reinforcement. The component 2 comprises a main portion 4 and at least one raised feature 6 extending from a surface plane s of said main portion 4. The at least one raised feature 6 is arranged to incur visually perceptible damage when the component 2 is subject to an impact with an energy above a predetermined impact energy threshold and to resist an impact with an energy below the predetermined impact energy threshold.

The present application belongs to the technical field of helicopter dynamics test, and relates to a method and system for measuring the distance between blades of upper and lower rotors based on measured bending moment. The measured bending moment data in flight is used, and the existing blade measurement system is used. Without additional measurement equipment, the relationship between bending moment and deformation is used to establish the calibration relationship between the test bending moment and the deformation of different profiles of the blade, so as to obtain the distance between the blades of the upper and lower rotors in the real-time flight state. The method can relatively real-time monitor the distance between the blades of the upper and lower rotors in the full flight spectrum state without additional measurement equipment, and ensure flight safety. The method has the advantages of simplifying the measurement method of the blade tip distance, shorter feedback time compared with the light ray sensor and radar wave measurement, and better real-time effect.

The application provides a fixed-wing aircraft airspeed calibration flight test method based on a standard machine. The method comprises the following steps: determining the weight of the fixed-wing aircraft as a first weight, setting the aircraft to a specified take-off configuration, and taking off according to a take-off procedure; the standard machine flies to a specified airspace according to a plan; at an initial specified flight height, a tow cone is released to a specified length, and the tow cone can obtain an accurate static pressure value in the current area without being affected by a spoiler; the flap and slot wing configuration of the fixed-wing aircraft is set to a first flap and slot wing configuration; the aircraft is trimmed at a given speed; the speed in the above step is set to other speeds within the speed envelope of the first flap and slot wing configuration, and the steps are repeated; the standard machine obtains airspeed errors under the first flap and slot wing configuration and different speeds; and the airspeed errors under the first flap and slot wing configuration and different speeds are fitted to obtain an airspeed correction curve under the first weight, the first flap and slot wing configuration. More accurate airspeed errors can be obtained.

A folding wingtip aircraft (100) includes a wing (104) that includes a primary portion (106) and a wingtip (102). The folding wingtip aircraft (100) includes a folding wingtip (102) system configured to rotate the wingtip (102) relative to the primary portion (106). The folding wingtip aircraft (100) also includes a flight control computer (202). The flight control computer (202) is configured to, in response to a first determination that a wingtip (102) position of the wingtip (102) is not in a folded position when a second determination indicates that a last completed command by the folding wingtip (102) system was a fold command, set error codes for the folding wingtip (102) system. The flight control computer (202) is also configured to change a setting to indicate that the folding wingtip aircraft (100) is grounded based on the error codes.

The application provides a kind of jet flight control wind tunnel rudder effect test system and test method, belong to wind tunnel test technical field.Test system includes aircraft model and be set in the engine and air induction system of aircraft model interior, fuel tank, jet actuation system, flight control computer and balance;Establish the body axis coordinate system corresponding to aircraft model, the origin of body axis coordinate system is model moment reference point, balance is located in the lower region of aircraft model, the axis of balance is parallel with the x axis of body axis coordinate system, and the heart of balance is located on the z axis of body axis coordinate system, engine is located in the tail region of aircraft model, the thrust line of engine overlaps with the x axis of body axis coordinate system, and air induction system is connected with jet actuation system.In the application, airborne air supply and control are realized, jet is taken as a part of aircraft system energy, and the accuracy of the result obtained by test on the influence of jet on the comprehensive performance of aircraft is improved.

The present invention discloses a system and method for performing congestion control at an apparatus. The apparatus may establish at least one first connection with a first node and establish a plurality of second connections with a second node. After the at least one first connection and the plurality of second connections are established, the apparatus may transmit data packets to the second node through the plurality of second connections according to an outbound traffic policy when receiving the data packets from the first node through the at least one first connection. If a first criterion is satisfied, the apparatus may establish a plurality of third connections with the second node and transmit data packets to the second node through the plurality of third connections when receiving data packets from the first node through the at least one first connection.

The application discloses a method for constructing a standard oxygen concentration environment in an aircraft fuel tank, and comprises the following steps: step 1, constructing an external environment of the aircraft fuel tank, wherein the aircraft fuel tank is a fuel tank simulated through experiments; step 2, constructing an internal environment of the aircraft fuel tank; step 3, constructing an oil and gas environment of the fuel tank; and step 4, constructing a standard oxygen concentration environment of the fuel tank. The technical scheme of the embodiment of the application solves the problem that calibration conditions are lacked for an aircraft fuel tank oxygen concentration testing device.

This invention discloses an integral skid structure suitable for a flight test platform, including a front skid assembly and a rear skid assembly. The front skid assembly is bolted to the front cabin of the test platform, and a locating pin prevents the bolts from being sheared during the platform's taxiing. The rear skid assembly is bolted to the rear cabin of the test platform, and a flanged design limits the platform's position. The front skid assembly includes a front skid, a front pad, a front locating pin, and a front bolt assembly; the rear skid assembly includes a rear skid, a rear pad, a rear locating pin, and a rear bolt assembly. This invention ensures the smooth implementation of flight test platform experiments and provides reliable technical support for flight platform testing.

A testing method and apparatus for unmanned aerial vehicle (UAV) production includes the following steps: S1, Operating condition construction: controlling the power system of the UAV under test to operate to a preset target speed range; S2, Multi-dimensional data synchronous acquisition: generating a synchronous trigger signal based on the real-time speed signal of the power system; using the synchronous trigger signal to control a strobe light source to illuminate the rotating parts, and acquiring a freeze-frame visual image of the rotating parts during the illumination; simultaneously, acquiring acoustic signals during the operation of the power system; S3, Feature extraction: performing contour analysis on the freeze-frame visual image to extract visual feature values ​​reflecting physical deformation; performing spectral analysis on the acoustic signal to extract acoustic signature feature values ​​reflecting the mechanical operating state; S4, Fusion judgment: inputting the visual feature values ​​and acoustic signature feature values ​​into a preset multimodal fusion model to calculate the comprehensive health index of the power system, and determining whether the product is qualified based on the comprehensive health index.

The application provides a system and method for detecting excess of aircraft components and parts closed area, belonging to the technical field of aircraft detection. The application fixes the product to be detected on a vibration platform, the lower surface of the vibration platform is fixed on the base through a spiral spring, the exciting motor on the base converts rotation into vibration of the vibration platform through a rotating wheel and a connecting rod, a sound signal acquisition device is installed on the product to be detected, and whether excess exists is judged by collecting the sound signal; the application creatively uses the vibration principle to provide a new method for checking excess in the closed area of the aircraft. The detection system uses digitalization and automation technology to realize digitalization and automation of the excess detection of the aircraft, completely changes the manual detection mode of "knocking and listening", significantly improves the accuracy and safety of the excess detection in the closed area of the aircraft parts, and reduces the safety risk and quality risk of the excess detection in the closed area of the aircraft structure.

The application discloses an aircraft fuel servo valve high-temperature test system and relates to the field of test systems.The system comprises a normal-temperature oil supply module, a high-pressure pump, a main heating and heat exchange module, a controller and an oil return shunt module.The normal-temperature oil supply module comprises an oil tank and a low-pressure booster pump, and the high-pressure pump is used for outputting high-pressure and high-temperature fuel;the main heating and heat exchange module is arranged between the high-pressure pump outlet and the fluid inlet of the measured valve and is used for heating the fuel to a preset test temperature;the oil return shunt module inlet is communicated with the oil return port of the measured valve, the first outlet is communicated with the high-pressure pump inlet, and the second outlet is communicated with the oil tank.The controller is connected with the main heating and heat exchange module and the oil return shunt module, is used for collecting two temperature signals, adjusting the heating power and the two outlet flow distribution, and maintaining the oil temperature after the high-temperature oil return and the normal-temperature fuel are mixed within the safety threshold of the high-pressure pump.The application solves the problem that the test system cannot continuously and stably operate due to the fact that the high-pressure pump cannot long-term bear extremely high-temperature fluid.

A kind of wind tunnel model launching test system and test method belong to model wind tunnel dynamic test technical field.Solve the problem that cannot give consideration to gravity launching and with initial velocity launching function in prior art.Technical points: tail brace bending knife is arranged on wind tunnel test section, tail brace transmission is arranged on tail brace bending knife, test model and tail brace transmission are connected by tail brace support rod, the top of launching model is fixedly connected with test model, binocular vision position measuring system is arranged outside wind tunnel test section;elastic energy-absorbing protective pad is arranged on the wall of wind tunnel test section, blocking protection net is arranged behind tail brace bending knife.The present application realizes the switching of force direction and the adjustment of launching speed by electromagnetic control, and the launching action is responsive and controllable, and the high-speed camera is used to record the time sequence curve of the position and posture of the launching model, to provide complete and accurate dynamic data support for wind tunnel test.

The present application belongs to the technical field of flexible body aircraft performance testing, and particularly relates to a flexible flat wing stiffness test loading device. The loading device comprises a wing test piece tooling table, a top suspension cable beam, a counterweight loading device and a data monitoring and storage system; the counterweight loading device can realize higher precision equivalent loading of rigid wings and flexible wings under different working conditions through a pump liquid controller and a loading infusion bag, can greatly improve the loading weight through the loading infusion bag, and realizes stepless loading of the wings, improving the flexibility of the loading scheme design; a soft loading force transmission mode is realized through a loading clamping frame, a sliding clamp, a clamp sliding groove and a loading plate, and the loading weight of the loading infusion bag can be more gently and accurately loaded to the surface of the wing; the 3D printing structure of the loading plate can improve the precision and reliability of the bearing test.

The application discloses a kind of unmanned aerial vehicle granular field uniformity detection device and detection method, belong to agricultural engineering and farmland detection technical field.The device includes detection device body, optical fiber sensing module, embedded PCB circuit board.The detection method is laid out and horizontally calibrated device, optical fiber sensing module collects the light signal variation of granular interruption detection beam, signal processing, effective counting obtains the number of particles per unit time, combined with single particle average mass conversion real-time spreading amount;Host computer collects multiple sampling point data to calculate the variation coefficient of spreading amount, compares preset threshold to determine uniformity grade, and accordingly guides unmanned aerial vehicle operation parameter adjustment.The application realizes real-time accurate detection of spreading uniformity, has strong environmental adaptability, is simple and easy to operate, effectively improves the intelligentization and high-efficiency level of field spreading monitoring, and is suitable for various agricultural unmanned aerial vehicle granular spreading operation scene.

Training an encoder is provided. The method comprises inputting a current state of a number of aircraft into a recurrent layer of a neural network, wherein the current state comprises a reduced state in which a value of a specified parameter is missing. An action applied to the aircraft is input into the recurrent layer concurrently with the current state. The recurrent layer learns a value for the parameter missing from current state, and the output of the recurrent layer is input into a number of fully connected hidden layers. The hidden layers, according to the current state, learned value, and current action, determine a residual output that comprises an incremental difference in the state of the aircraft resulting from the current action.

A phased approach to predicting whether a component currently installed on a vehicle should be removed is provided. In a first phase, a prediction model is created using historical data and sample values that are associated with a vehicle component previously removed from the vehicle. In a second phase, the output of the prediction model is used to predict whether the selected vehicle component currently installed on the vehicle should be removed. Based on the prediction, a graphical user interface (GUI) is generated for output to a user. The GUI visually indicates the vehicle component predicted to be removed and provides a recommendation for addressing the predicted removal of the component.

The application discloses a kind of unmanned plane freedom degree attitude test analysis system and method, wherein a kind of unmanned plane freedom degree attitude test analysis system, freedom degree attitude simulation module, based on quaternion dynamics model, simulates the attitude change of unmanned plane;Sensor dynamic calibration module, automatically identifies and calibrates gyroscope zero offset, accelerometer scale error, magnetic compass hard iron / soft iron interference;Multi-throttle vector control module, respectively control motor throttle, simulate extreme working condition;Flight control algorithm verification module, real-time injection attitude error, assess convergence time, overshoot, steady-state error;Data fusion and visualization module, real-time fusion data, support multi-dimensional visualization;Automated test script module, support Excel process driven test;Authority and data security module, carry out hierarchical authority management, support test data encryption, operation log audit, parameter tamper prevention.

A synchronous motion cracked control surface follow-up loading test device is disclosed. This device, driven by an electric motor, overcomes the shortcomings of existing cracked control surface test loading devices, such as high manufacturing and testing costs, complex electromechanical components, and the limitation to single-sided control surface loading tests. It also addresses the problems of high manufacturing and maintenance costs leading to difficulties in conducting tests, high dependence on pump stations due to hydraulic actuators, and the inability to meet actual testing requirements for loading effects. The device includes a test bench, a cracked control surface motion simulation module, and a follow-up loading module. The test bench includes a main body and casters. The cracked control surface motion simulation module includes a simulated control surface and a control surface rotation... The invention comprises a hinge, a simulated control surface base, a control surface drive mechanism, and a tilt sensor. The control surface drive mechanism includes a linear motor module, a slide table connector, a transmission hinge, a connecting rod, and a control surface transmission hinge. The servo loading module includes a servo drive mechanism, a hinge transmission mechanism, a spring loading mechanism, a servo frame, and a servo frame hinge. The servo drive mechanism includes a motor mounting base, a servo motor and its reducer, and a coupling. The hinge transmission mechanism includes a transmission hinge base, bearings, a transmission shaft, a key, and a frame connector. The spring loading mechanism includes a frame mounting base, adjusting bolts, a movable base, a loading spring, a force sensor, and a control surface base. This invention is used for servo loading tests.

An arrayed high-energy laser counter-UAV (unmanned aerial vehicle) test method and apparatus are provided, relating to the technical field of UAV tests. The method includes the following steps: preparing a test apparatus and a UAV; conducting a laser counter-UAV test; observing an ablation process and analyzing an ablated morphology of the UAV to obtain a radius and ablation depth of a tested ablation area of the UAV; performing finite element simulation on a test process to obtain a radius and ablation depth of a simulated ablation area of the UAV; and comparing the radii and ablation depths of the simulated and tested ablation areas of the UAV, adjusting finite element simulation parameters to make a simulation result coincide with a test result, and deriving an ablation depth-time variation curve of a key point in the simulated ablation area.

The application discloses a kind of lightweight explainable aviation equipment abnormal state perception method and system, belong to aircraft health management and deep learning technical field.The application constructs 1+1 branch's LiteInception light weight feature extraction network, designs and aligns two-stage cascaded diagnosis process with aviation maintenance process: first stage adopts LiteInception and hybrid architecture of Transformer to execute fault detection binary classification, and priority is guaranteed recall rate;Second stage adopts pure LiteInception architecture to execute fault identification multi-classification to abnormal sample, and priority is guaranteed accuracy and explainability.The application significantly compresses model size, adapts edge device real-time inference needs, can output the diagnosis evidence chain that can be traced to specific sensor, specific time period, considers diagnosis efficiency, precision and compliance, and provides complete solution for general aircraft fault diagnosis.

The application belongs to force application distance measurement technology and relates to a force application feedback type distance measurement device and method. The device is applied to upper and lower automatic tilting devices of a coaxial helicopter and comprises a positioning and clamping assembly (1), a force application assembly (2), a base (3) and a dial gauge (5). The positioning and clamping assembly (1) is fixed to the upper surface of the base (3) and is used for positioning and clamping inner and outer rings of the upper and lower automatic tilting device assemblies. The force application assembly (2) is slidably connected to the upper surface of the base (3) and is used for digitally applying force to the outer ring yoke of the automatic tilting device assembly in the vertical direction. The dial gauge (5) is used for measuring the distance between the inner and outer rings of the automatic tilting device assembly after the force application requirement is reached.

This invention provides a control channel multiplexing and docking method, apparatus, computer equipment, and medium based on a high-speed network, comprising: configuring a logical channel connected to the control system for each test site and dynamically setting IP routing mapping relationships; converting analog feedback signals output by sensors at the first test site into digital feedback signals, encapsulating them, and transmitting them to the test control room via a high-speed real-time network; receiving a first data packet, decapsulating it, converting it into an analog feedback signal, and inputting it to the control system; converting analog control commands issued by the control system into digital control commands, encapsulating them, and transmitting them to the first test site via a high-speed real-time network; decapsulating a second data packet according to its destination IP address, converting it into an analog control command, and outputting it to the actuator; updating the IP routing mapping relationship and switching the logical channel connection when switching test sites. This enables the hardware resources of a control system to be reused.