Fixed-wing aircraft large-angle-of-attack landing control method based on dynamic reference trajectory

Abstract

The invention discloses a fixed-wing aircraft large-angle-of-attack landing control method based on a dynamic reference trajectory, and belongs to the technical field of fixed-wing aircraft autonomous landing, and the method comprises the following steps: establishing a landing control strategy based on a kinematic model and an attitude angle conversion model in the fixed-wing aircraft autonomous landing process, designing two-stage annular auxiliary line radiuses of an air flight stage of the fixed-wing aircraft, designing an expected track yaw angle, an expected track inclination angle and an expected roll angle of the fixed-wing aircraft, solving the expected yaw angle and the expected pitch angle, and constructing a control law of an attack angle when the fixed-wing aircraft is about to land; aiming at a flight stage and a taxiing stage of fixed-wing aircraft landing, a proper angular velocity controller and an air speed controller are designed based on a dynamic reference trajectory, so that the pose of the fixed-wing aircraft and a large attack angle when the fixed-wing aircraft is about to land are controlled, the dangerous conditions of collision and stall are avoided, and the safety of the fixed-wing aircraft is improved. And the accuracy and safety of autonomous landing control are improved.

Description

Technical Field

[0001] The present invention relates to the technical field of fixed-wing aircraft autonomous landing, and particularly to a large angle of attack landing control method for fixed-wing aircraft based on a dynamic reference trajectory. Background Art

[0002] Traditional fixed-wing aircraft landing methods have extremely high requirements for pilots. Pilots need to process a large amount of information in a short time under complex sea conditions and meteorological conditions, precisely control the speed, altitude and attitude of fixed-wing aircraft, and make correct decisions to ensure safe landing on the ship. Under such high pressure and high workload, human errors are difficult to completely avoid. With the progress of aviation technology, fixed-wing aircraft automatic landing technology is gradually moving towards practical application. However, although existing automatic landing systems have made certain progress in key technologies such as automatic landing, guidance and control, their development is still immature, and there are problems such as difficulty in achieving precise trajectory planning and attitude control.

[0003] Since fixed-wing aircraft need to control both position and attitude during the landing process, which involves multi-variable control problems with high complexity. At present, the research on how to precisely synchronize these two key variables to achieve stable landing is not deep enough and further exploration is needed. At the same time, the physical characteristics of fixed-wing aircraft have a significant impact on the landing trajectory. Therefore, when designing the landing trajectory, fixed-wing aircraft cannot be simplified as a particle model, but their actual body volume and size should be fully considered. In addition, considering the angle of attack and sideslip angle of fixed-wing aircraft and designing their flight trajectory and attitude are crucial for ensuring the safety of fixed-wing aircraft operation and landing. Summary of the Invention

[0004] To solve the technical problems raised in the above background, the present invention proposes a large angle of attack landing control method for fixed-wing aircraft based on a dynamic reference trajectory, designs the landing control method for the flight and taxiing stages of fixed-wing aircraft, establishes a reference plane with adjustable inclination and a dynamic trajectory based on the reference plane, and further designs a pose synchronization stabilizing controller to achieve the control of the pose of fixed-wing aircraft and the large angle of attack during landing, enabling fixed-wing aircraft to achieve precise landing and avoid the risks of stall and collision during landing.

[0005] The above technical objectives of the present invention are achieved through the following technical solutions:

[0006] A large angle of attack landing control method for fixed-wing aircraft based on a dynamic reference trajectory, comprising the following steps:

[0007] Step 1, based on the kinematic model and attitude angle conversion model during the autonomous landing process of fixed-wing aircraft, establish its landing control strategy and construct a dynamic reference trajectory model;

[0008] Step 2: Design the two-stage circular auxiliary line radius during the airborne flight phase of the fixed-wing aircraft;

[0009] Step 3: Design the desired track yaw angle, track tilt angle, and roll angle of the fixed-wing aircraft, solve the desired yaw angle and pitch angle, and construct the control law for the angle of attack when the fixed-wing aircraft is about to land;

[0010] Step 4: Design the angular velocity controller and airspeed controller during the landing process of the fixed-wing aircraft to achieve the control effect.

[0011] Furthermore, the said Step 1 includes the following contents:

[0012] (1) Establish the kinematic model for the autonomous landing of the fixed-wing aircraft as:

[0013]

[0014] where p n , p e , h represent the spatial coordinates of the fixed-wing aircraft in the inertial coordinate system, g is the gravitational constant, χ ∈ (-π, π] is the track yaw angle, γ ∈ (-π / 2, π / 2) is the track tilt angle, φ ∈ (-π / 2, π / 2) is the roll angle, represents the input airspeed, and are the control input quantities for controlling the attitude of the fixed-wing aircraft;

[0015] (2) Considering the angle of attack and sideslip angle of the fixed-wing aircraft, construct the attitude angle conversion model of the fixed-wing aircraft as:

[0016]

[0017] where ψ is the yaw angle, θ is the pitch angle, is the corresponding value of the roll angle after transformation from the body coordinate system to the track coordinate system, L kb is the transformation matrix for converting the track coordinate system to the body coordinate system, which is expressed as follows:

[0018]

[0019] where β represents the sideslip angle of the fixed-wing aircraft, and a normal value is given for simplifying subsequent derivations, and α represents the angle of attack of the fixed-wing aircraft;

[0020] (3) Establish the control strategy during the flight and taxiing phases of the fixed-wing aircraft during landing, specifically as follows:

[0021] The inertial coordinate system F I is translated along the positive direction of the p n axis by a distance p n0 to obtain a new coordinate system p n0 is the taxiing distance. Among them, in the coordinate system , the point is the origin; in the coordinate system F I , the point O is the origin. The landing control strategy of the fixed-wing aircraft is divided into flying along a dynamically adjustable desired inclination trajectory in the air to point O and taxiing on the ground to the final target point point;

[0022] (4) Construct a dynamically adjustable inclination dynamic reference trajectory model for the fixed-wing aircraft flight phase, specifically as follows:

[0023] Define relevant points and the inclination of the reference plane. The flight path of the fixed-wing aircraft is a dynamic trajectory based on the reference plane. Construct an annular auxiliary line when the fixed-wing aircraft flies to a certain position, and define the center coordinate p c of the annulus as:

[0024] p c = [0, r cosδ, r sinδ] T

[0025] where r represents the radius of the annulus, and δ represents the inclination between the reference plane and the p n op e plane;

[0026] Through two perpendicular relationships, the normal vector n of the trajectory reference plane can be obtained as:

[0027] n = [0, -sinδ, cosδ] T

[0028] Furthermore, using geometric relationships, the annular reference trajectory equation can be calculated as:

[0029]

[0030] Project the annular auxiliary line onto the p I plane of the inertial coordinate system F n op e , and an elliptical curve can be obtained. The trajectory equation of this ellipse is:

[0031]

[0032] The tangent vector c1 at a certain point on the annular curve is:

[0033] c1 = [p e tanδcosδ - r + p e cosδ, -p n cosδ, -p n cosδ] T

[0034] Among them, the direction of c1 is the same as the direction of the ground speed vector, and in the absence of crosswind, the ground speed is the same as the airspeed vector;

[0035] The tangent vector c2 of the projection point of this point on the ellipse is:

[0036] c2 = [p e -r cosδ, -p n cosδ, 0] T

[0037] Among them, the direction of c2 is the speed direction of the fixed-wing aircraft at the ground projection point;

[0038] The included angle τ between the two tangent vectors is:

[0039]

[0040] Furthermore, step 2 includes the following content:

[0041] The flight phase of the fixed-wing aircraft in the air is divided into two phases, the high-altitude and low-altitude flight phases, and the radii r of the circular auxiliary lines in the two phases are respectively expressed as r1 and r2;

[0042] During the high-altitude flight phase of the fixed-wing aircraft, to ensure the safety of operation, the radius r1 of the circular auxiliary line in the first phase is restricted by the approximate relationship of the sideslip angle as:

[0043]

[0044] Among them, r1 changes with the coordinates p n of the fixed-wing aircraft;

[0045] During the low-altitude flight phase of the fixed-wing aircraft, to enable the yaw angle of the fixed-wing aircraft to smoothly decrease when touching the ground and aligning with the runway, so as to ensure the comfort of landing, the radius r2 of the circular auxiliary line in the second phase is:

[0046]

[0047] Among them, [p n1 , p e1 , h1] T represents a specific point where the fixed-wing aircraft flies to in the inertial coordinate system, and at this position, it satisfies:

[0048] r1 = r2

[0049] Furthermore, step 3 includes the following content:

[0050] Introduce the attraction vector and the tangent vector of the fixed-wing aircraft at the trajectory projection point. To make the fixed-wing aircraft converge to the desired trajectory, its desired velocity vector V e is in the form of:

[0051] V e =[a(p e tanδsinδ - r + p e cosδ), -ap n cosδ, p e tanδ - h - ap n sinδ] T

[0052] where the parameter a is a positive constant;

[0053] Furthermore, construct the desired track yaw angle χ e and track tilt angle γ e as:

[0054]

[0055] where ||V e || is the Euclidean norm of the desired velocity vector;

[0056] Based on the coordinated turning radius condition, construct the desired roll angle φ e as:

[0057]

[0058] where σ3 and ξ ∈ (0,1) are positive constant parameters;

[0059] Based on the constructed attitude angle conversion model, substitute the desired track yaw angle χ e , track tilt angle γ e , roll angle φ e to solve for the desired yaw angle ψ e and pitch angle θ e as:

[0060]

[0061] Furthermore, when the fixed-wing aircraft is about to land, there is a certain deviation angle κ allowed between the desired yaw angle ψ e and the p n axis direction:

[0062] (1) When the desired yaw angle ψ e is the reverse deviation based on the p n axis direction, that is:

[0063] 2π ≥ ψ e≥2π - κ

[0064] where k is a positive constant, and its magnitude is determined by actual requirements and is affected by the runway length and runway width;

[0065] In this case, the angle of attack α is

[0066]

[0067] where t satisfies the relationship:

[0068] t = cosβ·χ e - sinβ·γ e

[0069] (2) When the desired yaw angle ψ e is a positive deviation based on the p n axis direction, that is:

[0070] 0 ≤ ψ e ≤ κ

[0071] In this case, the angle of attack α is:

[0072]

[0073] By setting appropriate sideslip angles β, parameters a, σ3, and ξ, and the parameters of the controllers u1 and u2, the magnitude of the angle of attack α of the fixed - wing aircraft when approaching landing is controlled to ensure the safety at the landing moment and avoid the collision risk.

[0074] Furthermore, step 4 includes the following content:

[0075] The landing process of the fixed - wing aircraft includes five stages: gliding, flaring, level flight, floating, and rolling. Considering reducing the model complexity while retaining its general flight characteristics, the in - air flight stage is simplified into a gliding process, a flaring process, and a level - flight process. In each process, different positive airspeeds V a1 , V a2 , V a3 are set according to actual requirements and satisfy:

[0076] min{V a1 , V a2 , V a3} ≥ V min

[0077] where V min represents the minimum airspeed, that is, the minimum speed at which the fixed - wing aircraft does not stall;

[0078] Design the roll - angular - velocity controller as:

[0079]

[0080] The designed track tilt angular velocity controller is as follows:

[0081]

[0082] where σ4 and σ5 are normal constant parameters. It should be noted that according to the kinematic characteristics of the fixed-wing aircraft, the control of the track yaw angular velocity is restricted by the roll angle;

[0083] During the taxiing phase, set the roll angular velocity controller and the track tilt angular velocity controller of the fixed-wing aircraft to 0, that is:

[0084] u1 = u2 = 0

[0085] Design an airspeed controller to decelerate it on the ground until it reaches the target position:

[0086] V a = -σ6(p n - p n0 )

[0087] where σ6 is a normal constant parameter, p n0 is the taxiing distance. It should be noted that the taxiing distance p n0 should be less than the runway length.

[0088] The present invention proposes a control strategy for the landing of a fixed-wing aircraft during the flight and taxiing phases. During the flight phase, by constructing a dynamically adjustable inclination trajectory, a control law for the angle of attack of the fixed-wing aircraft when approaching landing is established. The roll angular velocity controller and the track tilt angular velocity controller are designed to achieve the synchronous convergence of the position and attitude of the fixed-wing aircraft to the desired dynamic trajectory, and ensure the safety and accuracy of the fixed-wing aircraft landing; during the taxiing phase, by designing an airspeed controller to achieve the deceleration process during ground taxiing until finally reaching the target location.

[0089] In summary, the present invention mainly has the following beneficial effects:

[0090] 1. The landing control method for a fixed-wing aircraft proposed by the present invention designs the landing strategies for the flight and taxiing phases. By designing the roll angular velocity controller, the track tilt angular velocity controller, and the airspeed controller, it can finally effectively achieve the synchronous stabilization of the position and attitude of the fixed-wing aircraft, which has important practical application value in the precise control of fixed-wing aircraft and reduces the risk of the fixed-wing aircraft losing control during landing.

[0091] 2. The present invention introduces the influences brought by the angle of attack and sideslip angle during the flight of the fixed-wing aircraft. Based on the additional constraints of the two, it designs the dynamic flight trajectory of the fixed-wing aircraft and constructs the control law for the angle of attack when approaching landing, which has important reference value for ensuring the maneuverability of the fixed-wing aircraft and the safety during large angle-of-attack landing.

[0092] 3. The present invention fully considers the physical constraints brought by the body size of a fixed-wing aircraft, designs an adjustable inclination angle of the reference plane to reduce excessive trajectory inclination, provides a higher dynamic buffering capacity than usual, and enables the fixed-wing aircraft to taxi within a certain distance after landing until it reaches the desired position. Brief Description of the Drawings

[0093] Figure 1 is the flowchart of the method of the present invention;

[0094] Figure 2 is the model diagram of the desired dynamically adjustable reference trajectory with adjustable inclination angle;

[0095] Figure 3 is the simulated curve graph of airspeed and control input in the landing application embodiment of the fixed-wing aircraft;

[0096] Figure 4 is the simulated curve graph of position in the landing application embodiment of the fixed-wing aircraft;

[0097] Figure 5 is the simulated curve graph of attitude in the landing application embodiment of the fixed-wing aircraft;

[0098] Figure 6 is the simulated curve graph of the radius of the circular auxiliary line in the landing application embodiment of the fixed-wing aircraft;

[0099] Figure 7 is the simulated curve graph of the angle of attack in the landing application embodiment of the fixed-wing aircraft;

[0100] Figure 8 Simulated curve graph of the motion trajectory in the landing application embodiment of the fixed-wing aircraft. Detailed Embodiments

[0101] In order to make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions of the embodiments of the present invention in conjunction with the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all of the embodiments. Based on the described embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts fall within the scope of protection of the present invention.

[0102] The following embodiments are used to illustrate the present invention, but cannot be used to limit the scope of protection of the present invention. The conditions in the embodiments can be further adjusted according to specific conditions. Any simple improvement of the method of the present invention under the premise of the concept of the present invention falls within the scope of protection required by the present invention.

[0103] A large angle of attack landing control method for a fixed-wing aircraft based on a dynamic reference trajectory. When implemented specifically, the following steps should be followed:

[0104] (1) To accurately describe the flight path and position state characteristics of a fixed-wing aircraft in three-dimensional space, the kinematic model during the autonomous landing process of the fixed-wing aircraft is:

[0105]

[0106] Among them, p n , p e , h represent the spatial coordinates of the fixed-wing aircraft in the inertial coordinate system, g is the gravitational constant, χ ∈ (-π, π] is the track yaw angle, γ ∈ (-π / 2, π / 2) is the track tilt angle, φ ∈ (-π / 2, π / 2) is the roll angle, represents the input airspeed, and are the control input quantities used to control the attitude of the fixed-wing aircraft.

[0107] (2) Considering the angle of attack and sideslip angle of the fixed-wing aircraft, a conversion model of the attitude angle of the fixed-wing aircraft is constructed as:

[0108]

[0109] Among them, ψ is the yaw angle, θ is the pitch angle, is the corresponding value of the roll angle after transforming from the body coordinate system to the track coordinate system, L kb is the transformation matrix for converting the track coordinate system to the body coordinate system;

[0110] The transformation matrix L ka for converting from the track coordinate system to the airflow coordinate system is:

[0111]

[0112] Among them, ε represents the roll angle around the velocity vector;

[0113] The transformation matrix L ab for converting from the airflow coordinate system to the body coordinate system is:

[0114]

[0115] Among them, β represents the sideslip angle of the fixed-wing aircraft. For simplicity in subsequent derivations, a normal value is given, and α represents the angle of attack of the fixed-wing aircraft;

[0116] To simplify the attitude angle conversion model, take the roll angle ε around the velocity vector as 0. Then, the transformation matrix L kb for converting from the track coordinate system to the body coordinate system is:

[0117]

[0118] (3) Define the coordinate system F I , such as Figure 2 shown, translate a distance p along the positive direction of the p n axis to obtain a new coordinate system n0 In the coordinate system F , point O is the origin, and the position of the fixed-wing aircraft in F I coordinate system can be written as I At the same time, after the coordinate transformation, the landing of the fixed-wing aircraft is divided into flying along the dynamic trajectory in the air to the landing point and taxiing on the ground to the target point

[0119] Figure 2 In the coordinate system F I , the expected dynamic trajectory model is as Figure 2 shown. The flight trajectory of the fixed-wing aircraft during the flight phase can be further divided into two phases. The first phase is the high-altitude flight phase of the fixed-wing aircraft, and the second phase is the low-altitude flight phase of the fixed-wing aircraft. Point C is the center of the circular auxiliary line in the high-altitude flight phase, point C2 is the center of the circular auxiliary line in the low-altitude flight phase, the radius r represents the radius of the circular auxiliary line at a certain moment during the flight phase, and the size of this radius is a variable value. The angle δ represents the inclination angle between the trajectory reference plane and the plane p n op e , and this inclination angle is adjustable and affected by the physical size of the fixed-wing aircraft. In practice, due to the regulation of the glide angle, this inclination angle is generally about 0.052 radians;

[0120] The initial position coordinates of the fixed-wing aircraft are p F = [p n , p e , h] T . Draw a perpendicular line from F to the p n op e plane, and the intersection of this perpendicular line and the circular auxiliary line at a certain moment in the high-altitude flight phase is point A. The coordinates of point A can be calculated as p A = [p n , p e , p e tanδ] T . Define the expected position of the fixed-wing aircraft on this circular auxiliary line as point E, and there is p A = p E . Define the vector

[0121] Define the center coordinates p C of the circular auxiliary line as:

[0122] pC = [0, rcosδ, rsinδ] T

[0123] Define the vector n as the normal vector of the reference plane. According to its and vertical relationship, we can obtain:

[0124] n·J = 0

[0125] n·η = 0

[0126] Solve the above equations to obtain the vector n = [0, -sinδ, cosδ] T , which can be used as the normal vector of the dynamic trajectory reference plane;

[0127] Furthermore, the trajectory equation of the annular auxiliary line can be calculated as:

[0128]

[0129] Project the annular auxiliary line onto the p I plane of the inertial coordinate system F n op e , and an elliptical curve can be obtained. The trajectory equation of this ellipse is:

[0130]

[0131] The tangent vector c1 at a certain point on the annular curve is:

[0132] c1 = [p e tanδsinδ - r + p e cosδ, -p n cosδ, -p n sinδ] T

[0133] where the direction of c1 is the same as the direction of the ground speed vector. In the case of no crosswind, the ground speed is the same as the airspeed vector;

[0134] The tangent vector c2 of the projection point of this point on the ellipse is:

[0135] c2 = [p e -rcosδ, -p n cosδ, 0] T

[0136] where the direction of c2 is the velocity direction of the fixed-wing aircraft at the ground projection point;

[0137] The included angle τ between the two tangent vectors is:

[0138]

[0139] Among them, the value of τ is very small.

[0140] (4) The in-air flight phase of the fixed-wing aircraft is subdivided into two phases, the high-altitude and low-altitude flight phases. The radii r of the circular auxiliary lines in the two phases are denoted as r1 and r2 respectively.

[0141] During the high-altitude flight phase of the fixed-wing aircraft, to ensure the safety of operation, the magnitude of the sideslip angle β is generally within 0.087 radians. At this time, it can be approximately considered that:

[0142] β = τ

[0143] Furthermore, the radius r1 of the auxiliary line in the first phase is:

[0144]

[0145] Among them, r1 changes with the fixed-wing aircraft p n and varies.

[0146] During the low-altitude flight phase of the fixed-wing aircraft, to enable the yaw angle of the fixed-wing aircraft to smoothly decrease when it touches the ground and aligns with the runway, so as to ensure the comfort of landing, the fixed-wing aircraft should fly along a certain circular auxiliary line until it touches the ground.

[0147] Based on the circular trajectory equation, the calculated radius r2 of the circular auxiliary line in the second phase is:

[0148]

[0149] Among them, [p n1 , p e1 , h1] T represents a specific point where the fixed-wing aircraft flies to in the inertial coordinate system.

[0150] To achieve a smooth transition of the radius of the circular auxiliary line from the high-altitude flight phase to the low-altitude flight phase, so as to avoid a large change in the attitude angle of the fixed-wing aircraft, then at this specific point, it satisfies:

[0151] r1 = r2

[0152] Specifically, the critical flight altitude for dividing the high-altitude flight phase and the low-altitude flight phase is generally set at a certain altitude during the flare-out phase, and this altitude can ensure a smooth transition of the radii of the circular auxiliary lines in the two phases.

[0153] (5) Define the vector from the center of the circle to the projected position of the fixed-wing aircraft as:

[0154]

[0155] Construct the tangent vector μ of the circular auxiliary line at point A. By virtue of its properties of being perpendicular to the radial direction and the normal vector of the circular auxiliary line reference plane respectively, we can obtain:

[0156] μ·n = 0

[0157] μ·λ = 0

[0158] Solving the above equations can yield:

[0159] μ = [p e tanδsinδ - r + p e cosδ, -p n cosδ, -p n sinδ] T

[0160] To make the fixed-wing aircraft converge to the desired dynamic trajectory, its desired velocity vector V e is in the form of:

[0161] V e = aμ + ν

[0162] = [a(p e tanδsinδ - r + p e cosδ), -ap n cosδ, p e tanδ - h - ap n sinδ] T

[0163] where is the introduced attracting vector, and the parameter a is a positive constant;

[0164] Specifically, the parameter a is used to adjust the convergence rate of the fixed-wing aircraft to the desired dynamic trajectory. A larger value of a tends to make the velocity vector V e closer to being parallel to the horizontal direction, thus making the convergence direction of the fixed-wing aircraft smoother. However, an excessively large V e value may lead to too long a convergence time and the inability to fly accurately along the desired trajectory in a timely manner;

[0165] Furthermore, construct the desired track yaw angle χ e and track tilt angle γ e of the fixed-wing aircraft as:

[0166]

[0167] where, ||V e || is the Euclidean norm of the desired velocity vector. Specifically:

[0168]

[0169] Construct the desired roll angle φ of the fixed-wing aircraft according to the coordinated turning radius condition e as follows:

[0170]

[0171] where σ3 and ξ ∈ (0, 1) are positive constant parameters.

[0172] (6) Based on the constructed attitude angle conversion model, substitute the desired track yaw angle χ e , track tilt angle γ e , and roll angle φ e to solve for the desired yaw angle ψ e and pitch angle θ e of the fixed-wing aircraft as follows:

[0173]

[0174] Furthermore, when the fixed-wing aircraft is about to land, there is a certain allowable deviation angle κ between the desired yaw angle ψ e and the p n axis direction:

[0175] When the desired yaw angle ψ e is anti-offset based on the p n axis direction, i.e.:

[0176] 2π ≥ ψ e ≥ 2π - κ

[0177] where κ is a positive constant, and its magnitude is determined by actual requirements and is affected by the runway length and runway width;

[0178] In this case, the angle of attack α is:

[0179]

[0180] where t satisfies the relationship:

[0181] t = cosβ·χ e -sinβ·γ e

[0182] When the desired yaw angle ψ e is positive-offset based on the p n axis direction, i.e.:

[0183] 0 ≤ ψ e ≤ κ

[0184] In this case, the angle of attack α is:

[0185]

[0186] Specifically, to determine the magnitude of the angle of attack α during landing, it is first necessary to set the values of the deviation angle κ and the sideslip angle β, and secondly, it is necessary to substitute the desired track yaw angle χ e , the track tilt angle γ e and the roll angle φ e , thus further requiring the design of the parameters a of the desired velocity vector V e , the parameter σ3 of the desired roll angle φ e of ξ and the parameters of the controllers u1 and u2;

[0187] By designing the magnitude of the angle of attack α during landing, ensuring that it is less than the critical angle of attack to ensure the safety at the moment of landing at a large angle of attack and avoid the risk of collision.

[0188] (7) The landing process of a fixed-wing aircraft includes five stages: gliding, flaring, level flight, floating, and taxiing. Considering reducing the model complexity while retaining its general flight characteristics, the in-air flight stage is simplified into a gliding process, a flaring process, and a level flight process. In each process, the airspeed is set to different constant positive values V a1 , V a2 , V a3 ;

[0189] In order for a fixed-wing aircraft to fly along a dynamic reference trajectory in three-dimensional space and ensure that it has sufficient speed, thus the airspeeds V a1 , V a2 , V a3 satisfy:

[0190] min{V a1 , V a2 , V a3} ≥ V min

[0191] wherein, V min represents the minimum airspeed, that is, the minimum speed at which a fixed-wing aircraft does not stall;

[0192] Design the roll angular velocity controller as:

[0193]

[0194] Design the track tilt angular velocity controller as:

[0195]

[0196] where σ4 and σ5 are normal constant parameters. It should be noted that according to the kinematic characteristics of a fixed-wing aircraft, the control of the track yaw angular velocity is restricted by the roll angle;

[0197] During the taxiing phase, set the roll angular velocity controller and the flight path inclination angular velocity controller of the fixed-wing aircraft to 0, i.e.:

[0198] u1 = u2 = 0

[0199] Design an airspeed controller to decelerate and taxi on the ground until the target position is reached:

[0200] V a = -σ6(p n - p n0 )

[0201] where σ6 is a positive constant parameter. It should be noted that the taxiing distance p n0 should be less than the runway length.

[0202] (8) Verify the effect of a large angle of attack landing control method for a fixed-wing aircraft based on a dynamic reference trajectory according to the present invention;

[0203] Verify using a simulation example. The simulation process is as follows:

[0204] The initial pose, the desired pose of the landing point, the desired pose of the target point, and the trajectory inclination angle of the fixed-wing aircraft are as follows:

[0205]

[0206] δ = 0.05rad

[0207] where, represents the initial pose, represents the desired pose of the landing point, represents the desired pose of the target point, and δ represents the trajectory inclination angle;

[0208] The airspeed settings during the descent process, the flare process, and the level flight process during the landing are as follows:

[0209]

[0210] where h represents the flight altitude, V a1 represents the airspeed during the descent process, V a2 represents the airspeed during the flare process, V a3 represents the airspeed during the level flight process;

[0211] Other parameter settings are:

[0212] ξ = 0.6, a = 1, σ3 = 0.45, σ4 = 2, σ5 = 6, σ6 = 0.15, κ = 0;

[0213] With the aid of the designed roll angular velocity controller, the flight path inclination angular velocity controller, and the airspeed controller, Figure 8Shows the trajectory of a fixed-wing aircraft in three-dimensional space. It can be observed that the landing process of the fixed-wing aircraft is divided into a flight phase and a taxiing phase: In the flight phase, the fixed-wing aircraft flies along the dynamic reference trajectory to the landing point; in the taxiing phase, the fixed-wing aircraft transitions to ground taxiing, gradually decelerates, and finally stabilizes at the target position. Figure 3 Shows the changes in the airspeed and control input of a fixed-wing aircraft during the landing process. The airspeed of the fixed-wing aircraft maintains different constant airspeeds at different stages before reaching the landing point to approximate the actual landing scenario; subsequently, during the ground taxiing phase, the airspeed gradually decreases and finally converges to zero to complete stabilization.

[0214] Figure 4 Shows the change process of the position state of a fixed-wing aircraft in three-dimensional space, clearly showing that its position gradually converges to the origin over time. Figure 5 Shows the change process of the attitude angle of the fixed-wing aircraft. The simulation results show that when it is about to land, the yaw angle ψ is approximately 6.28 rad, which is to align the p-axis in the opposite direction to the runway, the pitch angle θ is approximately 0.31 rad, and the roll angle φ is approximately 0.75 rad. All are within the error tolerance range, which can ensure the safety of touchdown, and finally the attitude angle of the fixed-wing aircraft can converge to the target attitude over time. n Axis is deflected in the opposite direction to align with the runway, the pitch angle θ is approximately 0.31 rad, and the roll angle φ is approximately 0.75 rad. All are within the error tolerance range, which can ensure the safety of touchdown, and finally the attitude angle of the fixed-wing aircraft can converge to the target attitude over time. Figure 6 Shows the change in the radius of the circular auxiliary line of the dynamic reference trajectory during the landing process of the fixed-wing aircraft. The simulation curve shows that from 0 s to approximately 18.72 s, the fixed-wing aircraft is in the high-altitude flight phase and flies towards the reference plane under the action of the desired vector. The radius of the circular auxiliary line changes in a proportional relationship with the fixed-wing aircraft p n Is in a proportional relationship, but since the fixed-wing aircraft is not on the desired reference plane at this time, there is no actual circular auxiliary line. The influence of r at this time is only to affect the desired velocity vector V e ; From 18.72 s to approximately 24.85 s, the fixed-wing aircraft is still in the high-altitude flight phase but has flown into the reference plane. At this time, as the fixed-wing aircraft p n Increases, the radius of the circular auxiliary line continues to increase; from 24.85 s to 34.45 s, the fixed-wing aircraft enters the low-altitude flight phase, and the radius of the circular auxiliary line remains unchanged. Figure 7 Shows the magnitude of the angle of attack of the fixed-wing aircraft during the landing process. The simulation results show that by designing the parameter values, the angle of attack during the flight of the fixed-wing aircraft can be controlled to be less than the stall critical angle of attack to ensure the safety of the fixed-wing aircraft landing at a large angle of attack.

[0215] Although embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that, unless otherwise defined, the technical terms or scientific terms used in the present invention should have the ordinary meanings understood by those with ordinary skills in the field to which the present invention pertains. The terms "including" or "comprising" and the like used in the present invention mean that the elements or items appearing before the word cover the elements or items listed after the word and their equivalents, without excluding other elements or items. The terms "connected" or "coupled" and the like are not limited to physical or mechanical connections, and may also include electrical connections, whether direct or indirect. The terms "upper", "lower", "left", "right", etc. are only used to represent relative positional relationships, and when the absolute position of the object being described changes, the relative positional relationship may also change accordingly.

[0216] Although embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that various changes, modifications, substitutions, and variations can be made to these embodiments without departing from the principles and spirit of the present invention. The scope of the present invention is defined by the appended claims and their equivalents.

Claims

1. A large angle of attack landing control method for fixed-wing aircraft based on a dynamic reference trajectory, characterized in that It includes the following steps: Step 1: Based on the kinematic model and attitude angle conversion model during the autonomous landing process of a fixed-wing aircraft, establish its landing control strategy and construct a dynamic reference trajectory model; Step 2: Design the two-stage circular auxiliary line radius during the airborne flight phase of the fixed-wing aircraft; Step 3: Design the desired track yaw angle, track tilt angle, and roll angle of the fixed-wing aircraft, solve the desired yaw angle and pitch angle, and construct the control law for the angle of attack when the fixed-wing aircraft is about to land; Step 4: Design the angular velocity controller and airspeed controller during the landing process of the fixed-wing aircraft to achieve the control effect.

2. The large angle of attack landing control method for a fixed-wing aircraft based on a dynamic reference trajectory according to claim 1, wherein The specific content of Step 1 includes: (1) Establish the kinematic model of the fixed-wing aircraft's autonomous landing as where p n , p e , h represent the spatial coordinates of the fixed-wing aircraft in the inertial coordinate system, g is the gravitational constant, χ ∈ (-π, π] is the track yaw angle, γ ∈ (-π / 2, π / 2) is the track tilt angle, φ ∈ (-π / 2, π / 2) is the roll angle, represents the input airspeed, and are the control input quantities for controlling the attitude of the fixed-wing aircraft; (2) Considering the angle of attack and sideslip angle of the fixed-wing aircraft, construct the attitude angle conversion model of the fixed-wing aircraft as where ψ is the yaw angle, θ is the pitch angle, is the corresponding value after the roll angle is transformed from the body coordinate system to the trajectory coordinate system, and L kb is the transformation matrix for the transformation from the trajectory coordinate system to the body coordinate system, expressed as follows where β represents the sideslip angle of the fixed-wing aircraft. For simplicity in subsequent derivations, a normal value is given, and α represents the angle of attack of the fixed-wing aircraft; (3) Establish the control strategy during the flight phase and taxiing phase of the fixed-wing aircraft during the landing process; Inertial coordinate system F I Translate along the positive direction of the p n axis by a distance p n0 to obtain a new coordinate system p n0 is the sliding distance. Among them, in the coordinate system , the point is the origin; in the coordinate system F I , the point O is the origin. The landing control strategy of the fixed-wing aircraft is divided into flying along a dynamically adjustable trajectory with an expected inclination angle in the air to point O and taxiing on the ground to the final target point point; (4) Construct a dynamic reference trajectory model with adjustable inclination angle during the flight phase of the fixed-wing aircraft; Define relevant points and the inclination angle of the reference plane. The flight path of a fixed-wing aircraft is a dynamic trajectory based on the reference plane. Construct a circular auxiliary line when the fixed-wing aircraft flies to a certain position, and define the center coordinate p of the circle C be p C = [0, r cosδ, r sinδ] T where r represents the radius of the ring, and δ represents the inclination angle between the reference plane and the n op e plane; Through two perpendicular relationships, the normal vector n of the trajectory reference plane can be obtained as n = [0, -sinδ, cosδ] T Furthermore, using geometric relationships, the circular reference trajectory equation can be calculated as Project the circular auxiliary line onto the inertial coordinate system F I p of n op e plane, an elliptic curve can be obtained, and the trajectory equation of this ellipse is The tangent vector c1 at a certain point on the circular curve is c1 = [p e tanδsinδ - r + p e cosδ, -p n cosδ, -p n sinδ] T where the direction of c1 is the same as the direction of the ground speed vector. In the case of no crosswind, the ground speed is the same as the airspeed vector; The tangent vector c2 of the projection point of this point on the ellipse is c2 = [p e -r cosδ, -p n cosδ, 0] T where the direction of c2 is the direction of the speed of the projection point of the fixed-wing aircraft on the ground; The included angle τ between the two tangent vectors is 3. A fixed-wing aircraft high-angle-of-attack landing control method based on a dynamic reference trajectory according to claim 1, characterized in that The specific content of Step 2 includes: The airborne flight phase of the fixed-wing aircraft is divided into two phases, the high-altitude and low-altitude flight phases. The circular auxiliary line radii r in the two phases are respectively denoted as r1 and r2; During the high-altitude flight phase of the fixed-wing aircraft, to ensure the safety of operation, the first-stage circular auxiliary line radius r1 is constrained by the sideslip angle approximate relationship as where r1 varies with the coordinates p of the fixed-wing aircraft n and changes accordingly; During the low-altitude flight phase of the fixed-wing aircraft, to enable the yaw angle when the fixed-wing aircraft touches down and aligns with the runway to smoothly decrease and ensure the comfort of landing, the second-stage circular auxiliary line radius r2 is Among them, [p n1 , p e1 , h1] T represents that a fixed-wing aircraft flies to a specific point in the inertial coordinate system, and at this position, it satisfies r1＝r2。 4. A large angle of attack landing control method for a fixed-wing aircraft based on a dynamic reference trajectory according to claim 1, characterized in that The specific content of Step 3 includes: Introduce the attraction vector and the tangent vector of the fixed-wing aircraft at the trajectory projection point. To make the fixed-wing aircraft converge to the desired trajectory, its desired velocity vector V e is in the form of V e = [a(p e tanδsinδ - r + p e cosδ), -ap n cosδ, p e tanδ - h - ap n sinδ] T where the parameter a is a positive constant; Further, construct the desired track yaw angle χ e and track tilt angle γ e for where ||V e || is the Euclidean norm of the desired velocity vector; Construct the desired roll angle φ of a fixed-wing aircraft based on the coordinated turn radius condition e For where σ3 and ξ ∈ (0, 1) are positive constant parameters; Based on the constructed attitude angle conversion model, substitute the desired track yaw angle χ e , track tilt angle γ e , roll angle φ e to solve for the desired yaw angle ψ e , pitch angle θ e as When a fixed-wing aircraft is about to land, it is desirable that the yaw angle ψ e has a certain deviation angle κ n from the p-axis direction (1) When the desired yaw angle ψ e is based on the reverse bias in the p n axis direction, that is 2π ≥ ψ e ≥ 2π - κ where κ is a positive constant, and its magnitude is determined by actual requirements and is affected by the runway length and runway width; In this case, the angle of attack α is where t satisfies the relational expression t = cosβ·χ e -sinβ·γ e (2) When the desired yaw angle ψ e is positive based on the p n axis direction, i.e., 0 ≤ ψ e ≤ κ In this case, the angle of attack α is By setting appropriate sideslip angle β, parameters a, σ3 and ξ, and the parameters of controllers u1 and u2, control the magnitude of the angle of attack α when about to land to ensure the safety at the landing moment and avoid the risk of collision.

5. A large angle of attack landing control method for a fixed-wing aircraft based on a dynamic reference trajectory according to claim 1, characterized in that The specific content of Step 4 includes: The landing process of a fixed-wing aircraft includes five stages: gliding, flaring, level flight, floating, and rolling. For the purpose of reducing the model complexity while retaining its general flight characteristics, the airborne flight stage is simplified into a gliding process, a flaring process, and a level flight process. In each process, the airspeed is set to different constant positive values V a1 , V a2 , V a3 , and satisfy min{V a1 ,V a2 ,V a3}≥V min Among them, V min represents the minimum airspeed, that is, the minimum speed at which a fixed-wing aircraft does not stall; Design the roll angular velocity controller as Based on the kinematic characteristics of the fixed-wing aircraft, the control of the track yaw angular velocity is restricted by the roll angle; Design the track tilt angular velocity controller as where σ4 and σ5 are positive constant parameters; Under the action of this roll angular velocity controller and track tilt angle controller, the position of the fixed-wing aircraft in three-dimensional space can converge to the desired dynamic trajectory; During the taxiing phase, set the roll angular velocity controller and track tilt angular velocity controller of the fixed-wing aircraft to 0, that is, u1 = u2 = 0 Design an airspeed controller to decelerate a fixed-wing aircraft during ground taxiing until it reaches the target position V a = -σ6(p n -p n0 ) Among them, σ6 is a positive constant parameter, and p n0 is the sliding distance; it should be noted that the sliding distance p n0 should be less than the runway length.

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