72 results about "Experimental aircraft" patented technology

The embodiment of the invention provides an equivalent model design method and system for an unmanned aerial vehicle carrying mission load flight test, and the method comprises the steps: obtaining mission load demand parameters and test condition parameters, the demand parameters comprise a long side A1 and a short side A2 of a rectangular airspace, A1 > = A2, and a demand speed v0; the test parameters comprise a test airspace long side B1 and a test airspace short side B2, B1 is greater than or equal to B2, the test machine minimum speed v1 and the test machine maximum speed v2, and v2 is greater than or equal to v1; when A1 < = B1, A2 < = B2 and v1 < = v0 < = v2 are not satisfied, calculating an equivalent coefficient p, and if the airspace condition A1 < = B1 and A2 < = B2 are not satisfied, calculating and determining the equivalent coefficient p; if the airspace condition A1 < = B1 and A2 < = B2 are satisfied, but the speed v1 < = v0 < = v2 is not satisfied, determining an equivalent coefficient p; based on the reference point coordinates and the equivalent coefficient p, longitude and latitude are achieved through geocentric coordinate system conversion, and horizontal navigational speed mapping is carried out. According to the technical scheme, by simulating the flight profile required by the mission load, the problem that the unmanned aerial vehicle cannot fly out of the flight profile required by the mission load in a flight test carrying the mission load can be solved.

The invention discloses an anti-unbalance loading testing machine for aero-engine main shaft bearing test, which comprises a machine body, a core shaft and two test seats arranged in the machine body, the two test seats are coaxially and oppositely arranged, and test holes for the core shaft to penetrate through are formed in the two test seats in a penetrating manner; a test cavity for mounting an external to-be-tested bearing is formed between the inner peripheral wall of the test hole and the outer peripheral wall of the mandrel, and one of the test seats is provided with an axial loading structure for applying an axial load to the other test seat; a guide structure is arranged between the two test seats, the guide structure is used for being in linkage fit with the axial loading structure to correctly conduct the axial load so as to prevent unbalance loading, and the two test seats are both provided with radial loading structures used for applying radial loads to the corresponding test seats. The problem that a traditional aero-engine main shaft bearing test device is prone to unbalance loading when the bearing is subjected to axial loading and radial loading is solved.

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.

The invention relates to a four-point bending test method which comprises the following test steps: I, preparing a spare part, namely manufacturing a test sample piece which must be made of the same material as a composite material part on an airplane; iI, state inspection before testing: (1) vibration requirements, (2) electromagnetic requirements, (3) temperature and humidity requirements, and (4) setting of an upper clamp and a lower clamp of a testing machine; iII, loading a test sample; iV, testing the test sample piece, and driving the upper clamp to move downwards for loading, namely (1) preloading setting, (2) sample piece measurement and (3) loading setting in sequence; and V, comparing test results, and correcting the load and the deformation in consideration of the deviation between the actual size and the nominal size. The four-point bending test method has different test standards for aircraft shell parts of different types and specifications, data is corrected based on tested data set deviation, the data precision is improved, and then safe flight of an aircraft is guaranteed.

The invention belongs to the field of airplane structure strength tests, and particularly relates to a fuselage compensation loading method for an airplane strength test. The method comprises the following steps: acquiring fuselage deformation and lateral loading point mounting dimensions under different load values; the deflection angle of a lateral loading force line is calculated according to the fuselage deformation and the installation size of the lateral loading point; calculating a new lateral load instruction according to an original lateral load instruction and the deflection angle; calculating a new vertical load instruction according to the new lateral load instruction, the original vertical load instruction and the deflection angle; and according to the new lateral load instruction and the new vertical load instruction, applying the airplane strength test fuselage load. According to the method, the vertical and lateral load values are adjusted according to the fuselage deformation data and the geometric installation size of the loading point, and the problem that the loading precision is reduced due to lateral loading force line deflection can be solved.

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 invention discloses a device and method for testing the mechanical property of a high-temperature-resistant inclined coil spring, and relates to the technical field of hypersonic aircrafts, the device comprises a fixed support, a movable support and balls, and the inclined coil spring is arranged on the upper surface of the movable support; when an upper pressure head of the universal testing machine applies pressure to the inclined coil spring, static friction force between the inclined coil spring and the movable support is converted into static friction force and rolling friction force between the ball and the fixed support and between the ball and the movable support, and slippage of the inclined coil spring and the movable support is eliminated. The testing device is arranged in the heating furnace, the inclined coil spring can be heated through the heating furnace and maintained at the target temperature, and the mechanical property of the inclined coil spring can be effectively and accurately measured under the high-temperature condition.

The invention relates to the technical field of gear box fault experiments, and discloses a non-inertial gear box fault implantation test platform which comprises a bottom plate, a top plate, a gear box fault test mechanism and a plurality of telescopic adjusting pieces. Wherein the top plate is located above the bottom plate. The gearbox fault experiment mechanism is installed on the top plate. The two ends of each telescopic adjusting piece are movably connected with hinge seats, and the hinge seats at the two ends of each telescopic adjusting piece are connected with the bottom plate and the top plate correspondingly. The position and the inclination angle of the top plate are changed by controlling the extension or shortening of the telescopic adjusting piece, so that the working states of the gearbox fault test mechanism under different working conditions are simulated. According to the invention, the top plate can be inclined within a certain range through the cooperative work of the telescopic member, so that the diving and lifting states of the gearbox flying on an airplane are simulated, and the working stability of the gearbox is verified greatly.

The invention relates to the technical field of aircraft tire testing, and discloses aircraft tire dynamic cable passing equipment and a test method, and the aircraft tire dynamic cable passing equipment comprises a hub wheel device, a loading device and a cable passing device; the loading device comprises a loading head, a steering device, a test wheel and a force measuring device, the steering device is used for driving the test wheel to rotate, and the force measuring device is used for monitoring and controlling the radial load of the test wheel; the hub wheel device comprises a wheel body and a driving piece, the driving piece is used for driving the wheel body to rotate, and the inner side face of the wheel body makes contact with the testing machine wheel to drive the testing machine wheel to rotate; the cable passing device comprises a cable pushing device and a cable; the cable pushing device is used for pushing the cable to a contact area of the wheel body and the testing machine wheel; according to the invention, the position of the test wheel and the deflection angle and load of the test wheel are adjusted through the loading device, different test requirements can be met, the test wheel can rotate through the arrangement of the wheel body and the driving part, so that the state of the test wheel during cable passing better fits the real situation, and the test result is more real and accurate.

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 aerospace engines, and provides a method for predicting the fretting fatigue life of a high-temperature alloy of an aero-engine, and the method comprises the steps: constructing a fretting fatigue simulation test device; clamping the position of the connecting plate of the fretting fatigue simulation test device on a fatigue testing machine chuck, and applying a preset sine tensile cyclic load to the second ends of the two middle plates to obtain a fretting fatigue index; simulating a fretting fatigue process between the middle plate and the connecting plate based on the fretting fatigue index to obtain stress-strain information of the middle plate; and obtaining a fretting fatigue life result of the aero-engine based on the stress-strain information. The fretting fatigue simulation test device is constructed to simulate the aero-engine bolt structure, the fretting fatigue performance test is performed on the fretting fatigue simulation test device, the obtained fretting fatigue index reveals the influence of the fretting fatigue life of the material, the fretting fatigue process is simulated based on the fretting fatigue index, and the fretting fatigue performance of the material is improved. And accurate prediction of the fretting fatigue life of the high-temperature alloy of the aero-engine is realized.

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 invention discloses a method for quantitatively judging the operation stability of a dynamic derivative test vibration mechanism, relates to the technical field of fault prediction, and aims to solve the problem that no unified judgment index exists for judging whether a test mechanism has a fault or not in an existing high-speed wind tunnel dynamic derivative test. According to the method, a normalization idea is adopted, firstly, angular displacement voltage value data is subjected to normalization processing, and angular displacement voltage value data scales under different mechanisms or different working conditions are unified; then, calculating a single-cycle residual sum of squares between the sine curve and a standard sine curve; and finally, judging whether the mechanism breaks down or not by evaluating whether the residual sum of squares of a single period exceeds an interpretation index or not and whether the angular displacement amplitude is out of tolerance or not. The invention provides important support for improving the equipment performance, prolonging the service life and ensuring the safety in the aerospace field.

The invention discloses a thermal flow oxygen salt five-field combined cross-sea-area high-speed flight simulation working condition experimental testing device and method, and belongs to the field of material testing. The device is used in cooperation with a universal testing machine, a fan, an electromagnetic induction heating module, an oxygen supply module, a salt mist generation module, a temperature / deformation field measurement module and an environmental parameter measurement module are arranged inside and outside a high-speed airflow circulation loop, and a sample is loaded through the universal testing machine and placed in an experiment section in the high-speed airflow circulation loop. Wide-temperature-range closed-loop control of the temperature of the sample, construction of a comprehensive stress state and construction of a controllable high-speed airflow, adjustable oxygen concentration and salt mist concentration gas environment are realized, so that a real material application environment is simulated, and acquisition of data of a temperature field, a deformation field and an oxidation corrosion field of the sample can be completed in a reaction process; the problem that a cross-sea-area high-speed aircraft service condition environment full-state simulation test cannot be realized in the prior art is solved.

The present invention relates to a device and a usage method for a pull-off test of composite fasteners, belonging to the technical field of aircraft strength tests, capable of ensuring that the test piece and the loading axis of the testing machine meet the centering requirements during the assembly process and realizing the rapid installation of the test piece; the device includes a base, an upper cover plate, a connecting rod and a centering assembly; during centering, the centering assembly is centered and connected to the upper cover plate, and the lower section of the connecting rod is centered and connected to the centering assembly; the centering connection is specifically a connection without horizontal displacement; the upper section of the connecting rod is fixedly connected to the testing machine; the base and the upper cover plate are fixedly connected; the base is fixedly connected to the testing machine; the test piece includes a composite material plate and a lug fixedly arranged on its upper surface; during the test, the base is fixedly connected to the testing machine, and the base and the upper cover plate are fixedly connected; the composite material plate is located in the space between the base and the upper cover plate, and the lug passes through the upper cover plate and is connected to the lower section of the connecting rod; the upper section of the connecting rod is fixedly connected to the testing machine.

A hammering test system includes a hammering test device including a target, a flying unit, and a hammering test mechanism configured to conduct a hammering test on a test object, a surveying instrument including a scanner for acquiring point cloud data by scanning with scanning light, and configured to be capable of performing tracking and distance and angle measurements of the target, and an arithmetic processing unit including a point cloud data analyzing unit configured to calculate shape data by analyzing point cloud data acquired by the scanner, and a flight plan calculating unit configured to calculate a flight plan of the hammering test device based on the shape data calculated by the point cloud data analyzing unit, and the surveying instrument tracks the target of the hammering test device and makes distance and angle measurements when the hammering test mechanism conducts a hammering test.

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 high-temperature fatigue testing, and particularly relates to a high-temperature bending fatigue testing device and method.The main body structure of the testing device comprises a fatigue testing machine and a heating furnace which are connected into a whole, and a movable clamp and a static clamp which are arranged in the heating furnace; the fatigue performance and the fatigue life of the metal material are measured through a bending test; the specific technological process of the test method comprises the following steps: (1) determining the actual service working condition of a sample; (2) determining bending fatigue test key parameters; and (3) determining test time, carrying out a high-temperature fatigue bending test, reasonably determining high-temperature bending fatigue test parameters based on corrosion fatigue equivalent conversion, and rapidly reproducing fatigue damage of equipment under a long-term high-temperature service condition through a short-time test in a laboratory under the condition of not changing a high-temperature fatigue mechanism and a crack development process. And rapid screening and examination verification of the novel structure / material of the equipment are realized.

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.

The invention discloses a mechanical loading device for a special-shaped main force-bearing component of a large additive aircraft fuselage. The mechanical loading device comprises an upper clamping and fixing frame, a lower clamping and fixing frame, a first force transfer mechanism, a second force transfer mechanism, a first push-pull force reversing mechanism and a second push-pull force reversing mechanism, the upper clamping and fixing frame is connected with an upper chuck of the uniaxial fatigue testing machine; the lower clamping fixing frame is connected with a lower chuck of the uniaxial fatigue testing machine; the first force transmission mechanism is arranged at one end of the upper clamping fixing frame, the second force transmission mechanism is arranged at the other end of the upper clamping fixing frame, and the first force transmission mechanism and the second force transmission mechanism are distributed in a left-right mirror symmetry mode relative to the lower chuck. The first push-pull force reversing mechanism is arranged on the first force transmission mechanism, the second push-pull force reversing mechanism is arranged on the second force transmission mechanism, and the first push-pull force reversing mechanism and the second push-pull force reversing mechanism are distributed in a left-right mirror symmetry mode relative to the lower chuck. The two ends of the test piece are connected with the first push-pull force reversing mechanism and the second push-pull force reversing mechanism respectively, and the middle of the test piece is connected with the lower clamping fixing frame.

An unmanned helicopter automatic inclinator bearing testing machine disclosed by the present invention comprises a test platform, a main body rack, a gantry loading support and a loading force arm, the main body rack is installed on the test platform, two ends of the gantry loading support are fixed on the test platform and stretch across the main body rack, and a to-be-tested bearing is installed on the upper side surface of the main body rack. The loading force arm is installed on the upper side face of the main body rack and located at the upper end of a bearing to be tested, the gantry loading support is provided with an axial load servo loading hydraulic cylinder and an overturning force simulation servo loading hydraulic cylinder, and a push rod of the gantry loading support faces downwards and is matched with the loading force arm to apply axial load and overturning force to the bearing to be tested. The device can simulate the real working condition of the automatic inclinator bearing of the unmanned helicopter for testing, accurately applies an axial load and an overturning force, provides reliable data for the performance test of the bearing, and facilitates the improvement of the quality and reliability of the automatic inclinator bearing of the unmanned helicopter.