Aircraft landing deceleration system and control method thereof

Abstract

The invention provides an aircraft landing deceleration system and a control method thereof, and relates to the technical field of aircrafts. The speed reducing system comprises a speed reducing plate arranged on a fuselage in the middle of the aircraft and a steering engine arranged on the inner wall of the aircraft; an unfolding mechanism is arranged between the speed reducing plate and the driving end of the steering engine, hinged to the steering engine and fixedly connected with the speed reducing plate. The unfolding mechanism is used for driving the speed brake to unfold under the driving of the steering engine; a locking mechanism is further arranged in the aircraft, the locking mechanism is connected between a fuselage of the aircraft and the air brake, and the locking mechanism is used for locking the air brake when the aircraft flies normally; compared with the prior art, the unfolding mechanism is directly controlled by the steering engine, an independent driving scheme is adopted, the overall structure is simple, the occupied space is small, and layout and performance improvement of other equipment in the aircraft cannot be limited; and meanwhile, a special locking mechanism is arranged to lock the speed brake, so that the safety, the stability and the reliability during unfolding are further improved.

Description

Technical Field

[0001] The present invention generally relates to the field of aircraft technology, and in particular to an aircraft landing deceleration system and a control method thereof. Background Art

[0002] A high-speed aircraft is an aircraft that has the authority to fly at high speeds, usually an aircraft whose maximum speed exceeds the speed of sound; the main purpose of the design and development of high-speed aircraft is to cover long distances in a short time and improve the efficiency and convenience of air transportation; in actual applications, high-speed aircraft are mainly used for long-distance flights and some short-distance flights, which can significantly shorten the flight time, allowing passengers and cargo to reach their destination in a shorter time; as the speed of the high-speed aircraft increases, its speed needs to be reduced before landing, shortening the taxiing distance, reducing the wear and tear and time of the aircraft on the runway, so that it can land safely.

[0003] In today's aerospace field, high-speed aircraft are developing rapidly, but their deceleration systems have shortcomings that cannot be ignored; when facing the critical link of high-speed takeoff and landing, the existing high-speed aircraft deceleration systems expose the problem of poor reliability; under the harsh conditions of high-speed takeoff and landing, the stability of the deceleration system is directly related to the safety of the aircraft. Once a failure occurs, it may lead to serious consequences.

[0004] The existing deceleration system deployment mechanism generally adopts an actuator drive method, which relies on a hydraulic or pneumatic energy source to operate. This not only makes the structure of the entire system extremely complicated and increases the difficulty of maintenance and overhaul, but also takes up a lot of valuable space, limiting the layout and performance improvement of other aircraft equipment. In addition, due to the complexity of the hydraulic and pneumatic systems themselves, their reliability is relatively low, and leakage, blockage and other faults may occur during flight, further affecting the safety and stability of the aircraft. Summary of the invention

[0005] In view of the above defects or deficiencies in the prior art, it is desired to provide an aircraft landing deceleration system and a control method thereof that can solve the above technical problems.

[0006] In a first aspect, the present invention provides an aircraft landing deceleration system, comprising: A speed brake, the speed brake being arranged on the fuselage in the middle of the aircraft; A steering gear, the steering gear being arranged on the inner wall of the aircraft; An unfolding mechanism, the unfolding mechanism is arranged between the driving end of the steering gear and the speed brake, the unfolding mechanism is hinged to the steering gear and fixedly connected to the speed brake; the unfolding mechanism is used to drive the speed brake to unfold under the drive of the steering gear; A locking mechanism is located inside the aircraft and connected between the fuselage of the aircraft and the airbrake. The locking mechanism is used to lock the airbrake when the aircraft is in a normal flight state.

[0007] According to the technical solution provided by the present invention, the locking mechanism includes: a pin puller assembly and a locking groove, the pin puller assembly is fixedly connected to the speed brake, and the locking groove is arranged on the inner wall of the aircraft; the pin puller assembly has a first state and a second state, when in the first state, the aircraft is in a normal flight state, and the pin puller assembly is inserted into the locking groove to lock the speed brake; when in the second state, the aircraft is in a decelerated landing state, and the pin puller assembly is separated from the locking groove to release the speed brake.

[0008] According to the technical solution provided by the present invention, the pin pulling assembly comprises: A cylinder, wherein the interior of the cylinder has a first cavity; A piston rod is provided, wherein the piston rod penetrates the cylinder body along a first direction; a force-bearing portion is provided in the middle of the piston rod; when the pin pulling assembly is in a first state, one end of the piston rod is inserted into the locking groove An electric squib is arranged on the cylinder and is located on a side of the force-bearing portion close to the locking groove.

[0009] According to the technical solution provided by the present invention, a shear pin is provided on the cylinder and located on a side of the electric squib away from the locking groove, and the piston rod is provided with a first groove corresponding to the shear pin. When the pin pulling assembly is in the first state, one end of the shear pin is located in the first groove.

[0010] According to the technical solution provided by the present invention, a first sealing ring is arranged between the electric squib and the cylinder; and a second sealing ring is arranged between the piston rod and the cylinder.

[0011] According to the technical solution provided by the present invention, the deployment mechanism includes: a connecting member, the connecting member having a first connecting part, a second connecting part and a third connecting part, the first connecting part is hinged to the inner wall of the aircraft, the second connecting part is hinged to the driving end of the servo, and the third connecting part is fixedly connected to the speed brake.

[0012] According to the technical solution provided by the present invention, a heat sealing structure is provided at the connection between the speed reducer and the aircraft fuselage.

[0013] According to the technical solution provided by the present invention, a heat-resistant coating is provided on the side of the speed brake away from the aircraft.

[0014] In a second aspect, the present invention provides a control method for an aircraft landing deceleration system, which is used to control the aircraft landing deceleration system according to any one of the first aspects, and comprises the following steps: When the aircraft decelerates for landing, a first control instruction is issued, wherein the first control instruction is used to control the activation of the locking mechanism; Based on the first control instruction, the locking mechanism releases the speed brake; issuing a second control instruction, wherein the second control instruction is used to control the start of the steering gear; Based on the second control instruction, the steering gear drives the deployment mechanism to cause the speed brake to deploy, so that the aircraft can achieve decelerated landing under the action of the speed brake.

[0015] The beneficial effects of the present invention are: The present invention provides an aircraft landing deceleration system and a control method thereof, wherein the deceleration system comprises: a deceleration plate arranged on a fuselage in the middle of an aircraft, and a steering gear arranged on an inner wall of the aircraft; an unfolding mechanism is arranged between the deceleration plate and a driving end of the steering gear, the unfolding mechanism is hinged to the steering gear and fixedly connected to the deceleration plate; the unfolding mechanism is used to drive the deceleration plate to unfold under the drive of the steering gear; a locking mechanism is also arranged inside the aircraft, and the locking mechanism is connected between the fuselage of the aircraft and the deceleration plate, and the locking mechanism is used to lock the deceleration plate when the aircraft is in normal flight; when the aircraft enters a deceleration zone, the deceleration plate is locked. When the aircraft is in a high-speed landing state, the control system issues a first control instruction to control the locking mechanism to start so as to release the speed brake; subsequently, the control system issues a second control instruction to control the steering gear to start, and the steering gear drives the deployment mechanism to drive the speed brake to deploy, so that the aircraft can achieve a decelerated landing under the action of the speed brake; compared with the prior art, the deployment mechanism of the present invention is directly controlled by the steering gear and adopts a separate driving scheme, with a simple overall structure and a small occupied space, and will not limit the layout and performance improvement of other equipment in the aircraft; at the same time, a special locking mechanism is provided to lock the speed brake, so as to further improve the safety, stability and reliability during deployment. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings: Figure 1 is a schematic diagram of an aircraft landing deceleration system provided by the present invention; Figure 2 is a cross-sectional view of the locking mechanism provided by the present invention; Figure 3 The present invention is a flow chart of a control method for an aircraft landing deceleration system.

[0017] In the figure: 1. speed brake; 2. servo; 3. pin puller assembly; 31. cylinder; 32. piston rod; 33. electric squib; 34. shear pin; 35. first cavity; 36. first sealing ring; 37. second sealing ring; 4. connecting piece; 41. first connecting part; 42. second connecting part; 43. third connecting part; 5. heat sealing structure; 6. heat-resistant coating. DETAILED DESCRIPTION

[0018] The present invention will be further described in detail below in conjunction with the accompanying drawings and embodiments. It is to be understood that the specific embodiments described herein are only used to explain the relevant invention, rather than to limit the invention. It should also be noted that, for ease of description, only the parts related to the invention are shown in the accompanying drawings.

[0019] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0020] Example 1 Please refer to Figure 1 As shown, the present invention provides an aircraft landing deceleration system, comprising: A deceleration plate 1 is arranged on the fuselage in the middle of the aircraft; A steering gear 2, which is arranged on the inner wall of the aircraft; The deployment mechanism is arranged between the driving end of the steering gear 2 and the speed brake 1. The deployment mechanism is hinged to the steering gear 2 and fixedly connected to the speed brake 1. The deployment mechanism is used to drive the speed brake 1 to deploy under the drive of the steering gear 2. The locking mechanism is located inside the aircraft and connected between the fuselage of the aircraft and the airbrake 1 . The locking mechanism is used to lock the airbrake 1 when the aircraft is in a normal flight state.

[0021] Specifically, the speed brake 1 is arranged on the fuselage in the middle of the aircraft, and its rotational torque is provided by the steering gear 2 in the aircraft cabin. Compared with the traditional hydraulic or pneumatic actuator, the steering gear 2 does not need complicated pipelines and energy sources such as hydraulic pumps and air pumps to drive, and the system weight is reduced by 20% to 30%, and the space occupied is smaller; the steering gear 2 adopts high-precision gear sets and motors inside, which can work stably in the high-speed flight vibration environment, ensuring the reliability of the locking and deployment of the speed brake 1; at the same time, the steering gear 2 directly drives the speed brake 1 through electrical signals, and the response time is usually in the millisecond level, ensuring that the speed brake 1 is quickly deployed when landing and shortening the taxiing distance; and the steering gear 2 can adjust the deployment angle of the speed brake 1 in real time according to the flight status such as speed and altitude to optimize the deceleration effect; Working principle: When the aircraft enters the deceleration landing state, the control system issues a first control instruction to control the locking mechanism to start so as to release the deceleration plate 1; then, the control system issues a second control instruction to control the servo 2 to start, and the servo 2 drives the deployment mechanism to drive the deceleration plate 1 to deploy, so that the aircraft can achieve deceleration landing under the action of the deceleration plate 1; compared with the prior art, the deployment mechanism of the present invention is directly controlled by the servo 2 and adopts a separate drive scheme, with a simple overall structure and small occupied space, and will not limit the layout and performance improvement of other equipment in the aircraft; at the same time, a special locking mechanism is provided to lock the deceleration plate 1, so as to further improve the safety, stability and reliability during deployment.

[0022] In some embodiments, the locking mechanism includes: a pin puller assembly 3 and a locking groove, the pin puller assembly 3 is fixedly connected to the speed brake 1, and the locking groove is arranged on the inner wall of the aircraft; the pin puller assembly 3 has a first state and a second state, when in the first state, the aircraft is in a normal flight state, and the pin puller assembly 3 is inserted into the locking groove to lock the speed brake 1; when in the second state, the aircraft is in a decelerated landing state, and the pin puller assembly 3 is separated from the locking groove to release the speed brake 1.

[0023] Specifically, Figure 1 As shown, the locking mechanism includes: a pin puller assembly 3 and a locking groove; when the aircraft is in a normal flight state, the pin puller assembly 3 is in a first state, at which time the pin puller assembly 3 is inserted into the locking groove, that is, the entire locking mechanism is located inside the aircraft to lock the deceleration plate 1. The deceleration plate 1 cannot rotate under the action of the pin puller assembly 3, thereby realizing the locking function of the deceleration plate 1 and enhancing the reliability of the deceleration system; Specifically, when the aircraft is decelerating for landing, the control system sends a first control instruction to the pin puller assembly 3. After receiving the first control instruction, the pin puller assembly 3 separates from the locking groove to release the lock of the speed brake 1. At the same time, the speed brake 1 is deployed under the drive of the servo 2.

[0024] In some embodiments, the pin pulling assembly 3 includes: The cylinder 31 has a first cavity 35 inside; The piston rod 32 is arranged to penetrate the cylinder 31 along the first direction; a force-bearing portion is arranged in the middle of the piston rod 32; when the pin pulling assembly 3 is in the first state, one end of the piston rod 32 is inserted into the locking groove; The electric squib 33 is disposed on the cylinder 31 and is located on a side of the force-bearing portion close to the locking groove.

[0025] Specifically, in this embodiment, the first direction is parallel to the extension direction of the piston rod 32; Specifically, Figure 2As shown, when the aircraft is in a normal flight state, the piston rod 32 is in a first state, and one end of the piston rod 32 (i.e. Figure 2 The left end of the piston rod 32 is inserted into the locking groove to lock the deceleration plate 1; and when the aircraft needs to decelerate for landing, after the control system issues a first control command, the electric squib 33 is energized and ignited, generating high-pressure gas in the first cavity 35, and the high-pressure gas acts on the end surface of the force-bearing part to push the piston rod 32 to move away from the locking groove, so that the piston rod 32 can be pulled out from the inside; In some embodiments, the starting current of the electric squib 33 is 5A / 50ms, and at the moment of its operation, a gas pressure of 5.3-7.2MPa can be generated in a closed space with a volume of 5ml; In this embodiment, a first support is provided on the inner wall of the aircraft cabin, and the speed brake 1 is connected to the aircraft through the first support. In this embodiment, when the piston rod 32 is in the first state, one end of the piston rod 32 passes through the first support to be inserted into the locking groove.

[0026] In some embodiments, a shear pin 34 is provided on the cylinder 31 and on the side of the electric squib 33 away from the locking groove, and a first groove is provided on the piston rod 32 corresponding to the shear pin 34. When the pin pulling assembly 3 is in the first state, one end of the shear pin 34 is located in the first groove.

[0027] Specifically, a shear pin 34 is provided on the cylinder 31 and on the side of the electric squib 33 away from the locking groove. When the pin pulling assembly 3 is in the first state, one end of the shear pin 34 is located in the first groove; when the pin pulling assembly 3 is in the first state, the shear pin 34 serves as a physical locking piece, and fixes the piston rod 32 in place by its own strength, to prevent the pin pulling assembly from malfunctioning due to flight vibration, aerodynamic fluctuations or unexpected loads, thereby ensuring that the speed brake 1 is firmly locked during flight; when the electric squib 33 is ignited, the shear pin 34 is accurately sheared off, releasing the movement of the piston rod 32, thereby releasing the lock of the speed brake 1; the fracture threshold of the shear pin 34 has been strictly calculated (such as material selection and diameter optimization), and will only fail under specific conditions (such as when the electric squib 33 is powered on and ignited), thereby avoiding accidental unlocking due to external interference (such as mechanical shock), thereby improving system reliability.

[0028] In some embodiments, a first sealing ring 36 is disposed between the electric squib 33 and the cylinder 31 ; and a second sealing ring 37 is disposed between the piston rod 32 and the cylinder 31 .

[0029] Specifically, a first sealing ring 36 is arranged between the electric squib 33 and the cylinder 31; a second sealing ring 37 is arranged between the piston rod 32 and the cylinder 31. The functions of the first sealing ring 36 and the second sealing ring 37 are to ensure that the gas will not leak from the gap between the electric squib 33 and the cylinder 31, and between the piston rod 32 and the cylinder 31, so as to maintain sufficient thrust to push the piston rod 32 to move; and the sealing rings ensure that the gas pressure is concentrated on the cross section of the piston rod 32 to avoid energy loss and ensure the reliability of the pin pulling action; at the same time, during ultra-high-speed flight, the first sealing ring 36 and the second sealing ring 37 can block the high-temperature airflow from invading the interior of the cylinder 31, thereby protecting the electric squib 33 and the piston rod 32 from thermal damage.

[0030] In this embodiment, the first sealing ring 36 is a Gly ring for dynamic sealing; the second sealing ring 37 is an O-ring for static sealing; In this embodiment, two second sealing rings 37 are provided. The first second sealing ring 37 is referred to as Figure 2 The second sealing ring 37 located on the left side is Figure 2 The second sealing ring 37 is located on the right side.

[0031] In this embodiment, it is determined by experimental calculation that the thrust force after the electric squib 33 is ignited can push or move the piston rod 32, as follows: According to linear calculation, after the electric squib 33 is ignited and detonated, the initial pressure P is 31.1 MPa, and the piston area acted by the gas is The thrust after the electric squib 33 is detonated is: ; The damping force of the piston rod 32 movement mainly comes from the friction between the piston rod 32 and the first support. , the friction between the piston rod 32 and the cylinder 31 , the friction force of the two second sealing rings 37 and , Shear force to cut off the shear pin 34 , so the motion damping force of the piston rod 32 is: ; (a) ,in is the dynamic friction coefficient of contact between steels, which is taken as 0.2; is the pressure between the first support and the piston rod 32, ; Under the action of pneumatic force, the pressure between the piston rod 32 and the first support ; The rebound force of the packing on the speed brake 1 Related to the length of the packing, ; Wherein, L is half of the length of the packing of the speed reducer 1, and K is the elastic coefficient of the packing; in this application, , ; therefore ;and ; therefore

[0032] (b) ; (c) According to the formula The friction force of the second sealing ring 37 is calculated as follows: in, is the inner diameter of the sealing ring, is the cross-sectional diameter of the sealing ring, is the friction coefficient, which is taken as 0.15; p is the sealing medium pressure; In this embodiment, the size of the first second sealing ring 37 is: , , ; Therefore, the first second sealing ring 37 ; The dimensions of the second second sealing ring 37 are: , , ; Therefore, the second second sealing ring 37 ; (d) The shear pin 34 is made of 2A12 aluminum alloy, and the shear force of the shear pin 34 is ; It follows that ; Therefore, the thrust of the electric squib 33 after detonation is greater than the friction force of the piston rod 32, so the piston rod 32 can be smoothly pulled out of the locking groove.

[0033] In some embodiments, the deployment mechanism includes: a connecting member 4, the connecting member 4 has a first connecting portion 41, a second connecting portion 42 and a third connecting portion 43, the first connecting portion 41 is hinged to the inner wall of the aircraft, the second connecting portion 42 is hinged to the driving end of the servo 2, and the third connecting portion 43 is fixedly connected to the speed reducer 1.

[0034] Specifically, Figure 1As shown, the deployment mechanism includes: a connecting member 4. In this embodiment, the connecting member 4 is a crank, and the connecting member 4 has a first connecting portion 41, a second connecting portion 42 and a third connecting portion 43. The first connecting portion 41 is hinged to the inner wall of the aircraft; the second connecting portion 42 is hinged to the driving end of the steering gear 2, and the third connecting portion 43 is fixedly connected to the speed brake 1; when the telescopic rod of the steering gear 2 is retracted, the connecting member 4 rotates counterclockwise, thereby driving the speed brake 1 to deploy; when the telescopic rod of the steering gear 2 is extended, the connecting member 4 rotates clockwise, thereby driving the speed brake 1 to reset to fit with the fuselage of the aircraft; In this embodiment, the opening time of the locking mechanism is determined according to experimental calculation, as follows: ; ; in is the net thrust on the piston rod after the electric squib is detonated, m is the mass of the piston rod, a is the acceleration of the piston rod, S is the distance the piston rod moves, and t is the time required for the locking mechanism to open; In this embodiment, S = 10.5mm, m = 0.069Kg; Therefore, the opening time of the locking mechanism .

[0035] In some embodiments, a heat sealing structure 5 is provided at the connection between the air brake 1 and the aircraft fuselage.

[0036] Specifically, Figure 1 As shown, a heat sealing structure 5 is provided at the connection between the speed reducer 1 and the aircraft fuselage. The heat sealing structure 5 is a raised platform or annular structure. The heat sealing structure 5 is made of high temperature resistant materials, such as nickel-based alloys, ceramic-based composite materials or carbon-carbon composite materials, etc., which can withstand the high temperature generated by air friction during supersonic flight; Specifically, when the speed brake 1 is closed, the edge of the speed brake 1 and the contact surface of the heat sealing structure 5 are compressed by a pre-tightening force to form a physical sealing barrier; under high temperature, the thermal expansion coefficient of the heat sealing structure 5 matches the speed brake 1 to avoid sealing failure due to temperature difference; when the speed brake 1 is deployed, the heat sealing structure 5 is designed to allow dynamic gap adjustment to prevent mechanical interference while maintaining sealing performance.

[0037] In some embodiments, a heat protection coating 6 is provided on a side of the speed brake 1 away from the aircraft.

[0038] Specifically, a heat-resistant coating 6 is provided on the side of the speed brake 1 away from the aircraft. The heat-resistant coating 6 is a ceramic-based composite material or a carbon-carbon composite material with low thermal conductivity. It can effectively block the high temperature generated by air friction from being conducted to the inside of the speed brake 1, thereby protecting the internal structure from thermal damage.

[0039] Example 2 Please refer to Figure 3 As shown, a control method of an aircraft landing deceleration system provided by the present invention is used to control the aircraft landing deceleration system provided by any one of Embodiment 1, comprising the following steps: S1: When the aircraft decelerates to land, a first control instruction is issued, and the first control instruction is used to control the activation of the locking mechanism; S2: Based on the first control instruction, the locking mechanism releases the speed brake; S3: issuing a second control instruction, which is used to control the start of the steering gear; S4: Based on the second control instruction, the servo drives the deployment mechanism to cause the speed brake to deploy, so that the aircraft can achieve deceleration landing under the action of the speed brake.

[0040] Specifically, in this embodiment, the locking mechanism includes: a cylinder 31, a piston rod 32, an electric squib 33, and a shear pin 34; the deployment mechanism includes a connector 4; When the aircraft enters the deceleration landing state, the control system issues a first control instruction. At this time, the electric squib 33 responds to the first control instruction, is powered on and ignites to generate high-pressure gas in the first cavity 35. The high-pressure gas acts on the end face of the force-bearing part of the piston rod 32 to push the piston rod 32 to move away from the locking groove. At the same time, the shear pin 34 is accurately sheared off, releasing the movement of the piston rod 32, so that the piston rod 32 is pulled out of the locking groove, thereby releasing the lock of the deceleration plate 1; subsequently, the control system issues a second control instruction. At this time, the servo 2 responds to the second control instruction, and the telescopic rod of the servo 2 begins to shrink to drive the connecting member 4 to rotate counterclockwise, thereby driving the deceleration plate 1 to unfold, so that the aircraft achieves a deceleration landing under the action of the deceleration plate 1; after the aircraft lands, the control system issues a third control instruction. At this time, the servo 2 responds to the third control instruction and begins to extend to drive the connecting member 4 to rotate clockwise, thereby driving the deceleration plate 1 to retract and fit with the fuselage of the aircraft.

[0041] The above description is only a preferred embodiment of the present invention and an explanation of the technical principles used. Those skilled in the art should understand that the scope of the invention involved in the present invention is not limited to the technical solution formed by a specific combination of the above technical features, but should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the inventive concept. For example, the above features are replaced with the technical features with similar functions disclosed in the present invention (but not limited to) to form a technical solution.

Claims

1. An aircraft landing deceleration system, characterized in that: include: A speed brake (1), the speed brake (1) being arranged on a fuselage in the middle of the aircraft; A steering gear (2), the steering gear (2) being arranged on an inner wall of the aircraft; an unfolding mechanism, the unfolding mechanism being arranged between the driving end of the steering gear (2) and the speed brake (1), the unfolding mechanism being hinged to the steering gear (2) and fixedly connected to the speed brake (1); the unfolding mechanism is used to drive the speed brake (1) to unfold under the drive of the steering gear (2); A locking mechanism is located inside the aircraft and connected between the fuselage of the aircraft and the airbrake (1), and the locking mechanism is used to lock the airbrake (1) when the aircraft is in a normal flight state.

2. An aircraft landing deceleration system according to claim 1, characterized in that: The locking mechanism comprises: a pin pulling assembly (3) and a locking groove, wherein the pin pulling assembly (3) is fixedly connected to the air brake (1), and the locking groove is arranged on the inner wall of the aircraft; the pin pulling assembly (3) has a first state and a second state, wherein when in the first state, the aircraft is in a normal flight state, the pin pulling assembly (3) is inserted into the locking groove to lock the air brake (1); when in the second state, the aircraft is in a decelerated landing state, the pin pulling assembly (3) is separated from the locking groove to release the air brake (1).

3. An aircraft landing deceleration system according to claim 2, characterized in that: The pin pulling assembly (3) comprises: A cylinder (31), wherein the cylinder (31) has a first cavity (35) inside; A piston rod (32), the piston rod (32) being arranged to penetrate the cylinder (31) along a first direction; a force-bearing portion is arranged in the middle of the piston rod (32); when the pin pulling assembly (3) is in a first state, one end of the piston rod (32) is inserted into the locking groove; An electric squib (33), the electric squib (33) being arranged on the cylinder (31) and located on a side of the force-bearing portion close to the locking groove.

4. The aircraft landing deceleration system according to claim 3, characterized in that: A shear pin (34) is provided on the cylinder (31) and located on a side of the electric squib (33) away from the locking groove, and a first groove is provided on the piston rod (32) corresponding to the shear pin (34). When the pin pulling assembly (3) is in the first state, one end of the shear pin (34) is located in the first groove.

5. The aircraft landing deceleration system according to claim 4, characterized in that: A first sealing ring (36) is provided between the electric squib (33) and the cylinder (31); and a second sealing ring (37) is provided between the piston rod (32) and the cylinder (31).

6. The aircraft landing deceleration system according to claim 1, characterized in that: The deployment mechanism comprises: a connecting member (4), the connecting member (4) having a first connecting portion (41), a second connecting portion (42) and a third connecting portion (43), the first connecting portion (41) being hinged to an inner wall of the aircraft, the second connecting portion (42) being hinged to a driving end of the steering gear (2), and the third connecting portion (43) being fixedly connected to the speed brake (1).

7. The aircraft landing deceleration system according to claim 1, characterized in that: A heat sealing structure (5) is provided at the connection between the speed reducer (1) and the aircraft fuselage.

8. The aircraft landing deceleration system according to claim 1, characterized in that: A heat protection coating (6) is provided on a side of the speed brake (1) away from the aircraft.

9. A control method for an aircraft landing deceleration system, used to control the aircraft landing deceleration system according to any one of claims 1 to 8, characterized in that: The steps include: When the aircraft decelerates for landing, a first control instruction is issued, wherein the first control instruction is used to control the activation of the locking mechanism; Based on the first control instruction, the locking mechanism releases the speed brake; issuing a second control instruction, wherein the second control instruction is used to control the start of the steering gear; Based on the second control instruction, the steering gear drives the deployment mechanism to cause the speed brake to deploy, so that the aircraft can achieve decelerated landing under the action of the speed brake.

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