Tilt rotor vertical take-off and landing aircraft and vertical take-off and landing balance system

Abstract

The invention discloses a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system.The tilt-rotor vertical take-off and landing aircraft and the vertical take-off and landing balance system.The tilt-rotor vertical take-off and landing aircraft comprises an aircraft body, wings are symmetrically arranged on the two sides of the aircraft body, wingtip small wings are arranged at the tail ends of the wings, and the wingtip small wings are arranged on the aircraft body. A pair of tilting mechanisms symmetrically arranged relative to the fuselage is arranged at the front end of the fuselage, each tilting mechanism comprises a front propeller, the front propellers are connected with the fuselage through tilting motor structures, and a micro turbojet engine is arranged at the rear end of the fuselage and connected with the fuselage through a fixing frame. The vertical take-off and landing balance system comprises a hardware part and a software part. Smooth transition between a vertical take-off and landing mode and a horizontal flight mode is achieved through optimized mode switching control logic.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft, and more specifically to a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balancing system. Background Art

[0002] Traditional aircraft are generally divided into two categories: Multi-rotor aircraft: capable of vertical take-off and landing and hovering, suitable for short-distance missions and precise positioning, but with short endurance and low efficiency in horizontal flight. Fixed-wing aircraft: with longer flight time and higher flight efficiency, suitable for long-distance missions, but requiring a runway or launch system for take-off and landing, and unable to hover. In order to combine the advantages of these two types of aircraft, vertical take-off and landing fixed-wing aircraft came into being. It combines the vertical take-off and landing capabilities of multi-rotor aircraft and the cruising efficiency of fixed-wing aircraft, and can adapt to a wider range of application scenarios, such as military reconnaissance, logistics transportation, disaster relief, etc. However, when vertical take-off and landing fixed-wing aircraft switch between different flight modes (vertical take-off and landing and horizontal flight), the complexity of control is greatly increased, and the following technical problems need to be solved.

[0003] Independent PID controllers are configured for vertical takeoff and landing and horizontal flight respectively. When switching flight modes, the system will directly switch different controller parameters. Manual or semi-automatic switching: Some VTOL aircraft rely on the operator to manually trigger the flight mode switch and select the appropriate switching time based on external conditions. This method is mostly used in simpler mission scenarios. Tilt-rotor technology: By adjusting the angle of the rotor to achieve a seamless transition between vertical and horizontal flight, this method has high requirements on hardware and control algorithms. Due to power layout problems, the aircraft is subjected to a large pitch moment during the tilt transition process and is difficult to balance. The aircraft attitude is very prone to problems during the tilt process, resulting in poor aircraft reliability.

[0004] Existing tilt-rotor aircraft also have solutions for vertical take-off and landing using built-in ducted fans. However, due to the relatively small power of the ducted fan, it is difficult to support the vertical thrust of the aircraft during the actual vertical take-off and landing of the aircraft, and it is difficult to provide sufficient torque to balance the nose-up torque caused by the front tilt-rotor. Summary of the invention

[0005] The object of the present invention is to provide a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system to solve the problems raised in the above-mentioned background technology: independent PID controllers are configured for vertical take-off and landing and horizontal flight respectively. When switching the flight mode, the system will directly switch different controller parameters. Manual or semi-automatic switching: Some VTOL aircraft rely on the operator to manually trigger the flight mode switch and select the appropriate switching time according to external conditions. This method is mostly used in simpler mission scenarios. By setting a small turbojet engine at the tail of the aircraft, the shortcomings of the ducted fan with small thrust and insufficient power can be improved.

[0006] To achieve the above-mentioned purpose, the present invention provides a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, including an aircraft body, the aircraft body including a fuselage, a canopy is arranged above the fuselage, wings are symmetrically arranged on both sides of the fuselage, a pair of tilt mechanisms symmetrically arranged about the fuselage are arranged at the front end of the fuselage, the tilt mechanisms include a front propeller, the front propeller is connected to the fuselage through a tilt motor structure, a cockpit is arranged at the front end of the fuselage, a micro-turbojet engine is arranged at the rear end of the fuselage, and is connected to the fuselage through a small turbojet engine fixing frame, the micro-turbojet engine is provided with a speed measurement module, an engine air pressure sensor, a fuel flow meter and an engine temperature sensor, a propulsion motor is arranged at the rear end of the fuselage, and a tail fin is arranged at the tail end of the fuselage.

[0007] Preferably, a carbon fiber tube beam frame and a plurality of wing ribs are arranged inside the wing, the carbon fiber tube beam frame is arranged along the length direction of the wing, the wing ribs are arranged at intervals along the length direction of the wing, and a winglet is arranged at the wing tip.

[0008] Preferably, the tilt mechanism includes a worm gear arranged inside the box body, one end of the worm gear is connected to the output worm gear of the tilt motor, a worm gear structure is arranged in the middle section of the output worm gear, the output worm gear is meshed with the gear structure on the tilt gear rod shaft, a fixing box for fixing the tilt motor is arranged inside the box body, the other end of the output worm gear is rotatably connected to the box body through a bearing, both ends of the tilt gear rod respectively pass through the box body and are connected to the frame plates on both sides of the tilt frame, a tilt device fixing frame is arranged on the wing, and the front propeller is connected to the tilt device fixing frame through a tilt power structure; the tilt power structure includes a tilt frame, a servo motor is arranged on the mounting surface of the tilt frame, the output end of the servo motor is connected to the front propeller, the frame plates on both sides of the tilt frame are connected to the fixing box body through a tilt connection assembly, and the fixing box body is connected to the tilt device fixing frame.

[0009] Preferably, the interior of the fuselage is provided with a flight control device, a mission payload, a fuel tank and a battery; the flight control device is integrated with a displacement sensor, a phase sensor, an acceleration sensor, a gyroscope and a magnetometer; a pitot tube is provided at the front end of the fuselage; a camera and an optical flow sensor or a visual sensor are provided at the lower end of the fuselage; and a GPS module is provided on the wing.

[0010] Preferably, a vertical take-off and landing balancing system includes a hardware part and a software part, wherein the hardware part includes:

[0011] Gyroscope: used to measure the angular velocity of the aircraft and help determine attitude changes;

[0012] Acceleration sensor: measures the acceleration of the aircraft and is used to calculate the state of motion;

[0013] Magnetometer: provides heading information and assists in attitude estimation;

[0014] Pitot tube: measures the change of the horizontal forward speed of the aircraft and assists in the horizontal direction control;

[0015] GPS module: provides location and speed information, assists navigation and stabilization;

[0016] Propeller speed measurement module: measures propeller speed and assists in calculating propeller thrust;

[0017] Speed ​​measurement module: provides engine speed information and assists in calculating engine thrust;

[0018] Fuel flow meter: provides engine fuel information and assists in controlling engine speed;

[0019] Engine temperature sensor: provides the internal temperature of the engine and assists in calculating the engine thrust;

[0020] Engine air pressure sensor: provides air pressure at the front and rear ends of the engine, and calculates engine thrust based on the impulse difference;

[0021] Optical flow sensor or visual sensor: Provides auxiliary position and attitude information when GPS signal is weak or failed;

[0022] The software part includes an attitude estimation algorithm: fusing data from inertial sensors, magnetometers and GPS to estimate the attitude and velocity of the aircraft in real time;

[0023] PID control system: independent PID control for attitude and position, adjusting the actuator to maintain stability;

[0024] Mode switching control system: Switch different control parameters between vertical take-off and landing and horizontal flight modes to ensure the best control effect in different flight conditions.

[0025] Preferably, the PID control system includes a hierarchical control structure:

[0026] High-level control: responsible for the position and heading of the aircraft in three-dimensional space. This layer outputs the desired attitude angle;

[0027] Low-level control: According to the desired attitude provided by the high-level controller, the attitude of the aircraft is adjusted to achieve stable flight;

[0028] Position PID controller: independently manages the position information of the aircraft and generates the expected value required by the attitude controller;

[0029] Attitude PID controller: independently manages the attitude information of the aircraft and directly controls the actuators to adjust the attitude.

[0030] Preferably, the posture PID control system method is as follows:

[0031] Step 1: Target setting to obtain aircraft status

[0032] Input: The error between the current attitude data of the aircraft (provided by the attitude estimation algorithm) and the expected attitude angle;

[0033] Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode and select different control schemes;

[0034] Step 2: Control process

[0035] Error calculation: Calculate the difference between the current attitude and the expected attitude in real time based on the various parameters of the aircraft obtained in the previous step;

[0036] PID regulation:

[0037] Proportional (P): adjust the control strength according to the current error amount;

[0038] Integration (I): accumulates past errors and eliminates ongoing deviations;

[0039] Derivative (D): predicts the error change trend, provides damping effect and prevents overshoot;

[0040] Execution adjustment: the output signal of the PID controller is transmitted to the corresponding actuator to adjust the posture;

[0041] Step 3: Debugging and Optimization

[0042] Initial debugging: Start with proportional gain, increase gradually, and observe the system response;

[0043] Integral and differential adjustment: After the proportional regulation is stable, gradually adjust the integral and differential parameters to optimize the response speed and stability of the system;

[0044] Iterative optimization: Through actual flight tests, PID parameters are continuously adjusted to ensure stability under various flight conditions;

[0045] Step 4: Output

[0046] Output: According to the calculation results of the above steps, the control signal of the actuator (such as motor speed or rudder angle) is adjusted to reduce the attitude error.

[0047] Preferably, the method of using the position PID controller is as follows:

[0048] Step 1: Goal Setting

[0049] Input: The error between the current position and velocity data of the aircraft (provided by GPS and other sensors) and the desired position;

[0050] Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode, select different control schemes, and generate the desired attitude angle;

[0051] Output: passed to the attitude PID controller to achieve precise control of the position.

[0052] Step 2: Control process

[0053] Error calculation: Calculate the difference between the current position information and the expected position in real time;

[0054] PID regulation:

[0055] Proportional (P): adjust the moving direction and speed of the aircraft according to the position error;

[0056] Integration (I): accumulates position errors to ensure that the aircraft can eliminate continuous offsets;

[0057] Differential (D): Adjust the acceleration of the aircraft according to the rate of change of the position error to prevent position overshoot;

[0058] Attitude setting: convert the output of the position PID controller into the desired attitude angle (such as tilt angle) to guide the aircraft to move to the target position;

[0059] Step 3: Debugging and Optimization

[0060] Hierarchical debugging: First, ensure that the position controller can stably generate the desired posture, and then perform overall debugging in conjunction with the posture controller;

[0061] Parameter adjustment: dynamically adjust PID parameters according to the performance of the aircraft in different flight stages (such as take-off, cruising, and landing);

[0062] Environmental adaptation: Debugging under different environmental conditions (such as wind speed changes) to ensure the robustness of the position controller;

[0063] Step 4: Output

[0064] Output: Based on the correction values ​​generated by the above process, the above corrections are passed to different accessories of the aircraft according to different flight modes to achieve precise control of the position.

[0065] Preferably, the mode switching control system includes a hierarchical control system:

[0066] High-level control: The high-level control system is responsible for managing the aircraft's flight status judgment and mode switching logic. The main responsibilities are: judging whether the flight mode needs to be switched based on the aircraft's current status (such as speed, attitude, acceleration, etc.), issuing switching instructions at the appropriate time, and deciding when to start and end mode switching;

[0067] Low-level control: The low-level control system is responsible for specific attitude and position control. Different flight states will use different control strategies and parameters, dividing the aircraft into 3 modes:

[0068] State 1: Vertical take-off and landing mode

[0069] The rotor combined with the turbojet engine control logic is adopted, and the attitude and position of the aircraft are adjusted mainly by the speed difference of the front propeller and the thrust state of the small turbojet engine. At this time, the vertical force formula of the aircraft is as follows:

[0070] F 升 =ρπr 2 (n*P) 2 η+q ma (C 9 -C 0 )+A 9 (p 9 -p 0 )

[0071] Where η is the propeller efficiency, π is the circumference of a circle, P is the propeller pitch, ρ is the atmospheric density, r is the propeller radius, n is the propeller speed per second, and q ma A is the mass of air entering the engine per unit time, 9 is the cross-sectional area of ​​the engine tail nozzle outlet, p 0 is the external atmospheric pressure, p 9 It is the cross-sectional pressure of the engine tail nozzle.

[0072] State 2: Horizontal flight mode

[0073] The fixed-wing flight control logic is adopted, and the attitude is mainly controlled by adjusting the ailerons and tail wing, and the forward movement is achieved by relying on the thrust of the propeller. At this time, the lift formula of the aircraft is as follows:

[0074]

[0075] Where ρ is the atmospheric density, S is the wing area, V X For C y is the lift coefficient.

[0076] State 3: Tilt transition mode

[0077] When switching from vertical mode to horizontal mode, the lift size and lift source are analyzed according to the flight parameters obtained by the sensor. A weighted algorithm is used to perform weighted calculation according to the proportion of the lift obtained by the wing surface and the thrust of the front propeller and the small turbojet engine in the lift source. Then, the control of the two modes is weightedly superimposed to control the flight attitude of the aircraft. The control correction superposition formula is: A 0 =F 1 *A 1 / F Y +F 2 *A 2 / F Y , where A 0 is the total correction of the aircraft, F 1 It is the sum of the vertical thrust generated by the front propeller of the aircraft and the lift generated by the small turbojet engine. 2 is the aerodynamic lift generated by the aircraft wing, F Y A is the combined force of the vertical lift and engine thrust of the aircraft. 1 is the vertical mode correction of the aircraft, A 2 It is the correction value of the aircraft horizontal mode.

[0078] Preferably, the mode switching control system includes mode switching, and the specific steps are as follows:

[0079] Step 1: Mode switching condition judgment:

[0080] Switch from vertical mode to horizontal mode:

[0081] When the aircraft reaches a certain altitude and speed and needs to switch to forward flight, the system will determine whether the aircraft is in a suitable state for switching;

[0082] The conditions usually include: sufficient flight altitude, small pitch angle, and air speed within the range suitable for fixed-wing flight;

[0083] Switch from horizontal mode to vertical mode:

[0084] When the aircraft needs to slow down and enter a hovering state, such as preparing to land or hovering in a complex environment, the system will determine the timing of the switch;

[0085] The conditions usually include: the flight speed is reduced to a state where it is difficult for fixed-wing flight to maintain lift, and the flight altitude is moderate;

[0086] Step 2: Mode transition phase:

[0087] The core of mode switching is to gradually adjust the control parameters during the transition phase to ensure flight stability. The specific steps are as follows:

[0088] (1) Gradual adjustment of control gain

[0089] During the mode switching process, the controller will gradually adjust the gain of the PID controller or switch the control strategy;

[0090] When switching from vertical mode to horizontal mode: start to gradually reduce the control force in vertical mode and increase the control force in fixed-wing mode. X The source and proportion of the aircraft's lift are calculated based on the aircraft's front propeller rotation speed, tilt angle, and the speed and pressure difference of the small turbojet engine, and weighted control is performed accordingly, gradually reducing the speed difference control between the motors in the rotor mode and increasing the control of the rudder surface in the horizontal flight mode;

[0091] When switching from horizontal mode to vertical mode: reduce the control force of the rudder surface and gradually increase the control of the motor speed so that the aircraft can maintain a hovering state;

[0092] (2) Power distribution adjustment

[0093] Tilt-rotor: If the aircraft uses a tilt-rotor design, when switching modes, the inclination angle of the rotor needs to be adjusted to convert it from vertical lift to horizontal thrust, or from horizontal thrust to vertical lift.

[0094] Thruster start: When transitioning from vertical take-off and landing mode to horizontal mode, the system will start the thrusters to provide horizontal thrust while gradually reducing the motor output in the vertical direction;

[0095] Turbojet engine gradual shutdown: After switching to horizontal flight mode, the thrust output of the turbojet engine is gradually reduced, and the flight is maintained by relying on the lift of the fixed wing;

[0096] (3) Posture control adjustment

[0097] Attitude controller switching: In vertical mode, attitude control mainly controls roll, pitch and yaw by adjusting the speed of the motor. In horizontal mode, attitude control is transferred to control by the control of the rudder surface.

[0098] Smooth transition: When the control system switches modes, the attitude change must be smooth to prevent the aircraft from experiencing large attitude deviations during the switching process, which may lead to instability.

[0099] Step 3: Stabilization after transition

[0100] Control strategy lock: Once the mode switch is completed, the system will lock the new control strategy to ensure that the aircraft operates according to the corresponding control logic in the new mode;

[0101] Controller calibration: After the switch is completed, the system will calibrate the current attitude and position to ensure that the flight status after the switch is in line with expectations.

[0102] Compared with the prior art, the advantages of the present invention are:

[0103] (1) Smooth flight mode switching

[0104] This solution achieves a smooth transition between vertical take-off and landing mode and horizontal flight mode through optimized mode switching control logic. Compared with existing technologies, the switching process reduces attitude fluctuations and speed changes, thereby improving flight stability and safety. Specific effects include:

[0105] Seamless switching process: By gradually adjusting the control parameters and power distribution, the aircraft will not experience sudden changes in attitude or speed during switching, avoiding unstable flight.

[0106] Reduce shock and vibration: When switching, the system uses a smooth transition control algorithm to avoid attitude instability or body vibration during the transition process, thereby improving the overall flight stability.

[0107] (2) Improve flight stability and control accuracy

[0108] The control system ensures precise flight control in different flight modes through efficient attitude and position control:

[0109] Enhanced hovering stability: In vertical takeoff and landing mode, the improved PID control algorithm makes the aircraft hover more precise, and can maintain stability through faster response and adjustment even in the case of large changes in wind speed.

[0110] Improved cruise efficiency: In horizontal flight mode, the system optimizes the fixed-wing flight controller, which can better control the flight attitude and heading, making cruise flight more efficient and stable, and extending the flight time.

[0111] (3) Adapting to complex environments and mission requirements

[0112] Through adaptive control strategies, the system can automatically adjust flight modes and control parameters according to changes in the external environment (such as wind speed and airflow), enhancing the performance of the aircraft in complex environments:

[0113] Improved wind resistance: The mode switching system can intelligently adjust the attitude control gain in strong wind environments to ensure flight stability, especially during takeoff and landing, and is not easily affected by wind.

[0114] Responding to emergencies: The aircraft can flexibly switch flight modes according to mission requirements, such as quickly switching to hover mode for reconnaissance in emergency missions, or maintaining an efficient cruise mode for longer missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0115] Figure 1 It is a schematic diagram of the overall system of the present invention and a schematic diagram of the overall structure;

[0116] Figure 2 A top view of the internal structure of the aircraft of the present invention;

[0117] Figure 3 is a side cross-sectional view of the present invention;

[0118] Figure 4 It is a schematic diagram of the tilting mechanism of the present invention;

[0119] Figure 5 It is a schematic diagram of the overall system of a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balancing system of the present invention;

[0120] Figure 6 It is a schematic diagram of the aircraft mode of the present invention.

[0121] Reference numerals

[0122] 1. Aircraft body 11. Fuselage; 12. Canopy; 13. Flight control equipment; 14. Battery; 15. Mission payload; 16. Fuel tank; 17. Pitot tube; 18. Camera and optical flow sensor or visual sensor; 2. Wing; 21. Carbon fiber tube beam frame; 22. Aileron; 23. Wing rib; 24. Winglet; 25. GPS module; 26 Tilt device fixing frame; 3. Tilt mechanism; 31. Tilt motor; 32. Output worm; 33. Tilt gear rod; 34. Bearing; 35. Fairing; 36. Fixing box; 37. Tilt frame; 38. Front propeller; 39. Servomotor; 4. Propulsion motor; 5. Small turbojet engine; 51 Small turbojet engine fixing frame; 6. Tail. DETAILED DESCRIPTION

[0123] In order to make the purpose, technical scheme and advantages disclosed in the embodiments of the present invention clearer, the embodiments of the present invention are further described in detail in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the embodiments of the present invention and are not intended to limit the embodiments of the present invention. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. Examples of the embodiments are shown in the accompanying drawings, where the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions.

[0124] It should be noted that the terms "including" and "having" and any variations thereof are intended to cover non-exclusive inclusions. For example, a process, method, system, product or server that includes a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units that are not explicitly listed or inherent to these processes, methods, products or devices.

[0125] Like reference numerals and letters denote similar items in the following drawings, and thus, once an item is defined in one drawing, further definition and explanation thereof is not required in subsequent drawings.

[0126] In the description of the present invention, it should be noted that the terms "upper", "lower", "inside", "outside", etc. indicate directions or positional relationships based on the directions or positional relationships shown in the accompanying drawings, or are directions or positional relationships in which the product of the invention is usually placed when in use. They are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific direction, be constructed and operated in a specific direction, and therefore should not be understood as a limitation on the present invention.

[0127] In the description of the present invention, it is also necessary to explain that, unless otherwise clearly specified and limited, the terms "setting", "installation" and "connection" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be an indirect connection through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0128] Example

[0129] like Figure 1 , 2As shown, a tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system are characterized in that: it includes an aircraft body 1, the aircraft body 1 includes a fuselage 11, a cabin cover 12 is arranged above the fuselage 1, wings 2 are symmetrically arranged on both sides of the fuselage 11, a pair of tilt mechanisms 3 symmetrically arranged about the fuselage 11 are arranged on the front end of the fuselage 1, the tilt mechanism 3 includes a front propeller 38, and the front propeller 38 is connected to the fuselage 11 through a tilt motor structure, a cockpit is arranged at the front end of the fuselage 11, a micro-turbojet engine 5 is arranged at the rear end of the fuselage 11, and a speed measurement module, an engine air pressure sensor, a fuel flow meter and an engine temperature sensor are arranged in the micro-turbojet engine 5 connected to the fuselage 11 through a small turbojet engine fixing frame 51, a propulsion motor 4 is arranged at the rear end of the fuselage, and a tail 6 is arranged at the tail end of the fuselage 11.

[0130] like Figure 2 As shown, a carbon fiber tube beam frame 21 and a plurality of wing ribs 23 are arranged inside the wing 2, the carbon fiber tube beam frame 21 is arranged along the length direction of the wing 2, the wing ribs 23 are arranged at intervals along the length direction of the wing 2, and a winglet 24 is arranged at the wing tip of the wing 2.

[0131] like Figure 4 As shown, the tilt mechanism 3 includes a worm 32 arranged inside the box body, one end of the worm 32 is connected to the output worm 32 of the tilt motor 31, a worm gear structure is arranged in the middle section of the output worm 32, the output worm 32 is meshed with the gear structure on the axis of the tilt gear rod 33, a fixing box 36 for fixing the tilt motor 31 is arranged in the box body, the other end of the output worm 32 is rotatably connected to the box body through a bearing, the two ends of the tilt gear rod 33 respectively pass through the box body and are connected to the frame plates on both sides of the tilt frame 37, a tilt device fixing frame 26 is arranged on the wing 2, and the front propeller 38 is connected to the tilt device fixing frame 26 through a tilt power structure; the tilt power structure includes a tilt frame, a servo motor 39 is arranged on the mounting surface of the tilt frame 37, the output end of the servo motor 39 is connected to the front propeller 38, the frame plates on both sides of the tilt frame 37 are connected to the fixing box 36 box body through a tilt connection assembly, and the fixing box 36 box body is connected to the tilt device fixing frame 26.

[0132] like Figure 2 , 3 As shown, the interior of the fuselage 11 is provided with a flight control device 13, a mission payload 15, a fuel tank 16 and a battery 14. The flight control device 13 is integrated with a displacement sensor, a phase sensor, an acceleration sensor, a gyroscope, and a magnetometer. A pitot tube 17 is provided at the front end of the fuselage 11. A camera and an optical flow sensor or a visual sensor 18 are provided at the lower end of the fuselage 11. A GPS module 25 is provided on the wing 2.

[0133] like Figure 5 , 6 As shown, a vertical take-off and landing balance system includes a hardware part and a software part, wherein the hardware part includes:

[0134] Gyroscope: used to measure the angular velocity of the aircraft and help determine attitude changes;

[0135] Acceleration sensor: measures the acceleration of the aircraft and is used to calculate the state of motion;

[0136] Magnetometer: provides heading information and assists in attitude estimation;

[0137] Pitot tube: measures the change of the horizontal forward speed of the aircraft and assists in the horizontal direction control;

[0138] GPS module 25: provides position and speed information, assists navigation and stabilization;

[0139] Propeller speed measurement module: measures propeller speed and assists in calculating propeller thrust;

[0140] Speed ​​measurement module: provides engine speed information and assists in calculating engine thrust;

[0141] Fuel flow meter: provides engine fuel information and assists in controlling engine speed;

[0142] Engine temperature sensor: provides internal engine temperature and assists in calculating engine thrust;

[0143] Engine air pressure sensor: provides air pressure at the front and rear ends of the engine, and calculates engine thrust based on the impulse difference;

[0144] Optical flow sensor or visual sensor: Provides auxiliary position and attitude information when GPS signal is weak or failed;

[0145] The software part includes an attitude estimation algorithm: fusing data from inertial sensors, magnetometers and GPS to estimate the attitude and velocity of the aircraft in real time;

[0146] PID control system: independent PID control for attitude and position, adjusting the actuator to maintain stability;

[0147] Mode switching control system: Switch different control parameters between vertical take-off and landing and horizontal flight modes to ensure the best control effect in different flight conditions.

[0148] like Figure 5 , 6 As shown, the attitude PID control system method is as follows:

[0149] Step 1: Target setting to obtain aircraft status

[0150] Input: The error between the current attitude data of the aircraft (provided by the attitude estimation algorithm) and the expected attitude angle;

[0151] Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode and select different control schemes;

[0152] Step 2: Control process

[0153] Error calculation: Calculate the difference between the current attitude and the expected attitude in real time based on the various parameters of the aircraft obtained in the previous step;

[0154] PID regulation:

[0155] Proportional (P): adjust the control strength according to the current error amount;

[0156] Integration (I): accumulates past errors and eliminates ongoing deviations;

[0157] Derivative (D): predicts the error change trend, provides damping effect and prevents overshoot;

[0158] Execution adjustment: the output signal of the PID controller is transmitted to the corresponding actuator to adjust the posture;

[0159] Step 3: Debugging and Optimization

[0160] Initial debugging: Start with proportional gain, increase gradually, and observe the system response;

[0161] Integral and differential adjustment: After the proportional regulation is stable, gradually adjust the integral and differential parameters to optimize the response speed and stability of the system;

[0162] Iterative optimization: Through actual flight tests, PID parameters are continuously adjusted to ensure stability under various flight conditions;

[0163] Step 4: Output

[0164] Output: According to the calculation results of the above steps, the control signal of the actuator (such as motor speed or rudder angle) is adjusted to reduce the attitude error.

[0165] like Figure 5 , 6 As shown, the method of using the position PID controller is as follows:

[0166] Step 1: Goal Setting

[0167] Input: The error between the current position and velocity data of the aircraft (provided by GPS and other sensors) and the desired position;

[0168] Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode, select different control schemes, and generate the desired attitude angle;

[0169] Output: passed to the attitude PID controller to achieve precise control of the position.

[0170] Step 2: Control process

[0171] Error calculation: Calculate the difference between the current position information and the expected position in real time;

[0172] PID regulation:

[0173] Proportional (P): adjust the moving direction and speed of the aircraft according to the position error;

[0174] Integration (I): accumulates position errors to ensure that the aircraft can eliminate continuous offsets;

[0175] Differential (D): Adjust the acceleration of the aircraft according to the rate of change of the position error to prevent position overshoot;

[0176] Attitude setting: convert the output of the position PID controller into the desired attitude angle (such as tilt angle) to guide the aircraft to move to the target position;

[0177] Step 3: Debugging and Optimization

[0178] Hierarchical debugging: First, ensure that the position controller can stably generate the desired posture, and then perform overall debugging in conjunction with the posture controller;

[0179] Parameter adjustment: dynamically adjust PID parameters according to the performance of the aircraft in different flight stages (such as take-off, cruising, and landing);

[0180] Environmental adaptation: Debugging under different environmental conditions (such as wind speed changes) to ensure the robustness of the position controller;

[0181] Step 4: Output

[0182] Output: Based on the correction values ​​generated by the above process, the above corrections are passed to different accessories of the aircraft according to different flight modes to achieve precise control of the position.

[0183] like Figure 5 , 6 As shown, the mode switching control system includes the hierarchical control system:

[0184] High-level control: The high-level control system is responsible for managing the aircraft's flight status judgment and mode switching logic. The main responsibilities are: judging whether the flight mode needs to be switched based on the aircraft's current status (such as speed, attitude, acceleration, etc.), issuing switching instructions at the appropriate time, and deciding when to start and end mode switching;

[0185] Low-level control: The low-level control system is responsible for specific attitude and position control. Different flight states will use different control strategies and parameters, dividing the aircraft into 3 modes:

[0186] State 1: Vertical take-off and landing mode

[0187] The rotor combined with the turbojet engine control logic is adopted, and the attitude and position of the aircraft are adjusted mainly by the speed difference of the front propeller 38 and the thrust state of the small turbojet engine 5. At this time, the vertical force formula of the aircraft is as follows:

[0188] F 升 =ρπr 2 (n*P) 2 η+q ma (C 9 -C 0 )+A 9 (p 9 -p 0 )

[0189] Where η is the propeller efficiency, π is the circumference of a circle, P is the propeller pitch, ρ is the atmospheric density, r is the propeller radius, n is the propeller speed per second, and q ma A is the mass of air entering the engine per unit time, 9 is the cross-sectional area of ​​the engine tail nozzle outlet, p 0 is the external atmospheric pressure, p 9 It is the cross-sectional pressure of the engine tail nozzle.

[0190] State 2: Horizontal flight mode

[0191] The fixed-wing flight control logic is adopted, and the attitude is mainly controlled by adjusting the ailerons and tail wing, and the forward movement is achieved by relying on the thrust of the propeller. At this time, the lift formula of the aircraft is as follows:

[0192]

[0193] Where ρ is the atmospheric density, S is the wing area, V X For C y is the lift coefficient.

[0194] State 3: Tilt transition mode

[0195] When switching from the vertical mode to the horizontal mode, the lift size and lift source are analyzed according to the flight parameters obtained by the sensor, and a weighted algorithm is used to perform weighted calculation according to the proportion of the lift obtained by the wing surface and the thrust of the front propeller 38 and the small turbojet engine 5 in the lift source, and then the control of the two modes is weighted superimposed to control the flight attitude of the aircraft. The control correction superposition formula is: A 0 =F 1 *A 1 / F Y +F 2 *A 2 / F Y , where A 0 is the total correction of the aircraft, F 1 It is the sum of the vertical thrust generated by the front propeller 38 of the aircraft and the lift generated by the small turbojet engine 5, F 2 is the aerodynamic lift generated by the aircraft wing 3, F Y A is the combined force of the vertical lift and engine thrust of the aircraft. 1 is the vertical mode correction of the aircraft, A 2 It is the correction value of the aircraft horizontal mode.

[0196] like Figure 5 , 6 As shown, the mode switching control system includes mode switching, and the specific steps are as follows:

[0197] Step 1: Mode switching condition judgment:

[0198] Switch from vertical mode to horizontal mode:

[0199] When the aircraft reaches a certain altitude and speed and needs to switch to forward flight, the system will determine whether the aircraft is in a suitable state for switching;

[0200] The conditions usually include: sufficient flight altitude, small pitch angle, and air speed within the range suitable for fixed-wing flight;

[0201] Switch from horizontal mode to vertical mode:

[0202] When the aircraft needs to slow down and enter a hovering state, such as preparing to land or hovering in a complex environment, the system will determine the timing of the switch;

[0203] The conditions usually include: the flight speed is reduced to a state where it is difficult for fixed-wing flight to maintain lift, and the flight altitude is moderate;

[0204] Step 2: Mode transition phase:

[0205] The core of mode switching is to gradually adjust the control parameters during the transition phase to ensure flight stability. The specific steps are as follows:

[0206] (1) Gradual adjustment of control gain

[0207] During the mode switching process, the controller will gradually adjust the gain of the PID controller or switch the control strategy;

[0208] When switching from vertical mode to horizontal mode: start to gradually reduce the control force in vertical mode and increase the control force in fixed-wing mode. X The source and proportion of the aircraft's lift are calculated based on the rotation speed and tilt angle of the aircraft's front propeller 38 and the rotation speed and pressure difference of the small turbojet engine 5, and weighted control is performed accordingly, gradually reducing the speed difference control between the motors in the rotor mode and increasing the control of the rudder surface in the horizontal flight mode;

[0209] When switching from horizontal mode to vertical mode: reduce the control force of the rudder surface and gradually increase the control of the motor speed so that the aircraft can maintain a hovering state;

[0210] (2) Power distribution adjustment

[0211] Tilt-rotor: If the aircraft uses a tilt-rotor design, when switching modes, the inclination angle of the rotor needs to be adjusted to convert it from vertical lift to horizontal thrust, or from horizontal thrust to vertical lift.

[0212] Thruster start: When transitioning from vertical take-off and landing mode to horizontal mode, the system will start the thrusters to provide horizontal thrust while gradually reducing the motor output in the vertical direction;

[0213] Turbojet engine gradual shutdown: After switching to horizontal flight mode, the thrust output of the turbojet engine is gradually reduced, and the flight is maintained by relying on the lift of the fixed wing;

[0214] (3) Posture control adjustment

[0215] Attitude controller switching: In vertical mode, attitude control mainly controls roll, pitch and yaw by adjusting the speed of the motor. In horizontal mode, attitude control is transferred to control by the control of the rudder surface.

[0216] Smooth transition: When the control system switches modes, the attitude change must be smooth to prevent the aircraft from experiencing large attitude deviations during the switching process, which may lead to instability.

[0217] Step 3: Stabilization after transition

[0218] Control strategy lock: Once the mode switch is completed, the system will lock the new control strategy to ensure that the aircraft operates according to the corresponding control logic in the new mode;

[0219] Controller calibration: After the switch is completed, the system will calibrate the current attitude and position to ensure that the flight status after the switch is in line with expectations.

[0220] When in use, when the aircraft needs to take off and land vertically, the aircraft is first in a vertical mode ready to take off. First, the two front propellers 38 at the front rotate, and the small turbojet engine 5 starts to start, so that the aircraft can generate enough thrust in the vertical direction to take off. When the aircraft reaches a certain height, the tail propulsion motor 4 starts to gradually increase the horizontal speed of the aircraft, and then the front propeller 38 starts to tilt slowly, and the small turbojet engine 5 gradually reduces the speed, and finally the front propeller 38 completely tilts to the horizontal mode. During the tilting process, the aircraft mainly relies on the speed difference between the left and right front propellers 38, the different tilting angles of the left and right propellers 38, the small turbojet engine 5, the ailerons 22 on the wings 3, and the tail wing to control the aircraft. When the aircraft needs to take off and land on a taxi, the aircraft is in a horizontal mode, and the front propeller 38 and the propulsion motor 4 provide horizontal thrust for the aircraft. When the aircraft taxis to a certain speed, since the wing 3 has sufficient lift, even if the aircraft completes the taxiing takeoff, the aircraft can start the small turbojet engine 5 to obtain sufficient lift for takeoff in advance, which reduces the taxiing distance of the aircraft compared to a completely horizontal takeoff. When the aircraft lands vertically, the speed of the propulsion motor 4 is gradually reduced, the speed of the small turbojet engine 5 is gradually increased, and the front propeller 38 is gradually tilted, so that the aircraft first reaches a hovering state, and then the vertical mode control is gradually adopted to complete the landing of the aircraft.

[0221] The basic principles, main features and advantages of the present invention are shown and described above; those skilled in the art should understand that the present invention is not limited to the above embodiments, and the above embodiments and descriptions are only preferred examples of the present invention and are not used to limit the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, and these changes and improvements all fall within the scope of the present invention to be protected; the scope of protection claimed by the present invention is determined by the attached claims and their effects.

Claims

1. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, characterized in that: The invention comprises an aircraft body (1), wherein the aircraft body (1) comprises a fuselage (11), a cabin cover (12) is arranged above the fuselage (1), wings (2) are symmetrically arranged on both sides of the fuselage (11), a pair of tilting mechanisms (3) symmetrically arranged about the fuselage (11) are arranged at the front end of the fuselage (1), the tilting mechanism (3) comprises a front propeller (38), the front propeller (38) is connected to the fuselage (11) through a tilting motor structure, a cockpit is arranged at the front end of the fuselage (11), a micro-turbojet engine (5) is arranged at the rear end of the fuselage (11), and the micro-turbojet engine (5) is connected to the fuselage (11) through a small turbojet engine fixing frame (51), a speed measurement module, an engine air pressure sensor, a fuel flow meter and an engine temperature sensor are arranged in the micro-turbojet engine (5), a propulsion motor (4) is arranged at the rear end of the fuselage, and a tail wing (6) is arranged at the tail end of the fuselage (11).

2. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, characterized in that: A carbon fiber tube beam frame (21) and a plurality of wing ribs (23) are arranged inside the wing (2); the carbon fiber tube beam frame (21) is arranged along the length direction of the wing (2); the wing ribs (23) are arranged at intervals along the length direction of the wing (2); and a winglet (24) is arranged at the wing tip of the wing (2).

3. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, characterized in that: The tilt mechanism (3) comprises a worm (32) arranged inside a box body, one end of the worm (32) is connected to an output worm (32) of a tilt motor (31), a worm gear structure is arranged in the middle section of the output worm (32), the output worm (32) is meshed with a gear structure on the axis of a tilt gear rod (33), a fixing box (36) for fixing the tilt motor (31) is arranged inside the box body, the other end of the output worm (32) is rotatably connected to the box body through a bearing, and two ends of the tilt gear rod (33) respectively pass through the box body and the tilt frame (31). The wing (2) is provided with a tilt device fixing frame (26), and the front propeller (38) is connected to the tilt device fixing frame (26) through a tilt power structure; the tilt power structure includes a tilt frame, a servo motor (39) is provided on the mounting surface of the tilt frame (37), the output end of the servo motor (39) is connected to the front propeller (38), the frame plates on both sides of the tilt frame (37) are connected to the box body of the fixing box (36) through a tilt connection assembly, and the box body of the fixing box (36) is connected to the tilt device fixing frame (26).

4. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, characterized in that: A flight control device (13), a mission payload (15), a fuel tank (16) and a battery (14) are arranged inside the fuselage (11); a displacement sensor, a phase sensor, an acceleration sensor, a gyroscope and a magnetometer are integrated in the flight control device (13); a pitot tube (17) is arranged at the front end of the fuselage (11); a camera and an optical flow sensor or a visual sensor (18) are arranged at the lower end of the fuselage (11); and a GPS module (25) is arranged on the wing (2).

5. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system, characterized in that: A vertical take-off and landing balancing system includes a hardware part and a software part, wherein the hardware part includes: Gyroscope: used to measure the angular velocity of the aircraft and help determine attitude changes; Acceleration sensor: measures the acceleration of the aircraft and is used to calculate the state of motion; Magnetometer: provides heading information and assists in attitude estimation; Pitot tube: measures the change of the horizontal forward speed of the aircraft and assists in the horizontal direction control; GPS module (25): provides position and speed information, assists navigation and stabilization; Propeller speed measurement module: measures propeller speed and assists in calculating propeller thrust; Speed ​​measurement module: provides engine speed information and assists in calculating engine thrust; Fuel flow meter: provides engine fuel information and assists in controlling engine speed; Engine temperature sensor: provides internal engine temperature and assists in calculating engine thrust; Engine air pressure sensor: provides air pressure at the front and rear ends of the engine, and calculates engine thrust based on the impulse difference; Optical flow sensor or visual sensor: Provides auxiliary position and attitude information when GPS signal is weak or failed; The software part includes an attitude estimation algorithm: fusing data from inertial sensors, magnetometers and GPS to estimate the attitude and velocity of the aircraft in real time; PID control system: independent PID control for attitude and position, adjusting the actuator to maintain stability; Mode switching control system: Switch different control parameters between vertical take-off and landing and horizontal flight modes to ensure the best control effect in different flight conditions.

6. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system according to claim 5, characterized in that: The PID control system includes a hierarchical control structure: High-level control: responsible for the position and heading of the aircraft in three-dimensional space. This layer outputs the desired attitude angle; Low-level control: According to the desired attitude provided by the high-level controller, the attitude of the aircraft is adjusted to achieve stable flight; Position PID controller: independently manages the position information of the aircraft and generates the expected value required by the attitude controller; Attitude PID controller: independently manages the attitude information of the aircraft and directly controls the actuators to adjust the attitude.

7. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balancing system according to claim 6, characterized in that: The attitude PID control system method is as follows: Step 1: Target setting to obtain aircraft status Input: The error between the current attitude data of the aircraft (provided by the attitude estimation algorithm) and the expected attitude angle; Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode and select different control schemes; Step 2: Control process Error calculation: Calculate the difference between the current attitude and the expected attitude in real time based on the various parameters of the aircraft obtained in the previous step; PID regulation: Proportional (P): adjust the control strength according to the current error amount; Integration (I): accumulates past errors and eliminates ongoing deviations; Derivative (D): predicts the error change trend, provides damping effect and prevents overshoot; Execution adjustment: the output signal of the PID controller is transmitted to the corresponding actuator to adjust the posture; Step 3: Debugging and Optimization Initial debugging: Start with proportional gain, increase gradually, and observe the system response; Integral and differential adjustment: After the proportional regulation is stable, gradually adjust the integral and differential parameters to optimize the response speed and stability of the system; Iterative optimization: Through actual flight tests, PID parameters are continuously adjusted to ensure stability under various flight conditions; Step 4: Output Output: According to the calculation results of the above steps, the control signal of the actuator (such as motor speed or rudder angle) is adjusted to reduce the attitude error.

8. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balance system according to claim 6, characterized in that: The method of using the position PID controller is as follows: Step 1: Goal Setting Input: The error between the current position and velocity data of the aircraft (provided by GPS and other sensors) and the desired position; Calculation and judgment: Based on the data from each sensor, calculations are performed according to the prescribed empirical formula to obtain the actual flight parameters of the aircraft, determine the aircraft mode, select different control schemes, and generate the desired attitude angle; Output: passed to the attitude PID controller to achieve precise control of the position. Step 2: Control process Error calculation: Calculate the difference between the current position information and the expected position in real time; PID regulation: Proportional (P): adjust the moving direction and speed of the aircraft according to the position error; Integration (I): accumulates position errors to ensure that the aircraft can eliminate continuous offsets; Differential (D): Adjust the acceleration of the aircraft according to the rate of change of the position error to prevent position overshoot; Attitude setting: convert the output of the position PID controller into the desired attitude angle (such as tilt angle) to guide the aircraft to move to the target position; Step 3: Debugging and Optimization Hierarchical debugging: First, ensure that the position controller can stably generate the desired posture, and then perform overall debugging in conjunction with the posture controller; Parameter adjustment: dynamically adjust PID parameters according to the performance of the aircraft in different flight stages (such as take-off, cruising, and landing); Environmental adaptation: Debugging under different environmental conditions (such as wind speed changes) to ensure the robustness of the position controller; Step 4: Output Output: Based on the correction values ​​generated by the above process, the above corrections are passed to different accessories of the aircraft according to different flight modes to achieve precise control of the position.

9. A tilt-rotor vertical take-off and landing aircraft and a vertical take-off and landing balancing system according to claim 8, characterized in that: The mode switching control system includes the following layers of control system: High-level control: The high-level control system is responsible for managing the aircraft's flight status judgment and mode switching logic. The main responsibilities are: judging whether the flight mode needs to be switched based on the aircraft's current status (such as speed, attitude, acceleration, etc.), issuing switching instructions at the appropriate time, and deciding when to start and end mode switching; Low-level control: The low-level control system is responsible for specific attitude and position control. Different flight states will use different control strategies and parameters, dividing the aircraft into 3 modes: State 1: Vertical take-off and landing mode The rotor combined with the turbojet engine control logic is adopted, and the attitude and position of the aircraft are adjusted mainly by the speed difference of the front propeller (38) and the thrust state of the small turbojet engine (5). At this time, the force formula of the aircraft in the vertical direction is as follows: F 升 =ρπr 2 (n*P) 2 η+q ma (C9-C0)+A9(p9-p0) Where η is the propeller efficiency, π is the circumference of a circle, P is the propeller pitch, ρ is the atmospheric density, r is the propeller radius, n is the propeller speed per second, and q ma is the mass of air entering the engine per unit time, A9 is the cross-sectional area of ​​the engine tail nozzle outlet, p0 is the external atmospheric pressure, and p9 is the cross-sectional pressure of the engine tail nozzle. State 2: Horizontal flight mode The fixed-wing flight control logic is adopted, and the attitude is mainly controlled by adjusting the ailerons and tail wing, and the forward movement is achieved by relying on the thrust of the propeller. At this time, the lift formula of the aircraft is as follows: Where ρ is the atmospheric density, S is the wing area, V X For C y is the lift coefficient. State 3: Tilt transition mode When the vertical mode is switched to the horizontal mode, the lift size and lift source are analyzed according to the flight parameters obtained by the sensor, and a weighted algorithm is used to perform a weighted calculation according to the proportion of the lift obtained by the wing surface and the thrust of the front propeller (38) and the small turbojet engine (5) in the lift source. Then, the control of the two modes is weightedly superimposed to control the flight attitude of the aircraft. The control correction superposition formula is: A0 = F1*A1 / F Y +F2*A2 / F Y , where A0 is the total correction of the aircraft, F1 is the sum of the vertical thrust generated by the front propeller (38) of the aircraft and the lift generated by the small turbojet engine (5), F2 is the aerodynamic lift generated by the wing (3) of the aircraft, and F Y It is the resultant force of the vertical lift and the engine thrust acting on the aircraft, A1 is the vertical mode correction of the aircraft, and A2 is the horizontal mode correction of the aircraft.

10. A tiltrotor vertical take-off and landing aircraft and a vertical take-off and landing balancing system according to claim 9, characterized in that: The mode switching control system includes mode switching, and the specific steps are as follows: Step 1: Mode switching condition judgment: Switch from vertical mode to horizontal mode: When the aircraft reaches a certain altitude and speed and needs to switch to forward flight, the system will determine whether the aircraft is in a suitable state for switching; The conditions usually include: sufficient flight altitude, small pitch angle, and air speed within the range suitable for fixed-wing flight; Switch from horizontal mode to vertical mode: When the aircraft needs to slow down and enter a hovering state, such as preparing to land or hovering in a complex environment, the system will determine the timing of the switch; The conditions usually include: the flight speed is reduced to a state where it is difficult for fixed-wing flight to maintain lift, and the flight altitude is moderate; Step 2: Mode transition phase: The core of mode switching is to gradually adjust the control parameters during the transition phase to ensure flight stability. The specific steps are as follows: (1) Gradual adjustment of control gain During the mode switching process, the controller will gradually adjust the gain of the PID controller or switch the control strategy; When switching from vertical mode to horizontal mode: start to gradually reduce the control force in vertical mode and increase the control force in fixed-wing mode. X The aircraft's lift source and proportion are calculated based on the rotation speed and tilt angle of the aircraft's front propeller (38) and the rotation speed and pressure difference of the small turbojet engine (5), and weighted control is performed accordingly to gradually reduce the speed difference control between the motors in the rotor mode and increase the control of the rudder surface in the horizontal flight mode; When switching from horizontal mode to vertical mode: reduce the control force of the rudder surface and gradually increase the control of the motor speed so that the aircraft can maintain a hovering state; (2) Power distribution adjustment Tilt-rotor: If the aircraft uses a tilt-rotor design, when switching modes, the inclination angle of the rotor needs to be adjusted to convert it from vertical lift to horizontal thrust, or from horizontal thrust to vertical lift. Thruster start: When transitioning from vertical take-off and landing mode to horizontal mode, the system will start the thrusters to provide horizontal thrust while gradually reducing the motor output in the vertical direction; Turbojet engine gradual shutdown: After switching to horizontal flight mode, the thrust output of the turbojet engine is gradually reduced, and the flight is maintained by relying on the lift of the fixed wing; (3) Posture control adjustment Attitude controller switching: In vertical mode, attitude control mainly controls roll, pitch and yaw by adjusting the speed of the motor. In horizontal mode, attitude control is transferred to control by the control of the rudder surface. Smooth transition: When the control system switches modes, the attitude change must be smooth to prevent the aircraft from experiencing large attitude deviations during the switching process, which may lead to instability. Step 3: Stabilization after transition Control strategy lock: Once the mode switch is completed, the system will lock the new control strategy to ensure that the aircraft operates according to the corresponding control logic in the new mode; Controller calibration: After the switch is completed, the system will calibrate the current attitude and position to ensure that the flight status after the switch is in line with expectations.

Images