44 results about "Experimental aircraft" patented technology

The invention belongs to the field of aviation structure strength tests, and particularly relates to a testing machine pitching and rolling angle adjusting method and system for a full-aircraft drop test, and the method comprises the steps: adjusting the lengths of a front lifting rope l1 and a rear lifting rope l2 through a first servo motor; the lengths of the adjusted front lifting rope l1 and the adjusted rear lifting rope l2 are calculated through the cosine law according to the geometrical relationship; the lengths of the left lifting rope 13 and the right lifting rope 14 are adjusted through a second servo motor; and the lengths of the adjusted left lifting rope 13 and the right lifting rope 14 can be calculated through the cosine law according to the geometrical relationship. The pitching and rolling postures can be automatically adjusted, and the problems that an existing method is low in efficiency and poor in precision are solved. According to test experience, according to an existing posture adjusting method, replacement needs about half a day to one day every time, the number of times of putting every day does not exceed ten at most, and manual operation is needed on wings or a fuselage every time. And after automatic adjustment is achieved, the target can be achieved in several minutes through adjustment of the motor, efficiency is high, and safety is higher.

The invention relates to a rotor six-force-element loading method and device for a static test of a tilt-rotor aircraft. The testing device comprises a rack, a plurality of fixed pulleys installed on the rack, a motor fake part fixedly connected with a to-be-tested machine arm, a connecting rod system fixedly connected with the rotating end of the motor fake part, and a rope system, wherein one end of the rope system is connected with the connecting rod system, and the other end of the rope system penetrates through the fixed pulleys in a winding mode and then is used for loading a test load. The rack is provided with a test position used for placing a to-be-tested machine arm. A to-be-tested vehicle arm is installed on the motor fake part through a bolt. The weight is placed at the tail end of the rope, the function of loading a specified load is achieved, and then the purpose of checking the strength of the arm is achieved. The tilt-rotor aircraft arm rotor six-force-element static test device realizes tilt-rotor aircraft arm rotor six-force-element static test, is convenient and simple to operate, and is flexible and adjustable in loaded load.

The invention provides a high-speed aviation bearing testing machine for simulating a high-temperature environment, and belongs to the technical field of high-temperature bearing testing. The external annular heating cavity and the internal high-temperature heating cavity are symmetrically arranged, high-temperature gas is guided in in two directions, and cooperative temperature control of the double heating cavities is achieved; a physical isolation layer is formed by the inner side spacing cover and the end face heat insulation cavity, so that lubricating oil is prevented from being in contact with high-temperature gas; the internal and external high-temperature heating cavities are physically isolated from the bearing tool through the heat insulation layer, so that heat conduction is prevented from interfering with measured data. The device comprises a machine base, a base and an upper cover, the base and the upper cover are fixedly installed on the machine base, the upper cover and the base jointly form a closed cavity wrapping a test tool, the device further comprises a connecting shaft, an external annular heating cavity, a tested bearing assembly and an internal high-temperature heating cavity, and the connecting shaft is inserted into the closed cavity; an external annular heating cavity is formed in the right side area of the tail end of the connecting shaft, and an internal high-temperature heating cavity is formed in the axis area of the connecting shaft.

The invention discloses a non-contact measurement system and method for full-field deformation of a fuselage in an airplane crash test, and belongs to the technical field of airplane structure strength testing, and the method comprises the steps: making a high-quality speckle pattern on the surface of the fuselage, and laying cooperation mark points; a binocular camera flexible self-calibration method is used for completing calibration of the stereo measurement system; unification of global coordinates of multiple systems is assisted through unmanned aerial vehicle photogrammetry; a synchronous controller is adopted to trigger multiple sets of systems to synchronously collect images in the falling and collision process; obtaining the motion trail and speed of the key measuring point based on a mark point positioning tracking and stereoscopic vision principle; acquiring a full-field three-dimensional displacement field and a strain field based on a digital image correlation method guided by a control point and a stereoscopic vision principle; and deformation visualization representation is realized through image splicing and data fusion. According to the invention, non-contact, high-precision and full-field dynamic measurement of the three-dimensional deformation field in the falling and collision process of the whole large aircraft is realized, and technical support is provided for analysis and evaluation of falling adaptability of the aircraft structure and verification of anti-falling and anti-collision design.

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.

1. Name of the product in this design: Electronic Testing Machine (Universal). 2. Application of this design: This design is used to test the mechanical properties of non-metallic and metallic materials, and is widely used in metallurgy, chemical industry, construction, transportation, energy, aviation, aerospace, shipbuilding, weaponry and other fields. 3. The key design feature of this product is its shape. 4. The image or photograph that best illustrates the design's key points: a 3D model.

The invention discloses a testing device and a testing method for a cut-off end of a stringer of a wallboard structure of a high-load aircraft. The testing device and the testing method are suitable for testing the mechanical property of a wallboard test piece comprising a skin and at least one stringer cut-off end. According to the device, lateral supporting assemblies are arranged on the two sides of a test piece, lateral constraint is provided for the wallboard test piece, axial relative displacement is allowed, and the single stringer wallboard test piece is used for equivalently simulating simple support boundary conditions on the two sides of an actual stiffened wallboard; the lateral supporting assembly comprises an L-shaped base, a connecting flat plate and a sliding supporting piece, the L-shaped base is connected with the testing machine through bolts, and the sliding supporting piece abuts against the side edge of the wallboard testing piece to form lateral supporting; according to the optimal position of the sliding supporting piece, a single stringer wall plate test piece obtains the maximum bearing capacity and the damage form which are consistent with those of a plurality of stringer wall plate test pieces under the action of a tensile load in a lateral supporting mode.

This application specifically relates to a design and evaluation method for mechanical failure simulation schemes in aircraft lift enhancement device failure testing, comprising: Step 1, designing different mechanical failure simulation schemes to obtain mechanical failure simulation parts; Step 2, scoring different mechanical failure simulation parts and their schemes for rejection criteria, and rejecting mechanical failure simulation parts and their schemes with excessively low rejection criteria scores; Step 3, scoring the mechanical failure simulation parts and their schemes based on the test and according to the mechanical failure test simulation consistency evaluation index; Step 4, selecting the mechanical failure simulation part and its scheme with the highest score and manufacturing the mechanical failure simulation part; Step 5, conducting lift enhancement device failure tests based on the mechanical failure simulation part and its scheme with the highest score, and scoring the mechanical failure simulation parts and their schemes again according to the mechanical failure test simulation consistency evaluation index based on the lift enhancement device failure test results, to obtain the final mechanical score.

The present application belongs to the field of aero-engine surface engineering, and particularly relates to a high-strength steel part rapid hydrogen embrittlement detection method for simultaneously performing two surface treatment processes. The method comprises the following steps: first rough machining, and then fine machining to a specified size; processing two notches on the gauge section of the test bar; performing surface treatment on the first notch and its two sides according to one process, and performing surface treatment on the second notch and its two sides according to another process; after heat treatment to eliminate hydrogen embrittlement, clamping the test bar to a testing machine, and performing sustained stretching for 200 hours under a specified load. The present application can solve the problem that different parts of an aero product need to be subjected to two surface treatment processes, and both of the two surface treatment processes need to be subjected to hydrogen embrittlement detection, and the hydrogen embrittlement detection cycle is long. The present application simultaneously performs two surface treatment processes by using one hydrogen embrittlement test bar, instead of using two hydrogen embrittlement test bars for the original two surface treatment processes.

The application provides an intelligent measurement and control system and method for a friction and wear tester of typical motion pairs of an aircraft tail nozzle, comprising: a state monitoring module: obtaining the deflection angles of horizontal actuators and vertical actuators in a motion loading control module; an image acquisition module: acquiring motion images of a test piece, and obtaining pose data of the test piece based on the acquired motion images of the test piece; a digital twin system module: controlling the motion loading control module based on the obtained pose data of the test piece and the deflection angles, and realizing friction and wear tests of typical motion pairs of the aircraft tail nozzle under preset temperature, load and motion form test conditions. Based on the digital twin technology, the application realizes accurate control of temperature, load and motion form of the friction and wear test of the typical motion pairs of the aircraft tail nozzle and state monitoring of a test system.

The invention relates to the technical field of aeronautical material detection, in particular to an aeronautical composite material physical property quality detection device which comprises a machine body, a lifting seat installed on the machine body, a lower clamp installed on the lifting seat and an upper clamp installed on the top of the machine body. Fixing blocks are fixedly connected to two workbenches of the bidirectional sliding table, and storage shells are fixedly connected to the upper ends of the two fixing blocks. When the aviation carbon fiber composite material is subjected to physical property quality detection through a tension testing machine, two ends of each sample can be automatically bonded with a metal sheet through glue, the operation is simple, time and labor are saved, meanwhile, the glue can be uniformly smeared on the metal sheet so as to ensure the bonding effect between the metal sheet and the sample, and the production efficiency is improved. And the glue can be accelerated to solidify through proper hot air heating, so that the detection efficiency and the detection effect of the aviation carbon fiber composite material are improved.

The application discloses a kind of aircraft crash test machine body full-field deformation non-contact measurement system and method, belong to aircraft structure strength test technical field, comprising: high-quality speckle pattern is made on the surface of fuselage and is laid cooperation mark point;Using binocular camera flexible self-calibration method completes stereoscopic measurement system calibration;Through unmanned aerial vehicle photogrammetry auxiliary multi-system global coordinate unification;Synchronous controller triggers multiple systems synchronous acquisition crash process image;Based on mark point positioning tracking and stereovision principle obtains key measuring point motion trajectory and speed;Based on control point guided digital image correlation method and stereovision principle obtain full-field three-dimensional displacement field and strain field;Realize deformation visual representation by image stitching and data fusion.The application realizes the non-contact, high-precision, full-field dynamic measurement of three-dimensional deformation field in the whole machine crash process of large aircraft, provides technical support for aircraft structure crashworthiness analysis and evaluation and anti-crash design verification.

The invention belongs to the technical field of pipeline tests, and particularly relates to a high-temperature stress endurance testing device for an aviation pipeline connecting piece. The device comprises a tensile testing machine, a high-temperature environment box and a video extensometer, the tensile testing machine comprises a stand column and an upper moving platform sliding on the stand column, the upper moving platform is connected with an upper straight-through connector of a pipeline test piece through an upper switching unit, and a transverse plate at the bottom end of the stand column is connected with a lower straight-through connector of the pipeline test piece through a lower switching unit. Hydraulic oil flow channels are arranged in the upper adapter unit and the lower adapter unit, one end of each hydraulic oil flow channel is connected with a pipeline of the pipeline test piece, and the other end is connected with the pump set unit; the high-temperature environment box sleeves the pipeline test piece and is used for providing a high-temperature environment; the video extensometer is located outside the high-temperature environment box and collects deformation data of the pipeline test piece in the test process through a transparent temperature-resistant glass window of the high-temperature environment box. The application can more accurately simulate and reproduce the use condition of the aviation pipeline piece in the high-temperature environment.

The application belongs to the field of aircraft structure strength test, and is a flat wing section center of rigidity position determination method and device. The weight of a loading device is recorded first, and then the loading device is controlled to load the test wing after checking and judging that each system or device is normal. The test wing is first loaded to an initial value, and then is loaded to different loading points according to settings. The variable amount of the test wing is fed back through a load sensor, and the average value of the load sensor is obtained through calculation. Then, the center of rigidity position of the test wing corresponding to the section can be determined according to the load ratio. The center of rigidity positions of different sections of the test wing can be obtained by loading different loading points of the test wing respectively. After the test is completed, unloading and pressure relief are performed. The electric drive actuator has the performance characteristics of high positioning accuracy, fast response speed, accurate control and high synchronism, and is better than the hydraulic actuator.

The invention belongs to the technical field of aviation structure strength tests, and particularly relates to a structure and function integrated composite material thermal-mechanical coupling tensile test method. The method comprises the following steps: preparing a test piece, wherein the test piece comprises a test piece examination area in the middle and clamping ends located at the two ends of the test piece examination area in the loading direction; the test piece examination area comprises an upper composite material panel, a lower composite material panel and a functional module area located between the upper composite material panel and the lower composite material panel, a resistance wire is wound on a functional module, and two wire ends of the resistance wire are led out of the test piece; respectively connecting the clamping ends at the two ends of the test piece to a U-shaped test clamp through fasteners; the two U-shaped test clamps are installed on a chuck of a testing machine; applying a pair of in-plane tensile loads with opposite directions to the test piece; meanwhile, two wire ends of the resistance wire are connected to a power supply, so that the resistance wire is heated, a heat source is introduced into the test piece, and thermal coupling loading is realized. Coupling simulation of the internal heat source and the load is realized.

A fatigue performance loading test base for airport runway base layers includes a bottom plate with a fixing rod fixed to its center area for securing it to a testing machine. Bolt fixing fittings are provided at both ends of the bottom plate for fixing with bolts. An adjustment block is provided on each of the two upper edges of the bottom plate: a left adjustment block and a right adjustment block. The left and right adjustment blocks have the same structure and are symmetrically arranged. A bolt passes through the bottom of the right adjustment block, and the adjustment block is fixed to the bottom plate through the engagement of the bolt and the bolt fixing fitting. Compared to existing technologies, the technical advantage of this invention is that it features a strip-shaped hole, facilitating adjustment of the distance between the two supports, and making testing convenient even for large semi-circular specimens.

The invention belongs to the technical field of aircraft ground tests, and particularly relates to a large-tonnage tilting rotorcraft ground combined jacking device and method. Comprising a first jacking mechanism, a second jacking mechanism, a tail jacking mechanism and a control unit which are symmetrically arranged on the two sides correspondingly. The distance between the first jacking mechanism and the second jacking mechanism is determined according to the distance between the left wing nacelle and the right wing nacelle of the testing machine; the top ends of the first jacking mechanism and the second jacking mechanism are provided with a first displacement sensor and a second displacement sensor respectively, and the top ends of the first jacking mechanism and the second jacking mechanism can ascend and descend in the direction perpendicular to the ground. The tail jacking mechanism is arranged in the pit shaft relative to the central surfaces of the first jacking mechanism and the second jacking mechanism; a third displacement sensor is installed at the top end of the tail jacking machine, and the top end of the tail jacking machine can ascend and descend in the direction perpendicular to the ground. The first displacement sensor, the second displacement sensor and the third displacement sensor are in communication connection with the control unit. And rapid switching between the test and the test flight can be rapidly realized.

The application belongs to the field of aircrafts, and particularly relates to a multi-blade tilt-rotor aircraft ground vehicle-mounted wind blowing test mechanism and a test method. The test platform expands the existing aircraft ground test method, can more truly simulate various states of the aircraft in flight, and the response states are closer to the aircraft in flight. Meanwhile, compared with the wind tunnel test, the aircraft is in a moving state, and the air is in a natural state, which is closer to the real flight environment of the aircraft, and provides a more reliable test scheme for the aircraft ground test stage.

The invention provides a compression-shear tool, method and system for testing the bonding strength of a non-deployable surface antenna housing, and belongs to the field of structural strength testing. The tool comprises a base, a top seat and a plurality of angle-variable pressing blocks. The base is provided with a first profiling surface matched with the mounting surface of the antenna shell; the lower ends of the pressing blocks are provided with second profiling surfaces matched with the curved surface of the antenna housing, and different sets of pressing blocks have different inclination angles. During testing, the bonding assembly is fixed to the base, the pressure applying block with the corresponding inclination angle is selected to be installed on the top base, pressure is applied through the universal testing machine till the bonding face fails, and strength data are recorded; by replacing the pressing blocks with different inclination angles, a multi-direction stress state is simulated, and a series of bonding strength values are obtained. According to the invention, the problem that the bonding strength of the non-deployable curved radome cannot be accurately tested under the actual complex load is solved, and a reliable experimental basis can be directly provided for the structural design of an aircraft.

The application belongs to the technical field of strength test of single lug joint of vertical tail wall plate root of an airplane, and particularly relates to a radial compression test loading structure of a single lug joint shaft hole of a vertical tail wall plate root of an airplane, when strength test is performed by using the radial compression test loading structure, a test machine can apply a pressure load on a wall surface of a loading pressure platform, the pressure load can be transmitted to a shaft hole of a single lug joint of a single lug joint test piece through a support bushing and a loading shaft, a radial compression load is applied in the shaft hole of the single lug joint, in the process, two support bushings can support the single lug joint of the single lug joint test piece from two sides, two support plates can support a transition section of the single lug joint test piece from two sides, lateral deviation of the single lug joint test piece is prevented, and the two support bushings and the support plates can follow the loading pressure platform to follow the support, the stress environment of the single lug joint can be well ensured, and the accuracy of a test result can be ensured.

This application discloses a multi-physics field coupled wear testing device for adjustable stator bushings of aero-engines, belonging to the field of aero-engine rotor wear testing technology. To address the problems of existing test benches being unable to accurately simulate complex stress conditions of bushings, difficulty in decoupling load components, and insufficient equipment adaptability, this application includes a support unit, a test spindle unit, and a test housing unit. Both the test spindle unit and the test housing unit are mounted on the support unit. The bushing specimen to be tested is placed between the test spindle unit and the test housing unit. The relative motion between the test spindle unit and the test housing unit simulates the working state of the bushing specimen. The test spindle unit integrates a loading component for simulating the working conditions of the bushing specimen, and the test housing unit integrates an environmental simulation component for simulating the working environment of the bushing specimen. This application is mainly used as a device for wear testing of adjustable stator bushings of aero-engines.

The invention relates to the technical field of aircraft engine suspension testing, in particular to a combined type repeated stretching device for aircraft engine suspension, which comprises an upper clamp and a lower clamp, the upper clamp comprises an adapter shaft, a sleeve, a universal ball head, a third baffle ring and a to-be-tested piece connector, one end of the adapter shaft penetrates through the sleeve and extends out to be connected with a testing machine, and the other end of the adapter shaft is connected with the universal ball head; the third baffle ring is matched with the outer side of the bottom of the sleeve, a connecting shaft of the universal ball head penetrates through the third baffle ring and is connected with a connector of a to-be-tested piece, and the connector of the to-be-tested piece is detachably connected with the to-be-tested piece. The universal ball head is used as a connecting device, the axis can be automatically adjusted in the stretching process, it is ensured that the axis of a part coincides with the axis of a testing machine, and therefore the coaxiality in the testing process is guaranteed, and the accuracy and reliability of the test are improved.

The invention belongs to the field of aircraft design, and particularly relates to an electric vertical take-off and landing aircraft development cost estimation method. The development cost of the electric vertical take-off and landing aircraft is the sum of the aircraft body engineering and design cost, the development support and test cost, the flight test machine cost, the flight test operation cost, the development profit and the development stage financing cost. The development cost of the electric vertical take-off and landing aircraft is calculated through a formula related to overall parameters of the aircraft. The formula is obtained based on sample data through multiple linear regression. According to the method for estimating the development cost of the electric vertical take-off and landing aircraft, the development cost of the electric vertical take-off and landing aircraft can be rapidly and accurately calculated, and the method has a good guiding effect on the overall design of the electric vertical take-off and landing aircraft.

The invention discloses an aeronautical material anti-icing coating icing binding force testing device which comprises an upper clamping plate and a lower clamping plate of the same specification, the upper clamping plate is arranged on one side of the lower clamping plate, the lower surface of the upper clamping plate is used for bonding a first coating sample plate, and the upper surface of the lower clamping plate is used for bonding a second coating sample plate; the device further comprises a supporting part and a connecting part, the supporting part is arranged between the upper clamping plate and the lower clamping plate, and the connecting part is installed on the upper clamping plate and the lower clamping plate and used for being externally connected with a clamping mechanism of a testing machine. Flexible edge sealing elements are arranged on the side walls of the upper clamping plate and the lower clamping plate, and an open space capable of storing water is jointly defined by the flexible edge sealing elements, the first coating sample plate and the second coating sample plate. According to the invention, ice making can be smoothly and rapidly carried out between two coating test plates, the ice making difficulty and cost are obviously reduced, and the probability that an ice layer is easy to crack before a stretching process of a coating ice binding force test is reduced.

The application discloses a test mechanism unit for simulating follow-up thrust generated by an elongated body missile in a flight process; wherein the elongated body missile model comprises an aluminum rod and a test engine; in an initial state, a side column is connected with the second end of the test engine through an explosive bolt, the aluminum rod is bent, and the elongated body missile model is in a deformation state; after the explosive bolt is ignited and separated, the aluminum rod drives the test engine to move to the other side of the support frame, and at the same time, the test engine is ignited to generate follow-up thrust. The application further discloses an engine follow-up thrust vibration test system and method, a camera unit acquires a motion image of the elongated body missile model; a vibration test unit acquires vibration data; and a thrust test unit is used for acquiring the thrust of the elongated body missile model. The application solves the problem that a traditional engine ground static ignition test system cannot generate follow-up thrust, and realizes the simulation of the vibration condition of the elongated body missile after elastic deformation in flight.

The present application relates to the field of aerospace test, and more particularly to a test machine cabinet resource utilization rate evaluation method, comprising S1: establishing a multi-dimensional sub-resource utilization rate model, the multi-dimensional sub-resources including installation space resources, power resources, heat dissipation resources, information transmission resources and on-orbit work hour resources; S2: determining the weight of each sub-resource, the weight satisfying the sum of the weights of each sub-resource being 1; S3: calculating the utilization rate of each sub-resource based on the multi-dimensional sub-resource utilization rate model, and calculating the comprehensive resource utilization rate in combination with the weight of each sub-resource; S4: evaluating and grading the comprehensive resource utilization rate, which is divided into four grades of efficient utilization, general utilization, low efficient utilization and serious shortage, and corresponding improvement directions are executed according to different grades. Through multi-dimensional coupling modeling and dynamic weight management, the present application realizes accurate evaluation, efficient allocation and safe control of test machine cabinet resources.

The application provides a high-speed aviation bearing testing machine simulating high-temperature environment, and belongs to the technical field of high-temperature bearing testing; high-temperature gas is introduced in two directions through symmetrical layout of an external annular heating cavity and an internal high-temperature heating cavity, so that temperature is controlled in cooperation with the double heating cavities; a physical isolation layer is formed by the setting of an inner side interval cover and an end face heat insulation cavity, so that lubricating oil is blocked from contacting high-temperature gas; the internal and external high-temperature heating cavities and the bearing tooling are physically isolated through a heat insulation layer, so that heat conduction interference is avoided to measure data. The application comprises a machine base, a base and an upper cover fixedly installed on the machine base, the upper cover and the base jointly forming a closed cavity wrapping a test tooling, a connecting shaft, an external annular heating cavity, a bearing assembly to be tested and an internal high-temperature heating cavity, the closed cavity is inserted with the connecting shaft, the external annular heating cavity is arranged at a right side area of a tail end of the connecting shaft, and the internal high-temperature heating cavity is arranged at a shaft center area of the connecting shaft.

The application discloses a civil aircraft wing root large-size wall plate butt joint structure test device and a test method. The test device comprises a test piece, a support component and two connecting assemblies. The test piece has an outer wing side wall plate and a central wing side wall plate. The two connecting assemblies are clamped on the outer wing side wall plate and the central wing side wall plate of the test piece respectively, and the two connecting assemblies are connected to two chucks of a test machine respectively. The support component supports any one of the connecting assemblies to level the two connecting assemblies. The test device of the application can level the two connecting assemblies by adding the support component, reduces the influence of the gap and assembly error on the test, and guarantees the loading centration of the large-size test piece. The application solves the problems of misalignment and uneven stress distribution in the test process of the large-size wall plate butt joint structure, and can more accurately simulate the stress state of the wing root large-size wall plate butt joint structure.

The invention belongs to the field of hydrodynamic tests of aircrafts, and particularly relates to a wing body fusion assembly type hydrodynamic test full-aircraft model. Comprising a hydrodynamic test fuselage module (1), a wing body fusion module (2), a vertical tail module (3), a horizontal tail module (4) and a buoy module (5), the fuselage module (1) and the wing body fusion module (2) are connected through a butt joint clamping groove (11), the fuselage module (1) and the vertical tail module (3) are connected through a butt joint platform (10), and the vertical tail module (3) and the horizontal tail module (4) are connected through a butt joint groove (22) in the top of a vertical tail; and the buoy module (5) is placed in a buoy positioning hole (18) reserved in the wing body fusion module (2).

The invention provides a vacuum high-low temperature rotating ball-disc friction testing machine, which relates to the technical field of material friction performance testing, and comprises a vacuum chamber, and a motion system, a ball sample system, a disc sample system, a high-low temperature cold trap system and an automatic sample changing system which are arranged in the vacuum chamber. The equipment can automatically replace a ball sample and a disc sample without damaging a vacuum environment. The equipment can simulate a high-vacuum and wide-temperature-range extreme environment, realizes automatic high-precision friction test, and is suitable for performance research of materials in the fields of aerospace and the like.