Rudder piece hinge moment adjusting method, hinge moment adjustable rudder piece and control method

Abstract

The invention provides a rudder blade hinge moment adjusting method, a hinge moment adjustable rudder blade and a control method thereof, and solves the problem of overlarge hinge moment in the flight process of an aircraft by changing the relative position between the rudder blade and a rudder shaft. The rudder shaft is a movable rudder shaft and can move within a set range in the chord length direction of the rudder sheet; therefore, the relative position between the rudder piece and the rudder shaft is changed (namely, the position of the rudder shaft relative to the rudder piece is changed), the distance between the resultant force point and the rudder shaft is changed, the hinge moment is changed, and the problem that the hinge moment is too large in the flight process of an aircraft is solved.

Description

Technical Field

[0001] The invention relates to a rudder blade, in particular to a hinge torque adjustable rudder blade, belonging to the technical field of aircraft. Background Art

[0002] When the aircraft changes its attitude during flight, it is necessary to change the deflection angle of the air rudder blade to generate additional force on the rudder blade, thereby changing the attitude of the aircraft. The air rudder needs to deflect during flight, so it is connected to the aircraft by a rudder shaft. The aerodynamic force received by the air rudder during flight can be equivalent to a resultant force at the resultant point. The resultant force can be divided into three components along the chord length direction, the direction perpendicular to the chord length direction, and the direction perpendicular to the rudder surface. Among them, the components along the chord length direction and the direction perpendicular to the chord length direction are small and will not affect the design. The component in the direction perpendicular to the rudder surface is much larger than the components in the other two directions, and is an important part of changing the attitude of the aircraft and meeting the attitude control capability. The product of the component in the direction perpendicular to the rudder surface and the distance from the resultant point to the rudder shaft is the hinge torque. Since the position of the resultant point changes with the deflection of the air rudder, the hinge torque changes from time to time during flight. The hinge torque acts on the rudder shaft and cannot exceed the maximum output load of the servo, otherwise it will cause the servo to fail and destroy the control capability of the aircraft.

[0003] During the flight of an aircraft, there are very few conditions where the hinge torque exceeds the maximum output load of the servo, but as long as it exists, there must be a solution. During the design process, in order to prevent the servo from failing due to excessive hinge torque, two methods are usually used. One is to select a servo with a larger output power when selecting the servo to ensure that the hinge torque will not be overloaded. The greater the power of the servo, the higher the price will be, which will lead to a sharp increase in production costs; the other method is to modify and optimize the flight control program of the aircraft so that the aircraft will not have a large rudder deflection angle during flight, thereby reducing the size of the hinge torque. This method will not only increase the design difficulty, but also reduce the aircraft's attitude control ability, which will ultimately affect the aircraft's flight distance, maneuverability and other technical indicators. Summary of the invention

[0004] In view of this, the present invention provides a method for adjusting the hinge torque of a rudder blade, which solves the problem of excessive hinge torque during the flight of an aircraft by changing the relative position between the rudder blade and the rudder shaft.

[0005] The technical solution of the present invention is: a method for adjusting the hinge torque of a rudder blade, wherein the rudder shaft is a movable rudder shaft, which can move within a set range along the chord length direction of the rudder blade, thereby changing the position of the rudder shaft relative to the rudder blade, and then changing the distance from the aerodynamic resultant point to which the rudder blade is subjected during flight to the rudder shaft, so as to achieve a change in the hinge torque.

[0006] As a preferred embodiment of the present invention: the two extreme positions of the movement of the rudder shaft along the chord length direction are respectively the left extreme position and the right extreme position, the right extreme position ensures that when the rudder shaft is in the right extreme position, when the resultant force point is on the right side of the rudder shaft, the hinge torque does not exceed the maximum output load of the servo; the left extreme position ensures that when the rudder shaft is in the left extreme position, when the resultant force point is on the left side of the rudder shaft, the hinge torque does not exceed the maximum output load of the servo.

[0007] As a preferred embodiment of the present invention: two installation positions of the rudder shaft are set inside the rudder blade along the chord length direction of the rudder blade, and the rudder shaft can be switched between the two installation positions; when the rudder shaft is in one of the installation positions, it is ensured that the hinge torque does not exceed the maximum output load of the servo when the resultant force point is on the side of the installation position.

[0008] In addition, the present invention provides a hinge torque adjustable rudder blade, comprising: a rudder blade body, a linear drive mechanism, a rudder shaft and a cover plate;

[0009] A rectangular groove is provided on the rudder blade body at the installation position of the rudder shaft, serving as an installation cavity for the linear drive mechanism and the rudder shaft;

[0010] The installation cavity is provided with a rudder shaft moving groove and a linear drive mechanism installation groove; the rudder shaft is installed in the rudder shaft moving groove; the linear drive mechanism is installed in the linear drive mechanism installation groove; the actuating end of the linear drive mechanism is connected to the rudder shaft, and is used to drive the rudder shaft to move in the rudder shaft moving groove along the chord length direction of the rudder blade;

[0011] The cover plate is used to close the installation cavity.

[0012] As a preferred embodiment of the present invention: the linear drive mechanism is an electric linear drive mechanism powered by a battery on the aircraft; and the linear drive mechanism has a self-locking property.

[0013] As a preferred embodiment of the present invention, the upper portion of the rudder shaft is a cubic structure; the lower portion is a cylinder, and is fixedly connected to the rudder on the aircraft.

[0014] As a preferred embodiment of the present invention: the rudder shaft moving groove is formed by connecting the rudder shaft moving groove A arranged in the installation cavity and the rudder shaft moving groove B arranged on the cover plate; the linear drive mechanism installation groove is formed by connecting the linear drive mechanism installation groove A arranged in the installation cavity and the linear drive mechanism installation groove B arranged on the cover plate.

[0015] Furthermore, the present invention also provides a control method for the above-mentioned hinge torque adjustable rudder blade, wherein the rudder shaft moving groove ensures that: when the rudder shaft is located at the rightmost end position in the rudder shaft moving groove and the resultant force point is on the right side of the rudder shaft, the hinge torque never exceeds the maximum output load of the servo; when the rudder shaft is located at the leftmost end position in the rudder shaft moving groove and the resultant force point is on the left side of the rudder shaft, the hinge torque never exceeds the maximum output load of the servo;

[0016] During the flight of the aircraft, initially, the rudder shaft is located at the end of one end of the rudder shaft moving groove; when the difference between the maximum output load of the servo and the hinge torque of the rudder shaft is less than a set threshold, a rudder shaft position adjustment signal is sent to the flight control computer. After receiving the rudder shaft position adjustment signal, the flight control computer controls the linear drive mechanism to drive the rudder shaft to move to the end of the other end of the rudder shaft moving groove.

[0017] Beneficial effects:

[0018] (1) The rudder blade hinge torque adjustment method of the present invention solves the problem of excessive hinge torque during the flight of an aircraft by changing the relative position between the rudder blade and the rudder shaft. Compared with traditional solutions, it can reduce the production cost of the servo. Moreover, during the flight of the aircraft, there is only a small relative displacement between the rudder shaft and the rudder blade, which will not affect the aerodynamic shape and flight control program of the entire aircraft.

[0019] (2) The hinge torque adjustable rudder blade of the present invention adopts a linear drive mechanism to directly drive the rudder shaft to achieve relative displacement between the rudder shaft and the rudder blade, and has a simple structure.

[0020] (3) The linear drive mechanism of the present invention only needs to be powered during displacement, does not need to consume more electricity, does not need to add additional battery components, and has a simple structure and low cost.

[0021] (4) The linear drive mechanism in the present invention has a power-off self-locking function. Since a large hinge torque will only occur under special conditions during flight, there is no need to frequently start the linear drive mechanism to move the rudder shaft, which can greatly save battery power on the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 The overall layout and force analysis diagram of the rudder blade of the present invention; (x1-x2 in the figure is the range of change of the resultant force point F)

[0023] Figure 2 This is a diagram of the structure of the rudder blade of the present invention;

[0024] Figure 3 This is the end cover structure diagram.

[0025] Wherein: 1-rudder blade body; 2-threaded hole A; 3-linear drive mechanism; 4-rudder shaft; 5-cover plate; 6-rudder shaft moving groove A; 7-threaded hole B; 8-linear drive mechanism mounting groove A; 9-cover plate through hole; 10-linear drive mechanism mounting groove B; 11-rudder shaft moving groove B. DETAILED DESCRIPTION

[0026] The present invention will be further described in detail below in conjunction with the accompanying drawings and embodiments.

[0027] Embodiment 1:

[0028] The present embodiment provides a method for adjusting the hinge torque of a rudder blade. When the hinge torque is too large during the flight of an aircraft, the hinge torque can be reduced without affecting the appearance and aerodynamic parameters of the aircraft and significantly increasing the cost, thereby meeting the maximum output load of the servo and ensuring that the aircraft continues to work according to the predetermined flight control program.

[0029] The method for adjusting the hinge torque of the rudder blade proposed in this example is specifically as follows: the rudder shaft 4 is a movable rudder shaft, which can move within a set range along the chord length direction of the rudder blade 1 (the rudder shaft 4 can only move along the chord length direction, limiting displacement and rotation in other directions); thereby achieving a change in the relative position between the rudder blade 1 and the rudder shaft 4 (i.e., changing the position of the rudder shaft 4 relative to the rudder blade 1), and then changing the distance from the resultant force point to the rudder shaft, so as to achieve a change in the hinge torque and solve the problem of excessive hinge torque during the flight of the aircraft.

[0030] The two limit positions for the rudder shaft 4 to move along the chord length direction are the left limit position and the right limit position, wherein the right limit position ensures that when the rudder shaft is in the right limit position, when the resultant force point of the aerodynamic force received by the rudder blade is on the right side of the rudder shaft 4 (referring to the right side of the current position of the rudder shaft 4), the maximum output load of the servo always meets the requirement of the hinge torque (i.e., the hinge torque never exceeds the maximum output load of the servo); the left limit position ensures that when the rudder shaft is in the left limit position, when the resultant force point of the aerodynamic force received by the rudder blade is on the left side of the rudder shaft 4, the maximum output load of the servo always meets the requirement of the hinge torque (i.e., the hinge torque never exceeds the maximum output load of the servo). The above two limit positions can be designed according to the rudder shaft position design method.

[0031] It can be seen from this that when the rudder shaft 4 is in the right extreme position and the hinge torque exceeds the maximum output load of the servo, the resultant point of the aerodynamic force acting on the rudder blade must be on the left side of the rudder shaft 4; when the rudder shaft 4 is in the left extreme position and the hinge torque exceeds the maximum output load of the servo, the resultant point of the aerodynamic force acting on the rudder blade must be on the right side of the rudder shaft 4.

[0032] Initially, the rudder shaft 4 is at the extreme position on one side of its moving range (such as the right extreme position). During the flight of the aircraft, the servo monitors the instantaneous output load in real time (the hinge torque acting on the servo shaft will increase the output load of the servo, and the hinge torque and the servo output load are related to each other. It can be considered that the hinge torque is the same as the servo output load. Therefore, it is actually only necessary to monitor the output load of the servo to ensure that the output load does not exceed the maximum value); when the hinge torque does not exceed the maximum output load of the servo, the rudder shaft 4 does not move; when the hinge torque received by the rudder shaft 4 is close to the maximum output load of the servo (that is, when the difference between the maximum output load of the servo and the hinge torque received by the rudder shaft is less than the set threshold), the rudder shaft is controlled to move along the chord length direction to the extreme position on the other side (such as moving from the initial right extreme position to the left extreme position); thereby ensuring that the maximum output load of the servo always meets the requirements of the hinge torque (that is, the hinge torque never exceeds the maximum output load of the servo).

[0033] Therefore, this method can make the rudder shaft move to two positions in the resultant force point area (based on the limitations of the two left and right extreme positions, it can be known that the movement range of the rudder shaft must be within the change area of ​​the resultant force point), so that the hinge torque on the rudder shaft 4 is always within the output control capability of the servo.

[0034] Further, since the maximum hinge moment will only appear at the left and right ends of the rudder shaft 4, in theory, the rudder shaft 4 only needs to have two positions to meet the entire flight condition. Based on this, as an example, two installation positions of the rudder shaft 4 are set inside the rudder blade 1 along the chord length direction of the rudder blade 1 (which can be understood as the above-mentioned right limit position and left limit position), and the rudder shaft 4 can switch positions in the two installation positions. During the initial flight, the rudder shaft 4 is in one of the installation positions, and when the resultant force point is on the side where the installation position is located, the maximum output load of the servo always meets the requirements of the hinge moment; when the resultant force point is on the side where the other installation position is located, the rudder shaft 4 does not need to move without exceeding the output load of the servo; when the hinge moment is close to the maximum output load of the servo, the rudder shaft 4 is switched to another installation position to ensure that when the resultant force point is on the side where the current installation position is located, the maximum output load of the servo always meets the requirements of the hinge moment. Finally, the design requirements of the entire flight condition are met. Since the displacement change of the resultant force point on the rudder blade is small, the left and right displacement of the rudder shaft is small, which will not affect the overall appearance of the aircraft.

[0035] Embodiment 2:

[0036] This embodiment provides a hinge torque adjustable rudder blade capable of implementing the method of the above-mentioned embodiment 1.

[0037] The hinge torque adjustable rudder blade adopts a linear drive mechanism installed inside the rudder blade to drive the rudder shaft 4 to move, thereby realizing the change of the relative position between the rudder blade 1 and the rudder shaft 4, and solving the problem of excessive hinge torque during the flight of the aircraft; the method of built-in grooves in the rudder blade 1 and the cover plate 5 is adopted to solve the problem of internal installation space and fixing method of the solenoid valve 3 and the rudder shaft 4; the top of the rudder shaft 4 adopts a cubic structure, so that the rudder shaft 4 can only move along the chord length direction, limiting displacement and rotation in other directions.

[0038] like Figure 1 and Figure 2 As shown, the hinge torque adjustable rudder blade provided in this embodiment includes: a rudder blade body 1, a linear drive mechanism 3, a rudder shaft 4 and a cover plate 5.

[0039] On the rudder body 1, a rectangular groove is provided at the installation position of the rudder shaft 4, which serves as an installation cavity for the linear drive mechanism and the rudder shaft 4; the opening of the installation cavity is closed by a cover plate 5. Specifically, threaded holes 2 are distributed at the four corners of the installation cavity, which are used for threaded fixed connection with the cover plate 5, and the cover plate through holes 9 are processed at the four corners of the cover plate 5; thus, the cover plate 5 is fixed to the rudder body 1 by fasteners that cooperate with the cover plate through holes 9 and the threaded holes 2.

[0040] The installation cavity is provided with a rudder shaft moving groove and a linear drive mechanism installation groove; wherein the rudder shaft moving groove is used for the installation and movement of the rudder shaft 4; the movement direction of the rudder shaft 4 in the rudder shaft moving groove is along the chord length direction of the rudder blade 1; Figure 1 and Figure 2 As shown, the rudder shaft 4 can move left and right in the rudder shaft moving groove. The linear drive mechanism installation groove is used for installing the linear drive mechanism 3. The output mode of the actuating end of the linear drive mechanism 3 is linear movement. The actuating end of the linear drive mechanism 3 is connected to the rudder shaft 4 and is used to drive the rudder shaft 4 to move left and right in the rudder shaft moving groove.

[0041] As an example, the linear drive mechanism 3 is an electric push rod, and its actuating end is a push rod; the linear drive mechanism 3 can also be a linear motor (in this case, its actuating end is the motor output shaft) or an electromagnetic valve with linear motion of the valve body (in this case, its actuating end is the valve body of the electromagnetic valve).

[0042] The cross-sectional shape of the rudder shaft moving groove is the same as that of the rudder shaft 4, so that the rudder shaft 4 can only move relative to the rudder blade 1 in the rudder shaft moving groove along the chord length direction of the rudder blade 1. As an example, the upper part of the rudder shaft 4 is a cubic structure, which is used to make the rudder shaft 4 only displace in the chord length direction in the rudder shaft moving groove, and the displacement and rotation in other directions will not move due to the restriction of the cubic structure and the groove. Based on this, the groove at the upper end of the rudder shaft moving groove is deeper, which is used to place the top cubic part of the rudder shaft 4, and its function is to limit the relative displacement of the rudder shaft 4 and the rudder blade 1 in the vertical direction. The rudder shaft moving groove is used for the chord length direction between the rudder shaft 4 and the rudder blade 1 (see Figure 2). Figure 1 The rudder shaft 4 can move at the left and right ends of the rudder shaft moving groove to ensure that the hinge torque of the aircraft always meets the output load of the steering gear during the entire flight process. The side of the cube structure has a threaded hole B7, and the actuating end of the linear drive mechanism 3 has a thread, and the two can be fixedly connected through the thread; the lower part of the rudder shaft 4 is cylindrical and fixedly connected to the steering gear on the aircraft.

[0043] The linear drive mechanism installation groove is located on one side of the rudder shaft moving groove and is used to install the linear drive mechanism; the size of the linear drive mechanism installation groove is consistent with the size of the selected linear drive mechanism 3, so that the linear drive mechanism 3 will not move inside the linear drive mechanism installation groove.

[0044] The linear drive mechanism 3 is an electric linear drive mechanism, which is powered by a battery on the aircraft, and the output mode is linear movement. When powered on, the actuating end of the linear drive mechanism 3 moves linearly, thereby causing the rudder shaft 4 to move relatively in the chord length direction inside the rudder blade 1. In addition, the linear drive mechanism 3 should have a power-off self-locking function. When the rudder shaft 4 does not need to move, the power supply to the linear drive mechanism 3 is disconnected, and the linear drive mechanism 3 is self-locked to ensure that the position of the rudder shaft 4 remains unchanged. Since the hinge torque is large only under special working conditions during flight, there is no need to frequently start the linear drive mechanism 3 to move the rudder shaft 4, which can greatly save the battery power on the aircraft.

[0045] like Figure 2 and Figure 3 As shown, as an example, the rudder shaft moving groove is formed by connecting the rudder shaft moving groove A6 set in the installation cavity and the rudder shaft moving groove B11 set on the cover plate 5; the linear drive mechanism installation groove is formed by connecting the linear drive mechanism installation groove A8 set in the installation cavity and the linear drive mechanism installation groove B10 set on the cover plate 5.

[0046] The threaded holes, through holes, grooves, etc. in the above structure can be processed by machining; the structure is simple, the processing technology is good, the cost is low, and the reliability is high.

[0047] During installation, first separate the rudder blade 1 from the cover plate 5; thread and fix the actuating end of the linear drive mechanism 3 with the rudder shaft 4; place the rudder shaft 4 of the fixedly connected linear drive mechanism 3 in the rudder shaft moving groove A6 and the electromagnetic valve installation groove A8 on the rudder blade 1; initially, the rudder shaft 4 is at the rightmost end of the rudder shaft moving groove A6; install the cover plate 5 on the rudder blade 1 accordingly, and fix and connect them through the threaded hole 2 and the through hole 9. When the linear drive mechanism 3 is not powered on, the rudder shaft 4 is fixed in the rudder blade 1, and there is no relative position movement.

[0048] During the flight of the aircraft, the rudder shaft 4 is at the rightmost end of the rudder shaft moving groove and is fixed in position. At this time, when the resultant force point is on the right side of the rudder shaft, the maximum output load of the servo always meets the hinge moment requirement. When the resultant force point is on the left side, the rudder shaft 4 does not need to move without exceeding the maximum output load of the servo. However, when the aircraft needs to fly in harsh conditions, the resultant force point is on the left side of the rudder shaft, and the hinge torque on the rudder shaft 4 continues to increase, if the hinge torque increases to close to the maximum output load of the servo (that is, the difference between the two is less than the set threshold), a rudder shaft position adjustment signal is sent to the flight control computer. After receiving the rudder shaft position adjustment signal, the flight control computer immediately sends a power supply signal to the linear drive mechanism 3 to power the linear drive mechanism 3. The linear drive mechanism 3 drives the rudder shaft 4 to move to the leftmost position of the rudder shaft moving groove, so that the hinge torque is always within the output load capacity of the servo. After that, the linear drive mechanism 3 is powered off and locked, and the rudder shaft 4 remains fixed at the leftmost position of the rudder shaft moving groove. At this time, when the resultant force point is on the left side of the rudder shaft, the output load of the servo always meets the hinge torque requirement. In this state, when the resultant force point is on the right side, the rudder shaft 4 does not need to move when the hinge torque does not exceed the maximum output load of the servo. When the resultant force point is on the right side of the rudder shaft 4 and the hinge torque applied to the rudder shaft 4 continues to increase and approaches the maximum output load of the servo, a rudder shaft position adjustment signal is sent to the flight control computer, the linear drive mechanism 3 is energized to drive the rudder shaft 4 to move to the rightmost position of the rudder shaft moving groove, and then the linear drive mechanism 3 is powered off.

[0049] The above process is repeated to ensure that the hinge torque always meets the requirement during the entire flight (i.e., the hinge torque never exceeds the maximum output load of the steering gear). Since there are very few severe working conditions during the entire flight, the linear drive mechanism 3 is driven very few times and consumes very little electricity.

[0050] Although the present invention has been described in detail above by general description and specific embodiments, it is obvious to those skilled in the art that some modifications or improvements can be made to the present invention. Therefore, these modifications or improvements made without departing from the spirit of the present invention all belong to the scope of protection claimed by the present invention.

Claims

1. A method for adjusting the hinge torque of a rudder blade, characterized in that: The rudder shaft is a movable rudder shaft, which can move within a set range along the chord length direction of the rudder blade, thereby changing the position of the rudder shaft relative to the rudder blade, and then changing the distance from the aerodynamic force point received by the rudder blade during flight to the rudder shaft, so as to achieve a change in the hinge torque.

2. The method for adjusting the hinge torque of a rudder blade according to claim 1, characterized in that: The two extreme positions for the movement of the rudder shaft along the chord length direction are respectively the left extreme position and the right extreme position, wherein the right extreme position ensures that when the rudder shaft is at the right extreme position and the resultant force point is at the right side of the rudder shaft, the hinge torque never exceeds the maximum output load of the servo; and the left extreme position ensures that when the rudder shaft is at the left extreme position and the resultant force point is at the left side of the rudder shaft, the hinge torque never exceeds the maximum output load of the servo.

3. The method for adjusting the hinge torque of a rudder blade according to claim 1 or 2, characterized in that: Two mounting positions of the rudder shaft are arranged inside the rudder blade along the chord length direction of the rudder blade, and the rudder shaft can be switched between the two mounting positions; when the rudder shaft is in one of the mounting positions, it is ensured that the hinge torque does not exceed the maximum output load of the servo when the resultant force point is on the side where the mounting position is located.

4. A hinge torque adjustable rudder blade, characterized in that: include: Rudder blade body, linear drive mechanism, rudder shaft and cover plate; A rectangular groove is provided on the rudder blade body at the installation position of the rudder shaft, serving as an installation cavity for the linear drive mechanism and the rudder shaft; The installation cavity is provided with a rudder shaft moving groove and a linear drive mechanism installation groove; the rudder shaft is installed in the rudder shaft moving groove; the linear drive mechanism is installed in the linear drive mechanism installation groove; the actuating end of the linear drive mechanism is connected to the rudder shaft, and is used to drive the rudder shaft to move in the rudder shaft moving groove along the chord length direction of the rudder blade; The cover plate is used to close the installation cavity.

5. The hinge torque adjustable rudder blade according to claim 4, characterized in that: The linear drive mechanism is an electric linear drive mechanism powered by a battery on the aircraft; and the linear drive mechanism has a self-locking function.

6. The hinge torque adjustable rudder blade according to claim 4 or 5, characterized in that: The upper part of the rudder shaft is a cubic structure; the lower part is a cylinder and is fixedly connected to the steering gear on the aircraft.

7. The hinge torque adjustable rudder blade according to claim 4 or 5, characterized in that: The rudder shaft moving groove is formed by connecting the rudder shaft moving groove A set in the installation cavity and the rudder shaft moving groove B set on the cover plate; the linear drive mechanism installation groove is formed by connecting the linear drive mechanism installation groove A set in the installation cavity and the linear drive mechanism installation groove B set on the cover plate.

8. A control method for a hinge torque adjustable rudder blade, characterized in that: The rudder blade is a hinge torque adjustable rudder blade as described in any one of claims 4 to 7 above; The rudder shaft moving groove ensures that: when the rudder shaft is located at the rightmost end position in the rudder shaft moving groove and the resultant force point is on the right side of the rudder shaft, the hinge torque never exceeds the maximum output load of the steering gear; when the rudder shaft is located at the leftmost end position in the rudder shaft moving groove and the resultant force point is on the left side of the rudder shaft, the hinge torque never exceeds the maximum output load of the steering gear; During the flight of the aircraft, initially, the rudder shaft is located at the end of one end of the rudder shaft moving groove; when the difference between the maximum output load of the servo and the hinge torque of the rudder shaft is less than a set threshold, a rudder shaft position adjustment signal is sent to the flight control computer. After receiving the rudder shaft position adjustment signal, the flight control computer controls the linear drive mechanism to drive the rudder shaft to move to the end of the other end of the rudder shaft moving groove.

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