Decoupling control method for inverted v-shaped tail wing flapping-wing aircraft based on forward distance ratio scheduling

Abstract

This invention, a decoupled control method for a flapping-wing aircraft based on approach ratio scheduling, belongs to the field of biomimetic flapping-wing aircraft control technology. It solves the problems of poor control accuracy and stability in existing flapping-wing robot control technologies. The invention introduces an aerodynamic characteristic lookup table indexed by approach ratio and angle of attack. During flight, the approach ratio and angle of attack are calculated in real time, and aerodynamic parameters in the dynamic model are dynamically retrieved and updated according to the aerodynamic characteristic lookup table, thus solving the problem of inaccurate fixed-parameter modeling. Secondly, a decoupled hybrid gain coefficient function that varies with the main wing flapping frequency is introduced to dynamically allocate pitch and yaw control variables to the left and right tail servos, enabling control commands to more accurately match real aerodynamic characteristics, thereby significantly improving the stability and adaptability of attitude control. This invention is used to control biomimetic flapping-wing aircraft.

Description

Technical Field

[0001] This invention belongs to the field of biomimetic flapping-wing aircraft control technology. Background Technology

[0002] Birds in nature generate complex unsteady aerodynamic forces through flapping their wings during flight, providing important inspiration for the design of bird-inspired flapping-wing robots. In recent years, with the development of micromotors, lightweight materials, and embedded control technology, biomimetic flapping-wing aircraft have gradually become an important research direction in the field of flying robots. For small flapping-wing aircraft weighing around 210g, in order to reduce system complexity and structural weight, independent roll control mechanisms are usually not set up. Instead, an inverted V-shaped tail structure composed of two rudders is used, and pitch and yaw are coordinated through differential tail movement.

[0003] The following problems exist in the existing control technology for flapping-wing flying robots: First, most existing dynamic models are based on quasi-steady assumptions or fixed aerodynamic coefficients, only considering the influence of angle of attack while ignoring the nonlinear aerodynamic characteristics caused by the change in flapping frequency, resulting in large modeling errors and low control accuracy. Second, inverted V-tails usually use simple linear hybrid control to achieve pitch and yaw control, but the unsteady wash generated by the flapping of the main wing causes the aerodynamic gain of the tail to change with frequency, resulting in strong coupling of the pitch-yaw channel, causing attitude oscillation or heading drift, and poor control stability. Therefore, it is urgent to solve the problem of poor control accuracy and stability of small flapping-wing flying robots in complex unsteady environments. Summary of the Invention

[0004] The purpose of this invention is to solve the problems of poor control accuracy and stability in existing flapping-wing flight robot control technology. This invention provides a decoupling control method for flapping-wing aircraft with an inverted V-shaped tail based on advance ratio scheduling.

[0005] A decoupled control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling, wherein the flapping frequencies of a pair of main wings of the flapping-wing aircraft are the same, the method includes:

[0006] Acquire current flight status information of the flapping-wing aircraft, including main wing flapping frequency, flight speed, inertial measurement information, actual forces and torques acting on the flapping-wing aircraft;

[0007] The inertial measurement information is fused and processed to obtain the flight angle of attack. Inertial measurement information includes angular velocity, specific force, and flight altitude.

[0008] Calculate the approach ratio under the current flight conditions based on flight speed and main wing flapping frequency. ;

[0009] advance ratio and angle of attack The aerodynamic parameters corresponding to the index are hierarchically retrieved in the constructed aerodynamic feature lookup table. The aerodynamic parameters include aerodynamic coefficients and aerodynamic derivatives.

[0010] The retrieved aerodynamic parameters are used to update the aerodynamic parameters in the dynamic model of the flapping-wing aircraft;

[0011] The updated dynamic model calculates the pitch control variable based on the errors between the desired and actual forces, and the errors between the desired and actual torques. With yaw control And by combining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency, the deflection angles of the left and right tail wing servos are calculated;

[0012] Based on the deflection angle of the left and right tail servos, corresponding control commands are generated to control the left and right tail servos of the flapping-wing aircraft, thereby achieving attitude control of the flapping-wing aircraft.

[0013] Preferably, the constructed aerodynamic feature lookup table is implemented as follows:

[0014] S1-1. The flapping-wing aircraft is fixed in the test section inside the wind tunnel. The actual force and actual torque of the flapping-wing aircraft under different set flight conditions are obtained through wind tunnel experiments. At the same time, the wind tunnel wind speed under each flight condition is changed to make the wind tunnel wind speed 0, and the actual force and actual torque of the flapping-wing aircraft are collected.

[0015] Calculate the actual force difference between the actual force under each flight condition and the actual force under the flight condition where the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic force caused by wind speed.

[0016] Calculate the difference between the actual torque under each flight condition and the actual torque under the flight condition when the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic torque caused by wind speed.

[0017] The actual forces under each flight condition include lift. ,resistance and lateral force The actual forces under this flight condition include yaw moment. Pitch moment and rolling torque;

[0018] S1-2. Calculate the approach ratio for each flight condition based on the wind tunnel wind speed and main wing flapping frequency. ;

[0019] Based on the aerodynamic forces and aerodynamic moments under each flight condition, calculate the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient for each flight condition, and use all coefficients as aerodynamic coefficients.

[0020] For each aerodynamic coefficient under each flight condition, the partial derivatives of the approach ratio and angle of attack under that flight condition are obtained, and all the results are aerodynamic derivatives.

[0021] S1-3. Use the approach ratio and angle of attack for each flight condition as the index for that flight condition, and use the aerodynamic coefficients and aerodynamic derivatives of that flight condition as the aerodynamic parameters corresponding to the index to form an aerodynamic characteristic lookup table.

[0022] Preferably, the implementation method for different flight conditions is as follows:

[0023] During the wind tunnel experiment, the flapping frequency of the main wing of the flapping-wing aircraft fixed in the test section of the wind tunnel is controlled by the drive motor to flap in stages within the range of 1Hz to 7Hz. At each main wing flapping frequency, the wind tunnel wind speed and the prototype angle of attack are changed simultaneously. The wind tunnel wind speed is traversed within the wind tunnel wind speed range at the first preset step size, and the flight angle of attack is traversed within the flight angle of attack range at the second preset step size. The flight condition composed of the current main wing flapping frequency, the current wind tunnel wind speed, and the current flight angle of attack is obtained at each experimental moment.

[0024] Preferably, the wind tunnel wind speed range is from 0 m / s to 5 m / s; and the flight angle of attack ranges from -10 degrees to 30 degrees.

[0025] Preferably, the calculation of the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient for each flight condition is performed as follows:

[0026] Lift under each flight condition ,resistance Lateral force , yaw moment Pitch moment The lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient under this flight condition are calculated using quasi-steady aerodynamic methods, taking into account the aerodynamic forces and aerodynamic moments corresponding to the flight condition.

[0027] Preferably, , ;

[0028] The wind speed in the wind tunnel during the wind tunnel experiment. For flight speed, The wingspan is the length of the wing. The main wing flapping frequency.

[0029] Preferably, the method for obtaining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency is as follows:

[0030] S2-1, For situations involving main wing flapping:

[0031] During wind tunnel testing, the flapping-wing aircraft was fixed in the test section inside the wind tunnel. Under constant given wind speed and angle of attack, the main wing flapping frequency was controlled to flap in stages within a range greater than 0 Hz and less than or equal to 7 Hz. At each main wing flapping frequency, the deflection angles of the left and right tail fins of the flapping-wing aircraft were controlled time-divisionally.

[0032] At each main wing flapping frequency, when the left tail of the flapping wing aircraft is deflected and the right tail is not deflected, the left tail deflection angle range is traversed according to the preset third step length, and the left tail pitching moment under each deflection angle control at the main wing flapping frequency is collected.

[0033] At each main wing flapping frequency, when the right tail of the flapping wing aircraft is deflected and the left tail is not deflected, the right tail deflection angle range is traversed according to the preset third step length, and the pitching moment of the right tail under each deflection angle control at the main wing flapping frequency is collected.

[0034] S2-2. Take the ratio of the left and right tail wing pitching moments collected under the same deflection angle control at each main wing flapping frequency, and use this ratio as the distribution ratio required for the unit control torque at the current flapping frequency, and also as the decoupling hybrid gain coefficient at the current flapping frequency.

[0035] S3-2. Summarize the decoupling hybrid gain coefficients at all main wing flapping frequencies, and use polynomial fitting to obtain the decoupling hybrid gain coefficient function as the main wing flapping frequency changes.

[0036] Preferably, the deflection angle of the left tail fin is greater than or equal to -30 degrees and less than or equal to 45 degrees, with the downward deflection angle defined as a negative value and the upward deflection angle as a positive value;

[0037] The right tail fin deflection angle ranges from -30 degrees to 45 degrees, with downward deflection angle defined as negative and upward deflection angle as positive.

[0038] Preferably, the approach ratio is... and angle of attack The implementation method for hierarchically retrieving the corresponding aerodynamic parameters in the constructed aerodynamic feature lookup table, using the index as an index, is as follows:

[0039] First, calculate the approach distance ratio found in the aerodynamic feature lookup table. The total number of index matches that were successfully executed;

[0040] When the total number is 1, the advance ratio will be... The aerodynamic parameters corresponding to the index are output as the result;

[0041] When the total number is greater than or equal to 2, from all the advance ratios Among the aerodynamic parameters corresponding to the index, the flight angle of attack used as the index is taken. The aerodynamic parameters corresponding to the flight angle of attack with the smallest deviation are output as the result.

[0042] Preferably, the method for calculating the deflection angles of the left and right tail fin servos is as follows:

[0043] ;

[0044] in, and These are the deflection angles of the left and right tail servos, respectively. The decoupled hybrid gain coefficient function for the main wing flapping frequency varies with the main wing flapping frequency. The corresponding decoupled hybrid gain coefficient.

[0045] The beneficial effects of this invention are:

[0046] This invention proposes a decoupled control method for an inverted V-tail flapping-wing aircraft based on approach ratio scheduling. It introduces an aerodynamic characteristic lookup table indexed by approach ratio and flight angle of attack, constructed through wind tunnel experiments (covering all operating conditions including frequency, speed, and angle of attack). During flight, the approach ratio and angle of attack are calculated in real-time, and aerodynamic parameters in the dynamic model are dynamically retrieved and updated based on the aerodynamic characteristic lookup table, thus solving the problem of inaccurate fixed-parameter modeling. Secondly, a frequency-adaptive decoupled control matrix is ​​designed, introducing a decoupled hybrid gain coefficient function that varies with the main wing flapping frequency, thus controlling the pitch control quantity. With yaw control Dynamically assigning servos to the left and right tail wing servos effectively compensates for channel coupling caused by airflow washing. This invention achieves real-time correction of aerodynamic parameters and dynamic decoupling of pitch and yaw in flapping-wing aircraft without adding additional actuators, significantly improving the attitude control accuracy and stability of small flapping-wing robots in variable-frequency flapping and complex unsteady aerodynamic environments.

[0047] This invention can effectively reduce the impact of unsteady wash flow on tail control efficiency and weaken the coupling interference between pitch and yaw channels, thereby improving the attitude control stability and control accuracy of small flapping-wing aircraft in complex aerodynamic environments. Attached Figure Description

[0048] Figure 1 This is a schematic diagram illustrating the principle of the decoupling control method for an inverted V-tail flapping-wing aircraft based on approach ratio scheduling as described in this invention.

[0049] Figure 2 This is a schematic diagram of a 210g inverted V-shaped tail flapping-wing aircraft and its coordinate system. Detailed Implementation

[0050] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0051] It should be noted that, unless otherwise specified, the embodiments and features described in the present invention can be combined with each other.

[0052] The present invention will be further described below with reference to the accompanying drawings and specific embodiments, but this is not intended to limit the scope of the invention.

[0053] Specific Implementation Method 1: Combination Figure 1 This embodiment describes a decoupling control method for an inverted V-tail flapping-wing aircraft based on approach ratio scheduling. The flapping frequencies of the pair of main wings of the flapping-wing aircraft are the same. The method includes:

[0054] Acquire the current flight status information of the flapping-wing aircraft, including the main wing flapping frequency, flight speed, inertial measurement information, and the actual forces and torques acting on the flapping-wing aircraft;

[0055] The inertial measurement information is fused and processed to obtain the flight angle of attack. Inertial measurement information includes angular velocity, specific force, and flight altitude.

[0056] Calculate the approach ratio under the current flight conditions based on flight speed and main wing flapping frequency. ;

[0057] advance ratio and angle of attack The aerodynamic parameters corresponding to the index are hierarchically retrieved in the constructed aerodynamic feature lookup table. The aerodynamic parameters include aerodynamic coefficients and aerodynamic derivatives.

[0058] The retrieved aerodynamic parameters are used to update the aerodynamic parameters in the dynamic model of the flapping-wing aircraft;

[0059] The updated dynamic model calculates the pitch control variable based on the errors between the desired and actual forces, and the errors between the desired and actual torques. With yaw control And by combining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency, the deflection angles of the left and right tail wing servos are calculated;

[0060] Based on the deflection angle of the left and right tail servos, corresponding control commands are generated to control the left and right tail servos of the flapping-wing aircraft, thereby achieving attitude control of the flapping-wing aircraft.

[0061] This preferred embodiment achieves effective decoupled control of the pitch and yaw channels of an inverted V-tail flapping-wing aircraft by introducing a hierarchical aerodynamic parameter retrieval mechanism based on approach ratio scheduling and combining it with a decoupled hybrid gain function that varies with flapping frequency. This method utilizes flight conditions (speed, angle of attack, flapping frequency) to update the aerodynamic parameters in the existing dynamic model in real time, enabling control commands to more accurately match real aerodynamic characteristics, thereby significantly improving the stability and adaptability of attitude control. Simultaneously, the use of a lookup table method to replace complex aerodynamic calculations reduces the online computational burden on the flight control system, facilitating engineering implementation. Ultimately, this method ensures that the flapping-wing aircraft obtains reliable independent pitch and yaw control capabilities under different flapping frequencies and flight conditions (such as hovering and forward flight).

[0062] In practical applications ; For flight speed, The wingspan is the length of the wing. The main wing flapping frequency.

[0063] The inertial measurement information is fused and processed to obtain the flight angle of attack. This can be achieved using existing technologies; the inertial measurement information includes the angular velocity output by the three-axis gyroscope, the specific force output by the three-axis accelerometer, and the flight altitude output by the barometer. The angular velocity, specific force, and flight altitude are input into the extended Kalman filter for multi-source fusion, and the flight angle of attack is output.

[0064] Furthermore, the implementation of the constructed aerodynamic feature lookup table is as follows:

[0065] S1-1. The flapping-wing aircraft is fixed in the test section inside the wind tunnel. The actual force and actual torque of the flapping-wing aircraft under different set flight conditions are obtained through wind tunnel experiments. At the same time, the wind tunnel wind speed under each flight condition is changed to make the wind tunnel wind speed 0, and the actual force and actual torque of the flapping-wing aircraft are collected.

[0066] Specifically, the implementation methods for different flight conditions are as follows:

[0067] During the wind tunnel experiment, the flapping frequency of the main wing of the flapping-wing aircraft, fixed in the test section inside the wind tunnel, is controlled by a drive motor to flap in stages within the range of 1Hz to 7Hz. At each main wing flapping frequency, the wind tunnel wind speed and the prototype angle of attack are changed simultaneously. The wind tunnel wind speed is traversed within the first preset step size, and the flight angle of attack is traversed within the flight angle of attack range with a second preset step size. The resulting flight condition at each experimental moment is a combination of the current main wing flapping frequency, the current wind tunnel wind speed, and the current flight angle of attack. The wind tunnel wind speed range is from 0 m / s to 5 m / s, and the flight angle of attack range is from -10 degrees to 30 degrees. During the experiment, the wind tunnel wind speed is changed in stages with a first preset step size of 1 m / s, and the flight angle of attack is changed in stages with a second preset step size of 5 degrees.

[0068] Calculate the actual force difference between the actual force under each flight condition and the actual force under the flight condition where the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic force caused by wind speed.

[0069] Calculate the difference between the actual torque under each flight condition and the actual torque under the flight condition when the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic torque caused by wind speed.

[0070] The actual forces under each flight condition include lift. ,resistance and lateral force The actual forces under this flight condition include yaw moment. Pitch moment and rolling torque;

[0071] S1-2. Calculate the approach ratio for each flight condition based on the wind tunnel wind speed and main wing flapping frequency. ;in, ;

[0072] The wind speed in the wind tunnel during the wind tunnel experiment. The wingspan is the length of the wing. Main wing flapping frequency;

[0073] Based on the aerodynamic forces and aerodynamic moments under each flight condition, calculate the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient for each flight condition, and use all coefficients as aerodynamic coefficients.

[0074] Specifically, the calculation methods for the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient under each flight condition are as follows:

[0075] Lift under each flight condition ,resistance Lateral force , yaw moment Pitch moment The lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient under this flight condition are calculated by combining the aerodynamic force and aerodynamic moment corresponding to the flight condition with the quasi-steady aerodynamic method.

[0076] For each aerodynamic coefficient under each flight condition, the partial derivatives of the approach ratio and angle of attack under that flight condition are obtained, and all the results are aerodynamic derivatives.

[0077] S1-3. Use the approach ratio and angle of attack for each flight condition as the index for that flight condition, and use the aerodynamic coefficients and aerodynamic derivatives of that flight condition as the aerodynamic parameters corresponding to the index to form an aerodynamic characteristic lookup table.

[0078] In this preferred embodiment, wind tunnel experiments are conducted to measure the actual forces and moments of the flapping-wing aircraft under various flight conditions and a zero-wind-speed baseline. The differences are calculated to extract the aerodynamic forces and moments purely caused by wind speed, thereby obtaining key data such as lift, drag, lateral force, and yaw, pitch, and roll moments. Furthermore, based on the approach ratio and angle of attack for each condition, six corresponding aerodynamic coefficients and their partial derivatives (i.e., aerodynamic derivatives) with respect to the approach ratio and angle of attack are calculated. Finally, using the approach ratio and angle of attack as indexes, an aerodynamic characteristic lookup table is constructed by combining the aerodynamic coefficients and aerodynamic derivatives. This implementation ensures that aerodynamic parameters dynamically match the flight state, providing accurate and efficient aerodynamic data support for the real-time decoupled control of the flapping-wing aircraft.

[0079] When applying, Figure 2A small, biomimetic flapping-wing aircraft weighing approximately 210g was used as the research object for wind tunnel experiments. The aircraft is driven by a single motor, with the main wing connected to a gear and linkage mechanism to achieve periodic flapping, generating the main lift and thrust. The tail section features an inverted V-shaped tail configuration, consisting of a left tail control surface and a right tail control surface, symmetrically arranged in an inverted V shape at the rear of the fuselage. Each tail control surface is driven by an independent servo, with the left servo mechanically connected to the left tail control surface and the right servo mechanically connected to the right tail control surface. Both servos receive PWM control signals from the flight control system and deflect to the target angle according to the signal commands. Co-directional deflection of the left and right control surfaces (both upward or downward simultaneously) generates a pitch moment, causing the nose to rise or fall; differential deflection (one upward and one downward) generates a yaw moment, causing the nose to turn left or right. The flapping-wing aircraft does not have an independent roll control mechanism (such as ailerons or differential flapping wings), and mainly maintains roll stability indirectly through the following two methods: First, the main wing adopts an appropriate dihedral or sweep angle design, so that the aircraft generates a restoring moment when subjected to roll disturbances; second, when the aircraft experiences lateral drift (i.e., lateral position shift) or heading deviation, differential deflection of the tail fin generates lateral aerodynamic forces, thereby suppressing drift. The prototype structural schematic diagram and airframe coordinate system are as follows: Figure 2 As shown.

[0080] Furthermore, the method for obtaining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency is as follows:

[0081] S2-1, For situations involving main wing flapping:

[0082] During wind tunnel testing, the flapping-wing aircraft was fixed in the test section inside the wind tunnel. Under constant given wind speed and angle of attack, the main wing flapping frequency was controlled to flap in stages within a range greater than 0 Hz and less than or equal to 7 Hz. At each main wing flapping frequency, the deflection angles of the left and right tail fins of the flapping-wing aircraft were controlled time-divisionally.

[0083] At each main wing flapping frequency, when the left tail of the flapping wing aircraft is deflected while the right tail is not deflected, the left tail deflection angle range is traversed at the same preset third step length, and the left tail pitching moment under each deflection angle control at that main wing flapping frequency is collected; the left tail deflection angle range is greater than or equal to -30 degrees and less than or equal to 45 degrees, and the downward deflection angle is defined as a negative value and the upward deflection angle as a positive value.

[0084] At each main wing flapping frequency, when the right tail of the flapping wing aircraft is deflected while the left tail is not deflected, the right tail deflection angle range is traversed at the same preset third step length, and the pitching moment of the right tail under each deflection angle control at that main wing flapping frequency is collected; the right tail deflection angle range is greater than or equal to -30 degrees and less than or equal to 45 degrees, and the downward deflection angle is defined as a negative value and the upward deflection angle as a positive value.

[0085] S2-2. Take the ratio of the left and right tail wing pitching moments collected under the same deflection angle control at each main wing flapping frequency, and use this ratio as the distribution ratio required for the unit control torque at the current flapping frequency, and also as the decoupling hybrid gain coefficient at the current flapping frequency.

[0086] S3-2. Summarize the decoupling hybrid gain coefficients at all main wing flapping frequencies, and use polynomial fitting to obtain the decoupling hybrid gain coefficient function as the main wing flapping frequency changes.

[0087] Furthermore, the approach ratio and angle of attack The implementation method for hierarchically retrieving the corresponding aerodynamic parameters in the constructed aerodynamic feature lookup table, using the index as an index, is as follows:

[0088] First, calculate the approach distance ratio found in the aerodynamic feature lookup table. The total number of index matches that were successfully executed;

[0089] When the total number is 1, the advance ratio will be... The aerodynamic parameters corresponding to the index are output as the result;

[0090] When the total number is greater than or equal to 2, from all the advance ratios Among the aerodynamic parameters corresponding to the index, the flight angle of attack used as the index is taken. The aerodynamic parameters corresponding to the flight angle of attack with the smallest deviation are output as the result.

[0091] In this preferred embodiment, when using approach ratio and angle of attack as indexes to query the aerodynamic feature lookup table, the total number of matching approach ratio indices is first determined. If there is only one match, the aerodynamic parameters corresponding to that index are directly output; if there are two or more matches, the set of aerodynamic parameters with the smallest deviation between the flight angle of attack and the current index value is selected as the final output. This hierarchical retrieval strategy balances search efficiency and matching accuracy, ensuring that aerodynamic parameters closest to the actual operating conditions can be quickly and reasonably obtained under various flight conditions, thereby improving the accuracy of dynamic model updates and the reliability of control response.

[0092] Furthermore, the method for calculating the deflection angles of the left and right tail servos is as follows:

[0093] ;

[0094] in, and These are the deflection angles of the left and right tail servos, respectively. The decoupled hybrid gain coefficient function for the main wing flapping frequency varies with the main wing flapping frequency. The corresponding decoupled hybrid gain coefficient.

[0095] In this preferred embodiment, a frequency-adaptive decoupling control matrix is ​​designed, specifically a transformation matrix containing a decoupling hybrid gain coefficient function that varies with the main wing flapping frequency, which converts the pitch control quantity... With yaw control The linear combination is the deflection angle of the left and right tail servos. and In this matrix, the first row achieves the co-directional superposition of the two control variables to produce a pitch effect, and the second row achieves the anti-directional superposition to produce a yaw effect, while the gain coefficient... The pitch and yaw channels are dynamically adjusted according to the main wing flapping frequency, thus adaptively adjusting the mixing ratio based on the flight frequency. As the flapping frequency increases, the intensity of the unsteady wash generated by the main wing increases, altering the airflow structure near the tail. By dynamically adjusting the mixing gain coefficient, the control coupling problem caused by changes in tail aerodynamic efficiency can be compensated. This method effectively achieves decoupled control of the pitch and yaw motions of the flapping wing aircraft, allowing control commands from both channels to act independently on the inverted V-tail, avoiding attitude interference caused by coupling, and improving the accuracy and stability of attitude control at different flapping frequencies.

[0096] While the invention has been described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the invention. Therefore, it should be understood that many modifications can be made to the exemplary embodiments, and other arrangements can be designed without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that different dependent claims and features described herein can be combined in ways different from those described in the original claims. It is also understood that features described in conjunction with individual embodiments can be used in other described embodiments.

Claims

1. A decoupled control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling, wherein the flapping frequencies of a pair of main wings of the flapping-wing aircraft are the same, characterized in that... The method includes: Acquire the current flight status information of the flapping-wing aircraft, including the main wing flapping frequency, flight speed, inertial measurement information, and the actual forces and torques acting on the flapping-wing aircraft; The inertial measurement information is fused and processed to obtain the flight angle of attack. Inertial measurement information includes angular velocity, specific force, and flight altitude. Calculate the approach ratio under the current flight conditions based on flight speed and main wing flapping frequency. ; advance ratio and angle of attack The aerodynamic parameters corresponding to the index are hierarchically retrieved in the constructed aerodynamic feature lookup table. The aerodynamic parameters include aerodynamic coefficients and aerodynamic derivatives. The retrieved aerodynamic parameters are used to update the aerodynamic parameters in the dynamic model of the flapping-wing aircraft; The updated dynamic model calculates the pitch control variable based on the errors between the desired and actual forces, and the errors between the desired and actual torques. With yaw control And by combining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency, the deflection angles of the left and right tail wing servos are calculated; Based on the deflection angle of the left and right tail servos, corresponding control commands are generated to control the left and right tail servos of the flapping-wing aircraft, thereby achieving attitude control of the flapping-wing aircraft.

2. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 1, characterized in that, The implementation of the constructed aerodynamic feature lookup table is as follows: S1-1. The flapping-wing aircraft is fixed in the test section inside the wind tunnel. The actual force and actual torque of the flapping-wing aircraft under different set flight conditions are obtained through wind tunnel experiments. At the same time, the wind tunnel wind speed under each flight condition is changed to make the wind tunnel wind speed 0, and the actual force and actual torque of the flapping-wing aircraft are collected. Calculate the actual force difference between the actual force under each flight condition and the actual force under the flight condition where the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic force caused by wind speed. Calculate the difference between the actual torque under each flight condition and the actual torque under the flight condition when the wind tunnel wind speed is changed to 0, and use this difference as the aerodynamic torque caused by wind speed. The actual forces under each flight condition include lift. ,resistance and lateral force The actual forces under this flight condition include yaw moment. Pitch moment and rolling torque; S1-2. Calculate the approach ratio for each flight condition based on the wind tunnel wind speed and main wing flapping frequency. ; Based on the aerodynamic forces and aerodynamic moments under each flight condition, calculate the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient for each flight condition, and use all coefficients as aerodynamic coefficients. For each aerodynamic coefficient under each flight condition, partial derivatives are taken with respect to the approach ratio and angle of attack under that flight condition. All results are aerodynamic derivatives. S1-3. Use the approach ratio and angle of attack for each flight condition as the index for that flight condition, and use the aerodynamic coefficients and aerodynamic derivatives of that flight condition as the aerodynamic parameters corresponding to the index to form an aerodynamic characteristic lookup table.

3. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 2, characterized in that, The implementation methods for different flight conditions are set as follows: During the wind tunnel experiment, the flapping frequency of the main wing of the flapping-wing aircraft fixed in the test section of the wind tunnel is controlled by the drive motor to flap in stages within the range of 1Hz to 7Hz. At each main wing flapping frequency, the wind tunnel wind speed and the prototype angle of attack are changed simultaneously. The wind tunnel wind speed is traversed within the wind tunnel wind speed range at the first preset step size, and the flight angle of attack is traversed within the flight angle of attack range at the second preset step size. The flight condition composed of the current main wing flapping frequency, the current wind tunnel wind speed, and the current flight angle of attack is obtained at each experimental moment.

4. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 3, characterized in that, The wind tunnel wind speed range is from 0 m / s to 5 m / s; the flight angle of attack range is from -10 degrees to 30 degrees.

5. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 2, characterized in that, The calculation method for the lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient under each flight condition is as follows: Lift under each flight condition ,resistance Lateral force , yaw moment Pitch moment The lift coefficient, drag coefficient, lateral force coefficient, yaw moment coefficient, pitch moment coefficient, and roll moment coefficient under this flight condition are calculated using quasi-steady aerodynamic methods, taking into account the aerodynamic forces and aerodynamic moments corresponding to the flight condition.

6. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 2, characterized in that, ， ； The wind speed in the wind tunnel during the wind tunnel experiment. For flight speed, The wingspan is the length of the wing. The main wing flapping frequency.

7. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 1, characterized in that, The method for obtaining the decoupled hybrid gain coefficient function that varies with the main wing flapping frequency is as follows: S2-1, For situations involving main wing flapping: During wind tunnel testing, the flapping-wing aircraft was fixed in the test section inside the wind tunnel. Under constant given wind speed and angle of attack, the main wing flapping frequency was controlled to flap in stages within a range greater than 0 Hz and less than or equal to 7 Hz. At each main wing flapping frequency, the deflection angles of the left and right tail fins of the flapping-wing aircraft were controlled time-divisionally. At each main wing flapping frequency, when the left tail of the flapping wing aircraft is deflected and the right tail is not deflected, the left tail deflection angle range is traversed according to the preset third step length, and the left tail pitching moment under each deflection angle control at the main wing flapping frequency is collected. At each main wing flapping frequency, when the right tail of the flapping wing aircraft is deflected and the left tail is not deflected, the right tail deflection angle range is traversed according to the preset third step length, and the pitching moment of the right tail under each deflection angle control at the main wing flapping frequency is collected. S2-2. Take the ratio of the left and right tail wing pitching moments collected under the same deflection angle control at each main wing flapping frequency, and use this ratio as the distribution ratio required for the unit control torque at the current flapping frequency, and also as the decoupling hybrid gain coefficient at the current flapping frequency. S3-2. Summarize the decoupling hybrid gain coefficients at all main wing flapping frequencies, and use polynomial fitting to obtain the decoupling hybrid gain coefficient function as the main wing flapping frequency changes.

8. The decoupling control method for an inverted V-tail flapping-wing aircraft based on approach ratio scheduling according to claim 7, characterized in that, The deflection angle of the left tail fin is greater than or equal to -30 degrees and less than or equal to 45 degrees. The downward deflection angle is defined as a negative value and the upward deflection angle as a positive value. The right tail fin deflection angle ranges from -30 degrees to 45 degrees, with downward deflection angle defined as negative and upward deflection angle as positive.

9. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 1, characterized in that, advance ratio and angle of attack The implementation method for hierarchically retrieving the corresponding aerodynamic parameters in the constructed aerodynamic feature lookup table using the index is as follows: First, calculate the approach distance ratio found in the aerodynamic feature lookup table. The total number of index matches that were successful; When the total number is 1, the advance ratio will be... The aerodynamic parameters corresponding to the index are output as the result; When the total number is greater than or equal to 2, from all the advance ratios Among the aerodynamic parameters corresponding to the index, the flight angle of attack used as the index is taken. The aerodynamic parameters corresponding to the flight angle of attack with the smallest deviation are output as the result.

10. The decoupling control method for an inverted V-tail flapping-wing aircraft based on advance ratio scheduling according to claim 1, characterized in that, The method for calculating the deflection angles of the left and right tail fin servos is as follows: ; in, and These are the deflection angles of the left and right tail servos, respectively. The decoupled hybrid gain coefficient function for the main wing flapping frequency varies with the main wing flapping frequency. The corresponding decoupled hybrid gain coefficient.

Images