A tilt-rotor short take-off and landing aircraft and method of controlling the same

Abstract

This invention discloses a tilt-lift short takeoff and landing (STOVL) aircraft and its control method, comprising a fuselage, wing assembly, multi-layered distributed propulsion system, tilt drive assembly, and control unit. The wing is hinged to the fuselage via a pivot in the mid-rear section of the chord, and the tilt drive assembly drives the wing to rotate and adjust the angle of attack. The multi-layered distributed propulsion system includes a first rotor assembly located in the mid-rear section to the wingtip region of the wing and a second rotor assembly located on the lower surface of the wing near the fuselage. During cruise, the first rotor assembly is stopped and feathered, and the propulsion system tilts synchronously with the wing. This invention, through a layered distributed propulsion layout and synchronous wing tilting for synergistic lift enhancement, solves the problems of insufficient control torque at low speeds and high cruise energy consumption, achieving a balance between STOVL performance and cruise economy.

Description

Technical Field

[0001] This invention relates to the field of short takeoff and landing (STOVL) aircraft technology, specifically to a tilt-blow STOVL aircraft and its control method. Background Technology

[0002] Short takeoff and landing (STOL) aircraft, due to their low runway dependence and flexible takeoff and landing capabilities, have broad application prospects in fields such as civilian aerial photography, cargo transport, and emergency rescue. Existing STOL aircraft typically employ lift-enhancing devices such as Fuller flaps to improve takeoff and landing performance. However, at high angles of attack, these devices often suffer from problems such as a significant rearward shift of the center of pressure leading to difficulties in pitch attitude trim, the horizontal stabilizer easily entering the stall zone, and large fuselage angles obstructing the pilot's view. Furthermore, traditional lift-enhancing configurations suffer from insufficient dynamic pressure on aerodynamic control surfaces at extremely low airspeeds, resulting in severe control torque attenuation and making the aircraft highly susceptible to attitude oscillations or even loss of control.

[0003] In existing short takeoff and landing (STOVL) schemes that use distributed electric propulsion (DEP) to enhance lift on wings or flaps, the rotor slipstream is often used to blow large flaps or trailing edge lift enhancement devices. When the flaps deflect at a large angle, the slipstream needs to change direction at a large angle, which can easily lead to momentum deflection losses. At the same time, large flap deflections will cause the wing's center of pressure to shift significantly aft, requiring the horizontal tail to generate a large trim moment, and may even generate negative lift to offset some of the lift gains.

[0004] Furthermore, if the rotor is primarily positioned below the wing chord or mainly blows across the lower wing surface, its contribution to boundary layer adhesion, circulation increase, and stall delay on the upper wing surface may be limited. At extremely low or near-zero airspeeds, relying solely on aerodynamic control surfaces such as flaps and elevators is insufficient to provide adequate roll and pitch control moments. Therefore, existing STOL or DEP-STOL solutions still have room for improvement in further reducing takeoff and landing speeds, enhancing attitude stability at extremely low speeds, and balancing cruise efficiency.

[0005] In terms of power layout, conventional layouts struggle to balance the low-speed attitude control torque requirements with the high-speed cruise drag reduction requirements. Furthermore, multi-rotor configurations typically require all rotors to operate continuously during the cruise phase, resulting in high aerodynamic drag and energy consumption. At the control logic level, traditional aircraft exhibit strong parameter coupling between throttle, attitude, and lift enhancement mechanisms (such as flaps and tilt angles), leading to high control complexity and a lack of effective safety fallback mechanisms for low-airspeed conditions, thus necessitating improvements in flight safety. Summary of the Invention

[0006] The purpose of this invention is to provide a tilt-lift short takeoff and landing (STOVL) aircraft and its control method. This invention utilizes a layered distributed propulsion layout and a synergistic lift-enhancing configuration with synchronous wing tilting, combined with a dual closed-loop decoupled control logic for airspeed and altitude, to achieve a balance between STOVL capability and cruise efficiency, while also improving low-speed flight stability and ease of control.

[0007] The technical solution of the present invention: a tilt-lift short takeoff and landing aircraft, comprising a fuselage, a wing assembly, a multi-layer distributed propulsion system, and a tilt drive assembly; The wing assembly includes a pair of symmetrically arranged wings, which are hinged to the fuselage via a pivot located at the mid-rear part in the chord direction; The tilt drive assembly is configured to drive the wing to rotate about the pivot to adjust the wing angle of attack; The multi-layered distributed propulsion system includes a first rotor assembly and a second rotor assembly; The first rotor assembly is located on the outer side of the wing, and the second rotor assembly is located on the inner side of the wing with its thrust line near the center of gravity of the aircraft. The multi-layered distributed propulsion system is connected to the wing and tilts synchronously with the wing.

[0008] The aforementioned tilt-lift short takeoff and landing aircraft has a first rotor assembly symmetrically arranged on the upper surface of the wing from the middle and outer span to the wingtip region, and is equipped with a stop feathering mechanism to adjust the blades to a low-drag state during cruise. The second rotor assembly is symmetrically arranged on the lower surface of the wing and includes two rotors, which operate continuously under all flight conditions.

[0009] The aforementioned tilt-blow short takeoff and landing aircraft has a flaperon aerodynamic control surface on the trailing edge of the wing. The flaperon aerodynamic control surface is located within the slipstream coverage area of ​​the multi-layer distributed propulsion system and is used to obtain enhanced aerodynamic control torque at low airspeed conditions.

[0010] The aforementioned tilt-lift short takeoff and landing aircraft also includes a control unit, which is connected to the tilt drive assembly, the multi-layer distributed propulsion system, and the aircraft's control surfaces. The control unit is configured as follows: The tilt angle of the wing is independently adjusted based on the deviation between the actual airspeed of the aircraft and the target airspeed. Based on the deviation between the actual altitude of the aircraft and the target altitude, the power output parameters of the multi-layer distributed propulsion system are independently adjusted; The tilt angle adjustment and the power output parameter adjustment are executed independently, without any hard mode switching logic.

[0011] The aforementioned tilt-boom short takeoff and landing aircraft, wherein the control unit is further configured as follows: In low-speed conditions where the airspeed is less than the preset threshold, the coordinated mode of flap aileron aerodynamic control and multi-layer distributed propulsion system thrust vector control is activated simultaneously. Roll attitude correction is dominated by differential upper rotor at the wingtip, pitch moment is compensated by differential thrust of upper and lower rotors, and yaw moment is adjusted by differential left and right rotors of the wing. When the airspeed exceeds a preset threshold during cruise, the first rotor assembly is stopped and put into feathering mode, while only the second rotor assembly remains operational.

[0012] The aforementioned tilt-blow short takeoff and landing aircraft's control unit is also configured to employ a cooperative control mode that combines manual throttle with altitude-closed-loop automatic throttle. The total throttle output is the sum of the highly closed-loop automatic throttle output and the manual throttle output; Furthermore, when manual throttle is engaged for an extended period, the integral accumulation range of the highly closed loop is frozen or limited to prevent integral saturation.

[0013] The aforementioned tilt-boom short takeoff and landing aircraft, wherein the control unit is further configured as follows: Preset minimum safe height threshold and safe exit threshold for values ​​exceeding this threshold; When the real-time flight altitude is lower than the minimum safe altitude threshold, the power increment is forcibly output and the altitude change rate damping mechanism is triggered. When the flight altitude recovers to above the safe exit threshold, the forced climb mode is smoothly exited, and normal altitude closed-loop control is restored. It also synchronously links the airspeed closed loop to reduce the wing tilt angle, thus forming a three-level safety interlock of altitude, airspeed, and angle of attack.

[0014] A control method for a tilt-blow short takeoff and landing (STOVL) aircraft, applied to the aforementioned aircraft, includes: An independent dual-closed-loop decoupled control architecture is constructed, which includes a first control loop and a second control loop; The first control loop independently adjusts the wing tilt angle based on the deviation between the aircraft's actual airspeed and the target airspeed. The second control loop independently adjusts the power output parameters of the multi-layer distributed propulsion system based on the deviation between the actual altitude of the aircraft and the target altitude. The first control loop and the second control loop operate independently of each other.

[0015] The aforementioned control method, wherein the independent adjustment of the wing tilt angle based on the deviation between the aircraft's actual airspeed and the target airspeed includes: The actual airspeed of the aircraft is collected in real time, and the actual airspeed is compared with the target airspeed to obtain the airspeed deviation. The airspeed deviation is linearly and smoothly driven to adjust the wing tilt angle, wherein the airspeed value is negatively correlated with the magnitude of the wing tilt angle; Configure airspeed data verification, rate of change limiting and multi-source redundancy fault tolerance mechanism. When abnormal airspeed data is detected, lock the current wing tilt angle or switch to the backup control law. Upon detecting a manual pitch operation, a pitch feedforward compensation strategy is triggered to optimize the airspeed closed-loop response characteristics, and the wing tilt angle is not affected by the manual pitch command.

[0016] The aforementioned control method further includes: A minimum safe height threshold and a height exit threshold are preset, wherein the height exit threshold is higher than the minimum safe height threshold; When the real-time flight altitude is lower than the minimum safe altitude threshold, the power increment is forcibly output and the altitude change rate damping mechanism is triggered. When the flight altitude recovers to above the altitude exit threshold, the forced climb mode is exited, and normal altitude closed-loop control is resumed. During a forced climb, when the actual airspeed is lower than the lower limit of safe airspeed, the wing tilt angle is automatically reduced to increase airspeed. When the angle of attack approaches the stall threshold, the maximum tilt angle is limited, thus achieving a three-level safety interlock of altitude, airspeed, and angle of attack.

[0017] Compared with the prior art, the present invention has the following beneficial effects: 1. This invention utilizes a layered distributed propulsion system that tilts synchronously with the wing. During takeoff and landing, the rotor thrust directly generates a vertical lift component related to the wing tilt angle. Simultaneously, a high-speed slipstream covers the upper wing surface, delaying airflow separation and increasing wing circulation. Compared to lift enhancement methods that rely on rotor slipstreams blowing large-angle flaps, this invention reduces slipstream deflection losses at large angles and minimizes the need for negative lift trim from the horizontal tail due to a significant shift in the pressure center, thereby improving overall aircraft lift efficiency. Compared to traditional flap-based lift enhancement methods, this invention's tilt-lift system involves the rotor and wing tilting together. The rotor directly provides the vertical lift component, and the slipstream adheres closely to the upper wing surface, avoiding airflow efficiency losses caused by large-angle flap deflection. Pitch control relies on differential thrust between the upper and lower rotor layers for vector control, eliminating the need for negative lift trim from the horizontal tail and preventing the offsetting effect of negative lift on overall lift gain, significantly improving overall aircraft lift efficiency. At the same time, it can avoid the problems of traditional flaps being prone to stalling at low speeds and high angles of attack, and the horizontal tail being unable to control due to insufficient dynamic pressure at low airspeeds, thus greatly improving the low-speed controllability and attitude stability of the entire aircraft.

[0018] 2. The first rotor assembly of this invention is located on the outer side of the upper surface of the wing to the wingtip region. It can provide effective roll control torque at low airspeeds, near-zero airspeeds, or when the dynamic pressure of the aerodynamic control surfaces is insufficient, utilizing a relatively long torque arm. The thrust differential between the upper and lower rotor assemblies of the first and second rotor assemblies can compensate for the pitching moment. Combined with the flaperon aerodynamic control surfaces located in the slipstream, this significantly enhances attitude stability at extremely low airspeeds. During the cruise phase, the upper rotor assembly of this invention is feathered, significantly reducing the aircraft's drag. Only the lower rotor assembly, closer to the center of gravity, remains operational, optimizing thrust efficiency during cruise and extending endurance.

[0019] 3. This invention employs a dual closed-loop decoupled control architecture for airspeed and altitude, separating airspeed control from altitude control and simplifying the control logic. Combined with a minimum safe altitude fallback mechanism and multi-source redundancy fault-tolerant design, it effectively prevents stall and altitude loss risks at low airspeed levels, improving flight safety. This invention enables controllable gliding landing with near-zero throttle input, maintaining roll stability entirely through wingtip rotor differential action, independent of airspeed and aerodynamic control effects, exhibiting extremely strong low-speed survivability and significantly enhancing survivability under power failure conditions. Attached Figure Description

[0020] Figure 1 This is a three-dimensional view of the overall structure of the tilt-lift short takeoff and landing aircraft of the present invention; Figure 2 This is a schematic diagram of the layout of a multi-layered distributed propulsion system according to an embodiment of the present invention; Figure 3 This is a schematic diagram of the wing tilting mechanism of the present invention; Figure 4 This is a tilt-isometric view of the wing tilt mode in a short takeoff and landing (STOVL) embodiment of the present invention. Figure 5 This is a schematic diagram illustrating the generation of the roll moment of an aircraft. Figure 6 This is a schematic diagram of the yaw moment generated by an aircraft.

[0021] Among them, 10 is the fuselage; 20 is the wing; 21 is the pivot; 30 is the tilt drive assembly; 40 is the multi-layer distributed propulsion system; 41 is the first rotor assembly; 42 is the second rotor assembly; and 50 is the flap aileron aerodynamic control surface. Detailed Implementation

[0022] The present invention will be further described below with reference to the accompanying drawings and embodiments, but this should not be construed as limiting the present invention.

[0023] Example 1: As Figures 1-6 As shown, this embodiment provides a tilt-lift short takeoff and landing (STOL) aircraft. The aircraft includes a fuselage 10, a wing assembly, a multi-layer distributed propulsion system 40, a tilt drive assembly 30, and a control unit.

[0024] The fuselage 10, as the main load-bearing structure of the aircraft, is made of lightweight composite materials and has internal space to accommodate control units and other avionics equipment. The wing assembly includes a pair of wings 20 symmetrically arranged on both sides of the fuselage 10. The wings 20 are hinged to the fuselage 10 via a pivot 21 located at the mid-rear part of the chord. Specifically, the pivot 21 serves as the rotation center of the wing 20 relative to the fuselage 10, and its position is located at the mid-rear part of the chord of the wing 20, for example, at 60% to 80% of the chord length. This arrangement allows the aerodynamic center of the wing 20 to move around the trailing edge when tilting, which is beneficial for maintaining the pitch moment balance of the entire aircraft under high angles of attack conditions during takeoff and landing, and avoids the problem of severe lift center drift caused by the traditional pivot being located at the leading edge.

[0025] The multi-layered distributed propulsion system 40 includes a first rotor assembly 41 and a second rotor assembly 42, adopting a modular layout with upper and lower layers. The first rotor assembly 41 is located in the mid-outer spanwise region of the wing 20 to the wingtip area, specifically on the upper surface of the wing 20 near the wingtip; the second rotor assembly 42 is located on the lower surface of the wing 20 near the fuselage 10. The first rotor assembly 41, located in the wingtip region, has a long moment arm, which can provide sufficient roll and yaw control moments when the dynamic pressure of the control surfaces is insufficient at low speeds; the second rotor assembly 42 is set close to the fuselage center of gravity, and its thrust line is close to the overall center of gravity, which can provide efficient cruise thrust and reduce the interference of thrust changes on pitch attitude. The first rotor assembly 41 is configured to stop feathering during cruise, that is, to adjust the blade angle to be parallel to the airflow direction or at a low drag angle, thereby significantly reducing aerodynamic drag during cruise flight and improving range economy. Since the first rotor assembly 41 is located on the upper surface of the wing, its slipstream is more likely to act on the boundary layer of the upper surface of the wing and the aerodynamic control surface 50 of the trailing edge flap and aileron, which is beneficial to increasing the proportion of attached flow on the upper surface of the wing and the blow-up gain; since the second rotor assembly 42 is located on the lower surface of the wing and close to the center of gravity of the whole aircraft, it can serve as the main propulsion source during the cruise phase, reducing cruise attitude coupling.

[0026] The multi-layer distributed propulsion system 40 is connected to the wing 20 and configured to tilt synchronously with the wing 20. Specifically, the first rotor assembly 41 and the second rotor assembly 42 can be fixedly mounted on the wing 20 through structural components such as motor mounts and connecting arms. When the tilt drive assembly 30 drives the wing 20 to rotate around the pivot 21, the multi-layer distributed propulsion system 40 rotates synchronously. This synchronous tilt mechanism allows the slipstream of the propulsion system to always cover the wing surface. During takeoff and landing, the high-speed slipstream blows away airflow separation on the wing surface, increasing wing circulation and thus producing a "lift-boosting" effect, significantly improving the short takeoff and landing capability of the aircraft. In this embodiment, the first rotor assembly is located in the wingtip region and has a long moment arm. The left and right differential can provide roll control moment. This roll control is independent of airspeed and aerodynamic control surfaces, and remains effective at low speeds, near-zero airspeeds, and when the throttle input is close to zero, completely solving the problems of insufficient dynamic pressure and reduced control effectiveness in traditional aircraft at low speeds. Pitch attitude control is achieved through differential thrust between the upper and lower rotors: increasing the thrust of the upper rotor and decreasing the thrust of the lower rotor generates a pitching moment; conversely, increasing the thrust of the lower rotor and decreasing the thrust of the upper rotor generates a pitching moment. Pitch control is achieved entirely through thrust vectoring, without relying on the horizontal tail. It maintains sufficient pitch trim capability at high angles of attack and low airspeeds, eliminating the risk of insufficient rudder effectiveness. In low-speed landing tests simulating near-zero pilot throttle input, reduced main propulsion throttle, or simulated power failure, roll attitude stability can be maintained by keeping the minimum control speed or differential control margin of the first rotor group 41 and through differential wingtip rotor control. It should be noted that this condition does not mean all rotors are completely stopped, but rather that the differential capability required for attitude control is retained even when the pilot's main throttle input or main propulsion output is reduced to near zero.

[0027] The tilt drive assembly 30 is configured to drive the wing 20 to rotate around the pivot 21 to adjust the wing angle of attack. The tilt drive assembly 30 includes a drive component, a transmission mechanism, and a lifting mechanism, and is equipped with a limit structure. The drive component serves as a power source and can specifically be an electric actuator, a hydraulic actuator, or a rotary servo motor. In a preferred embodiment, an electric actuator is used as the drive component, which features a compact structure, high control precision, and fast response speed. The transmission mechanism is configured to transmit the output motion of the drive component to the lifting mechanism. The transmission mechanism can be a linkage mechanism, a rack and pinion mechanism, or a wire rope pulley mechanism. The lifting mechanism is configured to drive the wing 20 to rotate around the pivot 21. Specifically, one end of the lifting mechanism is connected to the lower surface or near the leading edge of the wing 20, and the other end is connected to the internal structure of the fuselage 10. When the drive component extends or retracts, it drives the lifting mechanism through the transmission mechanism, thereby driving the wing 20 to tilt around the trailing edge pivot 21 to adjust the wing angle of attack. To ensure flight safety and prevent excessive wing tilt angles that could lead to aerodynamic instability or mechanical interference, the tilt drive assembly 30 is equipped with a limiting structure. This limiting structure can be a mechanical hard limit, such as a stop or protrusion at the wing's extreme rotation position. When the wing reaches its extreme angle, the stop contacts the fuselage structure, physically preventing further wing rotation. Alternatively, electronic limits can be used, such as a limit switch or potentiometer inside the drive component. When the drive component extends or retracts to a preset position, the drive power is automatically cut off. By setting this limiting structure, wing tilt overtravel can be effectively prevented, ensuring the structural integrity and aerodynamic stability of the aircraft under various operating conditions.

[0028] In a preferred embodiment, the wing adopts an integral high-wing structure, with the main spar or wing box load-bearing component in the middle of the wing serving as a pivot load-bearing component hinged to the upper support of the fuselage. Compared to a segmented tiltrotor structure, this integral high-wing structure maintains the continuity of the wing's main load-bearing components, reduces large-sized arc-shaped slots in the fuselage sidewalls or wing leading and trailing edges, and improves wing root load transfer efficiency and fuselage structural integrity. A lifting mechanism is provided between the wing's leading edge or mid-forward section and the fuselage. This lifting mechanism can be an electric linear actuator, a hydraulic actuator, or other linear actuators, which drives the wing to tilt around a pivot through extension and retraction.

[0029] When the wing levels off or reaches its cruise tilt angle, a contact support area is formed between the lower surface of the wing and the upper fuselage. This contact support area is equipped with elastic damping elements or flexible seals to absorb minor vibrations between the wing and fuselage and reduce aerodynamic gaps. For medium to large-scale implementations, a mechanical locking mechanism can also be provided to lock the wing to the fuselage after the wing reaches its cruise position. This allows the cruise flight load to be shared by the locking mechanism, the pivot bearing load, and the contact support area, thereby reducing the continuous load on the tilt drive assembly and improving cruise stiffness.

[0030] The control unit, as the core control hub of the aircraft, can be implemented using a flight control computer based on ArduPilotPlane firmware or an embedded control system built with an STM series microcontroller. The control unit is configured to perform dual closed-loop decoupled control of airspeed and altitude, manual / automatic coordinated control, and low-altitude fault-tolerant protection logic. Specifically, the control unit integrates a sensor data acquisition module, a main control computing module, and an actuator drive module. The sensor data acquisition module collects flight parameters such as airspeed, altitude, attitude angle, and angular velocity in real time; the main control computing module calculates control commands according to a preset control algorithm; and the actuator drive module drives the tilt drive assembly 30, the multi-layer distributed propulsion system 40, and the flap and aileron aerodynamic control surfaces 50. The airspeed and altitude dual-loop decoupled control logic refers to the control unit constructing two independent control loops: the first control loop (airspeed loop) adjusts the wing tilt angle only based on the deviation between the actual airspeed and the target airspeed, without interfering with throttle or power output; the second control loop (altitude loop) adjusts the power output parameters only based on the deviation between the actual altitude and the target altitude, without interfering with the wing tilt angle. This decoupled architecture fundamentally severs the complex coupling relationship between throttle, attitude, and flaps / tilt angle in traditional aircraft, significantly reducing the difficulty of operation. The manual-automatic coordinated control logic means that the control unit allows operator manual commands to be superimposed on the automatic control loop. For example, in altitude control, the operator's manual throttle command is superimposed as an increment on the automatic throttle output, ensuring the stability of automatic control while giving the operator the ability to intervene flexibly. The low-altitude fault-tolerant protection logic means that the control unit presets a minimum safe altitude threshold. When the aircraft altitude is detected to be below this threshold and the rate of descent is too large, a forced incremental power output and an altitude change rate damping mechanism are automatically triggered to prevent the aircraft from crashing, providing a safety net for flight safety. At low airspeed, the control unit simultaneously activates the coordinated mode of flap and aileron aerodynamic control and layered distributed propulsion system thrust vector control: the left and right differential motion of the first rotor group 41 at the wingtip becomes the dominant control means for roll attitude, and the flap and aileron aerodynamic control surface 50 provides auxiliary roll moment under the slipstream enhancement; the upper and lower layer thrust differential of the first rotor group 41 and the second rotor group 42 compensates for the pitch moment; and the left and right wing rotor thrust differential adjusts the yaw moment.

[0031] In a specific application scenario of this embodiment, the aircraft adopts a test configuration with a wingspan of 1200mm and a takeoff weight of 2kg. A hinged pivot 21 is located in the mid-rear section of the wing 20, and an electric push rod is positioned near the leading edge as a lifting mechanism. This mechanism drives the wing 20 to tilt smoothly and controllably around the trailing edge pivot 21. Combined with a mechanical limiting structure, the wing tilt angle range is limited, avoiding aerodynamic instability caused by over-tilt. The multi-layered distributed propulsion system 40 adopts a basic four-rotor configuration: two sets of rotors are symmetrically arranged near the fuselage on the lower surface of the wing 20 as the second rotor group 42, and two sets of rotors are symmetrically arranged from the mid-outer span to the wingtip on the upper surface of the wing 20 as the first rotor group 41. All four rotors have identical dimensions and rated power, naturally achieving vertical thrust balance. The structure is simple, highly reliable, and can fully meet the requirements of conventional takeoff, landing, and cruise operations. It should be understood that the above-mentioned dimensional parameters are only a preferred example of this embodiment. In other embodiments, the size, weight and number of rotors of the aircraft can be adaptively adjusted according to actual load requirements. For example, a 6-rotor or 8-rotor extended configuration can be adopted. The 6-rotor extended configuration is extended from the preferred 4-rotor basic configuration. The lower layer retains 2 sets of main rotors, and the upper layer adds 2 sets of smaller rotors. The flight control dynamically adjusts the rotation speed of each set of rotors in the upper layer to ensure that the total thrust of the upper and lower layers is equal and balanced, which is suitable for medium load and wind-resistant operation scenarios. The 8-rotor high-configuration extended configuration keeps the 2 sets of main rotors unchanged in the lower layer and evenly arranges multiple sets of small-power, small-size rotors in the upper layer. While improving system redundancy, it can make the slipstream more evenly distributed on the upper surface of the wing, which is suitable for high-load and strong wind-resistant high-level operation scenarios.

[0032] Furthermore, the first rotor assembly 41 includes a shutdown feathering mechanism configured to adjust to a low-drag state during cruise. When the aircraft enters the high-speed cruise phase, the first rotor assembly 41 stops rotating, and the shutdown feathering mechanism actuates, adjusting the blade angle to be parallel to the incoming flow direction or at the angle of minimum drag. This design is crucial, as it solves the problem of increased drag in traditional multi-rotor aircraft during the cruise phase due to the large rotor frontal area. By shutting down the feathering mechanism, the first rotor assembly 41 generates almost no additional aerodynamic drag during cruise, significantly reducing flight energy consumption and extending endurance.

[0033] The start-stop and feathering control logic of the first rotor group 41 is as follows: Real-time monitoring of the aircraft's airspeed, throttle opening, and wing tilt angle is used. The first rotor group 41 is automatically activated when any of the following conditions are met: 1. The actual airspeed is lower than the preset STOL mode switching threshold (e.g., 35 km / h); 2. The wing tilt angle is greater than the preset angle threshold (e.g., 15 degrees); 3. The flight control system receives a manual STOL mode command from the operator. During the activation of the first rotor group 41, the rotational speed of each rotor group is dynamically adjusted according to attitude control requirements, prioritizing the use of wingtip rotor differential to provide roll control torque.

[0034] The first rotor assembly (41) is stopped and feathered when all of the following conditions are met simultaneously: 1. The actual airspeed remains above the cruise mode holding threshold (e.g., 40 km / h) for more than 5 seconds; 2. The wing tilt angle is less than the preset angle threshold (e.g., 10 degrees); 3. The pitch angle change rate is less than the safety limit (e.g., ±5 degrees / second); 4. There is no operator's manual STOL mode command. The feathering action is performed by the stop feathering mechanism, adjusting the blade angle to be parallel to the incoming flow direction or at the angle of minimum drag, so that the first rotor assembly (41) generates almost no additional aerodynamic drag in cruise mode. When it is necessary to restart the first rotor assembly (41), the flight control system first releases the feathering lock, and then gradually increases the speed to the target value to avoid impact.

[0035] Through the aforementioned adaptive start-stop and feathering control logic, the aircraft can automatically optimize the working state of the power system according to the flight conditions without manual intervention. This ensures attitude stability and lift enhancement at low speeds, while also achieving extreme drag reduction and energy consumption optimization during cruise.

[0036] Regarding the specific number and distribution of the rotor groups, the first rotor group 41 includes at least two sets of rotors, symmetrically arranged on the upper surface of the wing 20; the second rotor group 42 includes two sets of rotors, symmetrically arranged on the lower surface of the wing 20. This symmetrical arrangement ensures thrust balance between the left and right wings, avoiding unexpected yaw or roll moments. As a preferred embodiment, the first rotor group 41 can be configured as two groups, located at the wingtips of the left and right wings respectively, utilizing long moment arms to provide maximum roll control moment; the second rotor group 42 can be configured as two groups, symmetrically distributed close to the fuselage center of gravity, providing the main cruise thrust. It should be understood that, depending on load requirements and redundancy requirements, the first rotor group 41 can also be expanded to four or more groups, evenly distributed on the upper surface of the wing, to enhance low-speed attitude control capabilities or provide higher safety redundancy.

[0037] Furthermore, the trailing edge of the wing 20 is equipped with flaperon aerodynamic control surfaces 50, which are located within the slipstream coverage of the multi-layer distributed propulsion system 40. During low-speed takeoff and landing, when the aircraft's airspeed is low and the dynamic pressure of the aerodynamic control surfaces is insufficient, the high-speed slipstream generated by the multi-layer distributed propulsion system 40 directly blows onto the flaperon aerodynamic control surfaces 50, ensuring that the control surfaces receive sufficient aerodynamic force even at low airspeeds. This significantly improves control surface efficiency and control response speed. This "blowing-assisted stabilization" mechanism, combined with the long-moment-arm thrust vector control of the first rotor assembly 41, achieves synergistic enhancement between aerodynamic control surfaces and thrust vector, completely solving the problem of control surface failure and attitude instability in traditional short takeoff and landing aircraft at extremely low airspeeds.

[0038] Example 2: This example provides a control method for a tilt-blow short takeoff and landing (STOVL) aircraft, which is applied to the aircraft described in the above example. The core of this method lies in constructing an independent dual-closed-loop decoupled control architecture, which fundamentally severs the complex coupling relationship between throttle, attitude, and lift enhancement mechanisms in traditional aircraft, significantly reducing the difficulty of operation and improving flight safety.

[0039] Specifically, the method first constructs an independent dual-loop decoupled control architecture, which includes a first control loop and a second control loop.

[0040] In this design, the first control loop is an airspeed closed loop, and the second control loop is an altitude closed loop. These two loops operate independently of each other in terms of control logic, without interfering with each other. In traditional aircraft control, changes in throttle directly affect airspeed and altitude, and changes in flaps or tilt angle can cause drastic changes in the lift coefficient, leading to altitude fluctuations. Pilots need to coordinate throttle, pitch attitude, and flap handles simultaneously, making operation difficult and prone to human error. This embodiment decouples airspeed control from altitude control, allowing the pilot to focus solely on altitude during takeoff and landing, while the system automatically maintains the preset airspeed; during cruise, the pilot only needs to set the target airspeed, and the system automatically maintains the altitude, achieving the simplicity of "single-parameter" control.

[0041] The first control loop independently adjusts the wing tilt angle based on the deviation between the aircraft's actual airspeed and the target airspeed. Specifically, the control unit collects the aircraft's actual airspeed in real time through sensors such as the pitot tube and compares it with the preset target airspeed to obtain the airspeed deviation. This airspeed deviation serves as the sole input variable. After PID calculation or fuzzy control algorithm, the output control command drives the tilt drive component 30 to adjust the tilt angle of the wing 20. For example, when the actual airspeed is lower than the target airspeed, the first control loop automatically reduces the wing tilt angle (i.e., reduces the wing angle of attack) to reduce drag and thus increase airspeed; conversely, it increases the tilt angle. During this process, the first control loop does not interfere with the power output parameters (throttle) and focuses solely on maintaining airspeed stability through aerodynamic configuration adjustments. This design utilizes the characteristics of tilt-lift aircraft: the wing tilt angle is directly related to the lift coefficient and drag coefficient, making it the optimal means of adjusting airspeed, thus avoiding the high-coupling problem caused by traditional aircraft adjusting airspeed by changing pitch attitude.

[0042] The second control loop independently adjusts the power output parameters based on the deviation between the aircraft's actual altitude and the target altitude. Specifically, the control unit acquires the aircraft's actual altitude in real time through sensors such as barometers or GPS, and compares it with the target altitude to obtain the altitude deviation. This altitude deviation serves as the sole input variable. After PID calculation, the output control command adjusts the total thrust output (i.e., total throttle) of the multi-layer distributed propulsion system 40. For example, when the actual altitude is lower than the target altitude, the second control loop automatically increases the power output to increase the vertical velocity and regain altitude; conversely, it decreases the power output. During this process, the second control loop does not interfere with the wing tilt angle, focusing only on maintaining altitude stability through the vertical component of the thrust vector. Because the multi-layer distributed propulsion system 40 tilts synchronously with the wing, its thrust direction changes with the wing angle of attack, but the second control loop, through a decoupling algorithm, always maps the total thrust output to a control effect on altitude, thus achieving independent altitude control.

[0043] In this design, the first and second control loops operate independently. This independence is reflected in the decoupling of control variables: airspeed deviation only drives tilt angle adjustment, and altitude deviation only drives throttle adjustment. Even if the pilot intervenes manually (such as suddenly increasing the throttle), the first control loop will quickly respond to the airspeed change and automatically adjust the wing tilt angle to maintain a constant airspeed, without the chain reaction of "increasing the throttle causing pitching up, increasing altitude, and then needing to push the stick to correct" that occurs in traditional aircraft. This decoupled architecture completely solves the technical problems of multi-variable coupling and complex control during the takeoff and landing phases of traditional short takeoff and landing aircraft, enabling the aircraft to maintain good handling stability even at extremely low airspeeds and high angles of attack.

[0044] Furthermore, the control method in this embodiment includes an airspeed loop control step, an altitude loop control step, and a safety fallback control step.

[0045] Regarding the airspeed loop control procedure, the actual airspeed of the aircraft is collected in real time, and the actual airspeed is compared with the target airspeed to obtain the airspeed deviation. Based on the airspeed deviation, the wing tilt angle is adjusted linearly and smoothly, where the airspeed value is negatively correlated with the wing tilt angle. Specifically, the airspeed control algorithm built into the control unit is not a simple on / off control or step control, but rather adopts a linear smoothing drive strategy. When the actual airspeed is lower than the target airspeed, the control unit outputs a command to control the tilt drive component 30 to gradually reduce the wing tilt angle 20 (i.e., reduce the angle of attack) to reduce flight drag and thus increase airspeed; conversely, it increases the tilt angle. This process uses a ramp function or low-pass filter algorithm for smoothing, avoiding sudden changes in the wing tilt angle that could cause drastic fluctuations in the lift coefficient, ensuring a smooth transition in flight attitude. For example, when the aircraft switches from cruise mode to short takeoff and landing mode, the wing tilt angle increases smoothly at a preset rate of change (e.g., 1 to 5 degrees per second), rather than abruptly jumping. Furthermore, this embodiment also incorporates airspeed data verification, rate of change limiting, and multi-source redundancy fault-tolerant mechanisms. Upon detecting abnormal airspeed data, it locks the current wing tilt angle or switches to a backup control law. Since airspeed is a core parameter of flight control, any blockage or data anomaly in the airspeed sensor (such as a pitot tube) directly threatens flight safety. Therefore, the control unit monitors the validity of airspeed data in real time, for example, checking whether the rate of change of airspeed exceeds physical limits (such as a sudden jump of 100 km / h). If an abnormal data anomaly is detected, the system automatically locks the wing tilt angle at the current safe position, no longer adjusting it with airspeed fluctuations to prevent the aircraft from stalling or diving due to erroneous signals. Simultaneously, the system can switch to a backup control law based on GPS ground speed or inertial navigation data to maintain basic flight capability.

[0046] Regarding the altitude loop control steps, the total power output parameters are obtained, which are the sum of the automatic altitude closed-loop output and the manual input superposition. This embodiment employs a manual-automatic coordinated control strategy, allowing the pilot's manual throttle commands to be superimposed on the automatic altitude closed-loop control. Specifically, the altitude closed loop outputs basic power output parameters based on real-time altitude and altitude change rate, with manual input serving as incremental superposition intervention. The total power output parameters are the sum of the automatic altitude closed-loop output and the manual input superposition, with manual input only acting as incremental superposition intervention. Simultaneously, integral anti-saturation logic is configured to freeze or limit the integral accumulation range of the altitude closed loop when prolonged manual throttle intervention is detected, preventing altitude jumps caused by integral saturation. For example, during short takeoff and landing (STOVL) operations, the pilot may need to flexibly adjust the throttle according to ground obstacle conditions. In this case, the manual throttle signal is directly superimposed on the automatic throttle, giving the pilot the final decision-making authority. To prevent control integral saturation problems caused by manual intervention, this embodiment configures integral anti-saturation logic to freeze or limit the integral accumulation range when prolonged manual input intervention occurs. In traditional PID control, if the pilot maintains a low throttle for an extended period, the automatic altitude closed loop will continuously accumulate integral terms due to persistent altitude deviations. Once the pilot releases the throttle, these accumulated integral terms can cause a sudden surge in throttle, resulting in a "jump" phenomenon. This embodiment monitors the duration and magnitude of manual input. When it determines that manual input has been in a prolonged state, it automatically freezes or limits the accumulation of integral terms in the altitude closed loop, ensuring that automatic control can smoothly take over after manual intervention ends, avoiding altitude oscillations.

[0047] Regarding the safety fallback control steps, this embodiment presets a minimum safe altitude threshold and an altitude exit threshold, with the altitude exit threshold being higher than the minimum safe altitude threshold. For example, the minimum safe altitude threshold can be set to 10 meters, and the altitude exit threshold to 12 meters, with a 2-meter hysteresis interval between them to prevent frequent triggering of protection logic at critical altitudes. When the real-time flight altitude is below the minimum safe altitude threshold, a forced power increment is output and the altitude change rate damping mechanism is triggered. When the aircraft drops below 10 meters due to sudden turbulence or operational error, the control unit will ignore the current altitude closed-loop control output or the pilot's small throttle command and forcibly increase the total thrust output of the multi-layer distributed propulsion system 40 to ensure that the aircraft quickly leaves the low-altitude danger zone. At the same time, the altitude change rate damping mechanism is triggered, i.e., limiting the aircraft's vertical descent rate to prevent grounding due to inertia. When the flight altitude recovers above the altitude exit threshold, the forced climb mode is exited, and normal altitude closed-loop control is restored. When the aircraft's altitude rises above 12 meters, the system determines that it is out of danger, smoothly disengages the forced power increment, and returns to normal automatic or manual control mode, avoiding sudden climb overshoot caused by the intervention of protection logic. This safety fallback logic provides a last line of defense for low-altitude operations, effectively preventing crashes caused by low-altitude altitude drops.

[0048] Furthermore, this embodiment defines an adaptive mode switching logic based on an airspeed threshold.

[0049] When the actual airspeed is lower than or equal to a preset low-speed threshold (e.g., ≤30km / h), the aircraft is determined to be in STOL (Slow-to-Low) operating condition, and the coordinated mode of flaperon aerodynamic control and multi-layer distributed propulsion system thrust vector control is automatically activated. In this mode: the differential motion of the upper rotor at the wingtip becomes the dominant control method for roll attitude, while the flaperon provides residual basic aerodynamic torque as an auxiliary; the differential thrust of the upper and lower rotors precisely compensates for pitch torque; and the differential motion of the left and right rotors completes yaw control. Through the above coordinated control, the inherent defects of insufficient dynamic pressure of aerodynamic control surfaces and lack of control torque under low-speed conditions are completely compensated.

[0050] When the actual airspeed exceeds a preset low-speed threshold (e.g., >30 km / h), the aircraft is determined to have entered normal cruise mode. It automatically shuts down all rotors of the first rotor assembly 41 and locks the feathering state, leaving only the lower second rotor assembly 42 operational. Simultaneously, the first control loop (airspeed closed loop) linearly and smoothly reduces the tilt angle of the wing 20 based on the airspeed deviation until the optimal cruise angle of attack (approximately 3-4 degrees) is reached.

[0051] The above-mentioned mode switching process is automatically completed by the flight control system without manual intervention, without abrupt changes or attitude disturbances, realizing adaptive and seamless switching between STOL low-speed mode and high-speed cruise mode.

[0052] Furthermore, this embodiment presets a minimum safe altitude threshold (e.g., 10 meters) and an altitude exit threshold (e.g., 12 meters). When the real-time flight altitude is lower than the minimum safe altitude threshold, the system triggers a forced climb mode, specifically including: Level 1 (Altitude Mandatory Protection): Ignoring the current altitude closed-loop control output and manual throttle command, forcibly increasing the total thrust output of the multi-layer distributed propulsion system 40, while triggering the altitude change rate damping mechanism to limit the vertical descent rate and ensure that the aircraft quickly leaves the low-altitude danger zone.

[0053] Level 2 (Airspeed Coordination Protection): During a forced climb, if the actual airspeed is lower than the preset safe airspeed limit (e.g., 25 km / h), the first control loop automatically reduces the wing tilt angle by 20° to reduce drag and increase airspeed, preventing the risk of stall caused by low-speed flight at high angles of attack.

[0054] Level 3 (Angle of Attack Limitation Protection): The flight control system monitors the aircraft's pitch and angle of attack in real time. When the angle of attack approaches the stall threshold, it automatically limits the maximum tilt angle of the wing 20 and outputs a warning command to suppress further pitch control, thus preventing stall triggering from an aerodynamic perspective.

[0055] When the aircraft climbs back to above the altitude exit threshold (e.g., 12 meters), the system smoothly exits the forced climb mode and resumes normal altitude and airspeed closed-loop control. Through the coordinated control of throttle, airspeed, and tilt angle, a three-level safety interlock of altitude, airspeed, and angle of attack is achieved, comprehensively avoiding flight risks such as low-altitude altitude loss, stall, and loss of control.

[0056] Furthermore, this embodiment also includes a decoupling logic for manual pitch control and feedforward compensation: manual pitch commands independently control the fuselage pitch attitude, and the flight control system outputs a small amount of trim compensation; the wing tilt angle is not affected by manual pitch commands, but is controlled independently by the airspeed closed loop; when a manual pitch operation is detected, the pitch feedforward compensation strategy is triggered to optimize the airspeed closed loop response characteristics.

[0057] Example 3: To verify the technical effectiveness of the tilt-blow short takeoff and landing (FPV) aircraft and its control method provided by this invention, the applicant manufactured a prototype with a wingspan of 1200mm and a takeoff weight of 2kg for full-scale FPV flight testing. This prototype adopted the structural layout and control logic described in the above examples. Specific test results are as follows: In ultra-low speed stability tests, the aircraft maintained pitch and roll stability at extremely low airspeeds of approximately 20 km / h, relying on the wingtip differential control of the first rotor assembly 41, the differential control of the upper and lower rotors, and the coordinated control of the slipstream flaps and ailerons. During the tests, when roll differential control was disabled, the aircraft rapidly experienced roll instability at extremely low airspeeds and was difficult to recover using conventional control surfaces; when roll differential control was restored, the aircraft quickly regained roll stability. This result verifies the necessity and effectiveness of the first rotor assembly 41 in low-speed stability enhancement. This data is significantly better than the minimum safe flight speed of traditional STOL aircraft, which is typically above 35 km / h. This is mainly due to the long moment arm formed by the first rotor assembly 41 in the wingtip region, which provides sufficient roll and yaw control moments through rotor differential control when the dynamic pressure of the aerodynamic control surfaces is insufficient, verifying the effectiveness of the layered distributed propulsion layout in low-speed stability enhancement. Meanwhile, the aircraft achieved a controllable taxiing landing with near-zero throttle input and maintained roll attitude stability entirely by relying on the differential operation of the wingtip motors, demonstrating the fault tolerance capability of the invention under extreme conditions.

[0058] In short takeoff and landing (STOVL) performance tests, the aircraft requires only 30% throttle and 17A of instantaneous operating current for takeoff, achieving a smooth liftoff at a ground speed of less than 20 km / h. This is thanks to the synchronized tilting design of the wing 20 and the multi-layer distributed propulsion system 40. During takeoff and landing, the tilting motion directly generates a vertical lift component, which, combined with the slipstream blowing away airflow separation on the wing surface, significantly improves the lift coefficient, thus achieving STOVL capability. Compared to traditional aircraft, this invention effectively reduces takeoff and landing speed and distance, and improves site adaptability.

[0059] In cruise economy tests, the aircraft maintained an optimal cruise angle of attack of 3-4 degrees, requiring only 15% throttle and a stable operating current of 7A to achieve cruise flight. This verifies the effectiveness of the first rotor assembly 41's feathering design under cruise conditions. This design significantly reduces cruise drag, and the second rotor assembly 42's close proximity to the center of gravity minimizes thrust transmission losses, resulting in a substantial reduction in overall cruise energy consumption and excellent range performance.

[0060] In the mode switching test, when the aircraft switched from STOL mode to high-speed cruise mode, the wing tilt angle automatically and smoothly decreased as airspeed increased. The transition between STOL and cruise modes was seamless, without any abrupt changes or attitude disturbances. This verifies the effectiveness of the dual closed-loop decoupled control architecture for airspeed and altitude. The first control loop linearly and smoothly adjusts the wing tilt angle according to the airspeed deviation, avoiding the complex operations and attitude jumps during mode switching of traditional aircraft, resulting in a smooth and consistent flight feel throughout the entire flight.

[0061] It should be understood that the above-mentioned flight test data only represent the specific performance of the present invention under specific test aircraft parameters, and are used to verify the feasibility of the technical solution. In practical applications, the specific airspeed, current, and angle parameters may be adjusted according to the size, payload, and power configuration of the aircraft, but all should be covered within the protection scope of the present invention.

[0062] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention, such as adjusting the number of rotor assemblies, changing the specific mechanical structure of the tilt drive assembly, or adaptively modifying the parameters in the control algorithm, should all be covered within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A tilt-lift short takeoff and landing (STOL) aircraft, characterized in that, It includes the fuselage (10), wing assembly, multi-layer distributed propulsion system (40) and tilt drive assembly (30); The wing assembly includes a pair of symmetrically arranged wings (20), which are hinged to the fuselage (10) via a pivot (21) located in the chord-direction rear section; The tilt drive assembly (30) is configured to drive the wing (20) to rotate about the pivot (21) to adjust the wing angle of attack; The multi-layer distributed propulsion system (40) includes a first rotor assembly (41) and a second rotor assembly (42); The first rotor assembly (41) is located on the outer side of the wing (20), and the second rotor assembly (42) is located on the inner side of the wing (20) with its thrust line close to the center of gravity of the aircraft. The multi-layer distributed propulsion system (40) is connected to the wing (20) and tilts synchronously with the wing (20).

2. The tilt-lift short takeoff and landing aircraft according to claim 1, characterized in that, The first rotor assembly (41) is symmetrically arranged on the upper surface of the wing (20) from the middle and outer span to the wingtip area, and is equipped with a stop feathering mechanism to adjust the blades to a low drag state during cruise. The second rotor assembly (42) is symmetrically arranged on the lower surface of the wing (20) and includes two rotors, which work continuously under all flight conditions.

3. The tilt-lift short takeoff and landing aircraft according to claim 1, characterized in that, The trailing edge of the wing (20) is provided with a flaperon aerodynamic control surface (50), which is located within the slipstream coverage of the multi-layer distributed propulsion system (40) and is used to obtain enhanced aerodynamic control torque under low airspeed conditions.

4. The tilt-lift short takeoff and landing aircraft according to claim 1, characterized in that, It also includes a control unit, which is connected to the tilt drive assembly (30), the multi-layer distributed propulsion system (40), and the control surfaces of the aircraft; The control unit is configured as follows: Based on the deviation between the actual airspeed of the aircraft and the target airspeed, the tilt angle of the wing (20) is independently adjusted; Based on the deviation between the actual altitude of the aircraft and the target altitude, the power output parameters of the multi-layer distributed propulsion system (40) are independently adjusted; The tilt angle adjustment and the power output parameter adjustment are executed independently, without any hard mode switching logic.

5. The tilt-lift short takeoff and landing aircraft according to claim 4, characterized in that, The control unit is also configured to: In low-speed conditions where the airspeed is less than the preset threshold, the coordinated mode of flap and aileron aerodynamic control and multi-layer distributed propulsion system thrust vector control is activated simultaneously. Roll attitude correction is dominated by differential control of the upper rotor at the wingtip, pitch moment is compensated by differential thrust of the upper and lower rotors, and yaw moment is adjusted by differential control of the left and right rotors of the wing. When the airspeed exceeds the preset threshold during cruise, the first rotor assembly (41) is stopped and put into feathering mode, while only the second rotor assembly (42) remains operational.

6. The tilt-lift short takeoff and landing aircraft according to claim 4, characterized in that, The control unit is also configured to employ a collaborative control mode that combines manual throttle with highly closed-loop automatic throttle. The total throttle output is the sum of the highly closed-loop automatic throttle output and the manual throttle output; Furthermore, when manual throttle is engaged for an extended period, the integral accumulation range of the highly closed loop is frozen or limited to prevent integral saturation.

7. The tilt-lift short takeoff and landing aircraft according to claim 4, characterized in that, The control unit is also configured to: Preset minimum safe height threshold and safe exit threshold for values ​​exceeding this threshold; When the real-time flight altitude is lower than the minimum safe altitude threshold, the power increment is forcibly output and the altitude change rate damping mechanism is triggered. When the flight altitude recovers to above the safe exit threshold, the forced climb mode is smoothly exited, and normal altitude closed-loop control is restored. It also synchronously links the airspeed closed loop to reduce the wing tilt angle, thus forming a three-level safety interlock of altitude, airspeed, and angle of attack.

8. A control method for a tilt-boost short takeoff and landing (STOVL) aircraft, applied to the aircraft according to any one of claims 1 to 7, characterized in that, include: An independent dual-closed-loop decoupled control architecture is constructed, which includes a first control loop and a second control loop; Through the first control loop, the tilt angle of the wing (20) is independently adjusted based on the deviation between the actual airspeed of the aircraft and the target airspeed; Through the second control loop, the power output parameters of the multi-layer distributed propulsion system (40) are independently adjusted based on the deviation between the actual altitude of the aircraft and the target altitude; The first control loop and the second control loop operate independently of each other.

9. The control method according to claim 8, characterized in that, The independent adjustment of the wing (20) tilt angle based on the deviation between the actual airspeed of the aircraft and the target airspeed includes: The actual airspeed of the aircraft is collected in real time, and the airspeed deviation is obtained by comparing the actual airspeed with the target airspeed. The tilt angle of the wing (20) is adjusted by linear smoothing of the airspeed deviation, wherein the airspeed value is negatively correlated with the size of the wing tilt angle; Configure airspeed data verification, rate of change limiting and multi-source redundancy fault tolerance mechanism, and lock the current wing tilt angle or switch to the backup control law when abnormal airspeed data is detected. Upon detecting a manual pitch operation, a pitch feedforward compensation strategy is triggered to optimize the airspeed closed-loop response characteristics, and the wing tilt angle is not affected by the manual pitch command.

10. The control method according to claim 8, characterized in that, The method further includes: A preset minimum safe height threshold and a height exit threshold are defined, wherein the height exit threshold is higher than the minimum safe height threshold. When the real-time flight altitude is lower than the minimum safe altitude threshold, the power increment is forcibly output and the altitude change rate damping mechanism is triggered. When the flight altitude recovers to above the altitude exit threshold, the forced climb mode is exited, and normal altitude closed-loop control is restored. During a forced climb, when the actual airspeed is lower than the lower limit of safe airspeed, the wing tilt angle is automatically reduced to increase airspeed. When the angle of attack approaches the stall threshold, the maximum tilt angle is limited, thus achieving a three-level safety interlock of altitude, airspeed, and angle of attack.

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