A control method for converting horizontal mode to vertical mode of vertical take-off and landing UAV

Abstract

The present invention belongs to the field of flight control law design, and is specifically a control method for converting a horizontal mode of a vertical take-off and landing unmanned aerial vehicle to a vertical mode, comprising an elevator control law and an engine throttle control law. The elevator control law is: during the mode conversion process, the elevator always maintains a maximum deflection angle, and the unmanned aerial vehicle is raised with the maximum control capability of the slipstream rudder or the vector nozzle. When the attitude of the unmanned aerial vehicle is close to a vertical state, the elevator control law becomes an attitude closed-loop control; the engine throttle control law is: the pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated by the numerical value of the pitch angular velocity. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the unmanned aerial vehicle during the mode conversion process. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the unmanned aerial vehicle during the mode conversion process.

Description

Technical Field

[0001] The invention belongs to the field of flight control law design, and in particular relates to a control method for converting a horizontal mode into a vertical mode of a vertical take-off and landing unmanned aerial vehicle. Background Art

[0002] Vertical take-off and landing aircraft use the vertical upward thrust provided by the engine to overcome gravity, thereby achieving vertical take-off and landing. This type of aircraft can get rid of the dependence on airport runways, and has the advantages of flexible departure, high attendance rate, and low maintenance cost. In addition, it can also take off and land on ships, and has unparalleled advantages over taxiing and landing carrier-based aircraft in improving amphibious combat capabilities. The main types of vertical take-off and landing aircraft have been developed to date: tail-seat type, tilt-rotor type, rotor type, jet engine thrust steering type, ducted fan type, and a hybrid of these types.

[0003] Among them, the tail-seat vertical take-off and landing UAV is an aircraft whose entire body tilts during take-off and landing. With its light weight, small size, and simple power system, it can take off and land in narrow spaces such as ship decks, mountains, canyons, and alleys. After carrying a payload, it can be used to perform reconnaissance, relay communications, and attack tasks, and has broad application prospects.

[0004] For tail-seat vertical take-off and landing UAVs that use ducted propulsion or turbojet thrust vectoring engines as power, their attitude control devices are the slipstream rudders of ducted propulsion or the vector nozzles of turbojet thrust vectoring engines. The control efficiency of the slipstream rudders or vector nozzles is related to the engine thrust, that is, the greater the thrust, the higher the control efficiency. In the process of converting the UAV from horizontal mode to vertical mode, it must have sufficient control efficiency to meet the requirements of attitude control and resistance to external disturbances, which requires the engine thrust to be large enough. However, when the engine thrust is large, the residual thrust of the UAV is too large, resulting in too much increase in the height of the UAV during the mode conversion process. The time taken to vertically hover and land on the ground will be longer, which is not conducive to the rapid landing and recovery of the UAV. In addition, for UAVs used for military purposes, if the altitude is too high during the mode conversion process, it is easy to expose its own target.

[0005] Therefore, how to ensure that the UAV control device has sufficient control efficiency and minimize the increase in the height of the UAV during the mode conversion process when the tail-seat vertical take-off and landing UAV converts from horizontal mode to vertical mode is of great significance for the tail-seat vertical take-off and landing UAV to achieve rapid landing and recovery and reduce the possibility of exposing its own target. This is also the technical problem to be solved by the present invention. Summary of the invention

[0006] In order to solve the above problems and shortcomings in the prior art, the present invention proposes a control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV.

[0007] In order to achieve the above-mentioned invention object, the technical solution provided by this application is as follows:

[0008] A control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV, comprising an elevator control law and an engine throttle control law;

[0009] The elevator control law is as follows: during the mode conversion process, the elevator always maintains the maximum deflection angle, and the UAV is lifted up with the maximum control capability of the slipstream control surface or the vector nozzle. When the attitude of the UAV approaches the vertical state, the elevator control law becomes an attitude closed-loop control.

[0010] The engine throttle control law is as follows: the pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated by the numerical value of the pitch angular velocity. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the UAV during the mode conversion process.

[0011] Furthermore, the elevator control law is specifically:

[0012] (1)

[0013] in, is the elevator command, is the maximum deflection angle of the elevator, is the pitch angle, measured by the onboard attitude sensor, is the pitch angle control target, and its value is , is the critical pitch angle of attitude closed-loop control, is the pitch angle proportional control parameter, is the pitch angle integral control parameter.

[0014] Preferably, the critical pitch angle of attitude closed-loop control is exist The reason is: if If the value is too small, the elevator control law will change to attitude closed-loop control too early, and the maximum control capability of the slipstream control surface or vector nozzle will not be fully utilized; If the value is too large, the elevator control law will change to attitude closed-loop control too late, and a large attitude overshoot may occur during the attitude change to the vertical state.

[0015] Furthermore, considering that the tail-seat vertical take-off and landing UAV changes from horizontal mode to vertical mode, the attitude of the UAV changes rapidly from horizontal state to vertical state, which may cause large normal overload and endanger safety. Based on the formula (1) of the elevator control law, the normal overload protection control quantity can be introduced, and the elevator control law is:

[0016] (2)

[0017] in, is the normal overload protection control quantity, is the normal overload protection control parameter, is the normal overload factor.

[0018] Furthermore, the normal overload factor Defined as:

[0019] (3)

[0020] in, is the normal overload, measured by the onboard inertial sensor, , Normal overload protection threshold; when normal overload When the normal overload protection is not enabled, When the normal overload protection is enabled, the normal overload protection control amount and the difference Linear positive correlation; when normal overload When the normal overload protection control value reaches the maximum value, and maintains constant.

[0021] Preferably, the threshold The value range is , threshold The value is ,in, is the maximum allowable normal overload of the UAV.

[0022] Furthermore, the engine throttle control law is specifically:

[0023] The pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated according to the numerical value of the pitch angular velocity:

[0024] (4)

[0025] in, is the engine throttle command, is the engine throttle command reference value, is the throttle integral control parameter, is the operating efficiency integration factor.

[0026] Preferably, the engine throttle command reference value is the engine throttle command at the initial moment of the horizontal mode conversion to the vertical mode, that is, the cruise trim throttle command at the end of the UAV cruise segment.

[0027] Preferably, at the initial moment when the horizontal mode is converted to the vertical mode, the integral control term of the engine throttle control law is The initial value of the integral is set to 0.

[0028] Furthermore, the control efficiency integration factor Defined as:

[0029] (5)

[0030] in, is the pitch angular velocity, measured by the onboard attitude sensor, , , , is the control efficiency integral threshold; when the pitch angle velocity When the pitch angular velocity is When , the integration step length is , is a positive value, the engine throttle command will continue to increase over time; when the pitch angular velocity When , the integral step reaches the maximum value and keeps The engine throttle command will continue to increase over time; when the pitch rate When , the integration step length is , is a negative value, the engine throttle command will continue to decrease over time; when the pitch angular velocity When , the integral step reaches the minimum value and keeps The engine throttle command will continue to decrease over time.

[0031] Preferably, the threshold The value range is 0° / s~10° / s; threshold The value range is 15° / s~20° / s; threshold The value is 25° / s~30° / s; threshold The value is 30° / s~35° / s; and should meet , .

[0032] Furthermore, based on the engine throttle control law formula (4) and formula (5), the protection against normal overload is considered. The specific method is: through the maximum allowable normal overload Calculate the maximum allowable pitch rate , the calculation formula is ,in, The unit is ° / s, is the gravitational acceleration, which is , is the aircraft speed in the body coordinate system The component on the axis, in m / s, is measured by the onboard navigation equipment.

[0033] Furthermore, the body coordinate system The axis is defined as: Axis to aircraft center of gravity It is the origin of coordinates, located in the symmetry plane of the aircraft and parallel to the axis of the fuselage, and the direction pointing to the nose is positive.

[0034] Furthermore, when considering normal overload protection, the threshold The value is The smaller value of 10° / s or 10° / s, the threshold The value is and 15° / s, whichever is smaller; threshold The value is and 25° / s, whichever is smaller; threshold The value is and 30° / s, whichever is smaller.

[0035] The beneficial effects of the present invention are:

[0036] 1. In the process of the tail-seat vertical take-off and landing UAV converting from the horizontal mode to the vertical mode, the elevator maintains the maximum deflection angle, and at the same time, the pitch angular velocity is introduced into the engine throttle control law to realize the adaptive control of the engine throttle on the control efficiency of the slipstream control surface or the vector nozzle during the mode conversion process. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the UAV during the mode conversion process.

[0037] 2. The control method for converting the horizontal mode to the vertical mode of the vertical take-off and landing UAV described in this method introduces the pitch angular velocity into the engine throttle control law, so as to realize the adaptive control of the engine throttle on the control efficiency of the slipstream control surface or the vector nozzle during the mode conversion process. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the UAV during the mode conversion process, which is conducive to the rapid landing and recovery of the UAV and improving the concealment of its own target.

[0038] 3. The control method for converting the horizontal mode to the vertical mode of a vertical take-off and landing UAV disclosed in this method adopts a PID control structure. The flight control law algorithm is simple and easy for designers to master. The onboard flight control software is convenient to develop and has high engineering application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 Schematic diagram of the control law structure of Example 1.

[0040] Figure 2 This is a schematic diagram of the control law structure of Example 2.

[0041] Figure 3 Schematic diagram of the value of the control efficiency integration factor.

[0042] Figure 4 Schematic diagram of the normal overload factor value. DETAILED DESCRIPTION

[0043] In order to make the purpose, technical solution and advantages of the present invention clearer, the technical solution in the embodiment of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiment of the present invention. Obviously, the described embodiment is to explain the present invention rather than to limit the present invention. Based on the embodiment of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0044] The specific implementation method of the present invention is described below with reference to the accompanying drawings and examples, but the present invention is not limited to this embodiment.

[0045] Example 1

[0046] As attached Figure 1 As shown, a control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV is composed of two parts: an elevator control law and an engine throttle control law. In the first part, during the conversion of the horizontal mode to the vertical mode of a tail-seat vertical take-off and landing UAV, the elevator maintains a maximum deflection angle; in the second part, the pitch angular velocity is introduced into the engine throttle control law to realize adaptive control of the engine throttle on the control efficiency of the slipstream control surface or the vector nozzle during the mode conversion process.

[0047] Part I: Elevator Control Law

[0048] During the horizontal mode conversion to vertical mode of a tail-seat vertical take-off and landing UAV, the attitude of the UAV changes rapidly from a horizontal state to a vertical state, and is in a large angle of attack range. The static stability torque is relatively strong, and the slipstream rudder or vector nozzle needs to overcome the static stability torque to make the attitude of the UAV rise. In addition, when there are external disturbance torques such as wind interference, the slipstream rudder or vector nozzle also needs to overcome the external disturbance, otherwise, the UAV will not be able to successfully perform the mode conversion. To this end, the scheme of the present invention is: during the mode conversion process, the elevator always maintains the maximum deflection angle, and the UAV is raised with the maximum control ability of the slipstream rudder or vector nozzle. When the attitude of the UAV is close to the vertical state, the elevator control law becomes an attitude closed-loop control. The elevator control law is:

[0049] (6)

[0050] in, is the elevator command, is the maximum deflection angle of the elevator, is the pitch angle, measured by the onboard attitude sensor, is the pitch angle control target, and its value is , is the critical pitch angle of attitude closed-loop control, is the pitch angle proportional control parameter, is the pitch angle integral control parameter.

[0051] Preferably, the critical pitch angle of attitude closed-loop control is exist The reason is: if If the value is too small, the elevator control law will change to attitude closed-loop control too early, and the maximum control capability of the slipstream control surface or vector nozzle will not be fully utilized; If the value is too large, the elevator control law will change to attitude closed-loop control too late, and a large attitude overshoot may occur during the attitude change to the vertical state.

[0052] Part II: Engine Throttle Control Law

[0053] The present invention uses the pitch angular velocity to judge whether the control efficiency of the slipstream control surface or the vector nozzle during the mode conversion process is sufficient. The specific judging principle is:

[0054] During the process of tail-seat vertical take-off and landing UAV converting from horizontal mode to vertical mode, if the control torque of the slipstream control surface or vector nozzle is greater than the sum of the static stability torque and the external disturbance torque, the aircraft will be able to continuously raise its head, and the pitch angular velocity will be a positive value. The larger the pitch angular velocity value, the stronger the aircraft's ability to raise its head, that is, the more sufficient the control efficiency is. If the control torque of the slipstream control surface or vector nozzle is less than the sum of the static stability torque and the external disturbance torque, the aircraft will continue to lower its head and cannot reach the vertical mode. The pitch angular velocity will be a negative value. The larger the absolute value of the pitch angular velocity, the more serious the aircraft's tendency to lower its head, that is, the less efficient the control is.

[0055] Preferably, the pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated according to the numerical value of the pitch angular velocity, so as to reduce the engine throttle as much as possible while ensuring that the aircraft has sufficient control efficiency. Specifically:

[0056] (7)

[0057] in, is the engine throttle command, is the engine throttle command reference value, is the throttle integral control parameter, is the operating efficiency integration factor.

[0058] Preferably, the engine throttle command reference value The value can be the engine throttle command at the initial moment of the horizontal mode conversion to the vertical mode, that is, the cruise trim throttle command at the end of the UAV cruise segment.

[0059] Preferably, at the initial moment when the horizontal mode is converted to the vertical mode, the integral control term of the engine throttle control law is The initial value of the integral can be set to 0.

[0060] Furthermore, as attached Figure 3 As shown, the operating efficiency integration factor Defined as:

[0061] (8)

[0062] in, is the pitch angular velocity, measured by the onboard attitude sensor, , , , is the control efficiency integral threshold; when the pitch angle velocity When the pitch angular velocity is When , the integration step length is , is a positive value, the engine throttle command will continue to increase over time; when the pitch angular velocity When , the integral step reaches the maximum value and keeps The engine throttle command will continue to increase over time; when the pitch rate When , the integration step length is , is a negative value, the engine throttle command will continue to decrease over time; when the pitch angular velocity When , the integral step reaches the minimum value and keeps The engine throttle command will continue to decrease over time.

[0063] Furthermore, the threshold , , , The value of cannot be too large or too small. If the threshold value is too small, the mode conversion process takes too long, the aircraft moves too far forward during the process, and the position of the aircraft when it reaches the vertical mode may deviate too far from the expected landing point; if the threshold value is too large, the mode conversion is too fast, and there may be a risk of slipstream control surface stall and normal overload exceeding the limit.

[0064] Preferably, the threshold The value range is 0° / s~10° / s; threshold The value range is 15° / s~20° / s; threshold The value is 25° / s~30° / s; threshold The value is 30° / s~35° / s; and should meet , .

[0065] Example 2

[0066] On the basis of Example 1, the protection against normal overload of the aircraft is considered in the elevator control law and the engine throttle control law.

[0067] As attached Figure 2 As shown, a control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV is composed of two parts: an elevator control law and an engine throttle control law. In the first part, during the conversion of the horizontal mode to the vertical mode of the tail-seat vertical take-off and landing UAV, the elevator maintains a maximum deflection angle, and at the same time, a normal overload protection control quantity is introduced to protect the normal overload; in the second part, the pitch angular velocity is introduced into the engine throttle control law to realize the adaptive control of the engine throttle on the control efficiency of the slipstream control surface or the vector nozzle during the mode conversion process. In addition, the engine throttle control law also protects the normal overload.

[0068] Part I: Elevator Control Law

[0069] During the process of the tail-seat vertical take-off and landing UAV changing from horizontal mode to vertical mode, the attitude of the UAV changes rapidly from horizontal state to vertical state, and it is in a large angle of attack range. The static stability torque is relatively strong, and the slipstream rudder or vector nozzle needs to overcome the static stability torque to make the attitude of the UAV rise. In addition, when there is an external disturbance torque such as wind interference, the slipstream rudder or vector nozzle also needs to overcome the external disturbance, otherwise, the UAV will not be able to successfully perform the mode conversion. To this end, the scheme of the present invention is: before the attitude of the UAV approaches the vertical state, the elevator always maintains the maximum deflection angle, and the UAV is raised with the maximum control ability of the slipstream rudder or vector nozzle. At the same time, considering that during the process of the tail-seat vertical take-off and landing UAV changing from horizontal mode to vertical mode, the attitude of the UAV changes rapidly from horizontal state to vertical state, and a large normal overload may occur, endangering safety, so it is necessary to introduce a normal overload protection control quantity. When the attitude of the UAV approaches the vertical state, the elevator control law becomes an attitude closed-loop control. The elevator control law is:

[0070] (9)

[0071] in, is the elevator command, is the maximum deflection angle of the elevator, is the pitch angle, measured by the onboard attitude sensor, is the pitch angle control target, and its value is , is the critical pitch angle of attitude closed-loop control, is the pitch angle proportional control parameter, is the pitch angle integral control parameter. is the normal overload protection control quantity, is the normal overload protection control parameter, is the normal overload factor.

[0072] Preferably, the critical pitch angle of attitude closed-loop control is exist The reason is: if If the value is too small, the elevator control law will change to attitude closed-loop control too early, and the maximum control capability of the slipstream control surface or vector nozzle will not be fully utilized; If the value is too large, the elevator control law will change to attitude closed-loop control too late, and a large attitude overshoot may occur during the attitude change to the vertical state.

[0073] Furthermore, as attached Figure 4 As shown, the normal overload factor Defined as:

[0074] (10)

[0075] in, is the normal overload, measured by the onboard inertial sensor, , Normal overload protection threshold; when normal overload When the normal overload protection is not enabled, When the normal overload protection is enabled, the normal overload protection control amount and the difference Linear positive correlation; when normal overload When the normal overload protection control value reaches the maximum value, and maintains constant.

[0076] Preferably, the threshold The value range is , threshold The value is ,in, The maximum permissible normal overload of the UAV is generally limited by the structural strength of the aircraft body, and the value will be clearly stated in the overall design of the aircraft.

[0077] Part II: Engine Throttle Control Law

[0078] The present invention judges whether the control efficiency of the slipstream control surface or the vector nozzle is sufficient during the mode conversion process by the pitch angular velocity. The specific judgment principle is: during the horizontal mode conversion of the tail-seat vertical take-off and landing UAV to the vertical mode, if the control torque of the slipstream control surface or the vector nozzle is greater than the sum of the static stability torque and the external disturbance torque, the aircraft will be able to continuously raise its head, and the pitch angular velocity is a positive value. The larger the value of the pitch angular velocity is, the stronger the aircraft's ability to raise its head is, that is, the more sufficient the control efficiency is; if the control torque of the slipstream control surface or the vector nozzle is less than the sum of the static stability torque and the external disturbance torque, the aircraft will continuously lower its head and cannot reach the vertical mode. The pitch angular velocity is a negative value. The larger the absolute value of the pitch angular velocity is, the more serious the aircraft's tendency to lower its head is, that is, the less effective the control efficiency is.

[0079] Preferably, the pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated according to the numerical value of the pitch angular velocity, so as to reduce the engine throttle as much as possible while ensuring that the aircraft has sufficient control efficiency. Specifically:

[0080] (11)

[0081] in, is the engine throttle command, is the engine throttle command reference value, is the throttle integral control parameter, is the operating efficiency integration factor.

[0082] Preferably, the engine throttle command reference value The value can be the engine throttle command at the initial moment of the horizontal mode conversion to the vertical mode, that is, the cruise trim throttle command at the end of the UAV cruise segment.

[0083] Preferably, at the initial moment when the horizontal mode is converted to the vertical mode, the integral control term of the engine throttle control law is The initial value of the integral can be set to 0.

[0084] Furthermore, as attached Figure 3 As shown, the operating efficiency integration factor Defined as:

[0085] (12)

[0086] in, is the pitch angular velocity, measured by the onboard attitude sensor, , , , is the control efficiency integral threshold; when the pitch angle velocity When the pitch angular velocity is When , the integration step length is , is a positive value, the engine throttle command will continue to increase over time; when the pitch angular velocity When , the integral step reaches the maximum value and keeps The engine throttle command will continue to increase over time; when the pitch rate When , the integration step length is , is a negative value, the engine throttle command will continue to decrease over time; when the pitch angular velocity When , the integral step reaches the minimum value and keeps The engine throttle command will continue to decrease over time.

[0087] Furthermore, the threshold , , , The value of cannot be too large or too small. If the threshold value is too small, the mode conversion process takes too long, the aircraft moves forward too far in this process, and the position of the aircraft when it reaches the vertical mode may deviate too far from the expected landing point; if the threshold value is too large, the mode conversion is too fast, and there may be a risk of slipstream control surface stall and normal overload exceeding the limit. Therefore, under the premise of not exceeding the maximum allowable normal overload, it is beneficial to select a larger threshold. The specific method is: through the maximum allowable normal overload Calculate the maximum allowable pitch rate , the calculation formula is ,in, The unit is ° / s, is the gravitational acceleration, which is , is the aircraft speed in the body coordinate system The component on the axis, in m / s, is measured by the onboard navigation equipment. For threshold , , , By determining the value range of , the protection against normal overload can be indirectly achieved through the threshold.

[0088] Furthermore, the body coordinate system The axis is defined as: Axis to aircraft center of gravity It is the origin of coordinates, located in the symmetry plane of the aircraft and parallel to the axis of the fuselage, and the direction pointing to the nose is positive.

[0089] Preferably, when considering normal overload protection, the threshold The value is The smaller value of 10° / s or 10° / s, the threshold The value is and 15° / s, whichever is smaller; threshold The value is and 25° / s, whichever is smaller; threshold The value is and 30° / s, whichever is smaller.

[0090] The above description is only a preferred embodiment of the present invention and does not constitute any form of hindrance to the present invention. Any simple modification or equivalent changes made to the above embodiments based on the technical essence of the present invention shall fall within the protection scope of the present invention.

Claims

1. A control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV, characterized in that: Contains elevator control law and engine throttle control law; The elevator control law is as follows: during the mode conversion process, the elevator always maintains the maximum deflection angle, and the UAV is lifted up with the maximum control capability of the slipstream control surface or the vector nozzle. When the attitude of the UAV approaches the vertical state, the elevator control law becomes an attitude closed-loop control. The engine throttle control law is as follows: the pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated by the numerical value of the pitch angular velocity. While ensuring that the aircraft has sufficient control efficiency, the engine throttle is reduced as much as possible, thereby reducing the increase in the height of the UAV during the mode conversion process; The elevator control law is specifically: （1） in, is the elevator command, is the maximum deflection angle of the elevator, is the pitch angle, measured by the onboard attitude sensor, is the pitch angle control target, and its value is , is the critical pitch angle of attitude closed-loop control, is the pitch angle proportional control parameter, is the pitch angle integral control parameter; Critical pitch angle for attitude closed-loop control exist Take values ​​within the range; The engine throttle control law is specifically: The pitch angular velocity is introduced into the engine throttle control law, and the engine throttle command is calculated according to the numerical value of the pitch angular velocity: （4） in, is the engine throttle command, is the engine throttle command reference value, is the throttle integral control parameter, is the manipulation efficiency integration factor; The engine throttle command reference value The value is the engine throttle command at the initial moment of the horizontal mode conversion to the vertical mode; At the initial moment of the horizontal mode conversion to the vertical mode, the integral control term of the engine throttle control law The initial value of the integral is set to 0; The operating efficiency integrating factor Defined as: （5） in, is the pitch angular velocity, measured by the onboard attitude sensor, , , , is the control efficiency integral threshold; when the pitch angle velocity When the pitch angular velocity is When , the integration step length is , is a positive value, the engine throttle command will continue to increase over time; when the pitch angular velocity When , the integral step reaches the maximum value and keeps The engine throttle command will continue to increase over time; when the pitch rate When , the integration step length is , is a negative value, the engine throttle command will continue to decrease over time; when the pitch angular velocity When , the integral step reaches the minimum value and keeps The engine throttle command will continue to decrease over time.

2. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 1, characterized in that: Based on the formula (1) of the elevator control law, the normal overload protection control variable is introduced, and the elevator control law is: （2） in, is the normal overload protection control quantity, is the normal overload protection control parameter, is the normal overload factor.

3. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 2, characterized in that: The normal overload factor Defined as: （3） in, is the normal overload, measured by the onboard inertial sensor, , Normal overload protection threshold; when normal overload When the normal overload protection is not enabled, When the normal overload protection is enabled, the normal overload protection control amount and the difference Linear positive correlation; when normal overload When the normal overload protection control value reaches the maximum value, and maintains constant.

4. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 3 is characterized in that: Threshold The value range is , threshold The value is ,in, is the maximum allowable normal overload of the drone.

5. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 1 is characterized in that: Threshold The value range is 0° / s~10° / s; threshold The value range is 15° / s~20° / s; threshold The value is 25° / s~30° / s; threshold The value is 30° / s~35° / s; and should meet , .

6. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 1 is characterized in that: Based on the engine throttle control law formula (4) and formula (5), the protection against normal overload is considered. The specific method is as follows: Calculate the maximum allowable pitch rate , the calculation formula is ,in, The unit is ° / s, is the gravitational acceleration, which is , is the aircraft speed in the body coordinate system The component on the axis, in m / s, is measured by the onboard navigation equipment.

7. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 6 is characterized by: The body coordinate system The axis is defined as: Axis to aircraft center of gravity It is the origin of coordinates, located in the symmetry plane of the aircraft and parallel to the axis of the fuselage, and the direction pointing to the nose is positive.

8. The control method for converting a horizontal mode to a vertical mode of a vertical take-off and landing UAV according to claim 6 is characterized by: When considering normal overload protection, the threshold The value is The smaller value of 10° / s or 10° / s, the threshold The value is and 15° / s, whichever is smaller; threshold The value is and 25° / s, whichever is smaller; threshold The value is and 30° / s, whichever is smaller.

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