3650 results about "Landing gear" patented technology

An electric tilt rotor aircraft and a control system thereof. The aircraft comprises a fuselage (1), wings (2), horizontal stabilizers (3) and a landing gear system (4), wherein the fuselage (1) comprises a fuselage body and a rudder (10) for controlling and adjusting a flight direction of the aircraft; the wings (2) each comprise an integral wing box section (21), an aileron (22), a wing left / right nacelle (23) and a supporting rod left / right nacelle (24); left and right supporting rods are integrated on front sides of the integral wing box sections (21); the wing left / right nacelle (23) is connected to the integral wing box section (21) by means of a tilting shaft (25), and the wing left / right nacelle (23) rotates around the tilting shaft (25); the horizontal stabilizers (3) each comprise a horizontal stabilizer box section (31) and a horizontal stabilizer left / right nacelle (32); the horizontal stabilizer left / right nacelle (32) is connected to the horizontal stabilizer box section (31) by means of the tilting shaft (25), and the horizontal stabilizer left / right nacelle (32) rotates around the tilting shaft (25); the landing gear system (4) uses a tricycle landing gear system, and comprises a nose landing gear (41) and two main landing gears (42); brake systems are integrated in the main landing gears (42); and the nose landing gear (41) plays a role in balancing and controlling direction. By using the electric tilt rotor aircraft, urban air traffic can be realized.

Systems and methods related to landing unmanned aerial vehicles (UAVs) are provided. In one example, a method includes receiving a UAV on a surface of a landing platform. The method may further include operating a positioning device disposed under the surface to locate the UAV. The method may further include operating the positioning device to move the UAV to a location and / or an orientation on the surface. The UAV may comprise landing gear having a plurality of legs, where each leg comprises a shock absorption system. The method may further include operating the shock absorption system during the receiving operation to reduce force received at stress areas of the UAV, and after the receiving operation, operating the shock absorption system to dampen movement by the UAV. Related devices and systems are also provided.

Systems and methods are provided for displaying roll rate on a vertical take-off and landing (VTOL) aircraft display. Aircraft roll data is received from at least one geospatial sensor of the VTOL aircraft. A roll rate scale associated with a desirable roll rate range of the VTOL aircraft is generated. An aircraft roll rate of the VTOL aircraft is determined based on the aircraft roll data. The VTOL aircraft display, including a graphical representation of the roll rate scale and the aircraft roll rate with respect to the roll rate scale, is generated for display on the display device.

A craft is provided comprising: a propulsion system; a fuselage coupled with the propulsion system, wherein the fuselage is configured to carry a payload and fuel for the propulsion system; and a pre-assembled wing coupled with the fuselage; wherein the craft is designed to take-off without using a runway and does not have landing gear to land.

The invention relates to an aircraft nose landing gear taxiing and shimmy simulation test bed. The test bed comprises a test device main body structure, a power execution control mechanism, a vibration excitation mechanism and a vibration detection system. The method solves the problems of determination, contrast evaluation and detection analysis of the shimmy influence factors of the nose landing gear of the aircraft. The torsion steering engine is added, so that the turning condition of a ground plane can be simulated. The height-adjustable supporting cross beam can adapt to undercarriages of different sizes only by replacing the undercarriage clamp instead of replacing the whole frame, so that the use cost is greatly reduced. And more accurate data are provided for the design of the nose landing gear of the aircraft.

A ground detection system for an aircraft based on capacitive sensing. The system comprises an electrode mounted to an aircraft landing gear. As the aircraft landing gear approaches the ground, the electrode and the ground together form a variable capacitor having a capacitance that depends on a distance of the electrode from the ground. Signal processing electronics are used to measure the capacitance and, on the basis of the measured capacitance, determine a distance of the aircraft from the ground.

An unmanned aerial vehicle includes a body. The body has an upper side and a lower side opposite the upper side. At least one rotor is coupled to the body to generate lift in an upward direction. A mechanism is mounted to the lower side of the body. The mechanism includes a landing gear, a storage system, a robotic arm housing, a counterweight housing, a robotic arm, an actuator, and an end effector. The actuator is coupled to the robotic arm to move the robotic arm between a retracted position and an extended position. The end effector is coupled to the distal end of the robotic arm. The end effector includes a gripper with a pair of opposing compressible elements for gripping an object between the pair of compressible elements.

The invention provides an aircraft landing gear dismounting and mounting device, belongs to the technical field of aviation equipment, and solves the problem that in the prior art, landing gear dismounting and mounting equipment is difficult to meet the requirements for efficient, accurate and flexible dismounting and mounting in the modern aviation maintenance field. Comprising a frame, a moving frame is arranged on the inner side of the frame, and the moving frame is driven by a first driving device; a moving plate is arranged on the frame and is driven by a second driving device; an arc-shaped guide rail is arranged on the movable plate, a supporting frame is arranged on the arc-shaped guide rail, a lifting device and a fixing device are arranged on the supporting frame, and a clamp is arranged on the top of the lifting device. According to the undercarriage disassembling and assembling device, through cooperative operation of the first driving device, the second driving device and the lifting device, all-directional flexible adjustment of the tool on the X axis, the Y axis and the Z axis is achieved. Due to the design of the arc-shaped guide rail, the supporting frame can flexibly rotate within a certain angle range, and the adaptability and flexibility of the device are further enhanced.

A glider system capable of sustained long-term flight includes a fixed-wing glider having a dual fuselage system. The top surface of the glider is covered with a plurality of photovoltaic cells capable of powering onboard sensors, propeller systems, and / or onboard processers. The glider includes wing spars with a composite material covering an inner foam core, with the type of foam core or use of reinforcement materials differs along the length or width of the wing spars. The glider is able to downward facing sacrificial winglets to provide a landing system without the need for bulkier landing gear.

An undercarriage, a middle large rotor and a middle large motor are connected to the lower portion of a fuselage body and sequentially connected to the top of a middle small tower, and the right front portion, the left front portion, the left rear portion and the right rear portion of the fuselage body are correspondingly connected with a right front fuselage arm, a left front fuselage arm, a left rear fuselage arm, a right rear fuselage arm and a power assembly. The connecting line of the rotating centers of the rotors of the four power assemblies is square or rectangular, the rotating directions of the rotors of the power assemblies at the same opposite angles are the same, the rotating directions of the rotors of the power assemblies at different opposite angles are opposite, and the rotating faces of the two rotors opposite to the middle large rotor in rotating direction incline around the axes of the respective arms. The included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle beta, the rotating surfaces of the two rotors with the same rotating direction as the middle large rotor incline around respective arm axes, the included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle theta, beta is larger than theta, the reaction torque and the horizontal component torque of the rotors jointly drive the aircraft to steer, and the aircraft has the advantages of being long in endurance time, large in load and high in wind resistance. And a universal flight platform is formed.

A method and system for mounting an axle in a bore of a bogie mounted on a landing gear of an aircraft, the landing gear being in a gear-down position, the bore extending in a substantially horizontal direction referred to as the axle installation direction, between a first open end and a second open end are disclosed. The method includes a) supplying a mobile bearing structure (30) comprising means for moving the axle (12) which are able to move along a direction substantially parallel to the axle installation direction, b) installing the axle (12) on the means for moving the axle, c) aligning the axis (C) of the axle (12) with the axis (B) of the bore (14), and d) inserting the axle (12) into said bore (14) of the bogie (16) by moving the means for moving the axle (12) along the direction substantially parallel to the axle installation direction.

The invention discloses a rocket-boosted warplane ultra-short-distance launching take-off system and method. The system comprises a chassis, a high-precision track, a hydraulic lifting mechanism and a rocket sled. The rear end of the high-precision track is connected with the top of the chassis in a vertically swinging mode, the front end of the high-precision track is driven by a hydraulic lifting mechanism to ascend and descend, and the rear portion of the high-precision track is provided with a check block. The bottom of the rocket sled is provided with a sliding block used for being in sliding fit with the high-precision track. The mounting interface is used for being connected with a warplane undercarriage through an electronic lock, and the electronic lock is used for timing when a rocket engine is started, unlocking when the rocket engine reaches the working time and sending a separation success signal after being separated from a warplane. Small-overload ultra-short-distance take-off of the fighter is achieved, the volume of ground take-off facilities of the fighter is effectively reduced, the requirement for ground conditions is lowered, the flexibility and convenience of deployment are effectively improved, and the structure or a power system of the fighter is not greatly changed.

The utility model discloses a landing lever locking device and an unmanned aerial vehicle. The landing lever locking device comprises an unmanned aerial vehicle body. The foot stool assembly comprises a sliding rail seat, a lifting rod and a locking device, and the lifting rod is connected to the sliding rail seat in a sliding mode. The lock catch device comprises a lock box, a sliding block and a driver, the lock box is provided with a vertically-through box lock hole, the sliding block can be connected into the lock cavity in a left-right sliding mode, the sliding block is provided with a vertically-through lock catch hole, the lock catch hole comprises a sliding lock hole and a fastening hole communicated with the sliding lock hole, at least one clamping section is arranged on the side wall of the fastening hole, and the output end of the driver is connected to the sliding block. And the inner wall of the fastening hole is matched with the outer wall of the lifting rod. The lock box is arranged on the unmanned aerial vehicle body, the rising and falling rod is slidably connected to the lock box lock hole and the sliding lock hole, the inner wall of the fastening hole is matched with the outer wall of the rising and falling rod, and the sliding block can be inserted into the rising and falling rod in the radial direction under the driving condition of the driver and is inserted into and clamped to the outer wall of the rising and falling rod through the clamping section. The purpose of clamping the lifting rod can be achieved only by driving the sliding block to move rightwards through the driver.

The present invention discloses a load mounting device for an unmanned aerial vehicle, comprising a fuselage, two landing gears are arranged at the bottom of the fuselage, a fixed interface fixedly connected to the fuselage is arranged between the two landing gears; a fixed frame is arranged at the bottom of the fuselage, a connection interface is fixedly connected to the top of the fixed frame; a suspension rope is used to fix the load, the bottom end of the suspension rope is connected to the load, and the top end of the suspension rope is connected to the connection interface on the fixed frame through a movable interface. The present invention performs suspension fixing on the load by providing a fixed frame and a suspension rope, so that the center of gravity of different loads is always on the same vertical line, which does not affect the balance of the unmanned aerial vehicle, and is suitable for large-sized goods, especially some special-shaped goods. At the same time, through the automatic connection and separation of the fixed interface and the connection interface, no manual auxiliary operation is required, and the fixed interface and the connection interface can be separated only when the fuselage lands.

The invention discloses an MBSE-based aircraft undercarriage system model integration method, and belongs to the technical field of system engineering. The method comprises the following steps: constructing an aircraft undercarriage system architecture, a classification mapping of simulation design information and a GOPPRR-E element model, and a meta-model library; constructing an architecture model according to a demand-function-logic-physical process, and generating a Simulink simulation model according to the architecture; a simulation model semantic mapping rule is constructed according to an OSLC specification, and design information is analyzed to establish an integrated data model; and a Simulink simulation tool integration adapter is constructed, and a tool interoperation service based on Web is provided. According to the method, a multidisciplinary model can be accurately constructed, uniform expression of architecture and simulation models and semantic consistency of heterogeneous models are realized, two-way transmission of cross-domain information and interoperation between tools are supported, the problem of interoperation between architecture modeling and simulation modeling tools in aircraft undercarriage system design is effectively solved, and the design efficiency is improved.

The invention belongs to the field of aviation technology undercarriage design, and relates to an energy conversion valve used for an undercarriage folding and unfolding system. When the main system works normally, the pilot electromagnetic valve in the energy conversion valve is not electrified, the energy conversion valve is located at the working position of the main system, at the moment, the load subsystem communicates with the main system and isolates the standby system, and the landing gear cabin door / landing gear actuator cylinder is normally retracted and released; when the main system breaks down, the standby system supplies pressure, a pilot electromagnetic valve of the energy conversion valve is powered on, the load subsystem communicates with the standby system, the main system is isolated, and an undercarriage cabin door / undercarriage actuator cylinder is released emergently. The two positions of the valve element are reliably isolated, and a single function fault is not leaked to another system.

A shock strut assembly is disclosed herein. The shock strut assembly includes a mixed fluid chamber, a liquid chamber coupled to the mixed fluid chamber, a metering pin having a first end and a second end and a channel formed therein extending from the first end to the second end, the first end disposed in the liquid chamber and the second disposed in the mixed fluid chamber, and a pressure relief valve fluidly coupled to the liquid chamber and the channel in the metering pin, the pressure relief valve configured to allow a flow of a fluid from the liquid chamber to the channel in the metering pin.

The invention relates to a test device and method for testing the performance of an elevator speed governor, and the device comprises a test rack which is internally provided with a lifting channel for testing; the main rope is hung and wound on the traction machine; the undercarriage is arranged on one side of the traction machine and connected to the main rope, and the car frame is arranged below the undercarriage and connected with the undercarriage through a detacher; the counterweight frame is arranged on the other side of the traction machine and is connected to the main rope; the speed limiter is arranged at the top of the test rack; the rope gripper is arranged at the top of the test rack; and the data acquisition unit is used for acquiring the running speed of the lift car frame, the wheel speed of the speed limiter, the lifting force of the speed limiting rope, an electrical action signal and a mechanical action signal. The performance of the speed governor can be tested under the working conditions that the lift car frame freely falls when running at a rated speed and the lift car frame falls downwards at different accelerations in a static state.

The invention discloses an aircraft landing gear drop test device and method, and relates to the technical field of aircraft landing test devices.The aircraft landing gear drop test device comprises a platform, and a supporting structure is arranged on the platform; the rack is movably connected with the supporting structure so as to ascend and descend relative to the supporting structure. Moreover, a mounting seat is arranged on the rack so as to be connected with the undercarriage; the connecting seat is arranged on the platform and is arranged below the rack; the test box body is detachably connected to the connecting base, and the test box body is provided with a first cavity with an upward opening so as to contain a pavement material. According to the undercarriage drop test device and the undercarriage drop test method, different pavement materials can be loaded in the first cavity of the test box body, so that different pavement structures are formed, and the undercarriage drop test device can perform simulation experiments on various pavement working conditions.

The invention relates to a tilting rotor unmanned aerial vehicle, in particular to an unmanned aerial vehicle combining a flying wing type structure and a multi-shaft tilting rotor. A completely fused flying wing type fuselage is formed by a low-speed wing shape, and no other fuselage characteristics exist except that bilaterally symmetrical wings are directly connected and fused in the middle; a left tilting rotor frame and a right tilting rotor frame are mounted at the left end and the right end of the flying wing type fuselage respectively; at least one to at most six brackets extend from the front part and the rear part respectively, and symmetrical tilting rotor frames are mounted on the brackets; motors and paddles are mounted on all the tilting rotor frames; and an undercarriage is mounted below the flying wing type fuselage. During tilting, a single motor and a single blade are not tilted, but a plurality of power units formed by one or more motors and blades are tilted in groups through a tilting rotor frame. Therefore, the structure is simplified, and the safety and the maneuverability are improved. The landing gear has two types, one type is a foot type landing gear suitable for vertical take-off and landing in a soft site, and the other type is a wheel type landing gear suitable for short-track take-off and landing in a hardened site.

The invention discloses a method for simulating bottom bending tensile stress of an airport rigid pavement slab under the action of aircraft wheel loads. The method comprises the following steps: establishing an airport rigid pavement three-dimensional refined finite element model and an airport rigid pavement two-dimensional simplified finite element model; selecting aircraft wheel sliding load of the aircraft type main landing gear; obtaining an equivalent airplane wheel sliding load; obtaining an equivalent airplane wheel sliding load calculation formula under the random transverse action position of the airplane wheel; and obtaining the bottom bending-tensile stress value of the pavement slab under the airplane wheel sliding load. According to the method, by establishing the rigid pavement two-dimensional simplified finite element analysis model, rapid prediction of the airport rigid pavement slab bottom stress response under the aircraft wheel load effect is achieved. According to the method, the bending-tensile stress of the slab bottom is quickly and accurately calculated according to different pavement structure parameters and the transverse action position of the airplane wheel. The method effectively overcomes the defects of complex modeling, high calculation cost and the like of traditional rigid pavement refined simulation, has the characteristics of simplicity and convenience in modeling and quickness in calculation, and remarkably improves the airport engineering design and analysis efficiency.

The utility model discloses a road crack patrol unmanned aerial vehicle capable of spraying marks, and relates to the technical field of unmanned aerial vehicles. The utility model provides a road crack inspection unmanned aerial vehicle capable of spraying marks, which comprises an unmanned aerial vehicle frame, four rotating rotors which are distributed at equal intervals are rotatably connected to the top end of the unmanned aerial vehicle frame, undercarriages which are transversely symmetrical are fixedly connected to the bottom end of the unmanned aerial vehicle frame, and the road crack inspection unmanned aerial vehicle further comprises a mounting frame which is fixedly connected to the bottom end of the unmanned aerial vehicle frame. The bottom end of the mounting frame is rotatably connected with a spraying cylinder, the bottom end of the unmanned aerial vehicle frame is fixedly connected with a storage cylinder, the storage cylinder communicates with the spraying cylinder through a hose, the bottom end of the mounting frame is fixedly connected with a driving motor, and an output shaft of the driving motor is connected with the spraying cylinder. Rotor wings are rotated to drive the device to fly to inspect cracks on a road, when the cracks are detected, the spraying barrel rotates downwards to adjust the angle, then paint is sprayed to the cracks for marking, and accurate guidance is provided for follow-up maintenance work.

The invention provides a main landing gear cabin door mechanism, an aircraft and a folding and unfolding method. The main undercarriage cabin door mechanism comprises a left cabin door and a right cabin door, one end of the left cabin door is provided with a first retractable driving mechanism, one end of the right cabin door is provided with a second retractable driving mechanism, and a small cabin door is hinged to the side, close to the right cabin door, of the left cabin door; the small cabin door and the left cabin door are provided with third folding and unfolding driving mechanisms which can respectively drive the small cabin door and the left cabin door to rotate; in the first position, the left cabin door and the right cabin door are both in a vertical state, and the small cabin door and the left cabin door are folded and overlapped; and at the second position, the left cabin door, the small cabin door and the right cabin door are horizontally positioned on the same plane. The small cabin door can be independently retracted and folded, after the cabin door is opened, the folded small cabin door can reduce the windward area of the cabin door in the put-down state, the pneumatic load borne by the cabin door is reduced, and the design of a driving mechanism for retracting and releasing the cabin door is lighter.

The invention discloses an undercarriage retracting and releasing mechanism wear prediction method based on digital twinning, and the method comprises the steps: firstly building an undercarriage three-dimensional model containing geometric dimensions and material characteristics, building a multi-body dynamics simulation model, and simulating the stress and wear process of a pin shaft under the actual working condition; and adopting an Archard wear calculation model considering key parameters to calculate the wear of the pin shaft in the simulation model in real time. On the basis, uncertainty analysis of parameters in Archard is carried out, the parameters are considered to obey normal distribution, and a prediction model under multivariate uncertainty is established by setting reasonable mean values and standard deviations. And finally, based on the digital twin model, by using a particle filtering method, initializing a particle swarm, changing a particle weight by virtue of a state and an observation equation, re-sampling to approach a real state, iteratively updating a curve, calibrating model parameters and a state, improving prediction precision, improving prediction accuracy and reliability, discovering a potential wear problem in advance, and providing a basis for a maintenance plan.

The invention belongs to the technical field of alloy steel, and relates to composite reinforced high-toughness steel and a preparation method thereof. The steel comprises the following chemical components in percentage by weight: 6.0%-9.4% of Ni, 7.6%-8.9% of Co, 0.8%-3.8% of Cr, 1.0%-2.6% of Mo, 0-0.9% of W, 0.05%-0.60% of V, 0-0.10% of Nb and the balance of Fe. C and Al are added to serve as strengthening elements, and the balance is Fe and impurity elements. When 0.14% < = C < 0.20%, 1.2% < = Al < = 1.6%, 1.0% < = Mo + W / 2 < = 2.0%; when 0.20% < = C < = 0.35%, 0.6% < = Al < = 1.3%, and 2.0% < = Mo + W / 2 < = 3.0%. The alloy is subjected to vacuum induction melting or multi-time melting, homogenizing diffusion annealing and forging rolling to form a material, so that the tensile strength of the steel reaches 2200-2350 MPa or above, the yield strength reaches 2000-2150 MPa or above, and the steel has excellent strength, plasticity and toughness matching and can be used for components such as engine shafts, aircraft landing gears, high-performance shells, frame beams and molds.

The aircraft landing gear part supporting device comprises a first supporting structure and a second supporting structure, an aircraft landing gear part comprises a landing gear body and a central wing, the first supporting structure is connected with the central wing, and the second supporting structure is connected with the landing gear body. The first supporting structure and the second supporting structure jointly support the aircraft landing gear part, so that the maintenance space between the landing gear part and the ground is increased, the aircraft landing gear part supporting device can support the aircraft landing gear part of a general-purpose aircraft, the maintenance space between the landing gear part and the ground can be increased, and the service life of the aircraft landing gear part is prolonged. And maintenance training is facilitated.

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

The utility model discloses an undercarriage drop test stand, which comprises a vertical track support, a vertical pulley assembly is arranged on the vertical track support, the vertical pulley assembly moves up and down along the vertical track support, and a cantilever pulley track assembly is fixedly connected to the vertical pulley assembly. The cantilever pulley track assembly moves up and down along the vertical track support along with the vertical pulley assembly, a horizontal pulley assembly is connected to the cantilever pulley track assembly in a sliding mode, an undercarriage mounting base is arranged on the lower side of the horizontal pulley assembly, and the undercarriage mounting base directly faces an inclined plane base plate arranged on the ground. According to the invention, impact generated by vertical falling of the undercarriage can be simulated, horizontal displacement generated by falling contact with a slope can also be simulated, various simulation test environments are provided, and the damage degree of the undercarriage and the stability of a connection area of the undercarriage and an aircraft main body are judged according to recorded test data.

The invention provides a broadband sound absorption structure suitable for undercarriage cabin door noise reduction and a design method, and belongs to the field of aircraft noise reduction. The sound absorption structure comprises a plurality of perforated plate type Helmholtz cavity unit cells and a plurality of embedded pipe type Helmholtz cavity unit cells. In the perforated plate type unit cells, one or more first through holes are formed in the perforated plate. And in the embedded pipe type unit cells, the embedded pipes penetrate through the perforated holes in the perforated plates and extend into the cavities. Cavities of the two single cells face the same direction, and the perforated plate forms a curved surface conformal with an undercarriage cabin door. The effective length of the neck of the perforated plate is increased by using the embedded pipe, the low-frequency sound absorption performance is improved, the sound absorption performance and the sound absorption bandwidth of the middle-frequency band are improved by using the perforated plate type Helmholtz cavity, meanwhile, the curved surface conformal design is highly matched with the special-shaped cabin door structure, and the sound absorption performance is improved. The low-frequency and medium-frequency sound absorption performance can be guaranteed, and meanwhile the light and thin requirements of an aircraft can be met; and a pure passive mechanical structure is adopted, so that the control complexity and stability risks of an active noise reduction system can be avoided.

The invention discloses an aircraft front wheel electrical zero calibration method, device and equipment and a medium, relates to the technical field of aircraft front wheel control, is applied to ground comprehensive maintenance equipment, and comprises the steps that a target angle instruction is used for controlling a scale line of a rotating sleeve in a nose landing gear to be aligned with a scale line of a front supporting column outer cylinder; a front wheel shimmy damping instruction is sent to a front wheel operation control box of the aircraft, so that the front wheel operation control box controls front wheels to keep stable; the current angle information of the front wheel angle sensor is determined as current electrical zero position reference information; the front wheel and the rotating sleeve are controlled to rotate through a front wheel rotating instruction, and if the deflection angle of the scale mark of the rotating sleeve is matched with the front wheel rotating instruction, it is judged that front wheel electrical zero calibration of the aircraft is completed. And the accuracy and simplicity of the electrical zero calibration of the airplane front wheel are improved.