Undercarriage capable of triggering wheel load signal and state management system

Abstract

The invention provides an undercarriage capable of triggering a wheel load signal and a state management system, and belongs to the technical field of undercarriage design, the undercarriage comprises a front undercarriage and two main undercarriages, the front undercarriage comprises a first outer cylinder, an inner cylinder, a first piston rod and a first triggering mechanism, the first outer cylinder is fixedly connected with a fuselage, and the inner cylinder is fixedly connected with the fuselage; the inner cylinder moves in the first outer cylinder in the vertical direction, the first piston rod transmits ground loads to the inner cylinder and the first outer cylinder, and the first triggering mechanism is installed on the upper portion of the nose landing gear. Each main undercarriage comprises a second outer cylinder, a second piston rod and a second trigger mechanism, the second outer cylinders are fixedly connected with the fuselage, the second piston rods move in the second outer cylinders in the vertical direction, oil cavities are formed between the second outer cylinders and the second piston rods, and the second trigger mechanisms are installed on the upper portions of the main undercarriages. The wheel load signal microswitch is higher in precision, higher in sensitivity, higher in reliability and higher in designability.

Description

Technical Field

[0001] The invention belongs to the technical field of landing gear design, and in particular relates to a landing gear and a state management system for triggering a wheel load signal. Background Art

[0002] The aircraft / helicopter ground signal is an important signal during the flight. Generally speaking, the landing gear system is the only part of the entire aircraft that is in contact with the ground. The wheel-borne signal is an important signal source during the flight of the aircraft / helicopter. The flight control system, avionics system and other subsystems determine the current position of the aircraft by collecting wheel-borne signals. How to accurately collect wheel-borne signals and how to design the triggering method of the wheel-borne signal micro switch are important contents in the design of the landing gear.

[0003] The wheel-load signal micro switch of the traditional landing gear is generally installed on the anti-torsion arm, such as Figure 1 As shown in the figure, when the aircraft lands, the buffer overcomes the oil damping force and the air spring force, the anti-torsion arm rotates relatively, the wheel-loaded signal micro switch contacts or disengages from the trigger plate, and the circuit is turned on and off to determine whether the landing gear is grounded. The disadvantage of this method is that the landing gear needs to overcome a large force, the anti-torsion arm needs to have a large angle change, and the corresponding buffer has a large compression amount to trigger the wheel-loaded signal micro switch, and the sensitivity is poor. In addition, the wheel-loaded signal micro switch is often installed on the outside of the fuselage and is easily affected by the external environment. During the operation of the landing gear, the sensor and the anti-torsion arm rotate relative to each other, and long-term use brings reliability problems such as wear.

[0004] The prior art method of setting the wheel load signal on the buffer support is shown in Figure 2 A complex mechanism was designed to achieve this function. First, the vertical displacement was converted to the horizontal direction through the trigger mechanism, and then the horizontal displacement was transferred to the vertical direction through the rocker arm. The disadvantage is that the structure is complex and difficult to implement, and it is not suitable for installing multiple micro switches. Summary of the invention

[0005] In order to solve the problem that the wheel-borne signal is set on the buffer pillar in the prior art, the structure is complex and difficult to implement, and it is not suitable for installing multiple micro switches, the present invention provides a landing gear and state management system for triggering the wheel-borne signal. Compared with the existing trigger mechanism, the wheel-borne signal micro switch has higher accuracy, higher sensitivity, better reliability and stronger designability. The technical solution is as follows:

[0006] In a first aspect, a landing gear for triggering a wheel-borne signal is provided, the landing gear comprising a front landing gear and two main landing gears,

[0007] The front landing gear includes a first outer tube 61, an inner tube 8, a first piston rod 171, and a first trigger mechanism, wherein the first outer tube 61 is fixedly connected to the fuselage, the inner tube 8 moves inside the first outer tube 61 in a vertical direction, the first piston rod 171 transfers the ground load to the inner tube 8 and the first outer tube 61, and the first trigger mechanism is installed on the upper part of the front landing gear;

[0008] Each main landing gear includes a second outer tube 62, a second piston rod 172, and a second trigger mechanism, wherein the second outer tube 62 is fixedly connected to the fuselage, the second piston rod 172 moves inside the second outer tube 62 in a vertical direction, an oil chamber is formed between the second outer tube 62 and the second piston rod 172, and the second trigger mechanism is installed on the upper part of the main landing gear.

[0009] Optionally, the first trigger mechanism is composed of a first upper end cover 2, a first spring 3, a stop cover 4, a first positioning pin 5, a self-lubricating support sleeve 7, a fixed piston 9, a plug cap 10, a plunger 11, a locking nut 12, a sealing ring 13, a sealing ring 14, a self-lubricating support sleeve 15, a first micro switch 1, and a stop ring 25;

[0010] Among them, the first upper end cover 2 is fixedly connected to the first outer cylinder 61 through the first positioning pin 5, the first outer cylinder 61 is fixedly connected to the fuselage, a first spring 3 is arranged on the upper end of the blocking cover 4, and a limiting boss is arranged on the first upper end cover 2 as a travel protection mechanism of the first micro switch 1 to prevent the first micro switch 1 from being damaged due to overtravel, a self-lubricating bushing 7 and a self-lubricating support sleeve 15 are installed between the first outer cylinder 61 and the inner cylinder 8 to reduce the friction force when the inner cylinder 8 moves, a round nut 24 is installed below the self-lubricating bushing 15, the round nut 24 is threadedly connected to the first outer cylinder 61, and the round nut 24 is used to fix the self-lubricating bushing 15, and the round nut 2 4 is used as a limit block to prevent the blocking cover 4 from being squeezed and deformed and damaging the first micro switch 1. The fixed piston 9 and the sealing ring 14 are installed on the upper part of the inner cylinder 8. The plunger 11 and the locking nut 12 are used to fix the first upper end cover 2 to provide pre-pressure for the first spring 3. The retaining ring 25 is installed between the inner cylinder 8 and the first outer cylinder 61 to prevent the inner cylinder 8 from escaping from the first outer cylinder 61. With the support of the self-lubricating support sleeve 7 and the self-lubricating support sleeve 15, the inner cylinder 8 makes a small movement inside the first outer cylinder 61 along the vertical direction. The displacement S1 of the inner cylinder 8 is equal to the rated stroke of the first micro switch 1, which is usually 2mm to 5mm.

[0011] Optionally, for the front landing gear, when the aircraft / helicopter touches the ground, the ground load is sequentially transmitted to the first piston rod 171, the inner cylinder 8, and the stop cover 4 through the wheel tire, and the stop cover 4 overcomes the pre-pressure K1 of the first spring 3 to compress the first micro switch 1, and the internal circuit of the first micro switch 1 is connected, and the first micro switch 1 sends a grounding signal to the electromechanical management computer;

[0012] When the airplane / helicopter takes off, under the action of gravity and the first spring 3, the cover 4, the inner cylinder 8, and the first piston rod 171 move downward, the cover 4 disengages from the first microswitch 1, the internal circuit of the first microswitch 1 is disconnected, and the first microswitch 1 sends a take-off signal to the electromechanical management computer.

[0013] Optionally, the preload K1 of the first spring 3 satisfies: K1>n*Mg, n is the maximum downward overload coefficient during flight, and Mg is the landing gear's own gravity; when landing, when the ground load N is greater than nMg+Mg, the first micro switch 1 can be triggered.

[0014] Optionally, the second trigger mechanism is composed of a second upper end cover 52, a second spring 53, a micro switch actuating piece 19, a second positioning pin 55, a second outer cylinder 56, a throttle valve 16, a second piston rod 172, a floating piston 18, a plunger 20, a sealing ring 21, a clamping nut 22, a self-lubricating bushing 23, a locking nut 26, and a second micro switch 51;

[0015] Among them, the upper end cover 52 is fixedly connected to the second outer cylinder 56 through the second positioning pin 55, the second outer cylinder 56 is fixedly connected to the fuselage, the second upper end cover 52 is installed on the upper part of the second spring 53 to compress the second spring 53, the plunger 20 is installed at the upper opening of the second outer cylinder 56, the micro switch trigger piece 19 is fixed to the plunger 20 through the clamping nut 22, and moves with the plunger 20. The plunger 20 is provided with a boss to ensure that the upper and lower end surface areas are different; a self-lubricating bushing 23 is installed between the second piston rod 172 and the second outer cylinder 56 to reduce the friction force when the second piston rod 172 moves in the vertical direction. The plunger 20 makes a small movement in the vertical direction, and the displacement S2 of the plunger 20 is equal to the rated stroke of the second micro switch 51, usually 2mm~5mm.

[0016] Optionally, for the main landing gear, when the aircraft / helicopter touches the ground, the landing gear is compressed, the oil pressure increases, and since the upper and lower pressure areas of the plunger 20 are different, there is an upward combined force, which overcomes the friction force and the spring force, pushes the plunger 20, and the micro switch trigger piece 19 moves upward, the internal circuit of the second micro switch 51 is connected, and the second micro switch 51 sends a grounding signal to the electromechanical management computer;

[0017] When the airplane / helicopter takes off, the landing gear extends, the oil pressure disappears, and under the action of the preload K2 of the second spring 53 and the gravity mg, the plunger 20 moves downward, the microswitch trigger piece 19 disengages from the second microswitch 51, the internal circuit of the second microswitch 51 is disconnected, and the second microswitch 51 sends a take-off signal to the electromechanical management computer.

[0018] Optionally,

[0019] The preload K2 of the second spring 53 satisfies: Where n is the maximum downward overload coefficient during flight, N is the ground load, A is the internal cross-sectional area of ​​the second outer cylinder 56, u is the friction coefficient, m is the mass of the plunger 20, and f is the friction between the sealing ring 21 and the plunger 20; when landing, when the ground load When the second micro switch 51 is triggered.

[0020] Optionally, the first trigger mechanism includes two first micro switches 1, and the second trigger mechanism of each main landing gear includes two second micro switches 51.

[0021] The take-off signal / grounding signal of the aircraft / helicopter is executed by majority voting: if more than half of the micro switches 1 send the take-off signal to the electromechanical management computer, it is judged to be in the take-off state at this time; if more than half of the micro switches 1 send the grounding signal to the electromechanical management computer, it is judged to be in the grounding state at this time.

[0022] In a second aspect, a state management system is provided, comprising a landing gear for triggering a wheel-load signal as described in any one of the first aspects, and an electromechanical management computer.

[0023] The beneficial effects of the present invention are at least:

[0024] 1) The micro switch trigger stroke can be designed, the required landing stroke is short and the sensitivity is high;

[0025] 2) The micro switch is installed inside the landing gear and is not easily affected by natural environments such as humidity, heat, salt spray, sand and dust, which can increase the service life of the switch;

[0026] 3) The triggering form of the micro switch is linear rather than rolling, which can reduce the wear problem between the micro switch and the trigger plate;

[0027] 4) Multiple switches can be easily installed to improve the reliability of the grounding signal. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 Schematic diagram of triggering micro switch in anti-twist arm rotation mode;

[0029] Figure 2 Schematic diagram for converting vertical displacement to horizontal and then to vertical;

[0030] Figure 3 It is a schematic diagram of the front landing gear structure of the present invention;

[0031] Figure 4 It is a schematic diagram of the main landing gear structure of the present invention. DETAILED DESCRIPTION

[0032] The present invention is further described in detail below through specific implementation modes and drawings.

[0033] The present invention provides a landing gear and state management system for triggering wheel-load signals, which can realize convenient triggering of wheel-load signals, and its triggering method is simple and direct, and has the advantages of high precision, high sensitivity, high reliability, and not being easily affected by the external environment.

[0034] See also Figure 3 and Figure 4 An embodiment of the present invention provides a landing gear for triggering a wheel-borne signal, the landing gear comprising a front landing gear and two main landing gears,

[0035] The front landing gear includes a first outer tube 61, an inner tube 8, a first piston rod 171, and a first trigger mechanism, wherein the first outer tube 61 is fixedly connected to the fuselage, the inner tube 8 moves inside the first outer tube 61 in a vertical direction, the first piston rod 171 transfers the ground load to the inner tube 8 and the first outer tube 61, and the first trigger mechanism is installed on the upper part of the front landing gear;

[0036] Each main landing gear includes a second outer tube 62, a second piston rod 172, and a second trigger mechanism, wherein the second outer tube 62 is fixedly connected to the fuselage, the second piston rod 172 moves inside the second outer tube 62 in a vertical direction, an oil chamber is formed between the second outer tube 62 and the second piston rod 172, and the second trigger mechanism is installed on the upper part of the main landing gear.

[0037] The first trigger mechanism is composed of a first upper end cover 2, a first spring 3, a blocking cover 4, a first positioning pin 5, a self-lubricating support sleeve 7, a fixed piston 9, a plug cap 10, a plunger 11, a locking nut 12, a sealing ring 13, a sealing ring 14, a self-lubricating support sleeve 15, a first micro switch 1, and a blocking ring 25;

[0038] The first upper end cover 2 is fixedly connected to the first outer cylinder 61 through the first positioning pin 5, and the first outer cylinder 61 is fixedly connected to the fuselage. The first spring 3 is arranged on the upper end of the blocking cover 4. The first upper end cover 2 is provided with a limit boss as a travel protection mechanism for the first micro switch 1 to prevent the first micro switch 1 from being damaged due to overtravel. The self-lubricating bushing 7 and the self-lubricating support sleeve 15 are installed between the first outer cylinder 61 and the inner cylinder 8 to reduce the friction force when the inner cylinder 8 moves. A round nut 24 is installed below the self-lubricating bushing 15. The round nut 24 is threadedly connected to the first outer cylinder 61. The round nut 24 is used to fix the self-lubricating bushing 15. At the same time, the round nut 24 As a limit block, it prevents the blocking cover 4 from being squeezed and deformed, and crushing the first micro switch 1. The fixed piston 9 and the sealing ring 14 are installed on the upper part of the inner cylinder 8. The plunger 11 and the locking nut 12 are used to fix the first upper end cover 2 to provide pre-pressure for the first spring 3; the retaining ring 25 is installed between the inner cylinder 8 and the first outer cylinder 61 to prevent the inner cylinder 8 from escaping from the first outer cylinder 61; the inner cylinder 8, supported by the self-lubricating support sleeve 7 and the self-lubricating support sleeve 15, makes a small movement inside the first outer cylinder 61 in the vertical direction, and the displacement S1 of the inner cylinder 8 is equal to the rated stroke of the first micro switch 1, usually 2mm~5mm.

[0039] For the front landing gear, when the aircraft / helicopter touches the ground, the ground load is transmitted to the first piston rod 171, the inner tube 8, and the stop cover 4 in sequence through the wheel tire. The stop cover 4 overcomes the preload K1 of the first spring 3 to compress the first micro switch 1. The internal circuit of the first micro switch 1 is connected, and the first micro switch 1 sends a grounding signal to the electromechanical management computer;

[0040] When the airplane / helicopter takes off, under the action of gravity and the first spring 3, the cover 4, the inner cylinder 8, and the first piston rod 171 move downward, the cover 4 disengages from the first microswitch 1, the internal circuit of the first microswitch 1 is disconnected, and the first microswitch 1 sends a take-off signal to the electromechanical management computer.

[0041] The second trigger mechanism is composed of a second upper end cover 52, a second spring 53, a micro switch actuating piece 19, a second positioning pin 55, a second outer cylinder 56, a throttle valve 16, a second piston rod 172, a floating piston 18, a plunger 20, a sealing ring 21, a clamping nut 22, a self-lubricating bushing 23, a locking nut 26, and a second micro switch 51;

[0042] The upper end cover 52 is fixedly connected to the second outer cylinder 56 through the second positioning pin 55, and the second outer cylinder 56 is fixedly connected to the fuselage. The second upper end cover 52 is installed on the upper part of the second spring 53 to compress the second spring 53. The plunger 20 is installed at the upper opening of the second outer cylinder 56. The microswitch trigger plate 19 is fixed to the plunger 20 through the clamping nut 22 and moves with the plunger 20. The plunger 20 is provided with a boss to ensure that the upper and lower end surface areas are different; a self-lubricating bushing 23 is installed between the second piston rod 172 and the second outer cylinder 56 to reduce the friction force when the second piston rod 172 moves in the vertical direction. The plunger 20 makes a small movement in the vertical direction. The displacement S2 of the plunger 20 is equal to the rated stroke of the second microswitch 51, usually 2mm~5mm.

[0043] For the front landing gear, when the aircraft / helicopter touches the ground, the ground load N is transmitted to the first piston rod 171, the inner cylinder 8, and the cover 4 in sequence through the wheel tire. The first spring 3 is set at the upper end of the cover 4, and the spring has a certain preload (K1>n*Mg). The cover 4 overcomes the preload of the first spring 3 and compresses the first micro switch 1. The internal circuit of the first micro switch 1 is connected, and the first micro switch 1 sends the grounding signal to the electromechanical management computer. The ground load N is greater than (nMg+Mg). At the same time, a limit boss is set on the outer cylinder 2 as a travel protection mechanism for the micro switch to prevent the micro switch 1 from being damaged due to overtravel. The displacement S1 is equal to the rated travel of the micro switch 1, usually 2mm~5mm.

[0044] When the aircraft / helicopter takes off, under the action of gravity and the preload of the first spring 3, the baffle cover 4, the inner cylinder 8 and the first piston rod 171 move downward, the baffle cover 4 is disengaged from the first micro switch 1, a baffle ring 25 is installed between the inner cylinder 8 and the first outer cylinder 61 to prevent the inner cylinder 8 from escaping from the outer cylinder 6, the internal circuit of the micro switch 1 is disconnected, and the aircraft obtains a take-off signal.

[0045] For the main landing gear, when the aircraft / helicopter touches down, the ground load N is transmitted to the second piston rod 172 through the wheel tire, and the pressure in the oil chamber of the landing gear increases. The plunger 20 is provided with a boss to ensure that the upper and lower end surfaces have different areas. The plunger moves upward under the action of the oil pressure, and the resultant force is in the upward direction, with a magnitude of Where d is the plunger diameter, u is the friction coefficient between the seal ring 21 and the plunger, which is usually 0.1, m is the plunger mass, and K2 is the spring preload. The preload K2 of the second spring 53 satisfies When the ground load When the internal circuit of the second micro switch 51 is connected, the second micro switch 51 sends a ground signal to the electromechanical management computer.

[0046] When the aircraft / helicopter takes off, the landing gear extends, and a round nut 26 is arranged between the second outer tube 56 and the second piston rod 172 to prevent the second piston rod 172 from coming out. The floating piston 18 contacts the throttle valve 16 under the pressure of the air chamber, and the pressure in the oil chamber disappears. At this time, the plunger 20 moves downward, and the internal circuit of the second micro switch 51 is disconnected. The second micro switch 51 sends a take-off signal to the electromechanical management computer. The ground load N only needs to satisfy The second micro switch 51 can be triggered to send a ground signal, so the triggering sensitivity of the micro switch is very high.

[0047] The first trigger mechanism includes two first micro switches 1, and the second trigger mechanism of each main landing gear includes two second micro switches 51. The take-off signal / grounding signal of the aircraft / helicopter is executed by majority voting: if more than half of the micro switches 1 send the take-off signal to the electromechanical management computer, it is judged that it is in the take-off state at this time; if more than half of the micro switches 1 send the grounding signal to the electromechanical management computer, it is judged that it is in the grounding state.

[0048] The key points of the present invention are as follows:

[0049] 1) In general, the inner tube and the outer tube of the front landing gear buffer are fixedly connected. The present invention releases the vertical freedom of the inner tube and the outer tube to create a movable amount S1;

[0050] 2) The movable amount S1 is equal to the design stroke of the micro switch;

[0051] 3) For the front landing gear, the design of the spring stiffness takes into account the effects of friction and aircraft maneuvers, and is related to the overload factor during aircraft maneuvers;

[0052] 4) The present invention provides a method for calculating the main starting spring force;

[0053] 5) The present invention can conveniently install multiple micro switches, determine the grounding signal by majority voting, and increase the reliability of the micro switches without adding additional mechanisms.

[0054] The above only expresses the implementation mode of the present invention, and its description is relatively specific and detailed, but it cannot be understood as limiting the scope of the patent. It should be pointed out that for ordinary technicians in this field, several modifications and improvements can be made without departing from the concept of the present invention, which all belong to the protection scope of the present invention. In addition, the parts not detailed in the present invention are all conventional technologies.

Claims

1. A landing gear for triggering a wheel-load signal, characterized in that: The landing gear consists of a nose landing gear and two main landing gears. The front landing gear comprises a first outer tube (61), an inner tube (8), a first piston rod (171), and a first trigger mechanism, wherein the first outer tube (61) is fixedly connected to the fuselage, the inner tube (8) moves inside the first outer tube (61) in a vertical direction, the first piston rod (171) transfers the ground load to the inner tube (8) and the first outer tube (61), and the first trigger mechanism is installed on the upper part of the front landing gear; Each main landing gear comprises a second outer tube (62), a second piston rod (172), and a second trigger mechanism, wherein the second outer tube (62) is fixedly connected to the fuselage, the second piston rod (172) moves inside the second outer tube (62) in a vertical direction, an oil chamber is formed between the second outer tube (62) and the second piston rod (172), and the second trigger mechanism is installed on the upper part of the main landing gear.

2. The landing gear according to claim 1, characterized in that: The first trigger mechanism is composed of a first upper end cover (2), a first spring (3), a blocking cover (4), a first positioning pin (5), a self-lubricating support sleeve (7), a fixed piston (9), a blocking cap (10), a plunger (11), a locking nut (12), a sealing ring (13), a sealing ring (14), a self-lubricating support sleeve (15), a first micro switch (1), and a blocking ring (25); The first upper end cover (2) is fixedly connected to the first outer tube (61) through the first positioning pin (5), the first outer tube (61) is fixedly connected to the body, a first spring (3) is arranged on the upper end of the blocking cover (4), a limiting boss is arranged on the first upper end cover (2) as a travel protection mechanism of the first micro switch (1), a self-lubricating bushing (7) and a self-lubricating support sleeve (15) are installed between the first outer tube (61) and the inner tube (8), a round nut (24) is installed below the self-lubricating bushing (15), and the round nut (24) is threadedly connected to the first outer tube (61) The fixed piston (9) and the sealing ring (14) are installed on the upper part of the inner cylinder (8); the plunger (11) and the locking nut (12) are used to fix the first upper end cover (2) to provide pre-pressure for the first spring (3); the retaining ring (25) is installed between the inner cylinder (8) and the first outer cylinder (61); the inner cylinder (8) is supported by the self-lubricating support sleeve (7) and the self-lubricating support sleeve (15) and moves slightly inside the first outer cylinder (61) along the vertical direction, and the displacement S1 of the inner cylinder (8) is equal to the rated stroke of the first micro switch (1).

3. The landing gear according to claim 2, characterized in that: For the front landing gear, when the aircraft / helicopter touches the ground, the ground load is transmitted to the first piston rod (171), the inner cylinder (8), and the stop cover (4) in sequence through the wheel tire, and the stop cover (4) overcomes the preload K1 of the first spring (3) to compress the first micro switch (1), the internal circuit of the first micro switch (1) is connected, and the first micro switch (1) sends a grounding signal to the electromechanical management computer; When the aircraft / helicopter takes off from the ground, under the action of gravity and the first spring (3), the blocking cover (4), the inner cylinder (8) and the first piston rod (171) move downward, the blocking cover (4) is disconnected from the first micro switch (1), the internal circuit of the first micro switch (1) is disconnected, and the first micro switch (1) sends a take-off signal to the electromechanical management computer.

4. The landing gear according to claim 3, characterized in that: The preload K1 of the first spring (3) satisfies: K1>n*Mg, where n is the maximum downward overload coefficient during flight and Mg is the weight of the landing gear itself; when landing, when the ground load N is greater than (nMg+Mg), the first micro switch (1) can be triggered.

5. The landing gear according to claim 2, characterized in that: The second trigger mechanism is composed of a second upper end cover (52), a second spring (53), a micro switch actuating plate (19), a second positioning pin (55), a second outer cylinder (56), a throttle valve (16), a second piston rod (172), a floating piston (18), a plunger (20), a sealing ring (21), a clamping nut (22), a self-lubricating bushing (23), a locking nut (26), and a second micro switch (51); The upper end cover (52) is fixedly connected to the second outer tube (56) through a second positioning pin (55), the second outer tube (56) is fixedly connected to the body, the second upper end cover (52) is installed on the upper part of the second spring (53) to compress the second spring (53), the plunger (20) is installed at the upper opening of the second outer tube (56), the micro switch trigger plate (19) is fixed to the plunger (20) through a clamping nut (22), and moves with the plunger (20). The plunger (20) is provided with a boss to ensure that the upper and lower end surface areas are different; a self-lubricating bushing (23) is installed between the second piston rod (172) and the second outer tube (56), the plunger (20) makes a small movement in the vertical direction, and the displacement S2 of the plunger (20) is equal to the rated stroke of the second micro switch (51).

6. The landing gear according to claim 5, characterized in that: For the main landing gear, when the aircraft / helicopter touches the ground, the landing gear is compressed and the oil pressure increases. Since the upper and lower pressure areas of the plunger (20) are different, there is an upward combined force, which overcomes the friction force and the spring force, pushes the plunger (20), and the micro switch trigger plate (19) moves upward, the internal circuit of the second micro switch (51) is connected, and the second micro switch (51) sends a grounding signal to the electromechanical management computer; When the aircraft / helicopter takes off, the landing gear extends, the oil pressure disappears, and under the action of the preload K2 of the second spring (53) and the gravity mg, the plunger (20) moves downward, the micro switch trigger plate (19) is disconnected from the second micro switch (51), the internal circuit of the second micro switch (51) is disconnected, and the second micro switch (51) sends a take-off signal to the electromechanical management computer.

7. The landing gear according to claim 6, characterized in that: The preload K2 of the second spring (53) satisfies: Wherein, n is the maximum downward overload coefficient during flight, N is the ground load, A is the internal cross-sectional area of ​​the second outer cylinder (56), u is the friction coefficient, m is the mass of the plunger (20), and f is the friction force between the sealing ring (21) and the plunger (20); when landing, when the ground load When the second micro switch (51) is triggered.

8. The landing gear according to claim 2 or 5, characterized in that: The first trigger mechanism comprises two first micro switches (1), and the second trigger mechanism of each main landing gear comprises two second micro switches (51). The take-off signal / grounding signal of the aircraft / helicopter is executed by majority voting: if more than half of the micro switches (1) send the take-off signal to the electromechanical management computer, it is judged that the aircraft is in the take-off state at this time; if more than half of the micro switches (1) send the grounding signal to the electromechanical management computer, it is judged that the aircraft is in the grounding state at this time.

9. A state management system, characterized in that: The invention comprises a landing gear for triggering a wheel-borne signal as described in any one of claims 1 to 8, and an electromechanical management computer.

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