A distributed multi-rotor tiltwing aircraft maneuvering strategy analysis method

Abstract

The present application relates to the field of aircraft flight mechanics and flight control, and particularly relates to a distributed overdrive tilting wing aircraft control strategy analysis method, comprising the following steps: step one: determining a rudder distribution scheme; the rudder distribution scheme is determined based on a minimum control energy fusion method; the control of multiple aerodynamic rudders is always linked without considering the flight conditions of the aircraft; in order to fully exert the efficiency of the linkage of multiple aerodynamic rudders, the weighted minimum control energy principle is used to distribute the control input, so as to use the rudders with high authority and high efficiency as much as possible under the same control input, and as little as possible to use the rudders with low authority and low efficiency, and finally make all the rudder angles as small as possible; step two: determining a flap aileron bias scheme; step three: calculating the lift propeller speed.

Description

Technical Field

[0001] This invention belongs to the field of aircraft flight mechanics and flight control design, and relates to a method for analyzing the control strategy of a distributed multi-rotor tilt-wing aircraft. Background Technology

[0002] Vertical takeoff and landing (VTOL) aircraft possess highly efficient vertical takeoff and landing capabilities, hovering, low-altitude, low-speed flight, and unique backward and lateral flight capabilities, enabling them to take off and land vertically in complex terrain. Conventional helicopters (single-rotor with tail rotor configuration) experience asymmetrical airflow across the rotor blades during forward flight, significantly limiting their maximum speed. Furthermore, the aerodynamic efficiency limitations of rotorcraft restrict their range and endurance. In contrast, traditional fixed-wing aircraft excel in high-speed and long-endurance flight. Tiltrotor aircraft, through a unique configuration that integrates the advantages of both rotorcraft and fixed-wing aircraft, combine the strengths of both, offering vertical takeoff and landing, hovering capabilities, high cruising speed, long range, and high payload capacity, representing one of the future directions for aircraft development. As the wings tilt, tilt-wing aircraft switch between helicopter mode, tilt-transition mode, and fixed-wing mode. Therefore, these aircraft have multiple control modes, and combinations of different control modes can achieve the same motion state, leading to complex control redundancy issues. Furthermore, based on configuration analysis, the potential control variables of this aircraft far exceed the pilot's control variables, significantly increasing the pilot's workload. Therefore, to solve the control redundancy problem and reduce pilot workload, a control strategy needs to be developed to guide the aircraft in switching between helicopter, fixed-wing, and tilt-transition modes, while simultaneously improving the aircraft's controllability and safety. However, existing literature and publicly available information do not disclose a control strategy for distributed multi-rotor tilt-wing aircraft. Summary of the Invention

[0003] Purpose of the invention: In order to solve at least one of the above-mentioned technical problems, this application provides a method for analyzing the control strategy of a distributed multi-rotor tilt-wing aircraft.

[0004] Technical solution

[0005] A method for analyzing the control strategies of a distributed overdriven tiltwing aircraft includes the following steps:

[0006] Step 1: Determine the control surface allocation scheme;

[0007] Step 2: Determine the flap / aileron offset scheme;

[0008] Step 3: Calculate the lift propeller speed.

[0009] Furthermore, in step one, the control surface allocation scheme is determined based on the minimum control energy fusion method. Without considering the flight conditions of the aircraft, it is assumed that the control of multiple aerodynamic control surfaces is always linked. In order to give full play to the effectiveness of the linkage of multiple aerodynamic control surfaces, the principle of weighted minimum control energy is adopted to allocate the control input. This is to achieve the goal of using the control surfaces with high efficiency and high authority as much as possible under the same control input, and avoiding the use of control surfaces with low efficiency and low authority as much as possible, so that the deflection angle of all control surfaces is minimized.

[0010] Furthermore, in step one, the control allocation coefficients of the main control surfaces of the three control channels—heading, lateral, and longitudinal—are calculated respectively.

[0011] The control distribution coefficients of the three-way control channels constitute the overall control surface distribution scheme.

[0012] Furthermore, in step one, the calculation process for the control allocation coefficient of any control channel rudder surface is as follows:

[0013] Step 1A: Define the objective function I for weighted minimum energy control, as shown in the following formula:

[0014]

[0015] Where: δ1 and δ2 are the control variables of the two control surfaces, |δ1| max and |δ2| max This represents the maximum control value for the corresponding control surface;

[0016] Step 1B: Calculate the control torque M generated by the control surface, using the following formula:

[0017] M = M δ1 δ1+M δ2 δ2

[0018] Where: M is the total operating torque, M is the total operating torque, M δ1 and M δ2 These represent the control effects of the main control surfaces δ1 and δ2, respectively.

[0019] Step 1 C: Let M c To obtain the desired control torque, the minimum values ​​of δ1 and δ2 are found according to the objective function I; the process is as follows:

[0020] Using δ2, δ1 can be expressed as follows:

[0021] The weighted objective function is then expressed as:

[0022]

[0023] make: The following result is obtained:

[0024]

[0025]

[0026]

[0027] Where: K1 and K2 are the control allocation coefficients for control surfaces δ1 and δ2.

[0028] Furthermore, in the longitudinal direction, δ1 represents the differential collective pitch of the fore and aft thrust propellers; δ2 represents the deviation of the fore and aft flaps and ailerons.

[0029] In the lateral direction, δ1 is the differential collective pitch of the left and right thrust propellers; δ2 is the deviation of the left and right flaps and ailerons.

[0030] When navigating, δ1 represents the differential collective pitch of the diagonal thrust propeller; δ2 represents the deviation of the left and right flaps and ailerons.

[0031] Furthermore, in step two, the process for determining the flap / aileron offset scheme is as follows:

[0032] Step 2A: Set different flap / aileron offset schemes for the front and rear wings, and calculate the longitudinal control stick values ​​for different offset schemes; the deflection angle of the front and rear flaps / ailerons is in the range of 0 to 20°, and the flaps / ailerons on each wing are offset synchronously;

[0033] Step 2B: Based on the Simulink module in Matlab, establish the flight dynamics model of the aircraft. Based on the aforementioned control surface allocation scheme, calculate the longitudinal control stick position, pitch control effectiveness, and required power for different offset schemes.

[0034] Step 2C: Based on the longitudinal joystick input, pitch control efficiency, and power requirements of different offset schemes, determine the optimal offset scheme:

[0035] Select 3 to 5 flaperon offset schemes with positive longitudinal strut values ​​and the smallest values;

[0036] Select 1 to 3 flaperon offset schemes that require the least power;

[0037] Then select the flaperon offset scheme that has the greatest pitch control effect.

[0038] Furthermore, in step two, the flaperon offset schemes for the aircraft in helicopter mode, fixed-wing mode, and tilt-over mode are determined respectively.

[0039] In helicopter mode: Select the flaperon offset scheme where the flaperon does not deflect;

[0040] In high-speed fixed-wing flight mode: select the flaperon offset scheme where the flaperon does not deflect;

[0041] In low-speed fixed-wing flight mode: Determine the flaperon offset scheme according to steps 2A to 2D.

[0042] In tilt-over mode: Determine the flap and aileron offset scheme according to steps 2A to 2D.

[0043] Furthermore, in step three, the process is as follows:

[0044] Step 3A: Based on the wing tilt angle range of the aircraft in helicopter mode, fixed-wing mode, and tilt transition mode, set the front and rear wing tilt angles:

[0045] The wing tilt angle ranges from 0 to 90°, with the front and rear wings tilting synchronously at the same angle.

[0046] Step 3B: Set the thrust / lift propeller ratio range throughout the entire flight:

[0047] Based on the maximum speed range of the thruster and lifter and the folding requirements of the lifter, the thrust / lifter thrust ratio range is set to 0.1 to 0.5.

[0048] Step 3C: Based on the Simulink module in Matlab, and according to the control surface allocation scheme determined in Step 1 and the flap and aileron offset scheme determined in Step 2, refine the flight dynamics model of the aircraft, set the forward speed, perform trim calculations, and obtain the required power.

[0049] Without considering transmission efficiency, the total power required by the aircraft is equal to the power required by the rotor, expressed as:

[0050]

[0051] Among them, P TLJ P SLJ η represents the power required for the thrust propeller and lift propeller, and η is the rotor efficiency.

[0052] The fourth step is to select the thrust / lift propeller ratio corresponding to the minimum required power as a suitable range, and then calculate the lift propeller speed.

[0053] An aircraft that is manipulated by a manipulation strategy determined by the method.

[0054] Beneficial effects

[0055] This invention employs a distributed multi-rotor power system and a tilt-wing mechanism. The tilt-wing mechanism controls the flight mode of the aircraft, including multi-rotor vertical mode, fixed-wing high-speed cruise mode, and tilt-wing transition mission mode, thereby improving the aircraft's performance in terms of long range, high forward speed, and high payload.

[0056] It adopts a distributed multi-rotor power system and structures such as front and rear wings, ailerons, flaps, and rudders. By formulating control strategies for helicopter mode, fixed-wing mode, and tilt transition mode, it completes the control switching between different modes and realizes the conversion between different modes.

[0057] Employing a distributed propulsion system with extremely high power redundancy, it can still safely perform missions even in the event of partial power failure, control surface jamming, or damage, which is beneficial to the survivability and combat performance of individual aircraft. Attached Figure Description

[0058] Figure 1 This is a schematic diagram of a distributed overdrive tiltrotor aircraft in multi-rotor mode for vertical takeoff and landing.

[0059] Figure 2 This is a schematic diagram of the fixed-wing mode cruise flight of a distributed overdriven tiltrotor aircraft;

[0060] In the picture, 1 is the fuselage, 2 is the main rotor, 3 is the canard, 4 is the aft wing, 5 is the flaps, 6 is the ailerons, and 7 is the vertical stabilizer. Detailed Implementation

[0061] The present invention will now be described in further detail with reference to the accompanying drawings.

[0062] A distributed overdrive tiltrotor aircraft includes a fuselage, a front wing, a rear wing, a vertical tail, four thrust propeller systems, at least four lift propeller systems, two tilting mechanisms, and flaps / ailerons. The front and rear wings are located at the front and rear of the fuselage, respectively, forming a tandem wing configuration. The front and rear wings together provide lift for the aircraft in fixed-wing mode. The vertical tail, located at the rear of the fuselage, is positioned similarly to the vertical tail of a traditional fixed-wing aircraft and can therefore be used as a rudder for yaw control. The thrust propeller and lift propeller systems are axially symmetrically distributed on the front and rear wings with respect to the fuselage axis. The thrust propeller systems are located closer to the outer edge of the wing, while the lift propeller systems are located closer to the inner edge. Both systems achieve angular changes in the power direction by tilting the wings through the tilting mechanisms. The tilting mechanisms are located within the fuselage and are used to achieve synchronous changes in the angle of the front and rear wings, thereby enabling different flight modes: helicopter mode, fixed-wing mode, and tilt transition mode. The flaps / ailerons are hinged at the trailing edge cuts of the front / rear wings, with the flaps located on the inner side of the wings and the ailerons on the outer side. The flaps / ailerons can rotate up and down around the hinge.

[0063] The helicopter mode of the distributed overdrive tiltrotor aircraft is used to achieve vertical takeoff and landing. In this flight mode, as mentioned above, the tilt mechanism and its supporting rotor system are both in the vertical direction, and both the thrust propeller system and the lift propeller system are operational.

[0064] The fixed-wing mode of the distributed overdrive tiltrotor aircraft is used to achieve high-speed forward flight. In this flight mode, as mentioned above, the tilt mechanism and its supporting rotor system are both in the horizontal direction, and only the thrust propeller system on the outer side of the wing is active, while the inner lift propeller is stationary.

[0065] The tilt transition mode of the distributed overdrive tiltwing aircraft is used to switch between helicopter mode and fixed-wing mode. In this flight mode, the tilt mechanism and its rotor system, as described above, are at a certain angle to the fuselage, and both the thrust propeller system and the lift propeller system are operational.

[0066] The control surfaces of the fore and aft wings, flaps / ailerons, rotor, and vertical tail of the distributed overdrive tiltrotor aircraft can all be used to change the flight state or mode during flight, achieving the same control effect, i.e., control redundancy exists. Therefore, this control strategy is determined through three steps: determining the control surface allocation scheme, the flap / aileron offset scheme, and the lift propeller speed scheme.

[0067] The control surface allocation scheme is determined using a minimum control energy fusion method. This method is a full-time linkage approach, disregarding the aircraft's flight conditions, ensuring that the control of multiple aerodynamic control surfaces is always linked. Furthermore, to fully leverage the effectiveness of the linkage between multiple aerodynamic control surfaces, the principle of weighted minimum control energy is used to allocate control inputs. This aims to maximize the use of high-efficiency, high-authority control surfaces under the same control input, while minimizing the use of low-efficiency, low-authority control surfaces, ultimately minimizing the deflection angle of all control surfaces.

[0068] The specific calculation steps for the aforementioned control surface allocation scheme are as follows:

[0069] First, let the objective function of weighted minimum energy control be:

[0070]

[0071] Where: δ1 and δ2 are the control variables of the two main control surfaces, |δ1| max and |δ2| max This represents the maximum control value corresponding to the control surface (a given value due to structural limitations). The control channels of the aircraft are divided into longitudinal, lateral, and directional channels. Under different control channels, δ1 and δ2 are respectively: In the longitudinal direction, δ1 is the differential collective pitch of the front and rear thrust propellers; δ2 is the differential deviation of the front and rear flaps and ailerons; In the lateral direction, δ1 is the differential collective pitch of the left and right thrust propellers; δ2 is the differential deviation of the left and right flaps and ailerons; In the directional direction, δ1 is the differential collective pitch of the diagonal thrust propellers; δ2 is the differential deviation of the left and right flaps and ailerons.

[0072] The sum of squares of the two main control surfaces in the objective function reflects energy control.

[0073] The second step is to assume that M is the total control torque generated by all control surfaces. Then M can be written as:

[0074] M = M δ1 δ1+M δ2 δ2

[0075] Where: M is the total operating torque, M δ1 and M δ2 These represent the control effectiveness of the main control surfaces δ1 and δ2, respectively.

[0076] Third step, let M c To obtain the minimum values ​​of δ1 and δ2 in order to minimize the objective function I, the desired control torque is obtained.

[0077] According to the formula in the second step, M... c =M δ1 δ1+M δ2 δ2,

[0078] Then δ1 can be expressed as:

[0079] Therefore, the weighted objective function is expressed as:

[0080]

[0081] make The following result is obtained:

[0082]

[0083]

[0084]

[0085] Where K1 and K2 are the control allocation coefficients of the main control surfaces δ1 and δ2, respectively, with the same denominator but different numerators. As can be seen from the numerator, at different flight speeds and flight modes, the larger the control derivative of a certain main control surface, the larger the corresponding control allocation coefficient will be. As the flight speed and flight mode change, the control derivative of the other main control surface increases, and its corresponding control allocation coefficient also increases.

[0086] Fourth, based on the Simulink module in Matlab and the aforementioned minimum control energy fusion method, a flight dynamics model of this configuration aircraft is established. The wing tilt angle and forward velocity are set, and trim calculations are performed to obtain δ1, δ2, and M. c Finally, we obtain K1 and K2;

[0087]

[0088]

[0089] Therefore, the control allocation coefficients of each main control surface in different control channels are obtained, and the final control surface allocation scheme is determined. This control surface allocation scheme is applicable to the helicopter mode, fixed-wing mode, and tilt-transition mode of the aircraft.

[0090] Taking the longitudinal control channel as an example, with the wing tilt angle set at 15° (tilt transition mode) and the forward speed at 45 m / s, the following can be calculated:

[0091] δ1=1208N·m, δ2=287N·m,M c =1495 N·m

[0092] Where, |δ1| max For 50°, |δ2| max 20°

[0093] Therefore, we get K1 = 0.8 and K2 = 0.2.

[0094] In summary, under the flight conditions of 15° wing tilt angle (tilt transition mode) and forward speed of 45 m / s, it is recommended that the control allocation coefficients for the collective pitch differential of the front and rear thrusters and the deviation of the front and rear flaps and ailerons in the longitudinal control channel be 0.8 and 0.2, respectively (that is, the collective pitch differential of the front and rear thrusters plays a major role at this time).

[0095] The aforementioned flaperon offset scheme determines the deflection angle of the flaperon control surfaces for the front / rear wings at different wing tilt angles. The specific calculation steps are as follows:

[0096] The first step is to determine different flaperon offset schemes based on the wing deflection requirements of the aircraft in helicopter mode, fixed-wing mode, and tilt-transition mode:

[0097] The wing tilt angle ranges from 0 to 90°, with the front and rear wings tilting synchronously, meaning they have the same tilt angle; the flaperon deflection angle ranges from 0 to 20°, with the flaperons deflecting synchronously, meaning they have the same deflection angle.

[0098] Specific biasing schemes are shown below:

[0099]

[0100] The second step is to establish a flight dynamics model of the aircraft configuration based on the Simulink module in Matlab. Based on the aforementioned control surface allocation scheme, the trim calculation is used to obtain the longitudinal control stick amount, pitch control effectiveness and power required for different offset schemes.

[0101] The third step is to determine the flap and aileron offset scheme:

[0102] (1) Based on the calculation results of the longitudinal strut, select 3 to 5 flaperon offset schemes with positive longitudinal strut values ​​and the smallest values;

[0103] (2) Based on the power required calculation results, select 1 to 3 flaperon offset schemes with the minimum power required from (1);

[0104] (3) Based on the pitch control effect calculation results, select the flaperon offset scheme with the greatest pitch control effect from (2).

[0105] The following is an example of how to determine a specific bias scheme:

[0106] The longitudinal control stick values ​​obtained for each offset scheme in the tilt transition mode (wing tilt angle 5°) are as follows:

[0107] Front wing flap deflection / (° / °) 5 / 0 10 / 0 20 / 0 10 / 10 20 / 10 Longitudinal stick displacement / ° 2.2 4.2 8.6 -1.8 3.2 Required power / kw 378 345 362 359 386 Pitch control effectiveness 4.6 15.2 16.8 14.1 9.5

[0108] Based on the calculation results of the longitudinal strut quantity, schemes 1, 2, and 5 with positive longitudinal strut quantity and the smallest value are selected.

[0109] Based on the power demand calculation results, select 1 to 3 options with the smallest power demand from options 1, 2 and 5, i.e., select options 1 and 2;

[0110] Based on the calculation results of pitch control effectiveness, the option with the greatest pitch control effectiveness is selected from Scheme 1 and Scheme 2.

[0111] In summary, when the wing tilt angle is 5°, it is recommended to set the offset scheme of the front and rear wing flaps and ailerons to 10° / 0°.

[0112] The aforementioned lift propeller speed scheme involves performing trim calculations at different wing tilt angles to determine the optimal operating speed of the lift propeller at the corresponding wing tilt angle. The specific calculation steps are as follows:

[0113] The first step is to set the front and rear wing tilt angles based on the wing tilt angle ranges in helicopter mode, fixed-wing mode, and tilt transition mode of this aircraft configuration:

[0114] The wing tilt angle ranges from 0 to 90°, and the front and rear wings tilt synchronously, meaning they have the same tilt angle.

[0115] The following is an example of a specific wing tilt angle setting:

[0116] Wing incidence / ° 90 60 30 0

[0117] The second step is to set the thrust / lift propeller ratio range throughout the entire flight:

[0118] Based on the maximum speed range of the thruster and lifter and the folding requirements of the lifter, the thrust / lifter thrust ratio range is set to 0.1 to 0.5.

[0119] The following is an example of how to set the thrust / lift propeller ratio:

[0120] Thrust / lift propeller pull ratio 0.1 0.2 0.3 0.4 0.5

[0121] The third step involves using the Simulink module in Matlab to refine the flight dynamics model of the aircraft configuration based on the aforementioned control surface allocation scheme and flap / aileron offset scheme. The forward flight speed is then set, and trim calculations are performed to obtain the required power.

[0122] Without considering transmission efficiency, the total power required by the aircraft is the same as the power required by the rotor, expressed as:

[0123]

[0124] Among them, P TLJ P SLJ η represents the power required for the thrust propeller and lift propeller, and η is the rotor efficiency.

[0125] The power calculation results for the helicopter mode (wing tilt angle 90°) example are as follows:

[0126] Required power / kw 1338 1250 1217 1209 1394

[0127] The fourth step is to determine the lift propeller speed based on the minimum required power:

[0128] Based on the total power required by the aircraft under different wing tilt angles, different forward flight speeds, and different thrust / lift propeller thrust ratios obtained in step three, the thrust / lift propeller thrust ratio corresponding to the minimum required power is selected as the appropriate range, and the lift propeller speed is finally determined.

[0129] In the helicopter mode (wing tilt angle 90°) example scheme, the thrust / lift propeller ratio corresponding to the minimum required power is 0.4. At this time, the thrust propeller speed is 575 r / s, and the lift propeller speed can be obtained as 230 r / s.

[0130] In summary, in helicopter mode (wing tilt angle 90°), a lift rotor speed of 230 r / s is recommended.

[0131] The above description is merely a preferred embodiment of the present invention, used only to illustrate the invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the original intent of the invention, and these improvements and modifications should also be considered within the scope of protection of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for analyzing the control strategies of a distributed overdriven tiltwing aircraft, characterized in that: The method includes the following steps: Step 1: Determine the control surface allocation scheme; the calculation process for the control allocation coefficient of any control channel control surface is as follows: Step 1A: Define the objective function for weighted minimum energy control I The formula is as follows: in: and As the control parameters of the two control surfaces, and This represents the maximum control value for the corresponding control surface; Step 1B: Calculate the control torque M generated by the control surface, using the following formula: in: For the total control torque, and These represent the main control surfaces. and The control effect; Step 1 C: Let To achieve the desired control torque, according to the objective function I , obtain the corresponding and The minimum value is found; the process is as follows: use Will It is expressed as follows: The weighted objective function is then expressed as: make: The following result is obtained: in: , As rudder surface and Manipulation allocation coefficient; Step 2: Determine the flap / aileron offset scheme, the process is as follows: Step 2A: Set different flap / aileron offset schemes for the front and rear wings, and calculate the longitudinal control stick values ​​for different offset schemes; the deflection angle of the front and rear flaps / ailerons is in the range of 0 to 20°, and the flaps / ailerons on each wing are offset synchronously; Step 2B: Based on the Simulink module in Matlab, establish the flight dynamics model of the aircraft. Based on the aforementioned control surface allocation scheme, calculate the longitudinal control stick position, pitch control effectiveness, and required power for different offset schemes. Step 2C: Based on the longitudinal joystick input, pitch control efficiency, and power requirements of different offset schemes, determine the optimal offset scheme: Select 3 to 5 flaperon offset schemes with positive longitudinal stick values ​​and the smallest values; select 1 to 3 flaperon offset schemes with the lowest power requirements; and then select the flaperon offset scheme with the greatest pitch control effect. Step 3: Calculate the lift propeller speed, the process is as follows: Step 3A: Based on the wing tilt angle range in helicopter mode, fixed-wing mode, and tilt transition mode, set the front and rear wing tilt angles: The wing tilt angle ranges from 0 to 90°, with the front and rear wings tilting synchronously at the same angle. Step 3B: Set the thrust / lift propeller ratio range throughout the entire flight: Based on the maximum speed range of the thruster and lifter and the folding requirements of the lifter, the thrust / lifter thrust ratio range is set to 0.1 to 0.

5. Step 3C: Based on the Simulink module in Matlab, and according to the control surface allocation scheme determined in Step 1 and the flap and aileron offset scheme determined in Step 2, refine the flight dynamics model of the aircraft, set the forward speed, perform trim calculations, and obtain the required power. Without considering transmission efficiency, the total power required by the aircraft is equal to the rotor power required, expressed as: in, , Power required for thrust propellers and lift propellers For rotor efficiency; Select the thrust / lift propeller ratio corresponding to the minimum required power as an appropriate range, and then calculate the lift propeller speed.

2. The method according to claim 1, characterized in that: In step one, the control surface allocation scheme is determined based on the minimum control energy fusion method. Without considering the flight conditions of the aircraft, it is assumed that the control of multiple aerodynamic control surfaces is always linked. In order to give full play to the effectiveness of the linkage of multiple aerodynamic control surfaces, the principle of weighted minimum control energy is adopted to allocate the control input. This is to achieve the goal of using the control surfaces with high efficiency and high authority as much as possible under the same control input, and avoiding the use of control surfaces with low efficiency and low authority as much as possible, so that the deflection angle of all control surfaces is minimized.

3. The method according to claim 2, characterized in that: In step one, the control allocation coefficients of the main control surfaces of the three control channels (heading, lateral, and longitudinal) are calculated respectively. The control distribution coefficients of the three-way control channels constitute the overall control surface distribution scheme.

4. The method according to claim 3, characterized in that: In the longitudinal direction, This refers to the differential collective pitch of the front and rear thrust propellers; For the deviation movement of the front and rear flaps, ailerons, and rudders; When horizontal, For differential collective pitch of the left and right thrust propellers; For the deviation movement of the left and right flaps and ailerons; When heading, For diagonal thrust propeller collective pitch differential; The deviation of the left and right flaps and ailerons is due to the movement.

5. The method according to claim 4, characterized in that: In step two, the flaperon offset schemes for the aircraft in helicopter mode, fixed-wing mode and tilt transition mode are determined respectively. In helicopter mode: Select the flaperon offset scheme where the flaperon does not deflect; In high-speed fixed-wing flight mode: select the flaperon offset scheme where the flaperon does not deflect; In low-speed fixed-wing flight mode: Determine the flaperon offset scheme according to steps 2A to 2C. In tilt transition mode: Determine the flap and aileron offset scheme according to steps 2A to 2C.

6. An aircraft, characterized in that: The aircraft is manipulated using the manipulation strategy determined by any one of claims 1-5.

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