Aircraft nose landing gear structure and folding and unfolding method

Abstract

The invention discloses an aircraft nose landing gear structure and a folding and unfolding method, the aircraft nose landing gear structure comprises a fuselage connecting assembly, a strut assembly, an oil injection valve assembly, a shimmy damping assembly, a wheel assembly and a folding and unfolding control assembly, the fuselage connecting assembly is arranged at the upper section of the strut assembly, the top end of the strut assembly is connected with a fuselage, and the lower end of the strut assembly is connected with the wheel assembly; the supporting column assembly is arranged on the machine body and fixedly connected with the machine body through the machine body connecting assembly, the oil injection valve assembly is arranged on the supporting column assembly and higher than the horizontal position of the machine body connecting assembly, the upper portion of the anti-swing assembly is connected with the supporting column assembly, the lower portion of the anti-swing assembly is connected with the wheel assembly, and the two ends of the folding and unfolding control assembly are connected with the supporting column assembly and the machine body respectively. The first semicircular piece and the second semicircular piece are fixed through a tightening bolt, the friction plate is made of a carbon-carbon composite material, and vibration transmitted to a machine body is reduced. The split structure facilitates replacement of the friction plates, the whole shimmy damper does not need to be disassembled, and maintenance time is shortened.

Description

Technical Field

[0001] The invention relates to the field of aircraft landing, and in particular to an aircraft front landing gear structure and a retracting and extending method. Background Art

[0002] The landing gear is the most common landing device of an aircraft and one of the important load-bearing parts of an aircraft. It bears a large and complex force, and bears the ground reaction force generated during landing impact, ground rolling, and various control movements (such as braking, turning, etc.). In addition, the impact loads of landing impact and ground rolling have obvious dynamic characteristics. The landing gear includes the nose landing gear and the main landing gear. The nose landing gear refers to the landing device installed in front of the wing, which is generally composed of struts, load-bearing wheels and special small wheels. Its main function is to support the weight of the aircraft during landing, absorb part of the impact force, and transmit this impact load to other parts of the fuselage or the ground. It can also adjust the taxiing direction and increase the turning radius for easy parking. It is mainly used for landing control and stable support during low-speed forward periods. Most of the nose landing gear are independent, that is, their structure and design are not linked to or affect each other relative to the fuselage. The landing equipment; this is conducive to improving the flexibility and fault tolerance of the mechanical layout, while also improving safety and reducing the accident rate.

[0003] The existing front landing gear structure is mostly suitable for small aircraft or UAVs, cannot meet the use requirements of larger load-bearing models, and cannot be retracted or extended. Summary of the invention

[0004] The technical problem to be solved by the present invention is to overcome the defects of the prior art and provide a front landing gear structure and a retracting and extending method for an aircraft.

[0005] In order to solve the above technical problems, the present invention provides the following technical solutions: The present invention discloses a front landing gear structure and a retracting and extending method for an aircraft, and comprises a fuselage connection assembly, a strut assembly, an oil filling valve assembly, a sway reduction assembly, a wheel assembly and a retracting and extending control assembly, wherein the fuselage connection assembly is provided with an upper section of a strut assembly, the top end of the strut assembly is connected to a fuselage, the bottom end is connected to a wheel assembly, and the strut assembly is fixedly connected to the fuselage through the fuselage connection assembly, the oil filling valve assembly is arranged on the strut assembly and is higher than a horizontal position of the fuselage connection assembly, the top of the sway reduction assembly is connected to the strut assembly, and the bottom end is connected to the wheel assembly, and the two ends of the retracting and extending control assembly are respectively connected to the strut assembly and the fuselage.

[0006] As a preferred technical solution of the present invention, the fuselage connection assembly is sleeved on the support assembly, and the fuselage connection assembly as a whole is a transverse bracket, and both ends are fixedly connected to the fuselage by bolts.

[0007] As a preferred technical solution of the present invention, the strut assembly includes a nose landing gear strut and a filling and inflation sign. The interior of the nose landing gear strut is hollow and provided with an oil cavity filled with hydraulic oil. A piston rod is arranged below the nose landing gear strut.

[0008] As a preferred technical solution of the present invention, the oil injection valve assembly includes an actuating cylinder, an inlet pipe joint, an outlet pipe joint, and an inlet and outlet hydraulic pipe. One end of the actuating cylinder is fixed on the nose landing gear strut. The inlet pipe joint and the outlet pipe joint are fixed on the outer wall of the actuating cylinder. Both the inlet pipe joint and the outlet pipe joint are externally connected to hydraulic pipes, and the hydraulic pipes are connected to the fuselage oil pressure system.

[0009] As a preferred technical solution of the present invention, the anti-shimmy assembly includes a scissors link, a centering rod, and an anti-shimmy damper. The centering rod is arranged in the middle of the upper end of the scissors link. Two holes are arranged below the scissors link, and the centering rod passes through the holes after folding.

[0010] As a preferred technical solution of the present invention, the anti-shimmy damper is sleeved on the piston rod. The anti-shimmy damper is divided into a first semi-circular part and a second semi-circular part. Friction plates are arranged on the contact surfaces of the inner sides of the first semi-circular part and the second semi-circular part with the fuselage connection assembly. The two sides of the first semi-circular part and the second semi-circular part are fixed to each other by tightening bolts.

[0011] As a preferred technical solution of the present invention, the wheel assembly includes a front fork, a mounting shaft, and a tire. The upper end of the front fork is connected to the piston rod, and the lower end of the front fork is connected to the mounting shaft on both sides. The mounting shaft is sleeved with a tire.

[0012] As a preferred technical solution of the present invention, the retraction and extension control assembly includes a fixed ring, a first rod arm, and a second rod arm. The fixed ring is sleeved and fixed on the movable rod. The first rod arm is rotatably connected to the fixed ring. The first rod arm and the second rod arm are connected by a spring above and rotatably connected below.

[0013] As a preferred technical solution of the present invention, the nose landing gear retracts and extends the piston rod of the actuating cylinder to retract the landing gear into the wing wheel well or lower the landing gear.

[0014] Compared with the prior art, the beneficial effects of the present invention are as follows: 1. The first semi-circular part and the second semi-circular part are fixed by tightening bolts. The friction plate is made of carbon-carbon composite material, which reduces the vibration transmitted to the fuselage. The thickness of the inner friction plate is adjustable to compensate for the wear gap. The split structure facilitates the replacement of the friction plate without disassembling the entire anti-shimmy damper, reducing the maintenance man-hours; 2. The anti-shimmy component is designed with an up-and-down linkage (the strut is connected to the wheel assembly) to suppress the shimmy of the front wheels during taxiing. Referring to the principle of a hydraulic damping anti-shimmy damper, a hydraulic actuator is used to drive a four-bar linkage mechanism to ensure that the retraction and extension trajectory is controllable, and a mechanical lock (such as a strut-type down lock) is used to achieve double locking to prevent accidental deployment in the air. By integrating functions of oil and gas shock absorption, hydraulic retraction and extension, and mechanical locking, the reliability of the system and the convenience of maintenance are improved, which is applicable to heavy aircraft to meet the requirements of landing gear retraction and extension during flight. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The drawings are used to provide a further understanding of the present invention, and constitute a part of the specification. Together with the embodiments of the present invention, they are used to explain the present invention, but do not constitute a limitation to the present invention. In the drawings: Figure 1 is a schematic diagram of the overall structure of the present invention; Figure 2 is a schematic side view of the overall structure of the present invention; Figure 3 is a schematic diagram of the first partial structure of the present invention; Figure 4 is a schematic diagram of the second partial structure of the present invention; In the figures: 1, fuselage connection component; 2, strut component; 21, nose landing gear strut; 22, inflation sign; 3, oil filling valve component; 31, actuator; 32, liquid inlet pipe joint; 33, liquid outlet pipe joint; 34, inlet and outlet hydraulic pipe; 4, anti-shimmy component; 41, scissor link; 42, centering rod; 43, anti-shimmy damper; 431, first semi-circular part; 432, second semi-circular part; 433, friction plate; 44, hole; 5, wheel component; 51, front fork; 52, mounting shaft; 53, tire; 6, retraction and extension control component; 61, fixed ring; 62, first rod arm; 63, second rod arm; 7, piston rod. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0016] The following describes the preferred embodiments of the present invention with reference to the drawings. It should be understood that the preferred embodiments described herein are only used to illustrate and explain the present invention, and are not used to limit the present invention.

[0017] Among them, the same reference numerals in the drawings all refer to the same components.

[0018] Such as Figures 1-4As shown in the figure, the present invention provides an aircraft nose landing gear structure and a retraction and extension method, which includes a fuselage connection assembly 1, a strut assembly 2, an oil filling valve assembly 3, a shimmy damper assembly 4, a wheel assembly 5 and a retraction and extension control assembly 6. The fuselage connection assembly 1 is arranged on the upper section of the strut assembly 2. The top of the strut assembly 2 is connected to the fuselage, and the bottom end is connected to the wheel assembly 5 and is fixedly connected to the fuselage through the fuselage connection assembly 1. The oil filling valve assembly 3 is arranged on the strut assembly 2 and is at a horizontal position higher than that of the fuselage connection assembly 1. The upper part of the shimmy damper assembly 4 is connected to the strut assembly 2, and the lower part is connected to the wheel assembly 5. Both ends of the retraction and extension control assembly 6 are respectively connected to the strut assembly 2 and the fuselage. In the present invention, the nose landing gear is one of the main load-bearing components when the aircraft taxis and parks on the ground, bearing about 10%-15% of the aircraft body weight. It transmits the load to the fuselage structure through high-strength load-bearing struts (such as titanium alloy or special steel components) to ensure the stable parking of the aircraft and absorb the dynamic impact force during taxiing. The fuselage connection assembly 1 keeps the bracket assembly and the fuselage in a relatively stable state. As the main load-bearing structure, the fuselage connection assembly 1 connects the fuselage and the strut assembly 2 and needs to meet the requirements of static force systems. The position of the oil filling valve is higher than that of the fuselage connection assembly 1, which is convenient for filling hydraulic oil during maintenance and at the same time avoids blocking the oil circuit by ground contaminants. The shimmy damper assembly 4 suppresses the shimmy of the front wheels during taxiing through an up-and-down linkage design (connecting the strut and the wheel assembly 5). Referring to the principle of the hydraulic damping shimmy damper 43, a hydraulic actuator 31 is used to drive a four-bar linkage mechanism to ensure that the retraction and extension trajectory is controllable, and a mechanical lock (such as a strut-type down lock) is used to achieve double locking to prevent accidental deployment in the air. By integrating functions such as oil and gas shock absorption, hydraulic retraction and extension, and mechanical locking, the reliability of the system and the convenience of maintenance are improved, which is applicable to heavy-duty aircraft to meet the requirements for retraction and extension of the landing gear during flight.

[0019] In an alternative embodiment, the fuselage connection assembly 1 is sleeved on the strut assembly 2. The fuselage connection assembly 1 is an overall transverse bracket, and both ends are fixedly connected to the fuselage through bolts.

[0020] It should be noted that high-strength aluminum alloy or titanium alloy is used and fixed to the fuselage strengthening frame through bolts to disperse the landing gear load to the main fuselage structure. A mechanical lock mechanism (such as a hook-type up lock) can be integrated inside the bracket to ensure the stable storage of the landing gear during flight. The bracket adopts a honeycomb weight-reducing structure to reduce the weight while ensuring the strength, meeting the lightweight requirements of modern aircraft, and is applicable to the tricycle landing gear layout to ensure the ground taxiing stability and the in-air retraction and extension efficiency.

[0021] In an alternative embodiment, the strut assembly 2 includes a nose landing gear strut 21 and a filling and inflation identification plate 22. The inside of the nose landing gear strut 21 is hollow and is provided with an oil cavity, and the oil cavity is filled with hydraulic oil. A piston rod 7 is arranged below the nose landing gear strut 21.

[0022] It should be noted that through the interaction of compressed gas (nitrogen) and hydraulic oil in the nose landing gear, the impact energy during landing is absorbed, the vibration transmitted to the fuselage is reduced, the weight of the front part of the aircraft is borne during static and taxiing, and the main landing gear shares the load of the whole aircraft together. The inflation sign 22 is used to indicate the nitrogen pressure state to ensure that the shock absorption performance meets the airworthiness standards (such as a mirror height of 12 - 16 cm).

[0023] In an optional embodiment, the oil injection valve assembly 3 includes an actuating cylinder 31, an inlet liquid pipe joint 32, an outlet liquid pipe joint 33 and an inlet and outlet hydraulic pipe 34. One end of the actuating cylinder 31 is fixed on the nose landing gear strut 21. The inlet liquid pipe joint 32 and the outlet liquid pipe joint 33 are fixed on the outer wall of the actuating cylinder 31. Both the inlet liquid pipe joint 32 and the outlet liquid pipe joint 33 are externally connected to hydraulic pipes, and the hydraulic pipes are connected to the fuselage oil pressure system.

[0024] It should be noted that a non - balanced double - acting actuating cylinder 31 is adopted. When the piston rod 7 retracts, the landing gear is retracted, and when it extends, the landing gear is lowered. It matches the 21 MPa working pressure of the hydraulic system. The inner wall of the actuating cylinder 31 is sprayed with a ceramic coating to reduce the frictional resistance. The inlet / outlet liquid pipes are in parallel with the fuselage hydraulic system. When the main system fails, it can be switched to standby cold gas or electric drive to ensure the reliability of retraction and extension. A pressure sensor and a flowmeter are integrated to real - time feedback the oil fluid state to the cockpit display system, and an alarm is triggered when abnormal. The position of the oil injection valve needs to be calibrated regularly to prevent oil leakage or air bubbles from mixing in.

[0025] In an optional embodiment, the anti - shimmy assembly 4 includes a scissors link 41, a centering rod 42 and an anti - shimmy damper 43. The centering rod 42 is arranged in the middle of the upper end of the scissors link 41. There are two holes 44 provided below the scissors link 41. After folding, the centering rod 42 passes through the holes 44.

[0026] It should be noted that a geometrically invariant structure is formed to force the front wheel to remain in the neutral position during retraction and extension to prevent jamming. The centering rod 42 is connected to the strut and the wheel fork through hinges, allowing a ±50° deflection to adapt to the ground turning requirements. A rotary vane hydraulic damper is adopted, and the oil fluid flow rate is adjusted through a damping hole with a diameter of 0.8 mm to dynamically suppress shimmy. The damping coefficient can be automatically adjusted according to the taxiing speed (such as low damping at low speed and high damping at high speed). The holes 44 below the scissors link 41 allow the centering rod 42 to fold, reducing the space requirement for retraction and extension, being suitable for a compact wheel bay layout. The anti - shimmy efficiency is verified to be ≥85% through ground simulation tests and wind tunnel tests.

[0027] In an optional embodiment, the vibration damper 43 is sleeved on the piston rod 7. The vibration damper 43 is divided into a first semi-circular member 431 and a second semi-circular member 432. Friction plates 433 are provided on the contact surfaces of the inner sides of the first semi-circular member 431 and the second semi-circular member 432 with the fuselage connection assembly 1. The two sides of the first semi-circular member 431 and the second semi-circular member 432 are fixed to each other by tightening bolts.

[0028] It should be noted that the first semi-circular member 431 and the second semi-circular member 432 are fixed by tightening bolts. The friction plates 433 are made of carbon-carbon composite materials to reduce the vibration transmitted to the fuselage. The thickness of the inner friction plates 433 is adjustable to compensate for the wear gap. The split structure facilitates the disassembly and replacement of the friction plates 433 without disassembling the entire vibration damper 43, reducing the maintenance man-hours. The surfaces of the friction plates 433 are coated with high-temperature resistant ceramic coatings to withstand the 300°C high temperature generated during braking. The wear life of the friction plates 433 is ≥5000 takeoffs and landings, and the pre-tightening force of the tightening bolts needs to be calibrated regularly.

[0029] In an optional embodiment, the wheel assembly 5 includes a front fork 51, a mounting shaft 52, and a tire 53. The upper end of the front fork 51 is connected to the piston rod 7, and the two sides of the lower end are connected to the mounting shaft 52. The mounting shaft 52 is sleeved with the tire 53 through the upper part.

[0030] It should be noted that the titanium alloy forged front fork 51 is connected to the piston rod 7. The wheel axle integrates a brake disc and a temperature sensor to adapt to high-temperature and high-load working conditions. The surface of the wheel axle is treated by ion nitriding to improve the fatigue resistance. A double-wheel configuration can be adopted: to improve the ground floating performance and still be able to taxi safely when a single wheel fails. The tire 53 adopts a tubeless design and is equipped with a tire pressure monitoring system (TPMS). The front fork 51 is connected to the strut through a tapered bearing, allowing the front wheel to deflect ±50°, and matching the hydraulic steering system to achieve precise control.

[0031] In an optional embodiment, the retraction and extension control assembly 6 includes a fixed ring 61, a first rod arm 62, and a second rod arm 63. The fixed ring 61 is sleeved and fixed on the movable rod. The first rod arm 62 is rotatably connected to the fixed ring 61. The upper parts of the first rod arm 62 and the second rod arm 63 are connected by a spring, and the lower parts are rotatably connected.

[0032] It should be noted that the first lever arm 62 is hinged to the fixed ring 61, and the second lever arm 63 provides buffering through the spring to form a four-bar linkage to ensure that the retraction and release trajectory is controllable. The spring preload is adjustable to adapt to the load changes under different working conditions. The spring preload supports gravity emergency lowering, and the mechanical override cable can bypass the hydraulic system to unlock directly. The retraction and release actuator 31 is equipped with a backup electric motor to ensure that it can still be operated when the hydraulic pressure fails. The fixed ring 61 has a built-in Hall sensor to monitor the lever arm angle in real time and feed it back to the flight control computer. In the event of an abnormality, it triggers automatic correction. The system has passed DO-178C Level C certification, and the failure rate is ≤0.001 times / flight hour.

[0033] In an optional embodiment, the front landing gear is retracted and extended by the piston rod of the actuator 31 to retract the landing gear upward into the wing wheel bay or lower the landing gear.

[0034] The working principle of the present invention is as follows: the landing gear strut is installed with the fuselage through the fuselage connection assembly 1, connected with the front fork 51 through the piston rod 7, and fixed with the wheel and tire 53 through the mounting shaft 52. When the landing gear is loaded, the compression and extension movement of the landing gear strut is blocked by the oil in the strut. At the same time, the tire 53 can also carry part of the energy.

[0035] The scissor link 41 connects the landing gear strut and the front fork 51 through bolts, which effectively controls the lifting height of the landing gear strut and prevents the strut from bouncing up when the pressure inside the strut is too large, causing the nose to tilt and the center of gravity of the fuselage to move backward, which may cause the risk of damaging the tail.

[0036] The shimmy damper 43 is mounted on the landing gear strut and connected to the front fork 51 via the scissor link 41 to achieve the purpose of controlling the movement of the front wheel. Its built-in oppositely mounted friction plates 433 can increase the movement resistance and prevent the front wheel from shimmying left and right after being loaded.

[0037] The actuator 31 is connected to the landing gear retraction and extension control assembly 6. When the piston of the actuator 31 is shortened, the landing gear is retracted backward into the fuselage; when the piston of the actuator 31 is extended, the landing gear is lowered. There is a warning light when the landing gear is retracted and extended.

[0038] Finally, it should be noted that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art can still modify the technical solutions described in the aforementioned embodiments or replace some of the technical features therein by equivalents. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. An aircraft nose landing gear structure, characterized in that, It includes a fuselage connection component (1), a strut component (2), an oil filling valve component (3), an anti-shimmy component (4), a wheel component (5), and a retraction control component (6). The fuselage connection component (1) is arranged on the upper section of the strut component (2). The top of the strut component (2) is connected to the fuselage, and the bottom end is connected to the wheel component (5), and is fixedly connected to the fuselage through the fuselage connection component (1). The oil filling valve component (3) is arranged on the strut component (2) and is higher than the horizontal position of the fuselage connection component (1). The upper part of the anti-shimmy component (4) is connected to the strut component (2), and the lower part is connected to the wheel component (5). The two ends of the retraction control component (6) are respectively connected to the strut component (2) and the fuselage.

2. The structure of an aircraft nose landing gear according to claim 1, characterized in that, The fuselage connection component (1) is sleeved on the strut component (2). The fuselage connection component (1) is an overall horizontal bracket, and both ends are fixedly connected to the fuselage by bolts.

3. The structure of an aircraft nose landing gear according to claim 1, characterized in that, The strut component (2) includes a nose landing gear strut (21) and a filling and inflation sign (22). The inside of the nose landing gear strut (21) is hollow, provided with an oil chamber, and the oil chamber is filled with hydraulic oil. A piston rod (7) is arranged below the nose landing gear strut (21).

4. The structure of an aircraft nose landing gear according to claim 3, characterized in that, The oil filling valve component (3) includes an actuating cylinder (31), a liquid inlet pipe joint (32), a liquid outlet pipe joint (33), and an inlet and outlet hydraulic pipe (34). One end of the actuating cylinder (31) is fixed on the nose landing gear strut (21). The liquid inlet pipe joint (32) and the liquid outlet pipe joint (33) are fixed on the outer wall of the actuating cylinder (31). Both the liquid inlet pipe joint (32) and the liquid outlet pipe joint (33) are externally connected to hydraulic pipes, and the hydraulic pipes are connected to the fuselage oil pressure system.

5. The structure of an aircraft nose landing gear according to claim 3, characterized in that, The anti-shimmy component (4) includes a scissors link (41), an alignment rod (42), and an anti-shimmy damper (43). The alignment rod (42) is arranged in the middle of the upper end of the scissors link (41). Two holes (44) are arranged below the scissors link (41). After folding, the alignment rod (42) passes through the holes (44).

6. The structure of an aircraft nose landing gear according to claim 5, characterized in that, The anti-shimmy damper (43) is sleeved on the piston rod (7). The anti-shimmy damper (43) is divided into a first semi-circular part (431) and a second semi-circular part (432). Friction plates (433) are arranged on the contact surfaces of the inner sides of the first semi-circular part (431) and the second semi-circular part (432) with the fuselage connection component (1). The two sides of the first semi-circular part (431) and the second semi-circular part (432) are fixed to each other by tightening bolts.

7. A front landing gear structure of an aircraft according to claim 1, characterized in that, The wheel component (5) includes a front fork (51), a mounting shaft (52), and a tire (53). The upper end of the front fork (51) is connected to the piston rod (7), and the lower end is connected to the mounting shaft (52) on both sides. The mounting shaft (52) is sleeved with the tire (53) from above.

8. The structure of an aircraft nose landing gear according to claim 1, characterized in that, The retraction control component (6) includes a fixed ring (61), a first rod arm (62), and a second rod arm (63). The fixed ring (61) is sleeved and fixed on the movable rod. The first rod arm (62) is rotatably connected to the fixed ring (61). The upper parts of the first rod arm (62) and the second rod arm (63) are connected by a spring, and the lower parts are rotatably connected.

9. A method for retracting and extending the nose landing gear of an aircraft, characterized in that, The nose landing gear retracts the landing gear into the wing wheel well or lowers the landing gear by retracting and extending the piston rod of the actuator (31).

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