Transition section speed fusion and pitch control right transfer method for compound wing unmanned aerial vehicle

Abstract

The present application belongs to the technical field of unmanned aerial vehicle flight control, and particularly relates to a transition section speed fusion and pitch control right handover method for a compound wing unmanned aerial vehicle, which comprises the following steps: acquiring a forward ground speed, a vertical ground speed and a calibrated air speed output by an air speed tube in real time; in a first speed conversion interval, performing dynamic weighted fusion on the ground speed and the calibrated air speed with a first weight factor to output a continuous and smooth fused forward speed; independently and in parallel, solving a first expected pitch angle of a multi-rotor speed control path and a second expected pitch angle of a fixed-wing altitude control path; in a second speed conversion interval, performing dynamic gain mixed control on the two-path pitch angles with a fixed-wing gain weight to realize seamless and gradual handover of the pitch control right from the multi-rotor mode to the fixed-wing mode. The above method effectively solves the mutation problem caused by hard switching of the speed signal in the transition phase and the pitch control logic conflict problem, and improves the safety and stability of the transition flight.

Description

Technical Field

[0001] This invention belongs to the field of unmanned aerial vehicle (UAV) flight control technology, specifically relating to a method for speed fusion and pitch control handover during the transition phase of a compound wing UAV. Background Technology

[0002] Compound-wing UAVs (also known as vertical takeoff and landing fixed-wing UAVs) combine the flexibility of multi-rotor UAVs with the economic advantages of fixed-wing UAVs, such as long endurance and high-speed cruise. They are widely used in surveying, inspection, and logistics. Their core flight phases include: vertical takeoff and landing, multi-rotor hovering, level flight transition, and fixed-wing cruise. The level flight transition, in particular, refers to the phase where the aircraft accelerates from the low-speed flight of a multi-rotor to the fixed-wing stage where sufficient lift is generated to maintain altitude. This phase involves the most dramatic changes in aerodynamic characteristics and the most complex control logic within the entire flight envelope.

[0003] The existing technology has the following main technical shortcomings in handling this transition phase:

[0004] 1. The issue of conflicting sensor data reliability: Calibrated airspeed (CAS) measured by pitot tubes is easily interfered with by the strong downwash from multirotors during low speeds or hovering, resulting in severely distorted measurements that cannot be directly used for flight control. Furthermore, the ground speed calculated by the inertial navigation system (INS) cannot accurately reflect the true speed of the incoming airflow to the wings in windy conditions. Relying solely on ground speed for flight mode switching can easily lead to stall or energy overruns in headwinds or tailwinds. Existing solutions often employ hard threshold switching (i.e., abruptly switching the speed signal source at a fixed speed point). This hard threshold switching causes a step change in the speed error signal input to the closed-loop controller, leading to severe vibrations in the tail elevator or multirotor motors. This not only increases mechanical fatigue wear but also seriously jeopardizes the structural safety of the aircraft.

[0005] 2. Conflict in Pitch Attitude Control Logic: In multi-rotor flight mode, the aircraft controls longitudinal acceleration and deceleration by adjusting the pitch angle (raising or lowering the nose), with the pitch angle acting as the actuator for speed control. In fixed-wing flight mode, the pitch angle affects lift by changing the angle of attack, thus controlling altitude; the pitch angle acts as the actuator for altitude control. These two control logics are physically conflicting. If only one logic is used during the transition phase, it is impossible to simultaneously meet the dual requirements of acceleration and altitude maintenance, easily leading to dangerous situations such as sudden nose-up (risk of stall) or nose-down (risk of crash). Existing technology lacks a decoupling calculation and smooth handover mechanism for these two logics.

[0006] Currently, there is no existing technology that provides a systematic solution to both of the above problems simultaneously. Summary of the Invention

[0007] To address the aforementioned problems in the existing technology, this invention provides a method for speed fusion and pitch control handover during the transition phase of a compound-wing UAV. This method overcomes the technical defects of hard switching of speed signals and pitch control logic conflicts in the existing technology, and achieves a continuous and smooth transition of speed signals and a seamless dynamic handover of attitude control. It is used by compound-wing UAVs to achieve dynamic weighted fusion of sensor speed signals and smooth handover of pitch attitude control during the level flight transition phase from multi-rotor mode to fixed-wing mode.

[0008] To achieve the above objectives, the present invention adopts the following specific technical solution: In a first aspect, the present invention provides a method for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV), the method comprising the following steps: Step 1: Obtain the forward ground speed in the horizontal coordinate system output by the aircraft's inertial navigation system. and vertical ground speed And the calibrated airspeed measured by the pitot tube. ; Step 2, preset the first speed conversion range [ , According to the forward ground speed Real-time calculation of the first weighting factor And will calibrate airspeed With forward ground speed According to the first weighting factor The forward fusion velocity is calculated by performing a weighted summation. ; Step 3: Independently and in parallel calculate the first desired pitch angle used to control longitudinal acceleration and deceleration. θ mc and the second desired pitch angle used to control flight altitude ; Step 4, preset the second speed conversion range [ , According to the calibrated airspeed Real-time calculation of fixed-wing gain weights and multi-rotor gain weight ; the first desired pitch angle With multi-rotor gain weight The product of, and the second desired pitch angle With fixed-wing gain weight Sum the products and output the final total pitch angle command. And then send it to the attitude inner loop controller.

[0009] Furthermore, in step 2, the first weighting factor With fusion forward velocity The calculation formula is:

[0010]

[0011] Among them, when ≤ hour, =0, = The system is completely trusted at high speed; when ≥ hour, = 1, = The system fully trusts the calibrated airspeed; within the range [ , ]Inside, Follow Linearly increasing, This is a linear interpolation fusion of the two speed paths.

[0012] Furthermore, in step 3, the first desired pitch angle The solution steps are as follows: Obtain the total forward longitudinal acceleration demand command output by the flight control system. ; Determine the sign of the instruction: when At that time, the acceleration command is fully allocated to the propulsion system, and the multi-rotor pitch output... ;when At that time, the negative acceleration demand is converted into a pitch angle command for the multi-rotor system according to a preset ratio, which is used as the first desired pitch angle. .

[0013] Furthermore, in step 3, the second desired pitch angle The solution steps are as follows: Receive desired vertical velocity command The output reference vertical velocity is smoothed by a first-order command filter. and desired vertical acceleration ; Desired vertical acceleration Multiply by the feedforward gain factor Obtain the feedforward pitch angle compensation amount ; Calculate the reference vertical velocity actual vertical ground speed error , will the error Input proportional-integral controller, output feedback pitch angle compensation. ; feedforward pitch angle compensation With feedback pitch angle compensation amount The superimposed values, after being processed by the upper and lower limiters to prevent fixed-wing stall, output the second desired pitch angle. .

[0014] Furthermore, in step 4, the fixed-wing gain weights and final total pitch command The calculation formula is:

[0015]

[0016] when hour, , The pitch angle is entirely dominated by the multi-rotor speed control logic; when hour, , The pitch angle is completely handed over to the fixed-wing altitude maintenance logic, and the forward speed regulation is entirely undertaken by the propulsion motor, realizing the decoupling of the parallel power system.

[0017] Furthermore, the limiting range of the upper and lower limiters is determined based on the maximum and minimum safe angle of attack of the aircraft model to prevent wing stall or negative lift caused by excessive or insufficient pitch angle in fixed-wing mode.

[0018] Furthermore, the first speed transition range [ , [and the second speed conversion range] , The two speed conversion zones are independent of each other and configured independently according to the aerodynamic characteristics of the aircraft model; the first speed conversion zone is used to eliminate the speed range of the pitot tube from the interference of the rotor downwash; the second speed conversion zone is used to gradually control the altitude of the aircraft by the lift of the wing. The coverage of the two zones may partially overlap or be adjacent.

[0019] In a second aspect, the present invention provides a system for transition speed fusion and pitch control handover of a compound wing UAV for performing the transition speed fusion and pitch control handover method of the first aspect above. The system includes a state data acquisition module, an air-to-ground speed dynamic fusion module, a dual-mode pitch calculation module, and a pitch control gain mixing module. The state data acquisition module is used to acquire the forward ground speed and vertical ground speed in the horizontal coordinate system of the aircraft output by the aircraft's inertial navigation system, as well as the calibrated airspeed measured by the airspeed tube. The air-to-ground speed dynamic fusion module is used to calculate the first weighting factor in real time based on the set first speed conversion range and forward ground speed, and to obtain the fused forward speed by weighted summation of the calibrated airspeed and forward ground speed. The dual-mode pitch calculation module is used to independently and in parallel calculate the first desired pitch angle for controlling longitudinal acceleration and deceleration and the second desired pitch angle for controlling flight altitude. The pitch control weight gain mixing module is used to calculate the gain weights of the fixed wing and multi-rotor in real time according to the set second speed conversion range and calibration airspeed, and to perform a weighted summation of the first desired pitch angle and the second desired pitch angle to output the final total pitch angle command.

[0020] Thirdly, the present invention also provides a flight controller, which includes a memory, a processor, and a control program stored in the memory and capable of running on the processor. When the processor executes the control program, it is used to implement the method for speed fusion and pitch control handover during the transition phase of the compound wing UAV described in the first aspect.

[0021] Fourthly, the present invention also provides a computer-readable storage medium storing a computer program that, when executed by a processor, is used to implement the method for speed fusion and pitch control handover during the transition phase of the compound-wing unmanned aerial vehicle described in the first aspect.

[0022] Compared with the prior art, the technical solution of the present invention has the following beneficial effects: 1. Continuous and smooth speed signal: By dynamically weighting and fusing ground speed and calibrated airspeed within the first speed transition range, the problem of sudden changes in speed control commands caused by traditional hard threshold switching is completely eliminated, ensuring the continuity of control commands and flight stability across the entire speed range.

[0023] 2. Seamless handover of pitch control: Through the decoupled independent calculation of pitch angle in dual modes and the dynamic gain hybrid control mechanism, a smooth and gradual transition of the aircraft pitch control authority from "multi-rotor mode - speed control" to "fixed-wing mode - altitude control" is achieved, which fundamentally solves the physical conflict between the two control logics and effectively avoids the risk of stall and altitude drop faced by compound wing UAVs during the transition phase.

[0024] 3. Complete decoupling of the power system: After the handover of control is completed (during the pure cruise phase of the fixed wing), the pitch angle is completely dominated by the altitude control logic, and the forward acceleration and deceleration are completely managed independently by the propulsion motor, thus achieving complete decoupling of the parallel power system of the compound wing.

[0025] 4. Strong parameter adaptability: The upper and lower limits of each speed conversion range can be independently configured according to the aerodynamic characteristics of the specific aircraft model, the pitot tube installation position, and the influence range of the rotor downwash, making it highly versatile. Attached Figure Description

[0026] Figure 1 This is an overall flowchart of the method for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV) according to the present invention. Figure 2 This is a block diagram illustrating the principle of step 2; Figure 3 This is a block diagram illustrating the calculation principle of the second desired pitch angle in step 3. Figure 4 This is the principle block diagram for step 4. Detailed Implementation

[0027] 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.

[0028] Figure 1 This is an overall flowchart of the method for velocity fusion and pitch control handover during the transition phase of the compound-wing UAV of the present invention, showing the complete signal processing link from sensor data acquisition to final pitch command output. Figure 2 The diagram below shows the principle block diagram of the air-to-ground speed dynamic fusion module in step 2, illustrating the signal flow of the fused forward velocity output by the fusion of ground speed and calibrated air speed through weighting factors. Figure 3 The diagram shows the principle block diagram of the fixed-wing path pitch angle calculation module (altitude holding loop) in the dual-mode pitch calculation module in step 3, illustrating the internal structure and signal connection relationship of the first-order command filter, feedforward channel, PI feedback controller and limiter. Figure 4 The diagram below shows the principle block diagram of the pitch control weight gain mixing module in step 4, illustrating the linear variation law of fixed-wing gain weight with calibrated airspeed and the weighted superposition logic of dual-mode pitch angle commands.

[0029] In each figure, The forward ground speed is obtained by the inertial navigation system (INS). The vertical ground speed is obtained by the inertial navigation system (INS). CAS Airspeed is obtained from the pitot tube for airspeed calibration. To integrate the forward velocity, the output is dynamically weighted and integrated from the air-to-ground velocity, and then used by the outer loop speed controller. This is the lower limit of the first speed conversion interval, used to calculate the air-to-ground speed fusion weighting factor. This is the upper limit of the first speed conversion interval, used to calculate the air-to-ground speed fusion weighting factor. . This is the lower limit of the second velocity conversion interval, used to calculate the fixed-wing gain weight. . The upper limit of the second velocity conversion range is used to calculate the fixed-wing gain weight. . The first desired pitch angle is the pitch angle output generated by the multi-rotor mode speed control path logic. This is the second desired pitch angle, which is the pitch angle output generated by the fixed-wing mode altitude control path logic. The final total pitch angle command is generated by weighted synthesis of the desired pitch angles from both modes and then sent to the attitude inner loop controller. The feedforward pitch angle compensation is obtained by multiplying the desired vertical acceleration by the feedforward gain coefficient. The pitch angle compensation is calculated from the vertical velocity error using a proportional-integral (PI) controller. This is the lower limit of the pitch angle. The upper limit of the pitch angle is set based on the aircraft's safe angle of attack and is used to limit the pitch angle command to prevent stall. As the first weighting factor, from forward ground speed The weighted fusion of air-to-ground speeds is calculated by combining the threshold of the first speed range. For fixed-wing gain weights, determined by the calibrated airspeed CAS Based on the calculation of the second speed interval threshold, the value range is limited to [0,1]. The gain weights for multi-rotor aircraft are complementary to those for fixed-wing aircraft. The deceleration demand coefficient is used in the multi-rotor speed control path to map the negative longitudinal acceleration demand into a specific pitch angle command. This is the feedforward gain coefficient, used for feedforward pre-compensation calculations in the fixed-wing altitude control path. The proportional gain coefficient (proportional parameter) of the proportional-integral PI controller in the feedback branch. This refers to the integral gain coefficient (integral parameter) of the PI controller in the feedback branch. The longitudinal acceleration requirement command serves as the input judgment condition for the speed control path in multi-rotor mode. The desired vertical velocity command serves as the initial input command for the altitude control path in fixed-wing mode. The reference vertical velocity is the output of the desired vertical velocity command after being smoothed by a first-order command filter. The desired vertical acceleration is also the output of the first-order command filter. This refers to the vertical channel feedback error, which is the difference between the reference vertical velocity and the actual vertical ground velocity. ). The filtering time constant of a first-order command filter is reflected in the transfer function. middle.

[0030] Example 1 This invention provides a method for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV), such as... Figure 1 As shown, the method includes the following steps: Step 1: Obtain the forward ground speed in the horizontal coordinate system output by the aircraft's inertial navigation system. and vertical ground speed And the calibrated airspeed measured by the pitot tube. CAS ; Step 2, preset the first speed conversion range [ , According to the forward ground speed Real-time calculation of the first weighting factor And will calibrate airspeed CAS With forward ground speed According to the first weighting factor The forward fusion velocity is calculated by performing a weighted summation. ,like Figure 2 As shown; First weighting factor With fusion forward velocity The calculation formula is:

[0031]

[0032] Among them, when ≤ hour, =0, = The system is completely trusted at high speed; when ≥ hour, = 1, = CAS The system fully trusts the calibrated airspeed; within the range [ , ]Inside, Follow Linearly increasing, This is a linear interpolation fusion of the two speed paths.

[0033] Step 3: Independently and in parallel calculate the first desired pitch angle used to control longitudinal acceleration and deceleration. and the second desired pitch angle used to control flight altitude ; First desired pitch angle The solution steps are as follows: Obtain the total forward longitudinal acceleration demand command output by the flight control system. ; Determine the symbol of the instruction: when When the value is ≥0, the acceleration command is fully allocated to the propulsion system, and the multi-rotor pitch output... =0; when When the negative acceleration requirement is less than 0, the negative acceleration requirement is converted into a pitch angle command for the multi-rotor system according to a preset ratio, which is used as the first desired pitch angle. .

[0034] Second desired pitch angle The solution steps are as follows: Receive the desired vertical velocity command. The output reference vertical velocity is smoothed by a first-order command filter. and desired vertical acceleration ; Desired vertical acceleration Multiply by the feedforward gain factor Obtain the feedforward pitch angle compensation amount ; Calculate the reference vertical velocity actual vertical ground speed error z , will the error Input proportional-integral controller, output feedback pitch angle compensation. ; Adjust the feedforward pitch angle compensation amount With feedback pitch angle compensation amount The superimposed values, after being processed by the upper and lower limiters to prevent fixed-wing stall, output the second desired pitch angle. ,like Figure 3 As shown.

[0035] The limiting range of the upper and lower limiters is determined based on the maximum and minimum safe angle of attack of the aircraft model, in order to prevent wing stall or negative lift caused by excessive or insufficient pitch angle in fixed-wing mode.

[0036] Step 4, preset the second speed conversion range [ , According to the calibrated airspeed CAS Real-time calculation of fixed-wing gain weights and multi-rotor gain weight (1- ); the first desired pitch angle With multi-rotor gain weight (1- The product of ) and the second desired pitch angle With fixed-wing gain weight Sum the products and output the final total pitch angle command. And send it to the attitude inner loop controller, such as Figure 4 As shown.

[0037] Fixed-wing gain weight β and final total pitch command The calculation formula is:

[0038]

[0039] when hour, =0, The pitch angle is entirely dominated by the multi-rotor speed control logic; when hour, =1, The pitch angle is completely handed over to the fixed-wing altitude maintenance logic, and the forward speed regulation is entirely undertaken by the propulsion motor, realizing the decoupling of the parallel power system.

[0040] In the above steps, the first speed transition range With the second speed transition range They are independent of each other and configured independently according to the aerodynamic characteristics of the aircraft; the first speed conversion range is used to eliminate the speed range where the pitot tube is interfered with by the rotor downwash; the second speed conversion range is used for the speed range where the wing lift gradually takes over the altitude control. The coverage of the two can partially overlap or be adjacent.

[0041] Example 2 This invention provides a method for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV), the method comprising the following steps: Step 1, acquire real-time aircraft status data: periodically read the forward ground speed in the horizontal coordinate system calculated by the aircraft's inertial navigation system (INS) through the flight control system. and vertical ground speed and the calibrated airspeed output from the pitot tube measurement. CAS .

[0042] Step 2, perform dynamic weighted fusion of forward velocities: preset the first velocity conversion interval [ , According to the current forward ground speed Real-time calculation of the first weighting factor ; with the first weighting factor The mixing coefficient will be used to calibrate the airspeed. CAS With forward ground speed Perform weighted summation and output the fused forward velocity. This is used by the outer loop speed controller.

[0043] Step 3, Decouple and Parallel Calculation of Dual-Mode Desired Pitch Angle: To achieve coordinated control of forward acceleration and altitude maintenance during the level flight transition phase, a decoupled architecture is adopted, executing the following two logical paths in parallel: (1) Multi-rotor mode speed control path (first desired pitch angle) The forward acceleration allocation module receives the longitudinal acceleration request command. And perform polarity determination. When the instruction is a deceleration requirement (i.e. When the negative acceleration requirement is <0), the negative acceleration requirement is mapped to the pitch command of the multi-rotor system according to a preset ratio, and the first desired pitch angle is output. When the instruction is for acceleration (i.e. When ≥0), the acceleration command is fully allocated to the forward propulsion system for execution. The output is zero to avoid rotor attitude interfering with forward propulsion efficiency.

[0044] (2) Fixed-wing mode altitude control path (second desired pitch angle) The control strategy based on "feedforward + feedback" calculates the pitch angle used to maintain flight altitude: First, it receives the desired vertical velocity command. The reference vertical velocity is extracted through smoothing with a first-order command filter. and desired vertical acceleration Subsequently, the feedforward branch will input the desired vertical acceleration. The attitude pre-compensation is obtained by multiplying by the feedforward gain coefficient; the feedback branch uses the reference vertical velocity. actual vertical ground speed The error is taken as input, and the feedback correction is calculated by a proportional-integral (PI) controller. Finally, the feedforward and feedback values ​​are superimposed, and the anti-stall limiter performs interception processing according to the aircraft's safe angle of attack range, outputting a second desired pitch angle. .

[0045] Step 4, execute dynamic gain mixing for pitch control: preset the second velocity transition range. According to the current calibrated airspeed CAS Real-time calculation of fixed-wing gain weights Simultaneously, the multi-rotor gain weights are calculated as follows: ; the first desired pitch angle With the second desired pitch angle Multiply by the corresponding gain weights and sum them to output the final total pitch command. The information is sent to the attitude inner loop controller.

[0046] Example 3 This embodiment takes a certain type of compound wing UAV as an example. The UAV is equipped with a quadcopter power system for vertical take-off and landing, and a tail thrust motor for forward propulsion. The pitot tube is installed at the leading edge of the nose, and the inertial navigation system (IMU+GPS fusion) provides ground speed information.

[0047] The method for velocity fusion and pitch control handover during the transition phase of this compound-wing UAV includes the following steps: Step 1: Obtain real-time status data of the aircraft.

[0048] The flight control system reads the forward ground speed in the horizontal coordinate system of the aircraft at fixed intervals via the inertial navigation system bus. and vertical ground speed Simultaneously, read the calibrated airspeed from the airspeed sensor interface. CAS The above three signals serve as the initial inputs for subsequent fusion and control calculations. In this embodiment, the fixed period is 0.01s. Forward ground velocity. Vertical ground speed and calibrated airspeed CAS The units are all m / s; the vertical ground speed is positive downwards.

[0049] Step 2, dynamic weighted fusion of forward velocity.

[0050] Based on the physical fact that the pitot tube is severely affected by the rotor downwash airflow at low speeds, a lower limit for the first speed transition range is set. =10m / s, upper limit =15m / s. The first weighting factor is calculated in real time using the following formula. :

[0051] After amplitude limiting, the first weighting factor is... The output is limited to the [0,1] interval. The final output is the fused forward velocity. The calculation formula is as follows:

[0052] Therefore, when At m / s , The system completely trusts ground speed (at which point the pitot tube is severely interfered with); when At m / s , The system fully trusts the calibrated airspeed. At this point, the airspeed tube is out of the rotor downwash influence range, and the aerodynamic speed is accurately reflected by the airspeed. Within the 10~15m / s transition range, the two speed signals are weighted according to the first weighting factor. The values ​​are fused using continuous linear interpolation, achieving a smooth transition without abrupt changes. Fusion forward speed... The output is sent to the energy management module and the outer loop speed controller as the speed state quantity of the flight control system.

[0053] Step 3: Decouple and calculate the desired pitch angle in both modes in parallel.

[0054] To achieve coordinated control of forward acceleration and altitude maintenance during the level flight transition phase, this method employs a decoupled control logic architecture, executing the following two logic paths in parallel through the flight control processor: (1) Rotor mode speed control path - first desired pitch angle The solution.

[0055] The forward acceleration distribution module receives longitudinal acceleration demand commands in real time. This module determines the polarity of the command and calculates the first desired pitch angle using the following formula. :

[0056] in, This is the preset deceleration-angle mapping gain coefficient. When When the aircraft is determined to be in a deceleration demand state, the negative acceleration demand is linearly mapped to the multirotor system's pitch-up attitude command. (in >0 indicates the nose is raised); when At that time, it is determined that the aircraft is in an acceleration demand state, and the multi-rotor control path outputs... At this point, forward acceleration is performed independently by the propulsion system, thereby avoiding unnecessary pitch components from the rotor that could interfere with forward propulsion efficiency.

[0057] (2) Fixed-wing mode altitude control path - second desired pitch angle Solution This approach calculates the pitch angle requirement for maintaining flight altitude based on a combined "feedforward + feedback" control strategy. The specific calculation process is as follows: Command smoothing and reference value extraction: The height holding control loop receives the target vertical velocity command. The instruction is input to the transfer function. The output reference vertical velocity is smoothed using a first-order command filter. and desired vertical acceleration .in, The filtering time constant is s For the Laplace operator, for The derivative with respect to time.

[0058] Feedforward compensation calculation: The attitude pre-compensation amount of the feedforward branch is calculated according to the following formula. :

[0059] in, This is the feedforward gain coefficient.

[0060] Feedback closed-loop calculation: The feedback branch uses the reference vertical velocity actual vertical ground speed Error ( The attitude feedback correction is calculated using a proportional-integral (PI) controller as input. :

[0061] in, This is the proportional gain coefficient. This is the integral gain coefficient.

[0062] Command synthesis and safety limiting: The feedforward compensation and feedback correction are superimposed, and after sign inversion, the result is input to the stall prevention limiter. The final output is the second desired pitch angle. The expression is:

[0063] Where sat is the saturation limiting function. and The angle boundary value (e.g., -10° to +15°) is preset according to the safe angle of attack range of the aircraft to ensure that the wing does not stall due to excessive pitch in fixed-wing mode.

[0064] Step 4: Dynamic gain mixing and handover of pitch control.

[0065] To ensure a smooth transition of pitch control from multi-rotor to fixed-wing mode as the aircraft's speed increases, a lower limit for the second speed transition range is set. =15m / s, upper limit =25m / s (This range should be within the speed range where the wing can provide effective lift; the specific value can be determined according to the aerodynamic data of the aircraft model).

[0066] Based on the current calibrated airspeed CAS Real-time calculation of fixed-wing gain weights :

[0067] After being limited in width, The multi-rotor gain weight is... The final output is the total pitch command. The calculation formula is as follows:

[0068] Closed-loop tracking is performed via the attitude inner loop controller to drive the multirotor tilting control surfaces or flight control servos to achieve attitude control. When entering the pure fixed-wing cruise phase at m / s , The control of the aircraft's pitch angle is completely transferred to the fixed-wing altitude holding loop, and the forward speed regulation is entirely undertaken by the tail thrust motor, thus achieving complete decoupling of the compound wing parallel power system.

[0069] By cyclically executing steps 1 to 4 in the flight control system at a frequency of 100Hz, the system responds in real time to changes in flight status, ensuring the continuity and safety of the entire level flight transition process.

[0070] Example 4 This embodiment provides a system for performing the aforementioned method of velocity fusion and pitch control handover during the transition phase of a compound-wing UAV. The system includes a state data acquisition module, an air-to-ground speed dynamic fusion module, a dual-mode pitch calculation module, and a pitch control gain mixing module; wherein: The state data acquisition module is used to acquire the forward ground speed and vertical ground speed in the horizontal coordinate system of the aircraft output by the aircraft's inertial navigation system, as well as the calibrated airspeed measured by the airspeed tube, thereby realizing step 1.

[0071] The air-to-ground speed dynamic fusion module is used to calculate the first weighting factor in real time based on the set first speed conversion range and forward ground speed, and to obtain the fused forward speed by weighted summation of the calibrated airspeed and forward ground speed, thereby realizing step 2.

[0072] The dual-mode pitch calculation module is used to independently and in parallel calculate the first desired pitch angle for controlling longitudinal acceleration and deceleration and the second desired pitch angle for controlling flight altitude, thereby achieving step 3.

[0073] The pitch control weight gain mixing module is used to calculate the gain weights of the fixed wing and multi-rotor in real time according to the set second speed conversion range and calibration airspeed, and to perform a weighted summation of the first desired pitch angle and the second desired pitch angle to output the final total pitch angle command, thereby realizing step 4.

[0074] Example 5 This embodiment provides a flight controller, which includes a memory, a processor, and a control program stored in the memory and capable of running on the processor. When the processor executes the control program, it is used to implement the method for speed fusion and pitch control handover during the transition phase of the compound wing UAV in the above embodiment.

[0075] Example 6 This embodiment provides a computer-readable storage medium storing a computer program that, when executed by a processor, is used to implement the method for speed fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle as described in the above embodiment.

[0076] Obviously, those skilled in the art can make various modifications and variations to the embodiments of the present invention without departing from the spirit and scope of the invention. Therefore, if these modifications and variations fall within the scope of the claims of the present invention and their equivalents, the present invention also intends to include these modifications and variations.

[0077] In summary, the above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV), characterized in that, Includes the following steps: Step 1: Obtain the forward ground speed in the horizontal coordinate system output by the aircraft's inertial navigation system. and vertical ground speed And the calibrated airspeed measured by the pitot tube. ; Step 2, preset the first speed conversion range [ , According to the forward ground speed Real-time calculation of the first weighting factor And will calibrate airspeed With forward ground speed According to the first weighting factor The forward fusion velocity is calculated by performing a weighted summation. ; Step 3: Independently and in parallel calculate the first desired pitch angle used to control longitudinal acceleration and deceleration. and the second desired pitch angle used to control flight altitude ; Step 4, preset the second speed conversion range [ , According to the calibrated airspeed Real-time calculation of fixed-wing gain weights and multi-rotor gain weight ; the first desired pitch angle With multi-rotor gain weight The product of, and the second desired pitch angle With fixed-wing gain weight Sum the products and output the final total pitch angle command. And then send it to the attitude inner loop controller.

2. The method as described in claim 1, characterized in that, In step 2, the first weighting factor With fusion forward velocity The calculation formula is: Among them, when ≤ hour, =0, = The system is completely trusted at high speed; when ≥ hour, = 1, = The system fully trusts the calibrated airspeed; within the range [ , ]Inside, Follow Linearly increasing, This is a linear interpolation fusion of the two speed paths.

3. The method as described in claim 2, characterized in that, In step 3, the first desired pitch angle The solution steps are as follows: Obtain the total forward longitudinal acceleration demand command output by the flight control system. ; Determine the sign of the instruction: when When the value is ≥0, the acceleration command is fully allocated to the propulsion system, and the multi-rotor pitch output... =0; when When the negative acceleration requirement is less than 0, the negative acceleration requirement is converted into a pitch angle command for the multi-rotor system according to a preset ratio, which is used as the first desired pitch angle. .

4. The method as described in claim 3, characterized in that, In step 3, the second desired pitch angle The solution steps are as follows: Receive desired vertical velocity command The output reference vertical velocity is smoothed by a first-order command filter. and desired vertical acceleration ; Desired vertical acceleration Multiply by the feedforward gain factor Obtain the feedforward pitch angle compensation amount ; Calculate the reference vertical velocity actual vertical ground speed error , will the error Input proportional-integral controller, output feedback pitch angle compensation. ; feedforward pitch angle compensation With feedback pitch angle compensation amount The superimposed values, after being processed by the upper and lower limiters to prevent fixed-wing stall, output the second desired pitch angle. .

5. The method as described in claim 4, characterized in that, In step 4, the fixed-wing gain weights and final total pitch command The calculation formula is: when hour, =0, The pitch angle is entirely dominated by the multi-rotor speed control logic; when hour, =1, The pitch angle is completely handed over to the fixed-wing altitude maintenance logic, and the forward speed regulation is entirely undertaken by the propulsion motor, realizing the decoupling of the parallel power system.

6. The method as described in claim 5, characterized in that, The limiting range of the upper and lower limiters is determined based on the maximum and minimum safe angle of attack of the aircraft model, in order to prevent wing stall or negative lift caused by excessive or insufficient pitch angle in fixed-wing mode.

7. The method as described in claim 6, characterized in that, First speed transition zone [ , [and the second speed conversion range] , The two speed conversion zones are independent of each other and configured independently according to the aerodynamic characteristics of the aircraft model; the first speed conversion zone is used to eliminate the speed range of the pitot tube from the interference of the rotor downwash; the second speed conversion zone is used to gradually control the altitude of the aircraft by the lift of the wing. The coverage of the two zones may partially overlap or be adjacent.

8. A system for velocity fusion and pitch control handover during the transition phase of a compound-wing unmanned aerial vehicle (UAV) for performing the method described in any one of claims 1-7, characterized in that, It includes a status data acquisition module, an air-to-ground speed dynamic fusion module, a dual-mode pitch calculation module, and a pitch control weight gain mixing module; The state data acquisition module is used to acquire the forward ground speed and vertical ground speed in the horizontal coordinate system of the aircraft output by the aircraft's inertial navigation system, as well as the calibrated airspeed measured by the airspeed tube. The air-to-ground speed dynamic fusion module is used to calculate the first weighting factor in real time based on the set first speed conversion range and forward ground speed, and to obtain the fused forward speed by weighted summation of the calibrated airspeed and forward ground speed. The dual-mode pitch calculation module is used to independently and in parallel calculate the first desired pitch angle for controlling longitudinal acceleration and deceleration and the second desired pitch angle for controlling flight altitude. The pitch control weight gain mixing module is used to calculate the gain weights of the fixed wing and multi-rotor in real time according to the set second speed conversion range and calibration airspeed, and to perform a weighted summation of the first desired pitch angle and the second desired pitch angle to output the final total pitch angle command.

9. A flight controller, comprising a memory, a processor, and a control program stored in the memory and capable of running on the processor, characterized in that, When the processor executes the control program, it is used to implement the method as described in any one of claims 1-7.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by a processor, the computer program is used to implement the method as described in any one of claims 1-7.

Images