An aircraft landing gear fatigue test control system and method thereof

Abstract

This application belongs to the technical field of aircraft landing gear fatigue test design, and specifically relates to an aircraft landing gear fatigue test control system. Its design uses three emergency control two-way three-way valves to achieve accurate control of the landing gear retraction and extension. Moreover, one-way throttle valves are arranged on the retraction control pipeline of the retraction and extension actuator, the extension control pipeline of the retraction and extension actuator, and the emergency control pipeline of the retraction and extension actuator to achieve rapid retraction and extension control of the retraction and extension actuator. In addition, the design uses an emergency control pressure reducing valve for pressure reducing control to achieve rapid switching between two pressures in one system, without the need to design an additional landing gear emergency control branch, which can greatly reduce the complexity of the structure and operation. Also, the design uses a loading actuator to perform corresponding load loading during the normal retraction and extension of the landing gear, which is convenient and efficient. In addition, this application also relates to an aircraft landing gear fatigue test control method.

Description

Technical Field

[0001] This application belongs to the technical field of aircraft landing gear fatigue test design, and particularly relates to an aircraft landing gear fatigue test control system and method. Background Art

[0002] The retraction and extension of the aircraft landing gear are driven by a retraction and extension actuator, and are locked by an upper lock when retracted in place and by a lower lock when extended in place.

[0003] In order to verify whether the operating principles of the retraction and extension mechanism, upper lock, and lower lock of the aircraft landing gear, including locking and unlocking actions, are correct and whether the movement is flexible, an aircraft landing gear fatigue test is designed. However, in the current aircraft landing gear fatigue test, it is difficult to achieve fast and accurate control of the retraction and extension of the retraction and extension actuator, resulting in low test efficiency, poor accuracy of test results, and the need to design an additional landing gear emergency control branch, with complex structure and operation.

[0004] This application is proposed in view of the existence of the above technical defects.

[0005] It should be noted that the disclosure of the above background art content is only for assisting in understanding the inventive concept and technical solution of this application, and it does not necessarily belong to the prior art of this patent application. Without clear evidence indicating that the above content was publicly available at the filing date of this application, the above background art should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0006] The purpose of this application is to provide an aircraft landing gear fatigue test control system and method to overcome or mitigate at least one aspect of the known technical defects.

[0007] The technical solution of this application is as follows:

[0008] On the one hand, an aircraft landing gear fatigue test control system is provided, including:

[0009] An oil tank;

[0010] Three electromagnetic directional control valves are arranged in parallel and connected to the oil tank through an oil supply main pipe and an oil return main pipe. Among them, the first electromagnetic directional control valve is connected to the retraction and extension actuator through a retraction and extension actuator retraction control pipeline and a retraction and extension actuator extension control pipeline; the second electromagnetic directional control valve is connected to the upper lock through an upper lock locking control pipeline and an upper lock unlocking control pipeline; the third electromagnetic directional control valve is connected to the lower lock through a lower lock locking control pipeline and a lower lock unlocking control pipeline;

[0011] An emergency control pressure reducing valve is connected to the oil supply main pipe through an emergency liquid supply pipeline, and is connected to the retraction and extension actuator through a retraction and extension actuator emergency control pipeline, to the upper lock through an upper lock emergency control pipeline, and to the lower lock through a lower lock emergency control pipeline;

[0012] Three one-way throttle valves are arranged on the retraction control pipeline of the retraction and extension actuator, the extension control pipeline of the retraction and extension actuator, and the emergency control pipeline of the retraction and extension actuator;

[0013] Three emergency control two-way three-way valves are arranged on the emergency control pipeline of the retraction and extension actuator, the emergency control pipeline of the upper lock, and the emergency control pipeline of the lower lock;

[0014] The emergency switching two-way three-way valve is arranged on the extension control pipeline of the retraction and extension actuator;

[0015] The electric pump is arranged on the main oil inlet pipeline;

[0016] The loading actuator;

[0017] The controller is connected to three electromagnetic directional control valves, three emergency control two-way three-way valves, the emergency switching two-way three-way valve, the electric pump, as well as the upper lock end switch and the lower lock end switch, and controls to make the landing gear fatigue test control system of the aircraft have:

[0018] The normal retraction control working state of the landing gear. The controller controls the electric pump to start, controls the three emergency control two-way three-way valves to close, controls the emergency switching two-way three-way valve to close, and controls the three electromagnetic directional control valves to work, so that the lower lock unlocking control pipeline supplies liquid, controls the lower lock to unlock, and then the retraction control pipeline of the retraction and extension actuator supplies liquid, controls the retraction and extension actuator to drive the landing gear to retract. During the retraction of the landing gear, the loading actuator is controlled to apply a retraction load spectrum to the landing gear. After detecting that the landing gear is retracted in place, the upper lock locking control pipeline is controlled to supply liquid to lock the landing gear with the upper lock;

[0019] The normal extension control working state of the landing gear. The controller controls the electric pump to start, controls the three emergency control two-way three-way valves to close, controls the emergency switching two-way three-way valve to open, and controls the three electromagnetic directional control valves to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock to unlock, and then the extension control pipeline of the retraction and extension actuator supplies liquid, controls the retraction and extension actuator to drive the landing gear to extend. During the extension of the landing gear, the loading actuator is controlled to apply an extension load spectrum to the landing gear. After detecting that the landing gear is extended in place, the lower lock locking control pipeline is controlled to supply liquid to lock the landing gear;

[0020] The emergency extension control working state of the landing gear. The controller controls the electric pump to start, controls the three electromagnetic directional control valves to close, controls the emergency switching two-way three-way valve to close, and controls the three emergency control two-way three-way valves to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock to unlock, and then the extension control pipeline of the retraction and extension actuator supplies liquid, controls the retraction and extension actuator to drive the landing gear to extend. After detecting that the landing gear is extended in place, the lower lock locking control pipeline is controlled to supply liquid to lock the landing gear.

[0021] According to at least one embodiment of the present application, in the aircraft landing gear fatigue test control system described above, it further includes:

[0022] A fuel tank air filter, disposed on the fuel tank;

[0023] A fuel tank liquid level sensor, disposed on the fuel tank and connected to the controller;

[0024] A fuel tank temperature sensor, disposed on the fuel tank and connected to the controller;

[0025] A fuel tank liquid level gauge, disposed on the fuel tank;

[0026] A fuel tank drain valve, disposed on the fuel tank drain pipeline, and the inlet end of the fuel tank drain pipeline is connected to the fuel tank;

[0027] A fuel tank electric cooling heat exchanger, disposed on the fuel tank temperature circulation control pipeline, and the inlet end and the outlet end of the fuel tank temperature circulation control pipeline are connected to the fuel tank.

[0028] According to at least one embodiment of the present application, in the aircraft landing gear fatigue test control system described above, it further includes:

[0029] An oil inlet main pipe suction filter, disposed on the oil inlet main pipe and upstream of the electric pump;

[0030] An oil inlet main pipe manual butterfly valve, disposed on the oil inlet main pipe and between the electric pump and the oil inlet main pipe suction filter;

[0031] An oil inlet main pipe high-pressure oil filter, disposed on the oil inlet main pipe and downstream of the electric pump;

[0032] An oil inlet main pipe check valve, disposed on the oil inlet main pipe and downstream of the oil inlet main pipe high-pressure oil filter;

[0033] An oil inlet main pipe low-pressure accumulator, connected to the oil inlet main pipe through a pipeline, and an oil inlet main pipe low-pressure electromagnetic overflow valve is disposed on this pipeline, downstream of the oil inlet main pipe check valve;

[0034] An oil inlet main pipe high-pressure accumulator, connected to the oil inlet main pipe through a pipeline, and an oil inlet main pipe pressure sensor and an oil inlet main pipe shock-resistant pressure gauge are disposed on this pipeline, and a shock-resistant pressure gauge switch is disposed in front of the oil inlet main pipe shock-resistant pressure gauge, downstream of the oil inlet main pipe low-pressure accumulator; the oil inlet main pipe pressure sensor is connected to the controller;

[0035] An oil return main pipe flowmeter, disposed on the oil return main pipe;

[0036] An oil return main pipe oil return filter, disposed on the oil return main pipe and downstream of the oil return main pipe flowmeter;

[0037] A proportional overflow valve is connected to the inlet main pipe and the return main pipe through pipelines. The connection node with the inlet main pipe is between the low-pressure accumulator of the inlet main pipe and the high-pressure accumulator of the inlet main pipe, and the connection node with the return main pipe is upstream of the flowmeter of the return main pipe.

[0038] A stop valve is connected to the inlet main pipe and the return main pipe through pipelines. The connection node with the inlet main pipe is downstream of the high-pressure accumulator of the inlet main pipe, and the connection node with the return main pipe is upstream of the connection node between the proportional overflow valve and the return main pipe.

[0039] On the other hand, a method for controlling the fatigue test of an aircraft landing gear is provided, which is implemented based on any one of the above-mentioned aircraft landing gear fatigue test control systems, including:

[0040] Set the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear.

[0041] After the landing gear has been retracted for 2 s, determine whether the landing gear is being retracted and extended for the 50×nth time. If so, set the aircraft landing gear fatigue test control system to the emergency lowering control working state of the landing gear to lower the landing gear. After 24 s, reset the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear. If not, set the aircraft landing gear fatigue test control system to the normal lowering control working state of the landing gear to lower the landing gear. After 24 s, reset the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear.

[0042] Wherein,

[0043] n = 1, 2, 3... 600. Description of the Drawings

[0044] Figure 1 is a schematic diagram of the aircraft landing gear fatigue test control system provided by the embodiment of the present application;

[0045] Figure 2 is a schematic diagram of the method for controlling the fatigue test of an aircraft landing gear provided by the embodiment of the present application;

[0046] Wherein:

[0047] 1 - Fuel tank; 2 - Electromagnetic directional control valve; 3 - Receiving and releasing actuator; 4 - Upper lock; 5 - Lower lock; 6 - Emergency control pressure reducing valve; 7 - One-way throttle valve; 8 - Emergency control two-way three-way valve; 9 - Emergency switching two-way three-way valve; 10 - Electric pump; 11 - Loading actuator; 12 - Controller; 13 - Fuel tank air filter; 14 - Fuel tank level sensor; 15 - Fuel tank temperature sensor; 16 - Fuel tank level gauge; 17 - Fuel tank drain valve; 18 - Fuel tank electric cooling heat exchanger; 19 - Suction filter for inlet main pipe; 20 - Manual butterfly valve for inlet main pipe; 21 - High-pressure filter for inlet main pipe; 22 - Check valve for inlet main pipe; 23 - Low-pressure accumulator for inlet main pipe; 24 - Low-pressure electromagnetic relief valve for inlet main pipe; 25 - High-pressure accumulator for inlet main pipe; 26 - Pressure sensor for inlet main pipe; 27 - Shock-resistant pressure gauge; 28 - Shock-resistant pressure gauge switch; 29 - Flowmeter for return main pipe; 30 - Return filter for return main pipe; 31 - Proportional relief valve; 32 - Globe valve.

[0048] To better illustrate this embodiment, some components in the drawings are omitted, enlarged or reduced, which do not represent the dimensions of the actual product. In addition, the drawings are only for illustrative purposes and cannot be construed as a limitation of this patent. Detailed implementation manners

[0049] To make the technical solutions and their advantages of this application clearer, the technical solutions of this application will be further described clearly and completely in combination with the drawings. It can be understood that the specific embodiments described herein are only partial embodiments of this application, which are only used to explain this application rather than limit this application. It should be noted that for the convenience of description, only the parts related to this application are shown in the drawings, and other related parts can refer to the general design. Without conflict, the embodiments in this application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0050] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall have the ordinary meanings understood by those of ordinary skill in the art to which this application pertains. The words indicating directions such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", etc. used in the description of this application are only used to indicate relative directions or positional relationships, rather than implying that the device or component must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the object being described changes, its relative positional relationship may also change accordingly. Therefore, it should not be construed as a limitation to this application. The terms "first", "second", "third" and similar terms used in the description of this application are only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The similar words such as "a", "an" or "the" used in the description of this application should not be construed as an absolute limitation on the quantity, but should be understood as meaning at least one. The similar words such as "comprising" or "including" used in the description of this application are intended to mean that the elements or objects appearing before this word cover the elements or objects listed after this word and their equivalents, without excluding other elements or objects.

[0051] In addition, it should be noted that, unless otherwise clearly specified and limited, the similar words such as "installed", "connected" and "coupled" used in the description of this application should be understood in a broad sense. For example, the connection 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, and can also be the communication inside two components. Those skilled in the art can understand its specific meaning in this application according to the specific situation.

[0052] The following is a further detailed description of this application in conjunction with the Figures 1 to 2 accompanying drawings.

[0053] On the one hand, a fatigue test control system for an aircraft landing gear is provided, as Figure 1 shown, including:

[0054] An oil tank 1;

[0055] Three electromagnetic directional control valves 2, arranged in parallel, are connected to the oil tank 1 through an oil inlet main pipe and an oil return main pipe. Among them, the first electromagnetic directional control valve 2 is connected to a retraction actuator 3 through a retraction actuator retraction control pipeline and a retraction actuator extension control pipeline; the second electromagnetic directional control valve 2 is connected to an upper lock 4 through an upper lock locking control pipeline and an upper lock unlocking control pipeline; the third electromagnetic directional control valve 2 is connected to a lower lock 5 through a lower lock locking control pipeline and a lower lock unlocking control pipeline;

[0056] The emergency control pressure reducing valve 6 is connected to the inlet oil manifold through the emergency liquid supply pipeline, and is connected to the retraction and extension actuator 3 through the retraction and extension actuator emergency control pipeline, to the upper lock 4 through the upper lock emergency control pipeline, and to the lower lock 5 through the lower lock emergency control pipeline;

[0057] Three one-way throttle valves 7 are arranged on the retraction control pipeline of the retraction and extension actuator, the extension control pipeline of the retraction and extension actuator, and the emergency control pipeline of the retraction and extension actuator;

[0058] Three emergency control two-way three-way valves 8 are arranged on the emergency control pipeline of the retraction and extension actuator, the emergency control pipeline of the upper lock, and the emergency control pipeline of the lower lock;

[0059] The emergency switching two-way three-way valve 9 is arranged on the extension control pipeline of the retraction and extension actuator;

[0060] The electric pump 10 is arranged on the inlet oil manifold;

[0061] The loading actuator 11;

[0062] The controller 12 is connected to the three electromagnetic directional control valves 2, the three emergency control two-way three-way valves 8, the emergency switching two-way three-way valve 9, the electric pump 10, as well as the upper lock end switch and the lower lock end switch, and controls the aircraft landing gear fatigue test control system to have:

[0063] The normal retraction control working state of the landing gear. The controller 11 controls the electric pump 8 to start, controls the three emergency control two-way three-way valves 8 to close, controls the emergency switching two-way three-way valve 9 to close, and controls the three electromagnetic directional control valves 2 to work, so that the lower lock unlocking control pipeline supplies liquid, controls the lower lock 5 to unlock, and then the retraction control pipeline of the retraction and extension actuator supplies liquid, controls the retraction and extension actuator 3 to drive the landing gear to retract, controls the loading actuator 11 during the retraction of the landing gear, applies the retraction load spectrum to the landing gear, and after detecting that the landing gear is retracted in place, controls the upper lock locking control pipeline to supply liquid to lock the landing gear with the upper lock 4;

[0064] The normal extension control working state of the landing gear. The controller 11 controls the electric pump 8 to start, controls the three emergency control two-way three-way valves 8 to close, controls the emergency switching two-way three-way valve 9 to open, and controls the three electromagnetic directional control valves 2 to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock 4 to unlock, and then the extension control pipeline of the retraction and extension actuator supplies liquid, controls the retraction and extension actuator 3 to drive the landing gear to extend, controls the loading actuator 11 during the extension of the landing gear, applies the extension load spectrum to the landing gear, and after detecting that the landing gear is extended in place, controls the lower lock locking control pipeline to supply liquid to lock the landing gear;

[0065] In the working state of the landing gear emergency lowering control, the controller 11 controls the electric pump 8 to start, controls the three electromagnetic directional control valves 2 to close, controls the emergency switching two-way three-way valve 9 to close, and controls the three emergency control two-way three-way valves 8 to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock 4 to unlock, and further makes the retraction and extension actuator lowering control pipeline supply liquid, controls the retraction and extension actuator 3 to drive the landing gear to lower. After detecting that the landing gear is lowered in place, it controls the lower lock locking control pipeline to supply liquid to lock the landing gear.

[0066] For the aircraft landing gear fatigue test control system disclosed in the above embodiments, those skilled in the art can understand that its design uses three emergency control two-way three-way valves 8 to achieve accurate control of the retraction and lowering of the landing gear, and one-way throttle valves 7 are provided on the retraction control pipeline of the retraction and extension actuator, the lowering control pipeline of the retraction and extension actuator, and the emergency control pipeline of the retraction and extension actuator to achieve fast retraction and extension control of the retraction and extension actuator 3. In addition, the design uses an emergency control pressure reducing valve 6 for pressure reducing control to achieve fast switching between two pressures in one system, without the need to design an additional landing gear emergency control branch, which can greatly reduce the complexity of the structure and operation, and the design uses a loading actuator 11 to perform corresponding load loading during the normal retraction and lowering of the landing gear, which is convenient and efficient.

[0067] In some alternative embodiments, in the aircraft landing gear fatigue test control system described above, in order to facilitate the detection and control of the state of the oil in the fuel tank 1, the design further includes:

[0068] A fuel tank air filter 13, which is provided on the fuel tank 1;

[0069] A fuel tank liquid level sensor 14, which is provided on the fuel tank 1 and connected to the controller 12;

[0070] A fuel tank temperature sensor 15, which is provided on the fuel tank 1 and connected to the controller 12;

[0071] A fuel tank liquid level gauge 16, which is provided on the fuel tank 1;

[0072] A fuel tank drain valve 17, which is provided on the fuel tank drain pipeline, and the inlet end of the fuel tank drain pipeline is connected to the fuel tank 1;

[0073] A fuel tank electric cooling heat exchanger 18, which is provided on the fuel tank temperature circulation control pipeline, and the inlet end and outlet end of the fuel tank temperature circulation control pipeline are connected to the fuel tank 1.

[0074] In some alternative embodiments, in the aircraft landing gear fatigue test control system described above, in order to detect and control the state of the oil in the inlet main pipe and the return main pipe, the design further includes:

[0075] The oil inlet main pipe suction filter 19 is arranged on the oil inlet main pipe and is upstream of the electric pump 10;

[0076] The manual butterfly valve 20 of the oil inlet main pipe is arranged on the oil inlet main pipe and is between the electric pump 10 and the oil inlet main pipe suction filter 19;

[0077] The high-pressure oil filter 21 of the oil inlet main pipe is arranged on the oil inlet main pipe and is downstream of the electric pump 10;

[0078] The check valve 22 of the oil inlet main pipe is arranged on the oil inlet main pipe and is downstream of the high-pressure oil filter 21 of the oil inlet main pipe;

[0079] The low-pressure accumulator 23 of the oil inlet main pipe is connected to the oil inlet main pipe through a pipeline, and an electromagnetic overflow valve 24 for low pressure of the oil inlet main pipe is arranged on this pipeline. It is downstream of the check valve 22 of the oil inlet main pipe;

[0080] The high-pressure accumulator 25 of the oil inlet main pipe is connected to the oil inlet main pipe through a pipeline. A pressure sensor 26 for the oil inlet main pipe and a shock-resistant pressure gauge 27 for the oil inlet main pipe are arranged on this pipeline. And a shock-resistant pressure gauge switch 28 is arranged in front of the shock-resistant pressure gauge 27 for the oil inlet main pipe. It is downstream of the low-pressure accumulator 23 of the oil inlet main pipe; The pressure sensor 26 for the oil inlet main pipe is connected to the controller 12;

[0081] The flowmeter 29 of the oil return main pipe is arranged on the oil return main pipe;

[0082] The oil return filter 30 of the oil return main pipe is arranged on the oil return main pipe and is downstream of the flowmeter 29 of the oil return main pipe;

[0083] The proportional overflow valve 31 is connected to the oil inlet main pipe and the oil return main pipe through pipelines. The connection node with the oil inlet main pipe is between the low-pressure accumulator 23 and the high-pressure accumulator 25 of the oil inlet main pipe, and the connection node with the oil return main pipe is upstream of the flowmeter 29 of the oil return main pipe;

[0084] The stop valve 32 is connected to the oil inlet main pipe and the oil return main pipe through pipelines. The connection node with the oil inlet main pipe is downstream of the high-pressure accumulator 25 of the oil inlet main pipe, and the connection node with the oil return main pipe is upstream of the connection node between the proportional overflow valve 31 and the oil return main pipe.

[0085] In a specific embodiment, in the aircraft landing gear fatigue test control system disclosed in the above embodiment, the electric pump 10 can provide stepless pressure regulation from 0 to 35 MPa. In the normal landing gear retraction control working state and the normal landing gear lowering control working state, it outputs an oil pressure of 28 Mpa to the retraction and extension actuator 3, the upper locking device 4, and the lower locking device 5. In the emergency landing gear lowering control working state, due to the action of the emergency control pressure reducing valve 6, it outputs an oil pressure of 20 Mpa to the retraction and extension actuator 3, the upper locking device 4, and the lower locking device 5, and can realize the rapid retraction of the landing gear within 4 to 6 seconds and the rapid lowering within 5 to 7 seconds.

[0086] On the other hand, a method for controlling the fatigue test of an aircraft landing gear is provided, which is implemented based on any of the above-mentioned aircraft landing gear fatigue test control systems, such as Figure 2 shown below:

[0087] Before the test starts, it is detected whether the landing gear is in the lowered state and whether the three electromagnetic directional control valves 2 are closed. If not, the fault is eliminated;

[0088] After detecting that the landing gear is in the lowered state, the aircraft landing gear fatigue test control system is set to the normal retraction control working state of the landing gear to retract the landing gear, and it is detected whether the upper lock end switch of the landing gear inputs normally. If not, the landing gear is manually lowered to eliminate the fault. If so, the next step is carried out;

[0089] After the landing gear is retracted for 2 s, it is judged whether the landing gear is the 50×nth retraction and extension. If so, the aircraft landing gear fatigue test control system is set to the emergency lowering control working state of the landing gear to lower the landing gear, and it is detected whether the lower lock end switch of the landing gear inputs normally. If not, the fault is eliminated. If so, after 24 s, the aircraft landing gear fatigue test control system is reset to the normal retraction control working state of the landing gear to retract the landing gear; if not, the aircraft landing gear fatigue test control system is set to the normal lowering control working state of the landing gear to lower the landing gear, and it is detected whether the lower lock end switch of the landing gear inputs normally. If not, the fault is eliminated. If so, after 24 s, the aircraft landing gear fatigue test control system is reset to the normal retraction control working state of the landing gear to retract the landing gear;

[0090] Among them,

[0091] n = 1, 2, 3... 600.

[0092] The method for controlling the fatigue test of the aircraft landing gear disclosed in the above embodiment is implemented based on the aircraft landing gear fatigue test control system disclosed in the above embodiment, and the description is relatively simple. For specific relevant parts, reference can be made to the relevant description in the part of the aircraft landing gear fatigue test control system, and its technical effects can also be referred to the technical effects of the relevant part of the aircraft landing gear fatigue test control system, which will not be elaborated here.

[0093] The various embodiments in the specification are described in a progressive manner. Each embodiment focuses on the differences from other embodiments. The same or similar parts among the various embodiments can be referred to each other.

[0094] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of the present application is obviously not limited to these specific embodiments. Without departing from the principle of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the protection scope of the present application.

Claims

1. An aircraft landing gear fatigue test control system, characterized in that, Comprising: Fuel tank (1); Three electromagnetic directional control valves (2), arranged in parallel, connected to the fuel tank (1) through the inlet main pipe and the return main pipe. Among them, the first electromagnetic directional control valve (2) is connected to the retraction actuator (3) through the retraction actuator retraction control pipeline and the retraction actuator extension control pipeline; the second electromagnetic directional control valve (2) is connected to the upper lock (4) through the upper lock locking control pipeline and the upper lock unlocking control pipeline; the third electromagnetic directional control valve (2) is connected to the lower lock (5) through the lower lock locking control pipeline and the lower lock unlocking control pipeline; Emergency control pressure reducing valve (6), connected to the inlet main pipe through the emergency liquid supply pipeline, and connected to the retraction actuator (3) through the retraction actuator emergency control pipeline, connected to the upper lock (4) through the upper lock emergency control pipeline, and connected to the lower lock (5) through the lower lock emergency control pipeline; Three one-way throttle valves (7), arranged on the retraction actuator retraction control pipeline, the retraction actuator extension control pipeline, and the retraction actuator emergency control pipeline; Three emergency control two-way three-way valves (8), arranged on the retraction actuator emergency control pipeline, the upper lock emergency control pipeline, and the lower lock emergency control pipeline; Emergency switching two-way three-way valve (9), arranged on the retraction actuator extension control pipeline; Electric pump (10), arranged on the inlet main pipe; Loading actuator (11); Controller (12), connected to the three electromagnetic directional control valves (2), the three emergency control two-way three-way valves (8), the emergency switching two-way three-way valve (9), the electric pump (10), and the upper lock end switch and the lower lock end switch, controlling the aircraft landing gear fatigue test control system to have: Normal landing gear retraction control working state, the controller controls the electric pump to start, controls the three emergency control two-way three-way valves (8) to close, controls the emergency switching two-way three-way valve (9) to close, and controls the three electromagnetic directional control valves (2) to work, so that the lower lock unlocking control pipeline supplies liquid, controls the lower lock (5) to unlock, and then the retraction actuator retraction control pipeline supplies liquid, controls the retraction actuator (3) to drive the landing gear to retract. During the retraction of the landing gear, the loading actuator (11) is controlled to apply a retraction load spectrum to the landing gear. After detecting that the landing gear is retracted in place, the upper lock locking control pipeline is controlled to supply liquid, and the upper lock (4) locks the landing gear; Normal landing gear extension control working state, the controller controls the electric pump to start, controls the three emergency control two-way three-way valves (8) to close, controls the emergency switching two-way three-way valve (9) to open, and controls the three electromagnetic directional control valves (2) to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock (4) to unlock, and then the retraction actuator extension control pipeline supplies liquid, controls the retraction actuator (3) to drive the landing gear to extend. During the extension of the landing gear, the loading actuator (11) is controlled to apply an extension load spectrum to the landing gear. After detecting that the landing gear is extended in place, the lower lock locking control pipeline is controlled to supply liquid, and the landing gear is locked; In the emergency lowering control working state of the landing gear, the controller controls the electric pump to start, controls the three electromagnetic directional control valves (2) to close, controls the emergency switching two-way three-way valve (9) to close, and controls the three emergency control two-way three-way valves (8) to work, so that the upper lock unlocking control pipeline supplies liquid, controls the upper lock (4) to unlock, and further enables the retraction and extension actuator lowering control pipeline to supply liquid, controls the retraction and extension actuator (3) to drive the landing gear to lower. After detecting that the landing gear has been lowered in place, it controls the lower lock locking control pipeline to supply liquid to lock the landing gear.

2. The aircraft landing gear fatigue test control system according to claim 1, wherein it further comprises: a fuel tank air filter (13) provided on the fuel tank (1); a fuel tank liquid level sensor (14) provided on the fuel tank (1) and connected to the controller (12); a fuel tank temperature sensor (15) provided on the fuel tank (1) and connected to the controller (12); a fuel tank liquid level gauge (16) provided on the fuel tank (1); a fuel tank drain valve (17) provided on the fuel tank drain pipeline, and the inlet end of the fuel tank drain pipeline is connected to the fuel tank (1); a fuel tank electric cooling heat exchanger (18) provided on the fuel tank temperature circulation control pipeline, and the inlet end and outlet end of the fuel tank temperature circulation control pipeline are connected to the fuel tank (1).

3. The aircraft landing gear fatigue test control system according to claim 1, wherein it further comprises: a suction filter (19) for the inlet main pipeline provided on the inlet main pipeline and upstream of the electric pump (10); a manual butterfly valve (20) for the inlet main pipeline provided on the inlet main pipeline and between the electric pump (10) and the suction filter (19) for the inlet main pipeline; a high-pressure oil filter (21) for the inlet main pipeline provided on the inlet main pipeline and downstream of the electric pump (10); a check valve (22) for the inlet main pipeline provided on the inlet main pipeline and downstream of the high-pressure oil filter (21) for the inlet main pipeline; a low-pressure accumulator (23) for the inlet main pipeline connected to the inlet main pipeline through a pipeline, and an electromagnetic overflow valve (24) for the low pressure of the inlet main pipeline is provided on this pipeline, and it is downstream of the check valve (22) for the inlet main pipeline; a high-pressure accumulator (25) for the inlet main pipeline connected to the inlet main pipeline through a pipeline, and a pressure sensor (26) for the inlet main pipeline and a shock-resistant pressure gauge (27) for the inlet main pipeline are provided on this pipeline, and a shock-resistant pressure gauge switch (28) is provided in front of the shock-resistant pressure gauge (27) for the inlet main pipeline, and it is downstream of the low-pressure accumulator (23) for the inlet main pipeline; the pressure sensor (26) for the inlet main pipeline is connected to the controller (12); a flowmeter (29) for the return main pipeline provided on the return main pipeline; a return filter (30) for the return main pipeline provided on the return main pipeline and downstream of the flowmeter (29) for the return main pipeline; a proportional overflow valve (31) connected to the inlet main pipeline and the return main pipeline through pipelines, the connection node with the inlet main pipeline is between the low-pressure accumulator (23) and the high-pressure accumulator (25) for the inlet main pipeline, and the connection node with the return main pipeline is upstream of the flowmeter (29) for the return main pipeline; The globe valve (32) is connected to the oil inlet main pipe and the oil return main pipe through pipelines. The connection node with the oil inlet main pipe is downstream of the high-pressure accumulator (25) of the oil inlet main pipe, and the connection node with the oil return main pipe is upstream of the connection node between the proportional overflow valve (31) and the oil return main pipe.

4. A fatigue test control method for an aircraft landing gear, which is implemented based on the aircraft landing gear fatigue test control system described in any one of claims 1-3, characterized in that It includes: Set the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear. After the landing gear is retracted for 2 s, determine whether the landing gear is retracted and extended for the 50×nth time. If so, set the aircraft landing gear fatigue test control system to the emergency extension control working state of the landing gear to extend the landing gear. After 24 s, reset the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear. If not, set the aircraft landing gear fatigue test control system to the normal extension control working state of the landing gear to extend the landing gear. After 24 s, reset the aircraft landing gear fatigue test control system to the normal retraction control working state of the landing gear to retract the landing gear. Among them, n＝1,2,3……600。

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