A control method for a pterocarpa-inspired monoplane aircraft

Abstract

The application discloses a control method of a winged fruit imitating single-wing aircraft, and belongs to the technical field of flight control of bionic aircraft. First, the application establishes an aerodynamic force model of the winged fruit imitating single-wing aircraft based on the blade element theory, and constructs a six-degree-of-freedom flight dynamics equation in combination with a fuselage resistance model and a propeller thrust model; on the basis, state variables of the aircraft are defined and a state space model is established; subsequently, a PI controller and a PD controller are respectively designed for vertical direction height control and horizontal direction position control of the aircraft; finally, through periodic flap deflection control, the horizontal control quantity is converted into a periodic control signal related to the phase of aircraft autorotation, so that the adjustment and control of the flight trajectory of the aircraft are realized.

Description

Technical Field

[0001] This invention belongs to the field of biomimetic aircraft flight control technology, specifically relating to a control method for a winged monoplane aircraft. Background Technology

[0002] With the development of biomimetic aircraft technology, utilizing the flight mechanisms of plant seeds in natural dispersal processes for aircraft design has gradually become an important direction in aeronautical engineering research. Among these, the flight mechanism of maple seeds has provided important inspiration for the design of micro-aircraft. Researchers have proposed the concept of a winged fruit-inspired monoplane, which generates aerodynamic force through the rotation of a single wing structure around an axis, thereby achieving flight.

[0003] In their paper "Longitudinal flight dynamics modeling and a flightstability analysis of a monocopter," Tong et al. theoretically analyzed the motion characteristics of an aircraft by establishing a dynamic model, but their research mainly focused on aircraft stability and paid less attention to flight control methods. Cai et al., in their paper "Design and Optimization of a Samara-Inspired Lightweight Monocopter for Extended Endurance," proposed a winged monocopter structure. They improved flight efficiency by optimizing the wing shape and mass distribution, and experimentally verified the feasibility of the winged structure in micro-aircraft. However, this research mainly focused on aircraft structural design and aerodynamic performance optimization, with insufficient attention to aircraft flight control issues.

[0004] Existing research indicates that winged monoplanes exhibit significant spin motion during flight, with their aerodynamic forces displaying distinct periodic variations within the spin period. Traditional flight control methods are difficult to apply directly to this type of aircraft. Therefore, a novel control method for winged monoplanes is proposed to improve flight stability. Summary of the Invention

[0005] This invention proposes a control method for a winged fruit monoplane to address the problems of poor control and flight stability in existing winged fruit monoplanes. Addressing the characteristics of winged fruit monoplanes, which exhibit spin motion and periodic aerodynamic changes, this invention proposes a periodic control method based on spin phase to achieve stable control of the aircraft's altitude and horizontal position. Compared with traditional control methods, this invention features a simpler control structure, better stability, and can effectively improve the flight stability and trajectory tracking accuracy of winged fruit monoplanes.

[0006] The technical solution for achieving the present invention is: a control method for a winged monoplane aircraft, comprising the following steps:

[0007] Step 1: Establish a mathematical model of the winged monoplane aircraft system, then proceed to Step 2.

[0008] Step 2: Based on the mathematical model of the winged monoplane aircraft system, design the horizontal motion controller and the vertical motion controller, and proceed to Step 3.

[0009] Step 3: Establish the control model of the aircraft and set waypoint tasks to verify the flight control effect of the aircraft.

[0010] Compared with the prior art, the significant advantages of this invention are:

[0011] (1) A complete dynamic model applicable to winged monoplane aircraft was established, providing a theoretical basis for flight control design;

[0012] (2) A periodic flap control method based on spin phase is proposed, which can effectively utilize the spin-periodic aerodynamic characteristics of the aircraft.

[0013] (3) The altitude and horizontal position are decoupled by the PI-PD composite control structure, which improves flight stability and control accuracy;

[0014] (4) The control method has a simple structure and is easy to implement in the flight control system of small biomimetic aircraft. Attached Figure Description

[0015] Figure 1 This is a schematic diagram illustrating the principle of the control method for the winged monoplane aircraft of the present invention.

[0016] Figure 2 This is a schematic diagram of a winged monoplane.

[0017] Figure 3 The tracking process in the x-axis direction is the system output of the winged monoplane control method designed in this invention.

[0018] Figure 4 The tracking process in the y-axis direction is the system output of the winged monoplane control method designed in this invention.

[0019] Figure 5 The tracking process in the z-axis direction is the system output of the winged monoplane control method designed in this invention. Detailed Implementation

[0020] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0021] Combination Figure 1and Figure 2 The present invention provides a control method for a winged monoplane aircraft, comprising the following steps:

[0022] Step 1: Establish the mathematical model of the winged monoplane flight system, as follows:

[0023] Step 1-1: Based on blade element theory, divide the wing into multiple blade element units for aerodynamic analysis: Let the total length of the wing from the axis of rotation be l, the chord length of the blade element at the center r of the axis of rotation be c(r), and the width be dr. Let the wing rotate around the z-axis of the fuselage with an angular velocity... The aircraft rotates at a uniform angular velocity while the entire spacecraft experiences a descent velocity. Analyzing the blade element at point r, the relative airflow at that point is decomposed into tangential and axial velocity components, where the directions are consistent with the spin tangential direction. Furthermore:

[0024] (1),

[0025] in, The additional velocity generated by the induced flow, The velocity component along the axial direction, This represents the falling speed.

[0026] The relative airflow velocity at the blade unit is:

[0027] (2),

[0028] in, Let r be the relative airflow velocity at the blade unit.

[0029] Let the wing's installed geometric angle of attack be... The inflow angle induced by the flow at that cross section for:

[0030] (3),

[0031] Thus the effective angle of attack of the blades for:

[0032] (4),

[0033] in, For the flap deflection angle introduced later, Install geometric angle of attack on the wing, denoted as the inflow angle, r as the distance from the blade element to the center of the rotation axis, and t as time.

[0034] According to blade element theory, the lift of the blade element at radius r... and resistance They are respectively:

[0035] (5),

[0036] (6),

[0037] in, air density, , are the lift and drag coefficients, respectively, and t represents time.

[0038] In the body coordinate system, assuming lift is perpendicular to the relative airflow direction and drag is parallel to the relative airflow direction, then the force components in the x-axis direction are... Force components along the z-axis It can be written as:

[0039] (7),

[0040] (8),

[0041] The total aerodynamic force of the wing on the aircraft for:

[0042] (10),

[0043] in, Let x be the component of the total aerodynamic force of the wing in the x-direction. Let z be the component of the total aerodynamic force of the wing in the z-direction.

[0044] Based on the established aerodynamic distribution, for the airfoil element at spanwise position r, its position vector... for:

[0045] (11),

[0046] in, Let r be the position of the wing element in the x-direction. Let r be the position of the wing element in the y direction.

[0047] Aerodynamic vector for:

[0048] (12),

[0049] Aerodynamic torque of blade element unit as follows:

[0050] (13),

[0051] The pitching moment of the leaf element unit at the center of mass is calculated according to equations (10) to (13). Rolling torque and yaw moment :

[0052] (14),

[0053] (15),

[0054] (16),

[0055] Since the wing is basically in the plane of rotation ,and and Approaching 0, therefore:

[0056] (17),

[0057] (18),

[0058] (19),

[0059] in, The total rolling moment acting on the wing. The total pitching moment acting on the wing. This is the total yawing moment acting on the wing.

[0060] During its spin-descent descent, the fuselage of the winged monoplane acts like a slender, blunt body, generating air resistance through its interaction with the incoming airflow. The air resistance experienced by the fuselage can be modeled using a conventional drag model. Represented as:

[0061] (20),

[0062] in, air density; This refers to the fuselage drag coefficient; This refers to the equivalent frontal area of ​​the fuselage in the direction of the incoming flow. This represents the speed of the aircraft relative to the air.

[0063] For small-scale winged monoplanes, the flight speed is low, and the Reynolds number is around 10. 3 ~10 5 Within the range; at this point, the fuselage drag coefficient is related to the Reynolds number Re, and can be written in general form:

[0064] (twenty one),

[0065] (twenty two),

[0066] in, The characteristic length of the fuselage. It is the aerodynamic viscosity coefficient.

[0067] In actual calculations, an approximate constant can be obtained by consulting tables or using empirical formulas based on the fuselage shape and Reynolds number range. Thus we get:

[0068] (twenty three),

[0069] (twenty four),

[0070] in, Indicates an intermediate variable.

[0071] In the fuselage coordinate system, the relative airflow velocity vector of the fuselage is denoted as... :

[0072] (25),

[0073] in, This represents the component of the fuselage's relative airflow velocity along the x-axis. This represents the component of the fuselage's relative airflow velocity along the y-axis. This represents the component of the relative airflow velocity of the fuselage in the z-axis direction.

[0074] Air resistance of the fuselage Since the direction is opposite to the velocity direction, it can be equivalently decomposed onto three coordinate axes:

[0075] (26),

[0076] in , , Let be the equivalent drag coefficients along the x, y, and z directions, respectively. They are derived from the drag coefficients. The equivalent frontal area distribution of the fuselage in all directions is determined by the formula above; The form is designed to ensure that the resistance is always opposite to the direction of the velocity, while retaining the characteristics of secondary resistance; This represents the component of the air resistance along the x-axis. This represents the component of the air resistance along the y-axis. This represents the component of the air resistance along the z-axis.

[0077] Under the typical condition of spin fall, the vertical velocity component Much greater than the horizontal component, the dominant effect of fuselage drag is in the vertical direction:

[0078] (27),

[0079] and , The relatively small damping can be considered as linear damping of horizontal motion, which helps to suppress horizontal velocity divergence and improve control performance.

[0080] Since the fuselage's center of mass is close to the overall center of mass of the aircraft and its geometry is basically symmetric about its spin axis, the air resistance of the fuselage with respect to the aerodynamic moments about the x and y axes can be ignored; for the spin about the z axis, only an equivalent damping moment proportional to the angular velocity needs to be considered. :

[0081] (28),

[0082] in The equivalent spin damping coefficient of the fuselage.

[0083] Small propellers are placed near the wingtips to provide additional thrust on top of the spin descent, thereby adjusting the aircraft's spin speed and vertical motion characteristics.

[0084] Let the unit vector of the propeller axis direction in the body coordinate system be... The propeller thrust is Then its force vector in the body coordinate system for:

[0085] (29),

[0086] According to propeller aerodynamics, thrust An approximation using a thrust coefficient model is employed:

[0087] (30),

[0088] in, The area of ​​the propeller disk; The radius of the propeller; The propeller angular velocity; Thrust coefficient; vector of wingtip position relative to the center of mass. Recorded as:

[0089] (31),

[0090] in, , , These represent the distances from the wingtip to the center of mass in the x, y, and z directions, respectively.

[0091] The aerodynamic torque of the propeller on the aircraft for:

[0092] (32),

[0093] Steps 1-2: To facilitate controller design, define state variables as follows:

[0094] Define the state variable: aircraft position coordinates Describes the spatial coordinates of the aircraft's center of mass in the ground coordinate system; and its velocity in the ground coordinate system. velocity in the body coordinate system And satisfy Angular velocity in body coordinate system R0 represents the transformation matrix, u represents the velocity of the aircraft in the x-axis direction in the body coordinate system, v represents the velocity of the aircraft in the y-axis direction in the body coordinate system, w represents the velocity of the aircraft in the z-axis direction in the body coordinate system, p represents the attitude angular velocity of the aircraft in the x-axis direction in the body coordinate system, q represents the attitude angular velocity of the aircraft in the y-axis direction in the body coordinate system, s represents the attitude angular velocity of the aircraft in the z-axis direction in the body coordinate system, and T represents the transpose.

[0095] Steps 1-3: Based on the defined state variables, the established mathematical model of the winged monoplane flight system is transformed into six-degree-of-freedom flight dynamics equations, as follows:

[0096] According to Newton's second law and Euler's law of rigid body rotation, the translation equations and attitude dynamics equations in the body coordinate system are as follows:

[0097] (33),

[0098] in, yes The first derivative, yes The first derivative of , where m is the mass of the aircraft and I is the inertial matrix of the aircraft about its center of mass. Considering that the aircraft is a rigid body, then , This represents the moment of inertia about the x-axis; This represents the moment of inertia about the y-axis; This represents the moment of inertia about the z-axis. It is the product of inertia representing the coupling effect between the x-axis and z-axis. It is the product of inertia representing the coupling effect between the x-axis and y-axis. It is the product of inertia representing the coupling effect between the y-axis and the x-axis. It is the product of inertia representing the coupling effect between the y-axis and z-axis. It is the product of inertia representing the coupling effect between the z-axis and x-axis. It is the product of inertia representing the coupling effect between the z-axis and y-axis; The resultant torque acting on the aircraft;

[0099] The net force F acting on the aircraft consists of the following components:

[0100] (34),

[0101] in, For the overall aerodynamic force of the wing; For air resistance of the fuselage; For the thrust of the wingtip propeller; This represents gravity in the body coordinate system.

[0102] Gravity in the body coordinate system as follows:

[0103] (35),

[0104] in, The pitch angle of the aircraft. Let m be the roll angle of the aircraft, m be the mass of the aircraft, and g be the gravitational acceleration, taken as 9.8 N / kg.

[0105] Substituting equations (17)~(19), (27), (30), (33), and (36) into equation (34), we obtain the following six-degree-of-freedom flight dynamics equations:

[0106]

[0107] in, This represents the moment of inertia about the x-axis; This represents the moment of inertia about the y-axis; This represents the moment of inertia about the z-axis. It is the product of inertia representing the coupling effect between the x-axis and z-axis. It is the first derivative of u. It is the first derivative of v. It is the first derivative of w. Let x be the component of the total aerodynamic force of the wing in the x-direction. Let be the component of the total aerodynamic force of the wing in the z-direction. This represents the component of the air resistance along the x-axis. This represents the component of the air resistance along the y-axis. This represents the component of the air resistance along the z-axis. denoted as wingtip propeller thrust, and l as the total length from the wingtip to the axis of rotation. The total rolling moment acting on the wing. This is the total pitching moment acting on the wing.

[0108] Proceed to step 2.

[0109] Step 2: Based on the mathematical model of the winged monoplane aircraft system, design the horizontal motion controller and the vertical motion controller, as follows:

[0110] Step 2-1: Define vertical motion error ,in, For the desired height, This refers to the aircraft's current altitude; the system status is monitored under the control of the designed controller. The desired altitude command should be tracked as accurately as possible. Design the following PI controller to adjust the control quantity in the vertical direction. :

[0111] (37),

[0112] in, and These are proportional gain and integral gain, respectively. The proportional term is used to quickly respond to system errors, while the integral term is used to eliminate steady-state errors and improve the long-term tracking accuracy of the system. In order to improve the actual control effect of the controller, the gain parameter is tuned in the control law to meet the altitude control requirements under different flight conditions.

[0113] Step 2-2: Define horizontal motion error ,in, All are target horizontal positions. This represents the current actual horizontal position of the aircraft; to facilitate the monitoring of the system state under the control of the designed controller. The desired horizontal position command should be tracked as accurately as possible. .

[0114] Design the following PD controller. :

[0115] (38),

[0116] in, These are proportional gain and derivative gain, respectively. The proportional term is used to quickly correct position errors, while the derivative term is used to improve the responsiveness to error change trends, effectively suppressing overshoot and shortening the settling time.

[0117] In steps 2-3, the control quantity obtained in the horizontal motion controller The input will be used for horizontal motion control in the periodic control system; a continuous sine wave function is used for periodic modulation control of the flap deflection angle, resulting in higher smoothness and control accuracy; the innovative periodic control system design is as follows:

[0118] In order to make the aircraft from its current position Move to the target point Horizontal control quantity ,in, This is the control variable in the x-direction. The control quantity in the y-direction is converted into a periodic signal phase control quantity. :

[0119] (39),

[0120] Let the current spin angle of the spacecraft be... A sinusoidal periodic modulation model of the flap deflection angle controlled by a periodic control system. for:

[0121] (40),

[0122] in, This refers to the flap period offset angle; The deflection angle amplitude controls the intensity of the periodic disturbance; To control the directional phase offset, the positional distribution of the deflection waveform within the period is determined; the phase of the periodic signal... It controls the projection position of the periodic deflection waveform during the rotation period around the axis; by adjusting This increases or decreases the flap deflection angle near a certain azimuth angle in the cycle, thereby guiding the asymmetric aerodynamic force to act in the desired direction and achieving target tracking.

[0123] Proceed to step 3.

[0124] Step 3: The winged monoplane aircraft completes the waypoint tracking task under the control method of this invention: four waypoints are added in space as the target flight positions of the aircraft. The aircraft starts from the starting point, passes through three of the waypoints in sequence, and then returns to the starting point. The feasibility of the control method is demonstrated by the aircraft's performance in completing the waypoint tracking task.

[0125] Example

[0126] To evaluate the performance of the designed controller, the physical parameters of the winged monoplane aircraft system in the simulation are shown in Table 1:

[0127]

[0128] The desired instruction of the given system is for the aircraft to fly along the expected waypoints. Four waypoints are added in space as the target flight positions of the aircraft. The coordinates of the waypoints are (0, 0, 0), (3, 0, 0), (3, 3, 1.5) and (0, 3, 1.5). The aircraft starts from (0, 0, 0), passes through the three points in sequence, and returns to the starting position.

[0129] The tracking performance of the winged monoplane in the x, y, and z directions during waypoint tracking missions is as follows: Figure 3 , Figure 4 and Figure 5 As shown, X, Y, and Z represent the actual positions of the aircraft in the x, y, and z directions, respectively. , , Let x represent the target position of the aircraft in the x, y, and z directions. Waypoint tracking simulation results show that the designed control system enables the monoplane to sequentially complete waypoint flight tasks from (0, 0, 0) → (3, 0, 0) → (3, 3, 1.5) → (0, 3, 1.5) → (0, 0, 0) in a spin-flying state. In the x and y horizontal directions, the aircraft exhibits a fast response to step position commands, with a rise time of approximately 5–7 s, a peak overshoot of approximately 27%–33%, and reaches steady state within 15–20 s, with errors controlled within ±0.15 m and ±0.2 m, respectively. Due to the influence of spin aerodynamic coupling, the horizontal displacement exhibits certain periodic oscillations, but the oscillation amplitude decays over time, without cumulative drift or divergence. In the vertical direction, the peak overshoot during the ascent from 0 m to 1.5 m was approximately 0.4 m, and the aircraft hovered near the target altitude for about 10 seconds with a fluctuation amplitude of less than ±0.1 m. During the multi-axis combined maneuver, no significant instability was observed in the vertical altitude, indicating that the completed periodic control strategy can achieve three-dimensional waypoint tracking.

Claims

1. A control method for a winged monoplane aircraft, characterized in that, Includes the following steps: Step 1: Establish a mathematical model of the winged monoplane flight system, then proceed to Step 2; Step 2: Based on the mathematical model of the winged monoplane aircraft system, design the horizontal motion controller and the vertical motion controller, and proceed to Step 3; Step 3: Establish the control model of the aircraft and set waypoint tasks to verify the flight control effect of the aircraft.

2. The control method for the winged monoplane aircraft according to claim 1, characterized in that, In step 1, a mathematical model of the winged monoplane flight system is established, as follows: Step 1-1: As a type of monoplane aircraft, the winged nut monoplane system relies on the aerodynamic forces generated during its spin. An electric motor-driven propeller is mounted at the wingtip to provide power for the spin. The wing of the winged nut monoplane consists of a main wing and flaps. A servo motor drives the flaps to achieve periodic oscillation, resulting in uneven aerodynamic force distribution during the spin cycle and generating horizontal force. Based on the dynamic characteristics of the winged nut monoplane's flight process, a mathematical model of the winged nut monoplane system is established. Steps 1-2: To facilitate controller design, define state variables and convert the established mathematical model of the winged monoplane aircraft system into state-space equations.

3. The control method for the winged monoplane aircraft according to claim 2, characterized in that, Step 1-1, as follows: Based on blade element theory, the wing is divided into multiple blade element units for aerodynamic analysis: Let the total length of the wing tip from the axis of rotation be l, the chord length of the blade element at the center of the wing's rotation axis be c(r), and the width be dr. The wing rotates around the z-axis of the fuselage at an angular velocity... The aircraft rotates at a uniform angular velocity while the entire spacecraft experiences a descent velocity. Analyzing the blade element at point r, the relative airflow at that point is decomposed into tangential and axial velocity components, where the directions are consistent with the spin tangential direction. Furthermore: (1)， in, The additional velocity generated by the induced flow, The velocity component along the axial direction, The falling speed; The relative airflow velocity at blade unit r for: (2)， Let the wing's installed geometric angle of attack be... The inflow angle induced by the flow at that cross section for: (3)， Thus the effective angle of attack of the blades for: (4)， in, For the flap deflection angle introduced later, Install geometric angle of attack on the wing, denoted as the inflow angle, r as the distance from the blade element to the center of the rotation axis, and t as time. According to blade element theory, the lift of the blade element at radius r... and resistance They are respectively: (5)， (6)， in, air density, , These are the lift and drag coefficients, respectively, and t represents time. In the body coordinate system, assuming lift is perpendicular to the relative airflow direction and drag is parallel to the relative airflow direction, then the force components in the x-axis direction are... Force components along the z-axis It can be written as: (7)， (8)， The total aerodynamic force of the wing on the aircraft for: (10)， in, Let x be the component of the total aerodynamic force of the wing in the x-direction. Let z be the component of the total aerodynamic force of the wing in the z-direction; Based on the established aerodynamic distribution, for the airfoil element at spanwise position r, its position vector... for: (11)， in, Let r be the position of the wing element in the x-direction. Let r be the position of the wing element in the y direction; Aerodynamic vector for: (12)， in, Force component in the y-axis direction; Aerodynamic torque of blade element unit as follows: (13)， The pitching moment of the leaf element unit at the center of mass is calculated according to equations (10) to (13). Rolling torque and yaw moment : (14)， (15)， (16)， Since the wing is basically in the plane of rotation ,and and Approaching 0, therefore: (17)， (18)， (19)， in, The total rolling moment acting on the wing. The total pitching moment acting on the wing. The total yawing moment acting on the wing; During the spin-descent process, the fuselage of the winged monoplane acts as a slender, blunt body, generating air resistance through its interaction with the incoming airflow; the air resistance experienced by the fuselage can be modeled using a conventional drag model. Represented as: (20)， in, air density; This refers to the fuselage drag coefficient; This refers to the equivalent frontal area of ​​the fuselage in the direction of the incoming flow. This refers to the speed of the fuselage relative to the air. For small-scale winged monoplanes, the flight speed is low, and the Reynolds number is around 10. 3 ~10 5 Within the range; at this point, the fuselage drag coefficient is related to the Reynolds number Re, and can be written in general form: (21)， (22)， in, The characteristic length of the fuselage. The aerodynamic viscosity coefficient; and then: (23)， (24)， in, Indicates intermediate variables; In the fuselage coordinate system, the relative airflow velocity vector of the fuselage is denoted as... : (25)， in, This represents the component of the fuselage's relative airflow velocity along the x-axis. This represents the component of the fuselage's relative airflow velocity along the y-axis. This represents the component of the fuselage's relative airflow velocity along the z-axis. Air resistance of the fuselage The direction is opposite to the velocity direction, and it can be equivalently decomposed onto three coordinate axes: (26)， in, , , Let be the equivalent drag coefficients along the x, y, and z directions, respectively. They are derived from the drag coefficients. The equivalent frontal area distribution of the fuselage in all directions is determined by the formula above; The form is designed to ensure that the resistance is always opposite to the direction of the velocity, while retaining the characteristics of secondary resistance; This represents the component of the air resistance along the x-axis. This represents the component of the air resistance along the y-axis. This represents the component of the air resistance along the z-axis. Under the typical condition of spin fall, the vertical velocity component Much greater than the horizontal component, the dominant effect of fuselage drag is in the vertical direction: (27)， and , The value is relatively small and can be considered as linear damping of horizontal motion, which helps to suppress horizontal velocity divergence and improve control performance. Since the fuselage's center of mass is close to the aircraft's overall center of mass and its geometry is basically symmetric about its spin axis, the air resistance of the fuselage with respect to the aerodynamic moments about the x and y axes can be ignored; for the spin about the z axis, only an equivalent damping moment proportional to the angular velocity needs to be considered. : (28)， in, The equivalent spin damping coefficient of the fuselage; Small propellers are arranged near the wingtips to provide additional thrust on top of the spin descent, thereby adjusting the aircraft's spin speed and vertical motion characteristics; Let the unit vector of the propeller axis direction in the body coordinate system be... The propeller thrust is Then its force vector in the body coordinate system for: (29)， According to propeller aerodynamics, thrust An approximation using a thrust coefficient model is employed: (30)， in, For the propeller disk area, Where is the propeller radius. The propeller angular velocity, This is the thrust coefficient; The vector of the wingtip position relative to the center of mass Recorded as: (31)， in, , , These represent the distances from the wingtip to the center of mass in the x, y, and z directions, respectively. The aerodynamic torque of the propeller on the aircraft for: (32)。 4. The control method for the winged monoplane aircraft according to claim 3, characterized in that, In steps 1-2, to facilitate controller design, state variables are defined as follows: Define the state variable: aircraft position coordinates Describes the spatial coordinates of the aircraft's center of mass in the ground coordinate system; and its velocity in the ground coordinate system. velocity in the body coordinate system And satisfy Angular velocity in body coordinate system R0 represents the transformation matrix, u represents the velocity of the aircraft in the x-axis direction in the body coordinate system, v represents the velocity of the aircraft in the y-axis direction in the body coordinate system, w represents the velocity of the aircraft in the z-axis direction in the body coordinate system, p represents the attitude angular velocity of the aircraft in the x-axis direction in the body coordinate system, q represents the attitude angular velocity of the aircraft in the y-axis direction in the body coordinate system, s represents the attitude angular velocity of the aircraft in the z-axis direction in the body coordinate system, and T represents the transpose.

5. The control method for the winged monoplane aircraft according to claim 4, characterized in that, In steps 1-3, to facilitate controller design, the mathematical model of the winged monoplane flight system is transformed into six-degree-of-freedom flight dynamics equations based on the defined state variables, as follows: According to Newton's second law and Euler's law of rigid body rotation, the translation equations and attitude dynamics equations in the body coordinate system are as follows: (33)， in, yes The first derivative, yes The first derivative of , where m is the mass of the aircraft and I is the inertial matrix of the aircraft about its center of mass. Considering that the aircraft is a rigid body, then , This represents the moment of inertia about the x-axis; This represents the moment of inertia about the y-axis; This represents the moment of inertia about the z-axis. It is the product of inertia representing the coupling effect between the x-axis and z-axis. It is the product of inertia representing the coupling effect between the x-axis and y-axis. It is the product of inertia representing the coupling effect between the y-axis and the x-axis. It is the product of inertia representing the coupling effect between the y-axis and z-axis. It is the product of inertia representing the coupling effect between the z-axis and x-axis. It is the product of inertia representing the coupling effect between the z-axis and y-axis; The net force acting on the aircraft The resultant torque acting on the aircraft; The net force F acting on the aircraft consists of the following components: (34)， in, For the overall aerodynamic force of the wing; For the air resistance of the fuselage; For the thrust of the wingtip propeller; This represents gravity in the body coordinate system. Gravity in the body coordinate system as follows: (35)， in, The pitch angle of the aircraft. Let m be the roll angle of the aircraft, m be the mass of the aircraft, and g be the gravitational acceleration, taken as 9.8 N / kg. Substituting equations (17)~(19), (27), (30), (33), and (36) into equation (34), we obtain the following six-degree-of-freedom flight dynamics equations: (36)， in, This represents the moment of inertia about the x-axis; This represents the moment of inertia about the y-axis; This represents the moment of inertia about the z-axis. It is the product of inertia representing the coupling effect between the x-axis and z-axis. Representing variables The first derivative, Let x be the component of the total aerodynamic force of the wing in the x-direction. Let be the component of the total aerodynamic force of the wing in the z-direction. This represents the component of the air resistance along the x-axis. This represents the component of the air resistance along the y-axis. This represents the component of the air resistance along the z-axis. denoted as wingtip propeller thrust, and l as the total length of the wingtip from the axis of rotation. The total rolling moment acting on the wing. This is the total pitching moment acting on the wing.

6. The control method for the winged monoplane aircraft according to claim 5, characterized in that, In step 1, for the convenience of controller design, the following assumptions are made: Assumption 1: The aircraft is a rigid body, and the shape and mass distribution of the fuselage, wings and propeller structure do not change with aerodynamic forces during flight; Assumption 2: Gravitational acceleration is constant, and the direction of gravity does not change with attitude; Assumption 3: The ground is considered a plane, and the ground coordinates are inertial coordinates; Assumption 4: Aerodynamic loads are concentrated on the surfaces of the wing and fuselage; Assumption 5: Symmetric about the center of mass, product of inertia is negligible. ; Proceed to step 2.

7. The control method for the winged monoplane aircraft according to claim 6, characterized in that, In step 2, based on the mathematical model of the winged monoplane aircraft system, the horizontal motion controller and the vertical motion controller are designed, as follows: Step 2-1: Define vertical motion error ,in, For the desired height, This refers to the aircraft's current altitude; the system status is monitored under the control of the designed controller. The desired altitude command should be tracked as accurately as possible. ; Step 2-2: Define horizontal motion error ,in, All are target horizontal positions. This represents the current actual horizontal position of the aircraft; to facilitate the monitoring of the system state under the control of the designed controller. The desired horizontal position command should be tracked as accurately as possible. ; Steps 2-3: Adjust the horizontal control amount It will be input into the periodic control system for horizontal motion control; the flap deflection angle is periodically modulated using a continuous sine wave function, which has higher smoothness and control accuracy.

8. The control method for the winged monoplane aircraft according to claim 7, characterized in that, In step 2-1, the vertical motion error is defined. ,in, For the desired height, This refers to the aircraft's current altitude; the system status is monitored under the control of the designed controller. The desired altitude command should be tracked as accurately as possible. The details are as follows: Design the following PI controller to adjust the control quantity in the vertical direction. : (37)， in, and These are proportional gain and integral gain, respectively. The proportional term is used to quickly respond to system errors, while the integral term is used to eliminate steady-state errors and improve the long-term tracking accuracy of the system. In order to improve the actual control effect of the controller, the gain parameter is tuned in the control law to meet the altitude control requirements under different flight conditions.

9. The control method for the winged monoplane aircraft according to claim 8, characterized in that, In step 2-2, the horizontal motion error is defined. ,in, For the target horizontal position, This represents the current actual horizontal position of the aircraft; to facilitate the monitoring of the system state under the control of the designed controller. The desired horizontal position command should be tracked as accurately as possible. ; Design the following PD controller. : (38)， in, These are proportional gain and derivative gain, respectively. The proportional term is used to quickly correct position errors, while the derivative term is used to improve the responsiveness to error change trends, effectively suppressing overshoot and shortening the settling time.

10. The control method for the winged monoplane aircraft according to claim 9, characterized in that, In steps 2-3, the control quantity obtained in the horizontal motion controller The input will be used in the periodic control system for horizontal motion control; a continuous sine wave function is used for periodic modulation control of the flap deflection angle, resulting in higher smoothness and control accuracy; the periodic control system is designed as follows: In order to make the aircraft from its current position Move to the target point Horizontal control quantity ,in, This is the control variable in the x-direction. The control quantity in the y-direction is converted into a periodic signal phase control quantity. : (39)， Let the current spin angle of the spacecraft be... A sinusoidal periodic modulation model of the flap deflection angle controlled by a periodic control system. for: (40)， in, This refers to the flap period offset angle; The deflection angle amplitude controls the intensity of the periodic disturbance; To control the directional phase offset, the positional distribution of the deflection waveform within the period is determined; the phase of the periodic signal... It controls the projection position of the periodic deflection waveform during the rotation period around the axis; by adjusting This increases or decreases the flap deflection angle near a certain azimuth angle in the cycle, thereby guiding the asymmetric aerodynamic force to act in the desired direction and achieving target tracking.

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