Tilting type composite thrust system of small fixed-wing aircraft

Abstract

The invention relates to the field of unmanned aerial vehicles, and discloses a tilting type composite thrust system of a small fixed-wing aircraft, comprising: a physical bearing and power unit module comprising wings and six groups of tilting thrust units on the wings, each group of unit being composed of a turboprop engine and a variable pitch propeller and generating thrust; the thrust vector adjusting module comprises six sets of electric tilting joints and drives the tilting thrust unit to rotate so as to adjust the thrust direction. The state sensing and feedback module is provided with an angle sensor for monitoring the tilting angle in real time; the cooperation and synchronization control module comprises a synchronization controller which is connected with the angle sensor and the electric tilting joints and controls the six sets of electric tilting joints to be converted in batches according to the monitored angle. According to the invention, the integrated electric tilting joint is arranged, and the joint contains a main tilting shaft for wide-angle mode switching and an auxiliary fine tuning shaft for small-range active compensation, so that two adjusting mechanisms with different functions are integrated in a single structure.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of unmanned aerial vehicles, in particular to a fixed-wing small airplane tilting compound thrust system. BACKGROUND

[0002] In modern aviation applications, especially in complex terrain, space-limited areas to perform tasks such as material transportation, line inspection or emergency rescue, the dependence on the take-off and landing site of the aircraft becomes a significant constraint factor. Fixed-wing aircraft has the advantages of long range, high cruising speed and high energy efficiency, but it cannot operate without traditional long-distance runway. Therefore, developing an aircraft that can take off and land vertically like a helicopter and fly efficiently like a fixed-wing aircraft has become an inevitable trend to meet the needs of such application scenarios. The technical field to which the present application belongs is the core of achieving this goal - a tilting compound thrust system.

[0003] In the prior art, in order to realize the mode switching of vertical take-off and horizontal flight, various tilting mechanism schemes have been developed. These schemes successfully realize the large-scale deflection of the thrust direction between the vertical and the horizontal by setting a rotatable power unit on the wing. In the implementation mechanism, a high-power servo motor or hydraulic system is usually used to drive a rigid shaft to provide enough torque to flip the entire power unit. In terms of flight stability control, these schemes mostly rely on the traditional aerodynamic control surfaces (such as ailerons, elevators, rudders) of the aircraft itself. When the flight speed reaches a certain value, the aerodynamic force generated by the control surface is used to adjust and stabilize the flight attitude.

[0004] However, the above-mentioned prior art has exposed several inherent technical defects in practical application. First, the design idea of separating the large-scale tilting function from the high dynamic response stability augmentation function leads to redundant system structure, and the coexistence of independent tilting mechanism and additional stability augmentation surfaces or devices increases the total weight and mechanical complexity of the system, not only occupying the effective payload space, but also introducing more potential failure points; secondly, during the key phase of flight mode transition, independent control of multiple groups of tilting units cannot guarantee absolute motion synchronization. Any slight angle difference caused by motor response, mechanical transmission or uneven aerodynamic load will immediately be converted into an asymmetric thrust, which will generate a roll or yaw moment that interferes with the flight attitude, directly threatening the flight stability; finally, the control logic of tilting angle control and propeller pitch adjustment is generally disconnected in the prior art. These two key variables usually need to be coordinated independently by the flight control system. If the coordination is not proper, for example, the tilting angle changes ahead of the pitch angle, it is easy to cause a sudden drop in lift or a sudden change in thrust, resulting in a sharp fluctuation of the flight attitude. Therefore, the present application proposes a fixed-wing small airplane tilting compound thrust system to solve the above problems. SUMMARY

[0005] In view of the deficiencies of the prior art, the present application provides a fixed-wing small airplane tilting compound thrust system, which solves the problems of poor synchronization of tilting units, low transition efficiency in the mode transition process of the existing tilting rotor aircraft, and weak anti-side wind interference capability in the vertical take-off and landing mode.

[0006] To achieve the above object, the present application is implemented by the following technical scheme: a fixed-wing small airplane tilting compound thrust system, comprising:

[0007] A physical bearing and power unit module, which comprises a wing and six groups of tilting thrust units symmetrically arranged on the wing, the six groups of tilting thrust units are divided into three pairs, two groups in each pair are distributed symmetrically along the left and right of the wing, and each group of tilting thrust units has a vortex propeller engine and a variable pitch propeller for generating thrust;

[0008] A thrust vector adjustment module, which comprises six groups of electric tilting joints, each group of electric tilting joints connects a corresponding group of tilting thrust units and the wing, and is used to drive the overall rotation of the tilting thrust units to adjust the direction of the thrust;

[0009] A state sensing and feedback module, which comprises an angle sensor arranged on each group of electric tilting joints, for real-time monitoring of the tilting angle of each group of electric tilting joints;

[0010] A cooperative and synchronous control module, which comprises a synchronous controller, the synchronous controller is electrically connected with the angle sensor and the electric tilting joint, and is used to control the synchronous rotation of the three pairs of electric tilting joints according to the monitored tilting angle.

[0011] In a specific embodiment, the step of adjusting the size of the thrust by the physical bearing and power unit module comprises: the synchronous controller sends instructions to the pitch adjustment mechanism of the variable pitch propeller according to the current flight mode; the pitch adjustment mechanism drives the rotation of the blades of the variable pitch propeller to change the pitch angle; the size of the thrust is determined by a thrust coefficient, which is a function of the pitch angle and the forward ratio of the airplane.

[0012] Preferably, the structure of each group of electric tilting joints in the thrust vector adjustment module comprises: a double-axis hinge, the double-axis hinge has a main tilting shaft and an auxiliary fine adjustment shaft, the main tilting shaft is used to realize large-range tilting adjustment of 0 degrees to 90 degrees, and the auxiliary fine adjustment shaft is arranged perpendicularly to the main tilting shaft; a servo motor, which is used to drive the rotation of the main tilting shaft; a magneto-rheological damper, which is linked with the auxiliary fine adjustment shaft, and is used to drive the auxiliary fine adjustment shaft to make small-range angle fine adjustment.

[0013] Further, the state sensing and feedback module, the angle sensor is arranged on the main tilt axis and the auxiliary fine adjustment axis respectively to measure the tilt angle and the fine adjustment angle, and both of the angle sensors are electrically connected to the synchronous controller to feed back the motion posture of the six groups of electric tilt joints to the synchronous controller in real time.

[0014] In one embodiment, the method for controlling the synchronous rotation of the three pairs of electric tilt joints by the cooperative and synchronous control module includes: the synchronous controller receiving a target tilt angle from a flight control system; the synchronous controller collecting real-time tilt angles fed back by the six groups of angle sensors and calculating the angle error of each of the real-time tilt angles and the target tilt angle; the synchronous controller generating a control signal based on the angle error to independently drive the six groups of electric tilt joints and control the absolute value of the difference between the real-time tilt angles of each pair of electric tilt joints to be no more than 0.5 degrees.

[0015] Preferably, the cooperative and synchronous control module is further used to execute a pitch and tilt cooperative control strategy, which includes: in the vertical take-off and landing mode, the synchronous controller controlling the tilt angle of the electric tilt joints to be in the range of 85 degrees to 90 degrees and instructing the pitch angle of the variable pitch propeller to be in the large thrust setting value range; in the horizontal flight mode, the synchronous controller controlling the tilt angle of the electric tilt joints to be in the range of 0 degrees to 5 degrees and instructing the pitch angle of the variable pitch propeller to be in the high-efficiency cruising setting value range; in the transition mode, the synchronous controller controlling the tilt angle to be continuously changed in batches from 90 degrees to 0 degrees while instructing the pitch angle to be continuously changed from the large thrust setting value to the high-efficiency cruising setting value.

[0016] Further, the cooperative and synchronous control module is further used to execute a side wind interference compensation, and the steps are: the synchronous controller receiving a yaw posture error signal caused by the side wind interference and fed back by a body posture sensor; the synchronous controller calculating a compensation torque for compensation according to the yaw posture error signal; the synchronous controller converting the compensation torque into a control current instruction applied to the magnetorheological damper; the magnetorheological damper generating a controllable damping force under the control of the control current to drive the auxiliary fine adjustment axis to perform an angle compensation in the range of plus or minus 5 degrees.

[0017] As a specific implementation, the continuous change in batches in the transition mode includes: the synchronous controller triggering the driving of the three pairs of tilt thrust units to change the tilt angle in batches until all the tilt thrust units complete the change from 90 degrees to 0 degrees under a preset time or posture condition.

[0018] In a specific implementation, the set value interval of the pitch angle further comprises: the large force set value interval is 25 degrees to 30 degrees; and the high-efficiency cruising set value interval is 5 degrees to 15 degrees.

[0019] In a specific embodiment, the method for generating a control signal by the synchronous controller based on the angle error is: using a proportional, integral, and differential control algorithm, taking the angle error as an input, and generating a control signal for driving a servo motor in the electrically-driven tilting joint according to proportional, integral, and differential terms of the angle error.

[0020] The application provides a fixed-wing small airplane tilting compound thrust system.

[0021] 1. The application integrates the electrically-driven tilting joint, which contains a main tilting shaft for large-angle mode switching and an auxiliary fine-tuning shaft for small-range active compensation, to significantly improve the integration degree of the system and reduce the overall weight and size.

[0022] 2. The application uses a synchronous control method to synchronously control the three pairs of tilting thrust units on both sides of the wing with high precision, and controls the tilting angle of each pair of thrust units in real time to keep the angle difference within a very small range, thereby ensuring the high symmetry of the thrust vectors on both sides during the tilting process.

[0023] 3. The application establishes a certain pitch and tilting coordination control strategy, and the pitch angle of the variable-pitch propeller is synchronously and continuously adjusted according to the change of the tilting angle when the flight mode is converted, so that the transition process of the system from the vertical lift mode to the horizontal thrust mode is smooth and predictable. BRIEF DESCRIPTION OF DRAWINGS

[0024] Figure 1 It is a perspective view of the device of the application.

[0025] Figure 2Figure 1 is a side view of the electrically driven tilting joint of the present application;

[0026] Figure 3 Figure 2 is a perspective view of the electrically driven tilting joint of the present application;

[0027] Figure 4 Figure 3 is a block diagram of the system control module of the present application;

[0028] Figure 5 Figure 4 is a flow chart of the cooperative and synchronous control of the present application;

[0029] Figure 6 Figure 5 is a flow chart of the control of the side wind interference compensation of the present application.

[0030] Wherein, 1, wing; 2, tilting thrust unit; 21, turboprop engine; 3, variable pitch propeller; 4, electrically driven tilting joint; 41, double shaft hinge; 411, main tilting shaft; 412, auxiliary fine adjustment shaft; 42, magneto-rheological damper; 43, angle sensor; 44, servo motor; 5, synchronous controller. DETAILED DESCRIPTION

[0031] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the scope of protection of the present application.

[0032] Referring to the accompanying drawings Figure 1 , Figure 2 , Figure 3 and Figure 4 The present application provides a tilting compound thrust system for a fixed-wing light aircraft. The system can include a physical bearing and power unit module, a thrust vector adjustment module, a state sensing and feedback module, and a cooperative and synchronous control module.

[0033] The physical bearing and power unit module includes six groups of tilting thrust units 2 symmetrically arranged on the wing 1, and the six groups of tilting thrust units 2 are divided into three pairs, and the two groups in each pair are distributed symmetrically along the wing 1 left and right. The thrust vector adjustment module includes six groups of electrically driven tilting joints 4, and each group of electrically driven tilting joints 4 connects the corresponding tilting thrust unit 2 and the wing 1. The state sensing and feedback module includes an angle sensor 43 arranged on each group of electrically driven tilting joints 4. The cooperative and synchronous control module includes a synchronous controller 5, which is electrically connected with the angle sensor 43 and the electrically driven tilting joint 4.

[0034] The overall operation of the system is based on a closed-loop control logic. The synchronization controller 5 receives target commands, such as target tilt angle or flight mode switching commands, from the upper flight control system. At the same time, the angle sensor 43 installed on the electric tilt joint 4 monitors the current actual tilt angle of the tilt thrust unit 2 in real time and feeds back the angle signal to the synchronization controller 5.

[0035] After receiving the target command and the real-time angle signal, the synchronization controller 5 performs calculation according to the internal set control algorithm to generate a drive control signal. The drive control signal is sent to the actuator in the thrust vector adjustment module, i.e. the servo motor 44 and the magnetorheological damper 42 in the electric tilt joint 4, to drive the tilt thrust unit 2 to move. Through continuous target command input, real-time state feedback and control signal output, the system forms a complete closed-loop control loop, ensuring that the motion state of the tilt thrust unit 2 is consistent with the command of the upper flight control system.

[0036] The system supports the operation of the aircraft in multiple flight modes, mainly including vertical take-off and landing mode, horizontal flight mode and transition mode.

[0037] In the vertical take-off and landing mode, the thrust direction of the six groups of tilt thrust units 2 is perpendicular to the plane where the wing 1 is located, and the tilt angle is 90 degrees or in the interval close to 90 degrees, providing vertical lift for the aircraft.

[0038] In the horizontal flight mode, the thrust direction of the six groups of tilt thrust units 2 is parallel to the aircraft fuselage axis, and the tilt angle is 0 degrees or in the interval close to 0 degrees, providing forward thrust for the aircraft.

[0039] In the transition mode, the system controls three pairs of tilt thrust units 2 to convert between vertical take-off and landing mode and horizontal flight mode in batches, realizing the switching of flight state.

[0040] The specific structure and implementation principle of the physical load and power unit module, the thrust vector adjustment module, the state sensing and feedback module, and the coordination and synchronization control module are described in detail below.

[0041] Referring to Figs. 1 and 2, Figure 1 and Figure 2 The physical load and power unit module includes six groups of tilt thrust units 2 which are structurally identical and symmetrically installed on the wing 1. Each group of tilt thrust units 2 is composed of a turboprop engine 21 and a variable pitch propeller 3. The core function of this module is to provide all the required thrust for the aircraft in all flight stages.

[0042] The system adjusts the thrust size by controlling the pitch angle of the variable pitch propeller 3. The variable pitch propeller 3 includes blades and a pitch adjusting mechanism. The pitch adjusting mechanism is electrically connected with the synchronous controller 5, receives the control instruction sent by the synchronous controller 5, and drives the blades to rotate around the longitudinal axis of the blades to change the installation angle of the blades, i.e. the pitch angle.

[0043] The size of the thrust T generated by each group of tilting thrust units 2 is determined by the following aerodynamic formula:

[0044] T = C T · p · n 2 · D 4 ;

[0045] Wherein, T is the generated thrust; p is the atmospheric density at the current flight height; n is the rotation speed of the propeller; D is the diameter of the propeller; C T is the dimensionless thrust coefficient.

[0046] The thrust coefficient C T is not a constant value, and the value depends on the pitch angle β of the propeller and the advance ratio J of the aircraft. The advance ratio J is a dimensionless parameter representing the working state of the propeller, and is defined as:

[0047]

[0048] Wherein, V a is the forward flight speed of the aircraft.

[0049] In the specific flight process, the atmospheric density p is determined by the flight environment, and the diameter D of the propeller is a fixed value. The synchronous controller 5 can send instructions to the pitch adjusting mechanism to adjust the pitch angle β while adjusting the output power of the turboprop engine 21 to change the rotation speed n of the propeller according to the flight state and the control instruction. By actively changing the pitch angle β, the value of the thrust coefficient C T can be changed, so that the thrust T can be accurately and continuously adjusted under different advance ratio J working conditions, for example, J = 0 in hovering and J is a larger value in high-speed cruising.

[0050] Referring to the accompanying drawings Figure 2 and Figure 3 The core of the thrust vector adjusting module is six groups of electric tilting joints 4. Each group of electric tilting joints 4 is physically connected with a tilting thrust unit 2 and a wing 1, and the function is to drive the tilting thrust unit 2 to rotate as a whole according to the instruction of the synchronous controller 5, so as to change the direction of the thrust vector.

[0051] The internal structure of the electric tilting joint 4 includes a dual-axis hinge 41, a servo motor 44 and a magneto-rheological damper 42. The dual-axis hinge 41 is the mechanical motion base of the joint, which includes a main tilting axis 411 and an auxiliary fine-tuning axis 412. The main tilting axis 411 and the auxiliary fine-tuning axis 412 are arranged in a perpendicular geometric relationship with each other, and this arrangement mechanically decouples the two types of rotation. The main tilting axis 411 is used to achieve a large range of angle adjustment of the tilting thrust unit 2 between 0 degrees and 90 degrees, to complete the fundamental switching of the flight mode.

[0052] The driving of the main tilting axis 411 is completed by the servo motor 44. The servo motor 44 receives pulse or digital signals from the synchronization controller 5, and through the built-in reducer and transmission mechanism, it applies precise position control to the main tilting axis 411, enabling it to stably and accurately rotate to any target angle within the range of 0 degrees to 90 degrees.

[0053] The driving of the auxiliary fine-tuning axis 412 is completed by the magneto-rheological damper 42, which is used to achieve small-range, high-dynamic-response angle fine-tuning. The magneto-rheological damper 42 is a device filled with magneto-rheological fluid inside. The magneto-rheological fluid is a liquid suspended with micron-sized magnetic particles. In the absence of an external magnetic field, the liquid exhibits the characteristics of a low-viscosity Newtonian fluid. When the synchronization controller 5 applies a control current to the electromagnetic coil inside the magneto-rheological damper 42, the coil generates a controllable strength magnetic field.

[0054] Under the action of this magnetic field, the magnetic particles in the magneto-rheological fluid will rapidly arrange into a chain structure along the magnetic field lines, which hinders the flow of the liquid, causing its macroscopic viscosity and yield stress to change significantly and reversibly within milliseconds. This physical effect produces a precisely controllable damping torque directly related to the control current. This torque acts on the auxiliary fine-tuning axis 412 linked to the magneto-rheological damper 42, driving it to perform high-speed, accurate angle compensation within a range of plus or minus 5 degrees.

[0055] The output torque M mr of the magneto-rheological damper 42 can be expressed by the following model:

[0056]

[0057] where M mr is the total output torque of the magneto-rheological damper; M c (I c ) is the controlled torque, whose size is a function of the control current I c applied to the electromagnetic coil; is the viscous torque, whose size is the rotational angular velocity function; M0is the residual moment of the device, including viscous moment and friction moment under zero field. The synchronous controller 5 directly controls M c by adjusting the size of the control current I c (I c ), so as to realize the active and accurate torque exertion on the auxiliary fine adjustment shaft 412.

[0058] Referring to the accompanying drawings, Figure 2 and Figure 3 , the function of the state sensing and feedback module is to provide the real-time and accurate measurement data of the current state of the system for the cooperative and synchronous control module, which is the basis of the entire closed-loop control system.

[0059] The core component of this module is the angle sensor 43 arranged on each group of electrically driven tilting joints 4. Specifically, in order to comprehensively obtain the motion posture of the joint, the angle sensor 43 is arranged on the main tilting shaft 411 and the auxiliary fine adjustment shaft 412 respectively. This arrangement enables the synchronous controller 5 to simultaneously obtain the large-range tilting angle for flight mode switching and the small-range fine adjustment angle for active anti-disturbance.

[0060] In a specific embodiment, the angle sensors 43 can adopt absolute optical encoders or rotary transformers. Such sensors can provide absolute position information known upon power-on, without the need to search for a reference zero point, and have high resolution and high reliability. The measurement range of the sensor on the main tilting shaft 411 corresponds to a rotation of 0 degrees to 90 degrees, while the measurement range of the sensor on the auxiliary fine adjustment shaft 412 corresponds to a fine adjustment rotation of plus or minus 5 degrees or more.

[0061] Each angle sensor 43 converts the measured mechanical rotation angle into a digital electrical signal. This signal is sent to the synchronous controller 5 through a data bus, such as a controller area network (CAN) bus or an RS-485 bus. Therefore, the synchronous controller 5 can obtain the complete angle state of each group of electrically driven tilting joints 4 in real time, and the state information can be represented as a two-dimensional angle vector [θ main ,θ aux ].

[0062] Wherein, θ main is the real-time angle of the main tilting shaft 411; θ aux is the real-time angle of the auxiliary fine adjustment shaft 412.

[0063] In addition, the system also receives feedback signals from the body attitude sensor arranged on the body. The body attitude sensor can be an inertial measurement unit (IMU) which internally integrates a three-axis gyroscope and a three-axis accelerometer and can measure the roll, pitch and yaw three-axis attitude angle and angular velocity of the aircraft. The output signal of the body attitude sensor is sent to the synchronous controller 5, and especially when the crosswind interference compensation function is executed, the yaw attitude error signal provided by the sensor is the direct input basis for the controller to calculate the compensation moment.

[0064] Referring to the accompanying drawings Figure 5 The synchronous control module cooperates with the synchronous controller 5 as the core, and its function is to receive upper-layer instructions and lower-layer feedback, and generate accurate control instructions for the thrust vector adjustment module and the physical bearing and power unit module according to the preset control algorithm and strategy.

[0065] The primary function executed by the module is the synchronous control of the three pairs of tilting thrust units 2. In a specific operation process, the synchronous controller 5 receives a target tilting angle from the flight control system. At the same time, it collects the real-time tilting angle feedback from the angle sensor 43 on the main tilting shaft 411 of the six groups of electric tilting joints 4 through the data bus. The controller calculates the error between the real-time angle and the target angle of each group of tilting thrust units 2 respectively, and monitors the difference between the real-time angles of each pair of tilting thrust units 2, and controls it within the preset threshold of not more than 0.5 degrees.

[0066] In order to eliminate the angle error, the synchronous controller 5 independently runs a control algorithm for the servo motor 44 of each group. In a specific embodiment, the algorithm is a proportional, integral and differential PID control algorithm. The output control signal u(t) is determined by the following formula:

[0067]

[0068] Wherein, u(t) is the control signal generated at time t to drive the servo motor 44; e(t) is the angle error at this time; K p is the proportional gain coefficient; K i is the integral gain coefficient; K d is the differential gain coefficient. The controller calculates the proportional term, integral term and differential term of the angle error according to the algorithm to generate the final driving signal until the error converges to the allowed range.

[0069] The module also executes the pitch and tilting cooperative control strategy. The synchronous controller 5 internally stores or calculates a certain cooperative control law according to a function, which establishes a functional relationship between the tilting angle θ main of the main tilting shaft 411 and the pitch angle β of the variable pitch propeller 3, that is, β=f(θ main). When the system performs the flight mode conversion, the synchronization controller 5 controls the tilt angle θ main According to the function relationship, the corresponding pitch angle β at each moment is calculated synchronously while continuously changing from one value to another, and the pitch adjustment mechanism is instructed. For example, in the process of converting from the vertical take-off mode to the horizontal flight mode, under the batch conversion technique, θ main Smoothly decreases from 90 degrees to 0 degrees, and the controller instructs β to synchronously smoothly decrease from 28 degrees to 10 degrees.

[0070] In a specific embodiment, the batch conversion technique is mainly executed by the synchronization controller 5 in the cooperative and synchronization control module, and the specific implementation is as follows:

[0071] Grouping strategy: The physical bearing and power unit module includes six groups of tilt thrust units 2, which are divided into three pairs, and the two groups in each pair are distributed symmetrically left and right along the wing 1. The three pairs of tilt thrust units can be defined as the first pair (for example, the left and right groups closest to the fuselage), the second pair (the left and right groups in the middle), and the third pair (the left and right groups closest to the wing tip).

[0072] Conversion trigger: The start of the batch conversion is triggered by the synchronization controller 5 according to the preset time or attitude condition. For example, when the aircraft reaches a predetermined transition airspeed or a specific pitch attitude, the synchronization controller 5 starts the conversion program.

[0073] Batch conversion steps:

[0074] First batch conversion: The synchronization controller 5 first drives the first pair of tilt thrust units 2 to perform tilt angle conversion, for example, smoothly rotating from 90 degrees for vertical take-off to 0 degrees for horizontal flight. In this process, the second pair and the third pair of tilt thrust units 2 remain at 90 degrees and continue to provide vertical lift for the aircraft to maintain stability and lift.

[0075] Transition phase: When the tilt angle of the first pair of tilt thrust units 2 reaches a certain preset intermediate angle (for example, 45 degrees) or after a preset time interval, the synchronization controller 5 continues to the next step.

[0076] Second batch conversion: The synchronization controller 5 starts the tilt conversion of the second pair of tilt thrust units 2. At this time, the first pair of tilt thrust units 2 continues to rotate to 0 degrees, and the third pair of tilt thrust units 2 remains at 90 degrees.

[0077] Third batch conversion: When the tilt angles of the first pair and the second pair of tilt thrust units 2 both reach a certain preset condition, the synchronization controller 5 starts the tilt conversion of the third pair of tilt thrust units 2.

[0078] Complete the transition: The entire process continues until all three pairs of tilt thrust units 2 have completed the transition from 90 degrees to 0 degrees, and the aircraft has fully entered level flight mode.

[0079] Coordinated Control: During each stage of the batch conversion described above, the synchronous controller 5 consistently executes a coordinated control strategy for pitch and tilt. For thrust units undergoing tilting, their pitch angle is adjusted synchronously and smoothly as the tilt angle changes to ensure the continuity and efficiency of thrust output. For thrust units maintaining a 90-degree position, their pitch angle remains within the range of maximum thrust settings that provide vertical lift.

[0080] See attached document Figure 6 This module also performs crosswind interference compensation. When the aircraft encounters crosswind disturbances, the fuselage attitude sensors will detect the yaw attitude error signal ψ. error The signal is sent to the synchronization controller 5. Based on this signal and the aircraft's dynamic model, the controller calculates a compensation torque M required to counteract the disturbance. comp Subsequently, based on the previously established torque-current characteristic model of the magnetorheological damper 42, the controller applies the compensation torque M. comp Inverse calculation yields the control current command I applied to the six sets of magnetorheological dampers 42. c The current command drives the auxiliary fine-tuning shaft 412 to produce a precise angular deflection, causing the six thrust vectors to generate a lateral component. The torque formed by this component is equal in magnitude and opposite in direction to the crosswind interference torque, thereby maintaining the stability of the aircraft's yaw attitude.

[0081] To further clarify the collaborative working process of the technical solution of this invention, specific working scenario examples will be used for illustration below.

[0082] Example 1:

[0083] Crosswind interference compensation in vertical takeoff and landing mode:

[0084] This embodiment describes how the system actively compensates for crosswind interference when the aircraft is in a vertical takeoff and landing mode, such as hovering. In this mode, the main tilt axis 411 of the six tilt thrust units 2 is at an angle θ. main Controlled and stabilized at a 90-degree position by servo motor 44, it provides vertical lift to the aircraft.

[0085] When a crosswind acts on the fuselage, the fuselage attitude sensor inertial measurement unit detects a non-commanded yaw attitude angular velocity or attitude angle deviation and generates a yaw attitude error signal ψ. error The signal is transmitted to the synchronization controller 5.

[0086] Synchronous controller 5 receives ψ errorAfter that, according to the dynamics model of the aircraft and the current flight state, the required compensation moment M comp .

[0087] Subsequently, the synchronous controller 5 converts the calculated compensation moment M comp into precise control current commands I c for the six groups of magnetorheological dampers 42 according to the torque-current mathematical model of the magnetorheological dampers 42.

[0088] Under the action of the control current I c , the viscosity of the magnetorheological fluid inside the six groups of magnetorheological dampers 42 changes, generating a controlled damping torque. This torque drives the six groups of auxiliary fine-tuning shafts 412 to synchronously rotate a small angle. Since the main tilting shafts 411 are at 90 degrees at this time, the rotation of the auxiliary fine-tuning shafts 412 will cause the originally vertically upward thrust vector to have a horizontal component.

[0089] This horizontal thrust component generates a lateral action force and control moment equal in size and opposite in direction to the side wind interference moment, thereby actively counteracting the influence of the side wind and stabilizing the yaw attitude of the aircraft, maintaining its spatial position. The entire process is closed-loop, and the controller continuously monitors ψ error until it converges to the vicinity of the preset zero value.

[0090] Example Two:

[0091] Transition mode from vertical takeoff to horizontal flight:

[0092] This embodiment describes the complete process of smoothly transitioning the aircraft from the hovering state after vertical takeoff to horizontal cruise flight.

[0093] This process is triggered by a mode conversion instruction issued by the flight control system. After receiving the instruction, the synchronous controller 5 starts the preset pitch and tilt coordination control strategy and batch conversion program.

[0094] The specific steps of batch conversion are as follows:

[0095] First batch conversion: the synchronous controller 5 first sends instructions to the servo motors 44 of the first pair (e.g., the left and right groups closest to the fuselage) of tilt thrust units 2, causing them to drive the main tilting shafts 411 to continuously and smoothly decrease from 90 degrees to 0 degrees at a preset angular velocity. During this process, the main tilting shafts 411 of the second and third pairs of tilt thrust units 2 remain unchanged at 90 degrees.

[0096] Second batch conversion: When the angle of the first pair of tilting thrust units 2 reaches a certain preset intermediate angle (for example, 45 degrees), or after a preset time interval, the synchronous controller 5 starts the second pair of tilting thrust units 2 to perform the tilting conversion. At this time, the first pair continues to rotate to 0 degrees, and the third pair remains at the 90-degree position.

[0097] Third batch conversion: When the angles of the first pair and the second pair of tilting thrust units 2 both meet the preset conditions, the synchronous controller 5 starts the third pair of tilting thrust units 2 to perform the tilting conversion.

[0098] The cooperative control is executed in parallel throughout the process:

[0099] In each stage of the batch conversion, the synchronous controller 5 independently runs the synchronous control algorithm of the six sets of main tilting shafts 411. It continuously collects the feedback values of the six sets of angle sensors 43 and uses the proportional, integral, and derivative (PID) control algorithm to generate independent control signals to ensure that the absolute value of the difference between the real-time tilting angles of each pair of tilting thrust units is always less than 0.5 degrees, ensuring the attitude balance during flight.

[0100] At the same time, for the tilting thrust units that are being tilted, the synchronous controller 5 calculates the target value of the propeller pitch angle corresponding to the current tilting angle in real time according to the cooperative control law. For example, when the main tilting shaft 411 angle decreases from 90 degrees to 0 degrees, the pitch angle is simultaneously reduced from the high thrust setting value (for example, 28 degrees) to the high-efficiency cruise setting value (for example, 10 degrees). The controller sends the calculated target pitch angle command to the pitch adjustment mechanism of the variable pitch propeller 3. For the thrust units that remain at the 90-degree position, their pitch angles will remain in the high thrust setting value range.

[0101] The process continues until the angles of the six sets of tilting thrust units reach 0 degrees or the 0-5-degree range corresponding to high-efficiency cruise, and the pitch angles of all six sets of propellers also reach their high-efficiency cruise setting values. At this point, the aircraft completes the entire transition, the thrust vector is completely converted from vertical to horizontal, and the flight state is smoothly converted from lift support to aerodynamic lift support.

Claims

1. A fixed-wing microlight aircraft tilting compound thrust system, characterized in that, Comprise: A physical bearing and power unit module, comprising a wing (1) and six groups of tilting thrust units (2) symmetrically arranged on the wing (1), the six groups of tilting thrust units (2) are divided into three pairs, each group of tilting thrust units (2) has a turboprop engine (21) and a variable pitch propeller (3) for generating thrust; A thrust vector adjustment module, comprising six groups of electric tilting joints (4), each group of electric tilting joints (4) is connected to a corresponding group of tilting thrust units (2) for driving the tilting thrust units (2) to rotate as a whole to adjust the direction of the thrust; A state sensing and feedback module, comprising an angle sensor (43) arranged on each group of electric tilting joints (4) for real-time monitoring of the tilting angle of each group of electric tilting joints (4); A cooperative and synchronous control module, comprising a synchronous controller (5), the synchronous controller (5) is electrically connected with the angle sensor (43) and the electric tilting joint (4), for controlling three pairs of electric tilting joints (4) to rotate synchronously according to the tilting angle.

2. A fixed-wing microlight aircraft tilting compound thrust system according to claim 1, characterized in that, The physical bearing and power unit module for adjusting the size of the thrust step comprises: The synchronous controller (5) sends instructions to the pitch adjustment mechanism of the variable pitch propeller (3) according to the current flight mode; The pitch adjustment mechanism drives the rotation of the blades of the variable pitch propeller (3) to change the pitch angle; The size of the thrust is determined by the thrust coefficient, which is a function of the pitch angle and the forward ratio of the aircraft, by changing the pitch angle to optimize the thrust coefficient in different flight modes.

3. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 1, wherein, The structure of each group of electric tilting joints (4) in the thrust vector adjustment module comprises: A double-axis hinge (41) having a main tilting axis (411) and an auxiliary fine adjustment axis (412), the main tilting axis (411) is used to realize large range tilting adjustment of 0-90 degrees, and the auxiliary fine adjustment axis (412) is arranged perpendicular to the main tilting axis (411); A servo motor (44) for driving the rotation of the main tilting axis (411); A magneto-rheological damper (42) linked with the auxiliary fine adjustment axis (412) for driving the auxiliary fine adjustment axis (412) to make small range angle fine adjustment.

4. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 3, wherein, In the state sensing and feedback module, the angle sensor (43) is arranged in the following way: Respectively arranging the angle sensor (43) on the main tilting axis (411) and the auxiliary fine adjustment axis (412) for measuring the tilting angle and the fine adjustment angle; Both of the angle sensors (43) are electrically connected with the synchronous controller (5).

5. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 1, wherein, The method for controlling three pairs of electric tilting joints (4) to rotate synchronously in the cooperative and synchronous control module comprises: The synchronous controller (5) receives the target tilting angle from the flight control system; The synchronous controller (5) collects the real-time tilting angles fed back by the six groups of angle sensors (43), and calculates the angle error between each group of real-time tilting angles and the target tilting angle respectively; The synchronization controller (5) generates a control signal based on the angle error, independently drives six groups of the electric tilting joints (4), and controls the absolute value of the difference between the real-time tilting angles of each pair of the electric tilting joints (4) to be not greater than 0.5 degrees.

6. A fixed-wing ultralight aircraft tilting compound thrust system according to claim 1, wherein, The cooperative and synchronous control module is further used to perform a pitch and tilting cooperative control strategy, and the pitch and tilting cooperative control strategy comprises: In the vertical take-off and landing mode, the synchronization controller (5) controls the tilting angle of the electric tilting joint (4) to be in the interval of 85 degrees to 90 degrees, and instructs the pitch angle of the variable pitch propeller (3) to be in the large thrust setting value interval; In the horizontal flight mode, the synchronization controller (5) controls the tilting angle of the electric tilting joint (4) to be in the interval of 0 degrees to 5 degrees, and instructs the pitch angle of the variable pitch propeller (3) to be in the high-efficiency cruise setting value interval; In the transition mode, the synchronization controller (5) controls the tilting angle to be continuously changed in batches from 90 degrees to 0 degrees, while instructing the pitch angle to be continuously changed from the large thrust setting value to the high-efficiency cruise setting value.

7. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 3, wherein, The cooperative and synchronous control module is further used to perform side wind interference compensation, and the steps are: The synchronization controller (5) receives a yaw attitude error signal caused by side wind interference and fed back by a fuselage attitude sensor; The synchronization controller (5) calculates a compensation torque for compensation according to the yaw attitude error signal; The synchronization controller (5) converts the compensation torque into a control current instruction applied to the magnetorheological damper (42); The magnetorheological damper (42) generates a controllable damping force under the action of the control current, and drives the auxiliary fine adjustment shaft (412) to perform angle compensation within a range of plus or minus 5 degrees.

8. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 6, wherein, The batchwise continuous change in the transition mode comprises: The synchronization controller (5) triggers at a preset time or attitude condition, and drives three pairs of the tilting thrust units (2) to change the tilting angle in batches until all the tilting thrust units (2) complete the change from 90 degrees to 0 degrees.

9. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 6, wherein, The setting value interval of the pitch angle further comprises: The large thrust setting value interval is 25 degrees to 30 degrees; The high-efficiency cruise setting value interval is 5 degrees to 15 degrees.

10. The fixed-wing ultralight aircraft tilting compound thrust system according to claim 5, wherein, The method for the synchronization controller (5) to generate a control signal based on the angle error comprises: A proportional, integral, and differential control algorithm is adopted, the angle error is taken as an input, and the control signal for driving the servo motor (44) in the electric tilting joint (4) is generated according to the proportional term, integral term, and differential term of the angle error.

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