An aircraft landing deceleration system and its control method

Abstract

The present invention provides an aircraft landing deceleration system and its control method, which relates to the technical field of aircraft; the deceleration system includes: a speed brake provided on the fuselage in the middle of the aircraft, and a servo provided on the inner wall of the aircraft; an unfolding mechanism is provided between the speed brake and the driving end of the servo, the unfolding mechanism is hinged to the servo and fixedly connected to the speed brake; the unfolding mechanism is used to drive the speed brake to unfold under the drive of the servo; a locking mechanism is also provided inside the aircraft, and the locking mechanism is connected between the fuselage of the aircraft and the speed brake, and the locking mechanism is used to lock the speed brake when the aircraft is in normal flight; compared with the prior art, the unfolding mechanism of the present invention is directly controlled by the servo, adopts a separate driving scheme, has a simple overall structure, occupies a small space, and does 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, further improving the safety, stability and reliability during unfolding.

Description

Technical Field

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

[0002] A high-speed aircraft is an aircraft capable of having high-speed flight authority, usually referring to 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, improving the efficiency and convenience of air transportation; in practical applications, high-speed aircraft are mainly used for long-haul flights and some short-haul flights, which can significantly shorten the flight time, enabling passengers and goods to reach their destinations in a shorter time; as the speed of high-speed aircraft increases, its speed needs to be reduced before landing, shortening the taxiing distance, reducing the wear and time of the aircraft on the runway, and enabling it to land safely.

[0003] In today's aerospace field, high-speed aircraft are developing rapidly, yet their deceleration systems have drawbacks that cannot be ignored; when facing the critical link of high-speed takeoff and landing, the existing deceleration systems of high-speed aircraft 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, and once a failure occurs, serious consequences may result.

[0004] The existing deployment mechanisms of deceleration systems generally adopt the driving method of actuators. This driving method relies on hydraulic or pneumatic energy sources to operate, which not only makes the structure of the entire system extremely complex, increasing the difficulty of maintenance and repair, but also occupies a large amount of valuable space, restricting the layout and performance improvement of other equipment on the aircraft; in addition, due to the complexity of hydraulic and pneumatic systems themselves, their reliability is relatively low, and faults such as leakage and blockage 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:

[0007] A speed brake, the speed brake is arranged on the fuselage in the middle of the aircraft;

[0008] A servo, the servo is arranged on the inner wall of the aircraft;

[0009] Deployment mechanism, the deployment mechanism is arranged between the driving end of the servo and the speed brake, the deployment mechanism is hinged to the servo and fixedly connected to the speed brake; the deployment mechanism is used to drive the speed brake to deploy under the drive of the servo;

[0010] Locking mechanism, the locking mechanism is located inside the aircraft and is connected between the fuselage of the aircraft and the speed brake, and the locking mechanism is used to lock the speed brake when the aircraft is in a normal flight state.

[0011] According to the technical solution provided by the present invention, the locking mechanism includes: a pin pulling component and a locking groove, the pin pulling component is fixedly connected to the speed brake, and the locking groove is arranged on the inner wall of the aircraft; the pin pulling component has a first state and a second state. When in the first state, the aircraft is in a normal flight state, and the pin pulling component 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 pulling component is separated from the locking groove to release the speed brake.

[0012] According to the technical solution provided by the present invention, the pin pulling component includes:

[0013] Cylinder body, the inside of the cylinder body has a first cavity;

[0014] Piston rod, the piston rod penetrates through the cylinder body along the first direction; a force receiving part is arranged in the middle of the piston rod; when the pin pulling component is in the first state, one end of the piston rod is inserted into the locking groove

[0015] Electric blasting tube, the electric blasting tube is arranged on the cylinder body and is located on the side of the force receiving part close to the locking groove.

[0016] According to the technical solution provided by the present invention, a shear pin is arranged on the cylinder body on the side away from the locking groove of the electric blasting tube, and a first groove is arranged on the piston rod corresponding to the shear pin. When the pin pulling component is in the first state, one end of the shear pin is located in the first groove.

[0017] According to the technical solution provided by the present invention, a first sealing ring is arranged between the electric blasting tube and the cylinder body; a second sealing ring is arranged between the piston rod and the cylinder body.

[0018] According to the technical solution provided by the present invention, the deployment mechanism includes: a connecting piece, the connecting piece has 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.

[0019] According to the technical solution provided by the present invention, a thermal sealing structure is provided at the connection between the speed brake and the aircraft fuselage.

[0020] 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.

[0021] 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 aspect, and includes the following steps:

[0022] When the aircraft decelerates and lands, a first control instruction is issued, and the first control instruction is used to control the locking mechanism to start;

[0023] Based on the first control instruction, the locking mechanism releases the speed brake;

[0024] A second control instruction is issued, and the second control instruction is used to control the servo motor to start;

[0025] Based on the second control instruction, the servo motor drives the deployment mechanism to drive the speed brake to deploy, so that the aircraft realizes decelerated landing under the action of the speed brake.

[0026] The beneficial effects of the present invention are as follows:

[0027] The present invention provides an aircraft landing deceleration system and a control method thereof. The deceleration system includes: a speed brake provided on the fuselage in the middle of the aircraft, and a servo motor provided on the inner wall of the aircraft; a deployment mechanism is provided between the driving end of the speed brake and the servo motor, and the deployment mechanism is hinged to the servo motor and fixedly connected to the speed brake; the deployment mechanism is used to drive the speed brake to deploy under the drive of the servo motor; a locking mechanism is further provided inside the aircraft, and the locking mechanism is connected between the fuselage of the aircraft and the speed brake, and the locking mechanism is used to lock the speed brake when the aircraft is in normal flight; when the aircraft enters the decelerated landing state, the control system issues a first control instruction to control the locking mechanism to start to release the speed brake; subsequently, the control system issues a second control instruction to control the servo motor to start, and the servo motor drives the deployment mechanism to drive the speed brake to deploy, so that the aircraft realizes 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 servo motor 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 speed brake, further improving the safety, stability and reliability during deployment. Description of the Drawings

[0028] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments read in conjunction with the accompanying drawings:

[0029] Figure 1 FIG. is a schematic diagram of a landing deceleration system for an aircraft provided by the present invention;

[0030] Figure 2 FIG.

[0029] is a sectional view of a locking mechanism provided by the present invention;

[0031] Figure 3 FIG. Figure 1 is a flowchart of a control method for a landing deceleration system of an aircraft provided by the present invention.

[0032] In the figures: 1, speed brake; 2, servo; 3, pin pulling assembly; 31, cylinder body; 32, piston rod; 33, electric blasting tube; 34, shear pin; 35, first cavity; 36, first sealing ring; 37, second sealing ring; 4, connecting member; 41, first connecting portion; 42, second connecting portion; 43, third connecting portion; 5, heat sealing structure; 6, heat resistant coating. Detailed Embodiments

[0033] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It can be understood that the specific embodiments described herein are only for explaining the related invention and not for limiting the invention. Additionally, it should be noted that for the sake of description, only parts related to the invention are shown in the drawings.

[0034] It should be noted that, without conflict, the embodiments in 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 drawings and embodiments.

[0035] Embodiment 1

[0036] Please refer to Figure 1 shown in FIG. Figure 1 , a landing deceleration system for an aircraft provided by the present invention includes:

[0037] A speed brake 1, which is arranged on the fuselage in the middle of the aircraft;

[0038] A servo 2, which is arranged on the inner wall of the aircraft;

[0039] An unfolding mechanism, which is arranged between the driving end of the servo 2 and the speed brake 1. The unfolding mechanism is hinged to the servo 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 servo 2;

[0040] A locking mechanism, which is located inside the aircraft and connected between the fuselage of the aircraft and the speed brake 1. The locking mechanism is used to lock the speed brake 1 when the aircraft is in a normal flight state.

[0041] Specifically, the speed brake 1 is provided on the fuselage in the middle of the aircraft, and its rotational torque is provided by the servo 2 in the aircraft cabin. Compared with traditional hydraulic or pneumatic actuators, the servo 2 does not require complex pipelines and energy sources such as hydraulic pumps and air pumps, reducing the system weight by 20% - 30% and occupying less space; the internal of the servo 2 adopts a high-precision gear set and motor, 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 servo 2 directly drives the speed brake 1 through electrical signals, and the response time is usually in milliseconds, ensuring that the speed brake 1 can be deployed quickly during landing and shortening the taxiing distance; moreover, the servo 2 can adjust the deployment angle of the speed brake 1 in real time according to flight states such as speed and altitude to optimize the deceleration effect;

[0042] Working principle: When the aircraft enters the decelerated landing state, the control system issues a first control instruction to control the locking mechanism to start and release the speed brake 1; subsequently, 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 speed brake 1 to deploy, so that the aircraft realizes decelerated landing under the action of the speed brake 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 speed brake 1, further improving the safety, stability and reliability during deployment.

[0043] In some embodiments, the locking mechanism includes: a pin pulling component 3 and a locking groove. The pin pulling component 3 is fixedly connected to the speed brake 1, and the locking groove is arranged on the inner wall of the aircraft; the pin pulling component 3 has a first state and a second state. When in the first state, the aircraft is in the normal flight state, and the pin pulling component 3 is inserted into the locking groove to lock the speed brake 1; when in the second state, the aircraft is in the decelerated landing state, and the pin pulling component 3 is separated from the locking groove to release the speed brake 1.

[0044] Specifically, as Figure 1 shown, the locking mechanism includes: a pin pulling component 3 and a locking groove; when the aircraft is in the normal flight state, the pin pulling component 3 is in the first state. At this time, the pin pulling component 3 is inserted into the locking groove, that is, the entire locking mechanism is located inside the aircraft to lock the speed brake 1. Under the action of the pin pulling component 3, the speed brake 1 cannot rotate the shaft, realizing the locking function of the speed brake 1 and enhancing the reliability of the deceleration system;

[0045] Specifically, when the aircraft decelerates for landing, the control system sends a first control instruction to the pin pulling component 3. After receiving the first control instruction, the pin pulling component 3 is separated from the locking groove to release the locking of the speed brake 1. At the same time, the speed brake 1 is deployed under the drive of the servo 2.

[0046] In some embodiments, the pin pulling assembly 3 includes:

[0047] A cylinder body 31, the interior of the cylinder body 31 having a first cavity 35;

[0048] A piston rod 32, the piston rod 32 being disposed through the cylinder body 31 along a first direction; a force receiving portion is provided 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;

[0049] An electric detonator 33, the electric detonator 33 being provided on the cylinder body 31 and located on one side of the force receiving portion close to the locking groove.

[0050] Specifically, in this embodiment, the first direction is parallel to the extending direction of the piston rod 32;

[0051] Specifically, as Figure 2 shown, when the aircraft is in a normal flight state, the piston rod 32 is in the first state, and one end of the piston rod 32 (i.e., Figure 2 the left end in is inserted into the locking groove to lock the speed reducer 1; and when the aircraft needs to decelerate for landing, after the control system issues a first control instruction, the electric detonator 33 is energized and ignited, generating high-pressure gas in the first cavity 35, and the high-pressure gas acts on the end face of the force receiving portion to push the piston rod 32 to move in a direction away from the locking groove, realizing the withdrawal of the piston rod 32;

[0052] In some embodiments, the starting current of the electric detonator 33 is 5A / 50ms, and at the moment of its operation, it can generate a gas pressure of 5.3 - 7.2 MPa in a closed space with a volume of 5 ml;

[0053] In this embodiment, a first support is provided on the inner wall of the cabin of the aircraft, and the speed reducer 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.

[0054] In some embodiments, a shear pin 34 is provided on the cylinder body 31 and on the side of the electric detonator 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.

[0055] Specifically, a shear pin 34 is provided on the cylinder body 31 and on the side of the electric detonator 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 acts as a physical locking member and fixes the piston rod 32 in place by its own strength, preventing misoperation of the pin pulling assembly due to flight vibration, aerodynamic force fluctuation or accidental load, and ensuring that the speed reducer 1 is firmly locked during flight. When the electric detonator 33 is ignited, the shear pin 34 is precisely cut, releasing the movement of the piston rod 32, thereby releasing the lock of the speed reducer 1. The fracture threshold of the shear pin 34 is strictly calculated (such as material selection, diameter optimization) and will only fail under specific conditions (such as the electric detonator 33 being powered on and ignited), avoiding accidental unlocking due to external interference (such as mechanical impact) and improving the reliability of the system.

[0056] In some embodiments, a first sealing ring 36 is provided between the electric detonator 33 and the cylinder body 31; a second sealing ring 37 is provided between the piston rod 32 and the cylinder body 31.

[0057] Specifically, a first sealing ring 36 is provided between the electric detonator 33 and the cylinder body 31; a second sealing ring 37 is provided between the piston rod 32 and the cylinder body 31. The functions of the first sealing ring 36 and the second sealing ring 37 are to ensure that gas does not leak from the gaps between the electric detonator 33 and the cylinder body 31, and between the piston rod 32 and the cylinder body 31, maintaining sufficient thrust to push the piston rod 32 to move; and the sealing rings ensure that the gas pressure acts concentratedly on the cross-section of the piston rod 32, avoiding energy loss and ensuring 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 intrusion of high-temperature air flow into the interior of the cylinder body 31, protecting the electric detonator 33 and the piston rod 32 from heat damage.

[0058] In this embodiment, the first sealing ring 36 is a Gleitring for dynamic sealing; the second sealing ring 37 is an O-ring for static sealing.

[0059] In this embodiment, a total of two second sealing rings 37 are provided. The first second sealing ring 37 in the following text refers to Figure 2 the second sealing ring 37 located on the left side in Figure 2 and the second second sealing ring 37 refers to

[0060] the second sealing ring 37 located on the right side in.

[0061] According to linear calculation, after the electric detonator 33 is ignited and detonated, the initial pressure P is 31.1 MPa, and the piston area acted on by the gas is ; the thrust after the electric detonator 33 is detonated is: ;

[0062] The damping force during the movement of the piston rod 32 mainly comes from the frictional force between the piston rod 32 and the first support , the frictional force between the piston rod 32 and the cylinder body 31 , the frictional force of the two second sealing rings 37 and , the shearing force for shearing the shear pin 34 . Therefore, the movement damping force of the piston rod 32 is as follows: ;

[0063] (a) , where is the dynamic friction coefficient of the contact between steels, taking 0.2; is the pressure between the first support and the piston rod 32, ;

[0064] Under the action of the aerodynamic force, the pressure between the piston rod 32 and the first support ;

[0065] The resilience force of the packing acting on the retarder 1 is related to the length of the packing, ;

[0066] wherein, L is half of the length of the packing of the retarder 1, and K is the elastic coefficient of the packing; in this application , ;

[0067] Therefore ; and ;

[0068] Therefore

[0069] (b) ;

[0070] (c) Calculate the frictional force of the second sealing ring 37 according to the formula :

[0071] wherein, is the inner diameter of the sealing ring, is the cross-sectional diameter of the sealing ring, is the friction coefficient, taking 0.15; p is the pressure of the sealing medium;

[0072] In this embodiment, the dimensions of the first second sealing ring 37 are: , , ;

[0073] Therefore, the first second sealing ring 37 ;

[0074] The size of the second second sealing ring 37 is as follows: , , ;

[0075] Therefore, the second second sealing ring 37 ;

[0076] (d) The shear pin 34 is made of 2A12 aluminum alloy, and the shear force of the shear pin 34 ;

[0077] From this, it can be obtained that ;

[0078] , so the thrust after the electric detonator 33 detonates is greater than the frictional force of the piston rod 32, and therefore the piston rod 32 can be smoothly pulled out from the locking groove.

[0079] 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 steering gear 2, and the third connecting portion 43 is fixedly connected to the spoiler 1.

[0080] Specifically, as Figure 1 shown, the deployment mechanism includes: a connecting member 4. In this embodiment, the connecting member 4 is a crank. 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 spoiler 1; when the telescopic rod of the steering gear 2 contracts, the connecting member 4 rotates counterclockwise, thereby driving the spoiler 1 to deploy; when the telescopic rod of the steering gear 2 extends, the connecting member 4 rotates clockwise, thereby driving the spoiler 1 to reset to fit with the fuselage of the aircraft;

[0081] In this embodiment, the opening time of the locking mechanism is determined according to experimental calculations, specifically as follows:

[0082] ;

[0083] ;

[0084] where is the net thrust received by the piston rod after the electric detonator detonates, m is the mass of the piston rod, a is the acceleration of the piston rod movement, S is the distance the piston rod moves, and t is the time required for the locking mechanism to open;

[0085] In this embodiment, S = 10.5 mm, m = 0.069 Kg;

[0086] Therefore, the opening time of the locking mechanism 。

[0087] In some embodiments, a thermal sealing structure 5 is provided at the connection between the speed brake 1 and the aircraft fuselage.

[0088] Specifically, as Figure 1 shown, a thermal sealing structure 5 is provided at the connection between the speed brake 1 and the aircraft fuselage. The thermal sealing structure 5 is a raised platform or a ring structure. The thermal sealing structure 5 is made of a high-temperature resistant material, such as nickel-based alloy, ceramic matrix composite material, or carbon-carbon composite material, etc., and can withstand the high temperature generated by air friction during supersonic flight.

[0089] Specifically, when the speed brake 1 is closed, the contact surface between the edge of the speed brake 1 and the thermal sealing structure 5 is pressed tightly by a pre-tightening force to form a physical sealing barrier. At high temperatures, the thermal expansion coefficient of the thermal sealing structure 5 matches that of the speed brake 1 to avoid seal failure caused by temperature difference. When the speed brake 1 is deployed, the thermal sealing structure 5 is designed to allow dynamic clearance adjustment to prevent mechanical interference while maintaining the sealing performance.

[0090] In some embodiments, a thermal protection coating 6 is provided on the side of the speed brake 1 away from the aircraft.

[0091] Specifically, a thermal protection coating 6 is provided on the side of the speed brake 1 away from the aircraft. The thermal protection coating 6 is a ceramic matrix composite material or a carbon-carbon composite material, which has a low thermal conductivity and can effectively block the conduction of the high temperature generated by air friction to the inside of the speed brake 1, protecting the internal structure from heat damage.

[0092] Embodiment 2

[0093] Please refer to Figure 3 shown. A control method for an aircraft landing deceleration system provided by the present invention is used to control the aircraft landing deceleration system provided in any one of Embodiment 1, and includes the following steps:

[0094] S1: When the aircraft decelerates for landing, a first control instruction is issued, and the first control instruction is used to control the locking mechanism to start.

[0095] S2: Based on the first control instruction, the locking mechanism releases the speed brake.

[0096] S3: A second control instruction is issued, and the second control instruction is used to control the servo motor to start.

[0097] S4: Based on the second control instruction, the servo motor drives the deployment mechanism to drive the speed brake to deploy, so that the aircraft realizes decelerated landing under the action of the speed brake.

[0098] Specifically, in this embodiment, the locking mechanism includes: a cylinder body 31, a piston rod 32, an electric blasting tube 33, and a shear pin 34; the deployment mechanism includes a connecting member 4.

[0099] When the aircraft enters the decelerating landing state, the control system issues a first control instruction. At this time, the electric detonator 33 responds to the first control instruction, is energized and ignited to generate high-pressure gas in the first cavity 35. The high-pressure gas acts on the end face of the force-receiving 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 cut off, releasing the movement of the piston rod 32, realizing the extraction of the piston rod 32 from the locking groove, thereby releasing the locking of the speed brake 1. Subsequently, the control system issues a second control instruction. At this time, the steering gear 2 responds to the second control instruction, and the telescopic rod of the steering gear 2 starts to contract, driving the connecting member 4 to rotate counterclockwise, and then driving the speed brake 1 to unfold, so that the aircraft realizes decelerating landing under the action of the speed brake 1. When the aircraft lands, the control system issues a third control instruction. At this time, the steering gear 2 responds to the third control instruction and starts to extend, driving the connecting member 4 to rotate clockwise, and then driving the speed brake 1 to retract and fit with the fuselage of the aircraft.

[0100] The above description is only a preferred embodiment of the present invention and an explanation of the applied technical principles. 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 the specific combination of the above technical features, and 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 technical solution formed by mutually replacing the above features with the (but not limited to) technical features with similar functions disclosed in the present invention.

Claims

1. An aircraft landing deceleration system, characterized in that, Comprising: A spoiler (1), the spoiler (1) being provided on the fuselage in the middle of the aircraft; A servo (2), the servo (2) being provided on the inner wall of the aircraft; An unfolding mechanism, the unfolding mechanism being provided between the driving end of the servo (2) and the spoiler (1), the unfolding mechanism being hinged to the servo (2) and fixedly connected to the spoiler (1); the unfolding mechanism is used to drive the spoiler (1) to unfold under the drive of the servo (2); A locking mechanism, the locking mechanism being located inside the aircraft and connected between the fuselage of the aircraft and the spoiler (1), the locking mechanism being used to lock the spoiler (1) when the aircraft is in a normal flight state; the locking mechanism includes: a pin-pulling assembly (3) and a locking groove, the pin-pulling assembly (3) being fixedly connected to the spoiler (1), the locking groove being provided on the inner wall of the aircraft; the pin-pulling assembly (3) has a first state and a second state, 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 spoiler (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 spoiler (1); The pin-pulling assembly (3) includes: A cylinder body (31), the interior of the cylinder body (31) having a first cavity (35); A piston rod (32), the piston rod (32) being arranged to penetrate the cylinder body (31) along a first direction; a force-receiving portion is provided 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; An electric detonator (33), the electric detonator (33) being provided on the cylinder body (31) and located on the side of the force-receiving portion close to the locking groove; A shear pin (34) is provided on the cylinder body (31) and on the side of the electric detonator (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.

2. The aircraft landing deceleration system according to claim 1, wherein A first sealing ring (36) is provided between the electric detonator (33) and the cylinder body (31); a second sealing ring (37) is provided between the piston rod (32) and the cylinder body (31).

3. A landing deceleration system for an aircraft according to claim 1, characterized in that, The unfolding mechanism includes: 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 the inner wall of the aircraft, the second connecting portion (42) being hinged to the driving end of the servo (2), and the third connecting portion (43) being fixedly connected to the spoiler (1).

4. A landing deceleration system for an aircraft according to claim 1, characterized in that, A heat-sealing structure (5) is provided at the connection between the spoiler (1) and the fuselage of the aircraft.

5. A landing deceleration system for an aircraft according to claim 1, characterized in that, A heat-resistant coating (6) is provided on the side of the spoiler (1) away from the aircraft.

6. A control method for an aircraft landing deceleration system, used to control the aircraft landing deceleration system according to any one of claims 1-5, characterized in that, Including the following steps: When the aircraft decelerates and lands, a first control command is issued, and the first control command is used to control the locking mechanism to start; Based on the first control command, the locking mechanism releases the speed brake; A second control command is issued, and the second control command is used to control the servo to start; Based on the second control command, the servo drives the deployment mechanism to drive the speed brake to deploy, so that the aircraft realizes decelerated landing under the action of the speed brake.

Images