A redundant hydraulic control system for an aircraft landing gear

Abstract

The present invention relates to the technical field of aircraft, and particularly to a redundant hydraulic control system for an aircraft landing gear. To solve the problem of the emergency risk that the hydraulic retraction and extension system cannot operate due to drive failure, the technical solution of the present invention includes a hydraulic drive unit, a fuel tank, a directional control valve, a landing gear actuator, and a pneumatic emergency unit; the outlet of the hydraulic pump is connected to the first port of the directional control valve; the landing gear actuator has a lowering chamber and a retracting chamber; both the lowering chamber and the retracting chamber are connected to the directional control valve; the pneumatic emergency unit includes a gas cylinder, a switch valve, a shuttle valve, and a pneumatically controlled check valve; the shuttle valve has a first shuttle valve inlet, a second shuttle valve inlet, and a shuttle valve outlet, the gas cylinder is connected to the first shuttle valve inlet through the switch valve, the second shuttle valve inlet is connected to the second port of the directional control valve, and the shuttle valve outlet is connected to the lowering chamber; the pneumatically controlled check valve is connected to the end of the switch valve away from the gas cylinder, and the pneumatically controlled check valve is connected to the retracting chamber.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft, and in particular to a redundant hydraulic control system for an aircraft landing gear. Background Art

[0002] As an important system for an aircraft (including flyable devices such as airplanes and drones) to land, the working performance of its own landing gear and the working reliability of the retraction and extension actuating cylinder directly affect the use and safety of the aircraft. The landing gear of large fixed-wing drones generally adopts a nose-wheel type layout, which consists of a nose landing gear, a left main landing gear, and a right main landing gear. The existing landing gear retraction and extension systems of drones are divided into hydraulic retraction and extension systems and electric retraction and extension systems.

[0003] At present, the electric retraction and extension system adopted by the landing gear of drones has a small power density, limited output torque of ordinary electric servos, and a complex retraction and extension controller. The transmission lead screw is prone to wear and jamming, resulting in the inability to lower the landing gear, seriously affecting the safety of the aircraft.

[0004] However, the hydraulic retraction and extension system has an urgent risk that the retraction and extension system cannot operate due to a drive failure, which needs to be solved. Summary of the Invention

[0005] In view of the above analysis, the embodiments of the present invention aim to provide a redundant hydraulic control system for an aircraft landing gear, which can at least solve one of the above problems.

[0006] It includes: a hydraulic drive unit, a fuel tank, a directional control valve, a landing gear actuating cylinder, and a pneumatic emergency unit;

[0007] The hydraulic drive unit includes a hydraulic pump and a motor connected to each other. The inlet of the hydraulic pump is connected to one end of the fuel tank, and the outlet of the hydraulic pump is connected to the first port of the directional control valve;

[0008] The landing gear actuating cylinder has a lowering chamber and a retracting chamber;

[0009] The directional control valve has a first port, a second port, a third port, and a fourth port. The first port is connected to the outlet of the hydraulic pump, the second port is connected to the lowering chamber, the third port is connected to the retracting chamber, and the fourth port is connected to the other end of the fuel tank;

[0010] The pneumatic emergency unit includes a gas cylinder, a switch valve, a shuttle valve, and a pneumatically controlled check valve;

[0011] The shuttle valve has a first shuttle valve inlet, a second shuttle valve inlet, and a shuttle valve outlet. The gas cylinder is connected to the first shuttle valve inlet through the switch valve, the second shuttle valve inlet is connected to the second port of the directional control valve, and the shuttle valve outlet is connected to the lowering chamber;

[0012] The pneumatically controlled one-way valve is connected to one end of the switching valve away from the gas cylinder, and the pneumatically controlled one-way valve is connected to the retracting chamber.

[0013] In some embodiments, the hydraulic drive unit includes a first drive unit and a second drive unit. The first drive unit includes a first hydraulic pump and a first motor connected to each other, the second drive unit includes a second hydraulic pump and a second motor connected to each other, and the first hydraulic pump and the second hydraulic pump are connected in parallel.

[0014] In some embodiments, the gas cylinder contains gas with a pressure not lower than 12 Mpa.

[0015] In some embodiments, a filter screen is further included. The outlet of the hydraulic pump is connected to the first port of the reversing valve through the filter screen.

[0016] In some embodiments, the reversing valve includes a two-position four-way spool-type electromagnetic reversing valve, and the switching valve includes a two-position two-way spool-type electromagnetic switching valve.

[0017] In some embodiments, a first one-way valve is further included. The outlet of the first one-way valve is connected to the lowering chamber.

[0018] In some embodiments, an oil drain tank is further included. The oil drain tank is connected to the retracting chamber through the pneumatically controlled one-way valve.

[0019] In some embodiments, a throttle one-way valve is further included. The third port of the reversing valve is connected to the retracting chamber through the throttle one-way valve.

[0020] In some embodiments, a microswitch is further included. The microswitch includes a first microswitch and a second microswitch, which are triggered when the landing gear is retracted to a preset uppermost position. The microswitch further includes a third microswitch and a fourth microswitch, which are triggered when the landing gear is lowered to a preset lowermost position.

[0021] The present invention also provides a redundant hydraulic control method for an aircraft landing gear. Using the redundant hydraulic control system for an aircraft landing gear described in any of the above embodiments, when the landing gear actuator cannot work properly, the method includes the following steps:

[0022] The switching valve is opened, and the gas in the gas cylinder enters the lowering chamber through the switching valve and the shuttle valve. At the same time, the gas opens the pneumatically controlled one-way valve so that the oil in the retracting chamber is discharged through the pneumatically controlled one-way valve, causing the piston rod in the landing gear actuator to extend and the landing gear to lower.

[0023] Compared with the prior art, at least one of the beneficial effects that can be achieved by the present invention:

[0024] In the embodiment of the present invention, the pneumatic emergency unit ensures that the landing gear can still be lowered when the main driving mode for lowering the landing gear fails. The piston rod of the landing gear actuator is extended through the pneumatic method, enabling the aircraft to have the ability to resist the risk of equipment failure.

[0025] In the present invention, the above technical solutions can also be combined with each other to achieve more preferred combined solutions. Other features and advantages of the present invention will be described in the subsequent specification. Moreover, some advantages can be made obvious from the specification or understood by implementing the present invention. The objectives and other advantages of the present invention can be realized and obtained from the content specifically pointed out in the specification and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] The drawings are only for the purpose of showing specific embodiments and are not considered as a limitation to the present invention. Throughout the drawings, the same reference signs denote the same components.

[0027] Figure 1 It is a schematic diagram of the redundant hydraulic control system for the aircraft landing gear provided by the present invention.

[0028] REFERENCE SIGNS:

[0029] 1 - Hydraulic drive unit; 1a - First drive unit; 1b - Second drive unit; 1a1 - First hydraulic pump; 1a2 - First motor; 1b1 - Second hydraulic pump; 1b2 - Second motor; 3 - Safety valve; 4 - Oil tank; 5 - Filter screen; 6 - Directional control valve; 7 - Pressure sensor; 8 - First check valve; 9 - Shuttle valve; 10 - One-way throttle valve; 11 - Landing gear actuator; 12 - Microswitch; 121 - First microswitch; 122 - Second microswitch; 123 - Third microswitch; 124 - Fourth microswitch; 13 - Gas cylinder; 14 - Pressure gauge; 15 - Switch valve; 16 - Pneumatic check valve; 17 - Oil drain tank; 18 - Pneumatic emergency unit. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0030] The following will specifically describe the preferred embodiments of the present invention with reference to the drawings. The drawings form a part of the present invention and are used together with the embodiments of the present invention to explain the principle of the present invention, rather than to limit the scope of the present invention.

[0031] In the description of the embodiments of the present invention, it should be noted that unless otherwise clearly specified and limited, the term "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection. It can be a mechanical connection or an electrical connection. It can be directly connected or indirectly connected through an intermediate medium. For those of ordinary skill in the art, the specific meaning of the above terms in the present invention can be understood according to specific circumstances.

[0032] The terms "top", "bottom", "above", "below", and "on" used throughout the description are relative positions with respect to the components of the device, such as the relative positions of the top and bottom substrates inside the device. It is understood that the devices are multifunctional and independent of their orientation in space.

[0033] The general working surface of the present invention can be a flat surface or a curved surface, which can be inclined or horizontal. For the convenience of description, the embodiments of the present invention are placed on a horizontal plane and used on the horizontal plane, and "high and low" and "up and down" are defined accordingly.

[0034] A specific embodiment of the present invention discloses a redundant hydraulic control system for an aircraft landing gear, as Figure 1 shown, comprising: a hydraulic drive unit 1, a fuel tank 4, a directional control valve 6, a landing gear actuator 11, and a pneumatic emergency unit 18;

[0035] The hydraulic drive unit 1 includes a hydraulic pump and a motor connected to each other. The inlet of the hydraulic pump is connected to one end of the fuel tank 4, and the outlet of the hydraulic pump is connected to the first port of the directional control valve 6;

[0036] The landing gear actuator 11 has a lowering chamber 11a and a retracting chamber 11b;

[0037] The directional control valve 6 has a first port, a second port, a third port, and a fourth port. The first port is connected to the outlet of the hydraulic pump, the second port is connected to the lowering chamber 11a, the third port is connected to the retracting chamber, and the fourth port is connected to the other end of the fuel tank 4;

[0038] The pneumatic emergency unit 18 includes a gas cylinder 13, a switching valve 15, a shuttle valve 9, and a pneumatically controlled check valve 16;

[0039] The shuttle valve 9 has a first shuttle valve inlet, a second shuttle valve inlet, and a shuttle valve outlet. The gas cylinder 13 is connected to the first shuttle valve inlet through the switching valve 15. The second shuttle valve inlet is connected to the second port of the directional control valve 6, and the shuttle valve outlet is connected to the lowering chamber 11a;

[0040] The pneumatically controlled check valve 16 is connected to the end of the switching valve 15 away from the gas cylinder 13, and the pneumatically controlled check valve 16 is connected to the retracting chamber 11b.

[0041] Specifically, in some embodiments, the fuel tank 4 can be, for example, a pressurized fuel tank. The pressurized fuel tank increases the pressure of the hydraulic fluid to raise the working pressure of the hydraulic system. The pressurized fuel tank generally adopts a double-layer structure, with the inner layer being the hydraulic fuel tank and the outer layer being the pneumatic fuel tank. Through the action of air pressure, the hydraulic fluid in the inner layer is compressed, thereby increasing the working pressure of the hydraulic system.

[0042] The reversing valve 6 is a two-position four-way spool-type electromagnetic reversing valve, and the switching valve 15 is a two-position two-way spool-type electromagnetic switching valve. When the electromagnetic reversing valve is de-energized, it is in position a, and the first port, i.e., Figure 1 the shown P port, is connected to the second port A port, and the third port B port is connected to the fourth port T port; when the electromagnetic reversing valve is energized, it is in position b, and the first port P port is connected to the third port B port, and the second port A port is connected to the fourth port T port.

[0043] In some embodiments, the hydraulic drive unit 1 includes a first drive unit 1a and a second drive unit 1b. The first drive unit 1a includes a first hydraulic pump 1a1 and a first motor 1a2 connected to each other, and the second drive unit 1b includes a second hydraulic pump 1b1 and a second motor 1b2 connected to each other. The first hydraulic pump 1a1 and the second hydraulic pump 1b1 are connected in parallel.

[0044] In this embodiment, the first drive unit 1a and the second drive unit 1b jointly drive the landing gear actuator 11 to work. While providing sufficient power for the landing gear actuator 11, when any one of the motors or hydraulic pumps fails, the remaining drive units can complete the driving work alone, realizing the redundancy design of the hydraulic drive unit.

[0045] In some embodiments, the gas cylinder 13 contains gas with a pressure not lower than 12 Mpa. Preferably, the gas volume in the gas cylinder 13 is configured according to the size of the landing gear actuator 11. In this embodiment, there is 0.2 liters of high-pressure gas in the gas cylinder 13.

[0046] In some embodiments, it further includes an oil drain tank 17, and the oil drain tank 17 is connected to the retraction chamber 11b through the air-controlled check valve 16.

[0047] In some embodiments, it further includes a pressure gauge 14 for monitoring the gas pressure in the gas cylinder 13.

[0048] When the landing gear actuator 11 fails to work properly, the electromagnetic switch valve 15 of the pneumatic emergency unit 18 is opened from the normally closed state. The high-pressure gas in the gas cylinder 13 enters the lowering chamber 11a through the first shuttle valve inlet of the switch valve 15 and the shuttle valve 9. At the same time, the gas opens the air-controlled check valve 16, so that the oil in the retracting chamber 11b is discharged to the oil drain tank 17 through the air-controlled check valve 16, causing the piston rod in the landing gear actuator to extend and the landing gear to lower. The redundant design of the landing gear lowering function is realized, ensuring that the aircraft has the ability to lower the landing gear in the case of extreme failure risks.

[0049] In some embodiments, a first check valve 8 is further included, and the outlet of the first check valve 8 is connected to the lowering chamber 11a. The first check valve 8 is used for the evacuation device to evacuate and refill the oil through this first check valve 8 after the emergency lowering of the landing gear, which is connected to the lowering chamber of the landing gear actuator 11, to prevent the oil from being mixed with gas and affecting the retraction and extension of the landing gear.

[0050] In some embodiments, a safety valve 3 is further included, and the safety valve 3 is connected in parallel with the hydraulic pump.

[0051] In some embodiments, a filter screen 5 is further included, and the outlet of the hydraulic pump is connected to the first port of the reversing valve 6 through the filter screen 5.

[0052] In some embodiments, a pressure sensor 7 is further included, and the pressure sensor 7 is arranged between the outlet of the hydraulic pump and the reversing valve 6 for monitoring the working pressure of the driving oil.

[0053] In some embodiments, a throttle check valve 10 is further included, and the third port of the reversing valve 6 is connected to the retracting chamber 11b through the throttle check valve 10. When retracting the landing gear, the oil enters the retracting chamber through the check valve channel therein to avoid throttle loss. When lowering the landing gear, the oil passes through the throttle channel to realize the function of controlling the lowering speed of the landing gear.

[0054] In some embodiments, a micro switch 12 is further included. The micro switch 12 includes a first micro switch 121 and a second micro switch 122, and the first micro switch 121 and the second micro switch 122 are triggered when the landing gear is retracted to the preset uppermost position. The micro switch 12 further includes a third micro switch 123 and a fourth micro switch 124, and the third micro switch 123 and the fourth micro switch 124 are triggered when the landing gear is lowered to the preset lowermost position.

[0055] In the embodiments of the present invention, two micro switches are provided at both the starting position and the ending position of the landing gear to realize the redundant design of the in-place signal.

[0056] Preferably, in some embodiments, it further includes two signal connectors, and each signal connector is connected to a motor, two microswitches 12, an electromagnetic reversing valve 6, and a pressure sensor 7 to implement the redundancy design of the power supply and signal acquisition cables.

[0057] The present invention also provides a redundant hydraulic control method for an aircraft landing gear. By using the redundant hydraulic control system for an aircraft landing gear described in any of the above embodiments, when the landing gear actuator cannot work properly, the following steps are included:

[0058] The switch valve is opened, and the gas in the gas cylinder enters the lowering chamber through the switch valve and the shuttle valve. At the same time, the gas opens the gas-controlled one-way valve so that the oil in the retracting chamber is discharged through the gas-controlled one-way valve, causing the piston rod in the landing gear actuator to extend and the landing gear to lower.

[0059] Under normal circumstances, when the flight control system issues a command to retract the landing gear, the on-board battery system supplies power to the reversing valve. After the electromagnetic reversing valve is energized and reverses from position a to position b, it supplies power to the motor. The high-pressure oil at the outlet of the motor hydraulic pump enters the P port of the reversing valve through the filter screen and flows to port B, and then enters the retracting chamber of the hydraulic landing gear actuator through the one-way valve in the one-way throttle valve. The hydraulic oil in the lowering chamber of the landing gear actuator flows from port A of the electromagnetic reversing valve to port T and enters the suction port of the hydraulic pump. Under the action of the high-pressure oil, the piston rod contracts in place. At this time, the microswitch sends a in-place signal to the flight control computer, the motor is powered off, and the electromagnetic reversing valve is powered off after a 5-second delay to complete the landing gear retraction action.

[0060] When the flight control system issues a command to lower the landing gear, the electromagnetic reversing valve is powered off and works in position a. It supplies power to the motor. The high-pressure oil at the outlet of the hydraulic pump flows from the P port of the electromagnetic reversing valve to port A and enters the lowering chamber of the landing gear actuator. The hydraulic oil in the retracting chamber of the landing gear actuator flows from port B of the electromagnetic reversing valve to port T and enters the suction port of the hydraulic pump. Under the action of the high-pressure oil in the lowering chamber of the landing gear actuator, and at the same time, the piston rod is subjected to the gravity of the landing gear and the thrust of the high-pressure oil in the lowering chamber, the landing gear lowers.

[0061] The above is only a preferred specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present invention should be covered by the protection scope of the present invention.

Claims

1. A redundant hydraulic control system for an aircraft landing gear, characterized in that, Comprising: A hydraulic drive unit, an oil tank, a directional control valve, a landing gear actuator and a pneumatic emergency unit; The hydraulic drive unit includes a hydraulic pump and a motor connected to each other. The inlet of the hydraulic pump is connected to one end of the oil tank, and the outlet of the hydraulic pump is connected to the first port of the directional control valve; The landing gear actuator has a lowering chamber and a retracting chamber; The directional control valve has a first port, a second port, a third port and a fourth port. The first port is connected to the outlet of the hydraulic pump, the second port is connected to the lowering chamber, the third port is connected to the retracting chamber, and the fourth port is connected to the other end of the oil tank; The pneumatic emergency unit includes a gas cylinder, a switch valve, a shuttle valve and a pneumatically controlled check valve; The shuttle valve has a first shuttle valve inlet, a second shuttle valve inlet and a shuttle valve outlet. The gas cylinder is connected to the first shuttle valve inlet through the switch valve. The second shuttle valve inlet is connected to the second port of the directional control valve, and the shuttle valve outlet is connected to the lowering chamber; The pneumatically controlled check valve is connected to the end of the switch valve away from the gas cylinder, and the pneumatically controlled check valve is connected to the retracting chamber.

2. The redundant hydraulic control system for the landing gear of an aircraft according to claim 1, wherein: The hydraulic drive unit includes a first drive unit and a second drive unit. The first drive unit includes a first hydraulic pump and a first motor connected to each other. The second drive unit includes a second hydraulic pump and a second motor connected to each other. The first hydraulic pump and the second hydraulic pump are connected in parallel.

3. The redundant hydraulic control system for the landing gear of an aircraft according to claim 1, characterized in that: The gas cylinder contains gas with a pressure not lower than 12 Mpa.

4. The redundant hydraulic control system for an aircraft landing gear according to claim 1, wherein: It further includes a filter screen. The outlet of the hydraulic pump is connected to the first port of the directional control valve through the filter screen.

5. The redundant hydraulic control system for the landing gear of an aircraft according to claim 1, wherein Comprising: The directional control valve includes a two-position four-way spool-type electromagnetic directional control valve, and the switch valve includes a two-position two-way spool-type electromagnetic switch valve.

6. The redundant hydraulic control system for an aircraft landing gear according to claim 1, characterized in that: It further includes a first check valve. The outlet of the first check valve is connected to the lowering chamber.

7. The redundant hydraulic control system for the landing gear of an aircraft according to claim 1, characterized in that: It further includes an oil drain tank. The oil drain tank is connected to the retracting chamber through the pneumatically controlled check valve.

8. The redundant hydraulic control system for an aircraft landing gear according to claim 1, characterized in that: It further includes a throttle check valve. The third port of the directional control valve is connected to the retracting chamber through the throttle check valve.

9. The redundant hydraulic control system for the landing gear of an aircraft according to claim 1, characterized in that: It further includes microswitches. The microswitches include a first microswitch and a second microswitch, which are triggered when the landing gear is retracted to a preset uppermost position. The microswitches further include a third microswitch and a fourth microswitch, which are triggered when the landing gear is lowered to a preset lowermost position.

10. A redundant hydraulic control method for an aircraft landing gear, characterized in that: Adopting the redundant hydraulic control system for the aircraft landing gear according to any one of claims 1-9, when the landing gear actuator fails to work properly, the following steps are included: The switch valve is opened. The gas in the gas cylinder enters the lowering chamber through the switch valve and the shuttle valve. At the same time, the gas opens the pneumatically controlled check valve, so that the oil in the retracting chamber is discharged through the pneumatically controlled check valve, causing the piston rod in the landing gear actuator to extend and the landing gear to lower.

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