89 results about "Experimental aircraft" patented technology

The invention belongs to a fatigue life determining technology for an operating system of an aircraft and relates to a fatigue test method for a full-aircraft main operating system of the aircraft. According to the fatigue test method, the fatigue coordination loading of an operating load and an operating displacement can be carried out, a fatigue test including the operating load and an operating displacement spectrum on the full-aircraft main operating system of the aircraft can be implemented, and meanwhile, the problem of simultaneous test of aircraft body fatigue on the same fatigue testing machine is taken into account, so that the fatigue properties of components and supporting parts of the full-aircraft main operating system and static performance indicators of the full-aircraft main operating system are all actually assessed; the actual loaded conditions of the main operating system of the aircraft under normal operating and emergency operating conditions are more truly reflected, the fatigue properties of the structure and mechanism of the main operating system of the aircraft are embodied; and the test cost is reduced, and the reliability of a test result is improved.

The invention belongs to a flight control system fatigue test technology and relates to a load simulator of a plane booster dummy assembly. The load simulator comprises a base, a front connector, a back connector, a middle connector and a distributing valve simulator, wherein the distributing valve simulator comprises an outer cylinder, a piston rod, an outer cylinder connector and a piston rod connector. According to the invention, controlling load and controlling run of a plane booster or a composite steering engine distributing valve can be simulated and provided, a requirement on a main control system fatigue test of a whole plane under a plane control load and control displace spectrum is met, at the same time, the load simulator is fixed according to an appointed airplane control surface position, and fatigue tests are carried out on a control system behind the plane booster or the composite steering engine and a plane body simultaneously, so that the main control system fatigue test of the whole plane and the plane body fatigue test can be carried out simultaneously on the same fatigue test plane, the fatigue performance of the plane main control system is truly reflected, and the progress of the main control system fatigue test of the whole plane is accelerated.

The invention discloses an improved actuation mechanism for a fatigue testing machine achieving in-situ imaging of a synchrotron radiation light source. The actuation mechanism is arranged on a cylinder-shaped base plate on a platform of the synchrotron radiation light source and used for applying a vertical reciprocating displacement load to a test sample on the fatigue testing machine achieving in-situ imaging of the synchrotron radiation light source. Compared with the last generation of fatigue testing machines, further improvement is conducted, friction generated by a cam actuation mode is reduced, therefore, energy consumption of the fatigue testing machine is reduced, and the noise reduction performance of the fatigue testing machine is further improved; the load applied to the test sample is further guaranteed, and compared with the last generation of the fatigue testing machines, the details in the test implementation process are considered more detailedly.

The invention discloses a rolling forced vibration dynamic derivative test device for a high-speed flying wing model under a large attack angle. The rolling forced vibration dynamic derivative test device comprises a rigid support device, a motor driving device, a motion conversion device, a simple harmonic motion angle measurement device, a simple harmonic motion transmission device, a five-component dynamic load measurement device and a dynamic derivative test model; and the rigid supporting device supports the whole test mechanism, and the tail end is directly installed on a wind tunnel bent blade arm. The motor driving device is a power source, provides continuous rotary motion output for the whole set of test mechanism and is mounted at the tail part of the rigid supporting device; the motion conversion device converts the continuous rotating motion into required simple harmonic motion, and the angle time history of the simple harmonic motion is measured through the simple harmonic motion angle measuring device; the simple harmonic motion transmission device transmits motion to the five-component dynamic load measuring device, the dynamic derivative test model is installed onthe five-component dynamic load measuring device, and the dynamic derivative test model is subjected to dynamic load in the simple harmonic motion process of the five-component dynamic load measuringdevice; and signal processing is performed on the measured dynamic load time history and the simple harmonic motion angle time history so as to obtain a required dynamic derivative value.

The invention belongs to an aircraft test piece constraint technology, and relates to a pole restraint device. The pole restraint device comprises a binaural joint, a connecting screw rod, a spiral pad, a load sensor, and a false actuating cylinder, wherein the binaural joint is fixed on a test piece constraint part and connected with the connecting screw rod in a hinge manner; the spiral pad is added to the threaded connection part between the load sensor and the false actuating cylinder, and the load sensor is connected with the control system. The pole restraint device disclosed by the invention can regulate the gesture of a testing machine by regulating the telescopic length of a connecting steel tube of the false actuating cylinder; and constraint of a 'two-force bar' structure in rigid connection under a shutdown state is still reliable and valid. Therefore, the existing testing equipment is effectively utilized, and reliable horizontal constraint is also achieved.

The invention relates to a dynamic load simulation test method of an airplane wheel bearing. A test machine comprises a dragging system and a loading system, wherein a tested bearing rotates through the dragging system according to the stipulated rotation speed, the loading system is used for providing a test load to the tested bearing, the tested bearing is enabled to simulate the use working conditions such as the stress working condition and test condition of the airplane wheel bearing in the flying, high-speed sliding, landing impact moment, laterodeviation landing and stopping processes of an airplane, the working condition of the airplane wheel bearing can be comprehensively reflected, the strength and service life can be accurately evaluated, the credible test evidence is provided for researching the airplane wheel bearing at a higher rotation speed and heavier load, and the research period can be effectively shortened.

A helicopter tail rotor booster fatigue test system is composed of a fatigue testing machine, a hydraulic station, a rotary direct-drive valve, an actuator tester, a digital storage oscilloscope, a crossover coupling, an upper clamp, a lower clamp and a helicopter tail rotor booster, wherein the upper clamp and the lower clamp are installed on the fatigue testing machine, the two ends of the helicopter tail rotor booster are installed on the upper clamp and on the lower clamp respectively, the hydraulic station is connected with the rotary direct-drive valve through an oil pipe, the rotary direct-drive valve is further connected with the helicopter tail rotor booster, the actuator tester is connected with the rotary direct-drive valve through a circuit, and the digital storage oscilloscope is connected with the actuator tester to display relevant information of the helicopter tail rotor booster. The helicopter tail rotor booster fatigue test system not only can simulate the external aerodynamic loading action but also can simulate the actuator hydraulic pressure action, and the testing result has great engineering application significance for helicopter tail rotor booster fatigue lifetime determination.

The invention provides a control law design method for civil fly by wire airplane PIO research, which belongs to the technical field of flying tests. The method of the invention is realized based on an air stimulation tester. Through analyzing an aerodynamic model and flight quality feature of the air stimulation tester at the test state point, target quality features having different levels in longitudinal and transverse direction can be selected based on Cooper-Harper criterion. A full-state feedback method is adopted to compute the feedback gain matrix of control law. With the adding of a time delay module and a steering engine rate saturation module, a full-quantity model in the air stimulation tester has classic civil fly by wire airplane responding feature and can trigger a driver tostimulate oscillation PIO. The invention is advantageous in that under the premise that flying safety is guaranteed, the oscillation PIO can be triggered by different drivers, test objects can be provided for civil fly by wire airplanes PIO research, and domestic gaps can be filled up.

The invention provides an aircraft test comprehensive climate environment simulation system and a simulation method, and belongs to the technical field of an aircraft test. The simulation system comprises a testing machine, a laboratory structure system and an environment simulation system. The simulation method comprises the following steps: S1, low-temperature environment simulation; S2, high-temperature environment simulation; S3, temperature daily circulation environment simulation; S4, humidity environment simulation; S5, solar radiation environment simulation; S6, rain / wind-blown rain environment simulation; S7, snowfall / snow blowing environment simulation; S8, fog environment simulation; and S9, icing / freezing rain environment simulation. The system solves the problems that an existing experiment facility is generally small in size and relatively single in function, and has the advantages that the functions are complete, all modules work in a matched mode, and the number of devices is reduced.

The invention belongs to the technical field of aircraft strength test design, and particularly relates to a deformation constraint optimization method for a strength test wing, which comprises the following steps: step 1, constructing a mechanical model of the test wing under the action of a water load, the mechanical model being a mechanical balance model of the wing under the action of the water load and a trimming load; 2, determining constraint conditions for deformation constraint optimization of the test wing according to the mechanical model; 3, converting multi-objective optimizationcontaining the two parameters into single-objective optimization based on the degree of influence of the deflection and the rotation angle on wing deformation by taking two optimization objectives ofthe minimum deflection and the minimum deflection angle at the boundary s of the area to be assessed as two optimization objectives; and 4, according to the constraint conditions in the step 2 and theoptimization target in the step 3, performing iterative solution by adopting an ant colony algorithm to obtain an optimal balancing load scheme. According to the invention, aircraft wing displacementand corner change caused by water load can be reduced at the same time, wing deformation is restrained, and the test loading precision is improved.

The invention discloses an ultrahigh-speed tensile testing machine driven by rocket engines. The ultrahigh-speed tensile testing machine comprises a workbench, a stretching device, a driving device, a reduction gear, and a guide slide rail, wherein the stretching device comprises a first clamp, a second clamp and a sliding plate; the driving device comprises the rocket engines; the guide slide rail is fixed at the upper end of the workbench; the first clamp is fixed at one end of the upper end of the workbench, and the reduction gear is fixed at the other end of the upper end of the workbench; the second clamp is fixed on the sliding plate and used for clamping the other end of a tested sample, and a connecting line between the first clamp and the second clamp is parallel to a length direction of the guide slide rail; the sliding plate slides linearly on the guide slide rail; and the rocket engines are symmetrically fixed at two ends of the sliding plate. The tensile testing machine provided by the invention can implement a short-time ultrahigh-speed tensile test for a material, can adjust the stretching stroke at the same time, and is suitable for performance analysis of short-time ultrahigh-speed loading of a material.

The invention provides an outdoor measuring method for dynamic characteristic of a flow field in rising and landing periods of an airplane and a smoke generating device for measurement in an airplane outdoor true environment based on an unmanned aerial vehicle. A human-driven testing airplane is replaced by a dynamically similar unmanned aerial vehicle, thereby reducing testing cost and risk. The outdoor measuring method is suitable for testing and comparison of a novel layout airplane in a designing and developing period. Compared with a wind tunnel static simulation test, the outdoor measuring method is advantageous in that the outdoor measuring method is a dynamic measuring method through placing the to-be-measured airplane and the measuring device in the outdoor true flight environment, thereby realizing truer testing environment and high reference value of obtained experiment data. Furthermore, according to the outdoor measuring method, the visual characteristic of the flow field of a smoke tunnel is combined with environment authenticity of a flight environment, and a new thought is supplied for verifying airworthiness compliance of a large-size transporter airplane in the rising and landing periods.

The invention discloses an aircraft wing rib collapse large-opening stiffened wall plate shear test equipment and method. The equipment comprises a rib-collapsing large-opening stiffened wallboard test piece with a plurality of rib dummy pieces, and the end face of each rib dummy piece is designed to be a semicircular cambered surface; a flat plate supporting clamp is arranged at the cambered surface end of each fake rib part, so that the cambered surface of each fake rib part is in linear contact with the plane of the flat plate, an out-of-plane simple supporting condition is provided for a non-opening area of the wall plate, the supporting state can provide the out-of-plane simple supporting condition along with deformation of the test piece after the test piece is loaded, and the supporting condition is real; and clamping sections are designed on four sides of the test piece, the clamping sections are clamped by loading clamping plates, the loading clamping plates are in butt joint with a testing machine through lugs, and a shear loading state of the test piece is realized by applying a diagonal tensile load. Compared with a traditional method, the method has the advantages that a box section does not need to be designed to install the test piece, the form of the test piece is simplified, the test scale is greatly simplified, test implementation is easy, the test period is shortened, and expenditure is reduced.

The invention belongs to the field of airplane strength design, and particularly relates to a method for verifying the influence of assembly stress on structure fatigue life. The method comprises thesteps:1, obtainig a force bearing structure of an airplane, and determining a fatigue key part of the force bearing structure; 2, acquiring an assembly gap value of the fatigue key part and a workingstress value corresponding to the assembly gap value, judging whether the influence of different assembly gap values on the working stress value is obvious or not, and if so, entering a step 3; 3, designing a test piece according to the force bearing structure, wherein the test piece comprises a simulation piece and a profile, and the profile is connected with the simulation piece through a fastener; 4, setting the distance between the section bar and the simulation piece as a preset gap value through a fastener, and installing two ends of the test piece on a fatigue testing machine to carry out a fatigue life test, so as to obtain the fatigue life of the test piece; and 5, replacing the test piece, adjusting the preset gap value, and repeating the step 4 to obtain a corresponding curve graph of the assembly gap value and the fatigue life of the test piece.

The invention belongs to the technical field of aero-engine part testing, and particularly relates to a small-size I-shaped sample testing device. The device comprises an I-shaped sample, a clamping block, a positioning block, a positioning pin and a switching section. The positioning grooves are formed in the left side and the right side of the first clamping block respectively, a T-shaped through groove is formed in the middle of the first clamping block, the positioning grooves are formed in the left side and the right side of a second clamping block respectively, the T-shaped through groove is formed in the middle of the second clamping block, and the T-shaped through groove of the first clamping block and the T-shaped through groove of the second clamping block form the I-shaped through groove, a first positioning pin is mounted in the positioning grooves in the left sides of the first clamping block and the second clamping block, and a second positioning pin is mounted in the positioning grooves in the right sides of the first clamping block and the second clamping block; an I-shaped sample is arranged in the I-shaped through groove; the positioning block is used for limitingthe I-shaped sample in the I-shaped through groove, and a first switching section and a second switching section are used for connecting the first clamping block and the second clamping block with atesting machine respectively. The device is simple, effective and high in universality.

The invention provides a method for evaluating abradability of a sealing coating and application of the method, and relates to the technical field of aerospace. The evaluation method comprises the steps of firstly, putting an alloy base material coated with the to-be-detected sealing coating and an alloy simulation blade on an abradable testing machine to carry out rub-impact detection; measuringand recording the arithmetic average roughness, the average peak-valley height and the maximum surface peak-valley height of the surface of the abrasion mark of the sealing coating; substituting the measured values into an abrasion mark roughness evaluation formula, and calculating to obtain an abrasion mark roughness score, the abrasion mark roughness evaluation formula being as follows: the abrasion mark roughness score = average peak valley height / (10 * arithmetic average roughness) + maximum surface peak valley height 2 / (20 * arithmetic average roughness); and then analyzing the abrasion mark roughness score to obtain an evaluation result. The evaluation method can accurately reflect the influence of the coating abrasion mark morphology on the aerodynamic loss in an engine flow channel.

The invention provides an undercarriage ejection rod folding and unfolding performance testing machine. The testing machine comprises an undercarriage assembly, a test rack assembly and a performancetest assembly, and the test rack assembly comprises a C-shaped test rack, a rectangular sleeve cup, a proximity switch, a stand column and a guide rod. The C-shaped test rack is fixedly connected withthe concrete foundation, the rectangular sleeve cup is formed by welding a cup wall plate and a bottom plate, the bottom plate of the rectangular sleeve cup faces upwards, and vertical cylindrical through holes are formed in the center and at the four vertex angles of the bottom plate of the rectangular sleeve cup respectively. The performance test assembly comprises an anti-impact base plate, arectangular lifting platform, a pull pressure sensor, a cylinder and a connecting shaft, the first end face of the pull pressure sensor is fixedly connected with the upper end of a cylinder piston rod, and the second end face is fixedly connected with a first end face of the connecting shaft. The second end face of the connecting shaft is fixedly connected with the lower surface of the rectangularlifting platform, and the anti-impact base plate is fixedly connected with upper surface of the rectangular lifting platform. According to the invention, the detection precision of the folding and unfolding time is improved, the labor intensity of workers is reduced, and the experimental data more fitting the reality is obtained.

The invention belongs to the technical field of impact fatigue test design of common materials of arresting hooks of carrier-based aircrafts, and particularly relates to a clamp for impact tensile fatigue test of a metal material, which comprises a hammer, a connecting cover plate, a support, a base, a lower chuck, a pressing strip, an adjusting screw and a locking nut. The connection matching mode of the hammer and the base can be changed according to the specific interface information of a testing machine. A lower clamping section of a test piece is connected with the lower chuck through threads, and then is locked by applying pretightening force through an upper nut and a lower nut. An upper clamping section of the test piece penetrates through a central round hole in the top of the support and is locked and fixed by applying pretightening force through double nuts. The cover plate is connected to two anvil blocks of the lower chuck, so that the contact area during impact is increased, and the stress balance of the two anvil blocks is facilitated. The support is fixed on the base through the pressing strip and the adjusting screw. During a test, an impact load applied by the hammer is transmitted to the test piece through the lower chuck, so that impact tensile loading of the test piece is realized. According to the method, the whole set of design and machining is completed, the clamp is successfully applied to a titanium alloy impact tensile fatigue test, and a foundation is laid for fatigue life evaluation of engineering structures bearing the impact load effect in the service process of a shipboard aircraft arresting hook, an ejection rod and the like.

The invention belongs to the field of flight tests, is used for testing height and speed parameters of testers, and in particular, relates to a height and speed testing structure and a method based on minimum gas path delay. In the flight tests, the height and speed testing of the testers is finished through measuring atmosphere dynamic and static pressure; and a gas path of about 20 m for leading static pressure and total pressure from a front branch air speed pipe to a testing point in a cabin exists to generate delay of about 0.1 s for height and speed testing. The space in the front branch air speed pipe is limited, so that the record of setting a testing sensor in the front branch air speed pipe is none before. The invention aims to provide the height and speed testing structure and method based on minimum gas path delay for realizing accurate real-time acquisition of height and speed parameters in the flight tests. Compared with a traditional front branch rod height and speed testing technology, the height and speed testing technology based on minimum gas path delay has the advantages of: reducing the gas path delay and removing the gas way blockage phenomenon.

The invention discloses a tension experiment mechanism for an aviation tray mooring ring. The tension experiment mechanism comprises a mounting platform, a clamping rod and a mooring ring tension sleeve, wherein the mounting platform is installed at a set position through the clamping rod, the mooring ring is installed on the upper surface of the mounting platform, and the plane where the mooringring is located is perpendicular to the mounting platform all the time; a rotatable mounting claw is arranged on the mooring ring, a mounting hole is formed in the mounting claw, and the angle of themounting position of the mooring ring can be adjusted by adjusting the included angle between the hole axis of the mounting hole and the plane where the mooring ring is located; the mooring ring tension sleeve is of a U-shaped structure, coaxial threaded through holes are formed in the two opposite sides of the open end of the mooring ring tension sleeve, the mooring ring is installed in the mooring ring tension sleeve through a rotating piece, and an external tension tester carries out a tension test on the mooring ring through the mooring ring tension sleeve. The mooring ring tension sleeveand the clamping rod are kept on the same vertical line.

The invention belongs to the technical field of aircraft strength tests, and particularly relates to an aircraft strength test debugging method and system. The method comprises the following steps: S1, acquiring a test debugging task of a server, wherein the test debugging task comprises a debugging manual, debugging contents and matters needing attention for debugging the aircraft strength test equipment; S2, projecting the test debugging task to a test machine or a test facility to form an information virtual image; S3, reading a two-dimensional code label preset on the surface of each test device, and uploading test device information carried by the two-dimensional code label to the server; S4, collecting real-time picture information of test equipment debugging performed by a user wearing the debugging terminal, uploading the real-time picture information to the server, and sharing the real-time picture information with remote equipment control personnel at the same time; and S5, obtaining a debugging picture of a remote equipment controller. According to the invention, the digital recording of the test equipment fault in the whole aircraft strength test debugging is realized, and the debugging efficiency is improved.

A low-temperature multifunctional high-precision aerospace bearing material friction performance simulation testing machine comprises a loading system, a friction force measuring system, a ball head mounting device, a rotary displacement device, a test piece mounting device, a low-temperature cooling device and a jet oil supply lubricating system. The loading system provides upward jacking vertical loading force, the friction force measuring system is arranged above the loading system, the rotary displacement device and the friction force measuring system are arranged side by side, and the rotary displacement device comprises a horizontal displacement mechanism and a vertical rotating mechanism arranged on the horizontal displacement mechanism; the test piece mounting device is fixedly arranged above the rotary displacement device, and a low-temperature cooling pipe outlet of the low-temperature cooling device corresponds to the oil receiving disc and is used for cooling a test environment; and an outlet of an oil supply pipe of the jet oil supply lubricating system is arranged corresponding to a contact surface where a bearing material is in contact for oil injection lubrication. The testing machine is compact in structure, diverse in function and capable of accurately measuring friction torque and abrasion performance under the two states.

The invention belongs to the technical field of airplane complete machine strength tests, and discloses a test data three-dimensional display method based on data fusion. The method comprises the following steps: S1, acquiring test data, wherein the test data comprises test loading data and test machine deformation data; s2, establishing a three-dimensional model of the testing machine; s3, displaying the test data at the corresponding position of the test three-dimensional model in real time, so that the accuracy, stability and symmetry analysis efficiency of test force feedback, pressure feedback and displacement feedback data can be greatly improved; and meanwhile, the visual display and playback capabilities of test force feedback, pressure feedback, displacement feedback data and testoperation information such as row number, load and rising and falling number can be greatly improved.

A wing model for static aeroelasticity wind tunnel test belongs to the technical field of aeroelasticity tests. In the wing model, the model steel joint and the spar frame are connected with the composite material skin. The piezometer wing ribs are arranged among the spar frame and the plurality of supporting wing ribs. The embedded piezometer tubes are arranged in the piezometer wing ribs, the lightweight filling foam is arranged among the spar frame and the plurality of supporting wing ribs. An outer surface of a frame segment formed by the lightweight filling foam, the plurality of supporting wing ribs, the piezometer wing ribs and the spar frame is covered with the composite material skin. The frame segment is assembled on the model steel joint to form the wing model for the static aeroelasticity wind tunnel test.

The invention relates to the technical field of flight tests for testing control surface parameters of testers, in particular to a non-contact control surface angle calibration method. The aircraft control surface calibration of all domestic aircraft manufacturers adopts a contact measuring mode of a special protractor; and the special protractor has no generality, is large in size and troublesome in operation, and is highly influenced by human factors in calibration results. The invention aims to provide the non-contact control surface angle calibration method for realizing online field calibration of flight test control surface angles. Compared with a traditional contact control surface calibration technology, the non-contact control surface calibration technology has the following advantages: calibration staff needs no aerial operation; and the calibration work is high in automation degree and efficiency.

The invention provides an undercarriage testing plane mounting support tool. The undercarriage testing plane mounting support tool comprises a testing plane-shaped frame upper beam, a test plane-shaped frame bottom seat, a testing plane-shaped frame left stand column, a testing plane-shaped frame right stand column, a main pivoting support A, a knuckle bearing A, a main pivoting support B, a knuckle bearing B, an undercarriage lock support, a retraction and release actuation cylinder support A, a copper sleeve A, a retraction and release actuation cylinder support B, a copper sleeve B, a mainpivoting support installing core shaft, a main pivoting support installing tool, a retraction and release actuation cylinder support installing core shaft, a retraction and release actuation cylindersupport installing tool A, a retraction and release actuation cylinder support installing tool B, a reference block A, a reference strip A, a reference strip B and a reference block B. The undercarriage testing plane mounting support tool can realize precise installation of an undercarriage testing plane, and substantially reduces labor intensity of workers.

The invention discloses a testing machine for researching the rubber ablation behavior of a tire tread at the moment of aircraft landing. The testing machine comprises a test bed, a main transmissionsystem, a friction system, a friction surface adjusting system, a radial impact system, a temperature control system, a visualization system and a signal acquisition and detection system. The workingcondition of the rubber grinding wheel in the friction system in the friction ablation test is changed through the main transmission system, the radial impact system, the friction surface adjusting system and the temperature control system, and the friction surface adjusting system simulates different pavement conditions contacted with the tire tread when an airplane touches the ground; real-timedata acquisition and monitoring are carried out on the friction ablation condition of the rubber tire tread in the friction system through the signal acquisition and detection system, and the data istransmitted to the computer through the data acquisition card for data processing and analysis so that the rubber friction ablation behavior test of the tire tread at the moment of aircraft landing iscarried out. The invention provides an effective test platform for researching the friction ablation behavior of the rubber tire under the instantaneous high-speed impact load.

The invention discloses an environmental test cabin comprising a hollow cylindrical test cabin body; the test cabin body is arranged on a mounting foundation in the horizontal direction, and the wall surface of the test cabin body is provided with an inlet and outlet for a to-be-tested machine for test and test equipment to enter and exit from the test cabin body. A sliding cabin door for sealing the inlet and outlet is arranged at the inlet and outlet, is connected to the test cabin body and is horizontally arranged in a sliding manner along the axial direction of the test cabin body so as to open or close the test cabin body. According to the sliding type cabin door, the design size of the sliding type cabin door can be effectively reduced, and then the construction cost of an environment test cabin is greatly reduced; the environmental test cabin is simple in overall structure and low in manufacturing cost, the driving force only needs to overcome the friction force generated when the sliding cabin door slides, and the driving force is small, so that the driving mechanism which is simple in structure, low in manufacturing cost and simple in driving operation can be selected to drive the sliding cabin door, and the defects of complex structure, large size, high cost, oil leakage and the like caused by hydraulic drive are overcome.