Guidance control method for extended-range aircraft with landing angle constraint

Abstract

The present invention discloses a guidance and control method for an extended-range aircraft with a landing angle constraint. In the method, in the middle guidance section, the pitch servo is controlled to steer and the angle of attack of the aircraft is increased so that the aircraft can glide farther. When entering the terminal guidance section, a guidance law based on a sliding surface is set, and under a given expected landing angle, the required overload of the aircraft is calculated in real time, and the flight of the aircraft is controlled accordingly. Finally, the aircraft is controlled to land at the expected landing angle and hit the target under the condition that the range of the aircraft is increased.

Description

Technical Field

[0001] The invention relates to the field of aircraft guidance control, and in particular to a guidance control method for an extended-range aircraft with landing angle constraints. Background Art

[0002] With the development of technology, guided aircraft have attracted much attention due to their high precision. Traditional guided aircraft use proportional guidance law to complete the task of hitting the target. Considering the range of the aircraft, traditional gun-launched aircraft experience high overload during launch, and the components have poor resistance to high overload. Therefore, in the inner ballistics, the aircraft components are damaged and fail, limiting the range of the aircraft. In the traditional scheme, the range can be appropriately increased by controlling the glide trajectory. However, for certain specific targets, the aircraft is required to attack the target from the top, which puts higher requirements on the glide trajectory. The traditional glide trajectory cannot meet the design needs of large drop angles.

[0003] Due to the contradictory relationship between the above-mentioned range and the large landing angle, the inventors have conducted in-depth research on existing aircraft guidance and control schemes in the hope of designing a new guidance and control method that can solve the above-mentioned problems. Summary of the invention

[0004] In order to overcome the above-mentioned problems, the inventors have conducted intensive research and designed a guidance and control method for an extended-range aircraft with a landing angle constraint. In this method, in the middle guidance section, the pitch servo is controlled to steer and the angle of attack of the aircraft is increased to enable it to glide farther. When entering the terminal guidance section, a guidance law based on a sliding surface is set. Under a given expected landing angle, the required overload of the aircraft is calculated in real time, and the flight of the aircraft is controlled accordingly. Finally, the aircraft is controlled to land at the expected landing angle and hit the target while the range of the aircraft is increased, thereby completing the present invention.

[0005] Specifically, the purpose of the present invention is to provide a guidance and control method for an extended-range aircraft with a landing angle constraint, in which:

[0006] After passing the highest point, the aircraft enters the gliding stage. In the gliding stage, the aircraft adjusts the angle of attack by deflecting the pitch rudder, so that the aircraft glides at a balanced angle of attack.

[0007] After entering the terminal guidance phase, the aircraft is guided and controlled by the guidance law with landing angle constraint.

[0008] Wherein, in the gliding section, the balance angle of attack is to control the angle of attack of the aircraft to be above a set value, and the set value is 18 to 25°, preferably 20°.

[0009] Wherein, in the gliding stage, the required pitch rudder deflection angle is obtained in real time by the following formula (1):

[0010]

[0011] Among them, α represents the angle of attack, δ represents the pitch rudder deflection angle, C g represents the center of mass of the aircraft, C ym Indicates the aircraft pressure center, C yd represents the pitch rudder pressure center; τ represents the lift-to-drag ratio;

[0012] Preferably, the lift-to-drag ratio τ is obtained by the following formula (II):

[0013]

[0014] Among them, F y Indicates the total lift, F x Indicates total spring resistance.

[0015] In the terminal guidance stage, the guidance law with landing angle constraint obtains the required overload through the following formula (III):

[0016]

[0017] Among them, a M Indicates that overload is required.

[0018] Indicates the relative speed of the projectile and target;

[0019] represents the line of sight angular velocity;

[0020] R represents the relative distance between the projectile and the target;

[0021] K represents the adaptive parameter;

[0022] s represents the non-singular terminal sliding surface;

[0023] β represents the design gain;

[0024] ρ represents the design gain;

[0025] e 1 Indicates the angle deviation;

[0026] e 2 Indicates the angular velocity deviation;

[0027] σ represents the aircraft lead angle.

[0028] Among them, the relative speed of the projectile Obtained in real time through the following formula (IV):

[0029] The projectile-to-target line-of-sight angular velocity Obtained in real time through the following formula (V):

[0030]

[0031] Among them, V a Indicates the aircraft speed.

[0032] Wherein, the non-singular terminal sliding surface s is obtained by the following formula (VI):

[0033] s=e 2 +β|e 1 |ρsign(e 1 ) (six).

[0034] sign(e 1 ) represents a symbolic function;

[0035] Wherein, the angle deviation e 1 =λ-λ d

[0036] The angular velocity deviation

[0037] Among them, λ represents the sight angle of the missile and the target, λ d Indicates the desired terminal angle.

[0038] The aircraft speed and position are obtained by the following formula (VII):

[0039]

[0040] Among them, P a Indicates the position of the aircraft;

[0041] V a Indicates the aircraft speed;

[0042] P G Indicates the aircraft position information obtained by GPS measurement;

[0043] P I Indicates the aircraft position information measured by INS;

[0044] V G Indicates the aircraft speed information measured by GPS;

[0045] V I Indicates the aircraft speed information measured by INS;

[0046] σ 1 Indicates the weight of the information obtained by GPS measurement;

[0047] σ2 Indicates the weight of the information measured by INS;

[0048] σ 1 Obtained by the following formula:

[0049] σ 1 =1-σ 2

[0050] σ 2 Obtained by the following formula:

[0051]

[0052] Where R 0 Indicates the initial relative distance between the projectile and the target when launching;

[0053] R′ represents the relative distance between the missile and the target obtained by coupling the aircraft position information measured by GPS and the target position information.

[0054] The beneficial effects of the present invention include:

[0055] (1) According to the guidance and control method of the extended-range aircraft with landing angle constraint provided by the present invention, the gliding distance is increased by controlling the gliding attitude in the mid-guidance stage, thereby greatly improving the range of the aircraft;

[0056] (2) According to the guidance and control method for an extended-range aircraft with angle of impact constraint provided by the present invention, a unique guidance law is set in the terminal guidance stage for guidance and control, so that the aircraft can hit the target at the desired angle of impact while increasing the range. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Figure 1 The trajectory diagram of the aircraft in Embodiment 1 of the present invention is shown;

[0058] Figure 2 The aircraft trajectory diagram according to embodiment 2 of the present invention is shown. DETAILED DESCRIPTION

[0059] The present invention will be further described in detail below through the accompanying drawings and embodiments. Through these descriptions, the characteristics and advantages of the present invention will become more clear and distinct.

[0060] The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise noted.

[0061] According to a guidance and control method for an extended-range aircraft with landing angle constraint provided by the present invention, the aircraft is a non-rolling aircraft. In this method,

[0062] After passing the highest point, the aircraft enters the gliding stage. In the gliding stage, the aircraft adjusts the angle of attack by deflecting the pitch rudder, so that the aircraft glides at a balanced angle of attack.

[0063] After entering the terminal guidance phase, the aircraft is guided and controlled by the guidance law with landing angle constraint.

[0064] Preferably, in the gliding section, the balance angle of attack is to control the angle of attack of the aircraft to be above a set value, and the set value is 18 to 25°, more preferably 20°; The inventors have found that by setting the critical value of the angle of attack, the range of the aircraft can be maximized, and the field of view angle limit of the seeker can also be guaranteed.

[0065] In the gliding stage, the required pitch rudder deflection angle is obtained in real time by the following formula (1):

[0066]

[0067] Among them, α represents the angle of attack, δ represents the pitch rudder deflection angle, C g represents the center of mass of the aircraft, C ym Indicates the aircraft pressure center, C yd represents the pitch rudder pressure center; τ represents the lift-to-drag ratio;

[0068] Preferably, the lift-to-drag ratio τ is obtained by the following formula (II):

[0069]

[0070] Among them, F y represents the total lift, and its calculation formula is Where α represents the angle of attack, δ z It indicates the detected rudder deflection angle. The angle of attack is obtained by the inertial navigation system, and the detected rudder deflection angle is obtained by the steering gear sensor. They represent the lift coefficient of the angle of attack and the lift coefficient of the rudder deflection angle, respectively. Both lift coefficients are obtained from the wind blowing experiment during design and are bound to the missile-borne computer before launch. x Indicates total resistance It represents the drag coefficient, which is obtained from the wind blowing experiment during design and is bound to the missile computer before launch. W represents the wing area, which is a known quantity during design.

[0071] In a preferred embodiment, the aircraft starts working according to the pre-launch programming time. Before launching, the experimenter will obtain the approximate time when the aircraft reaches the highest point according to the program, and bind the calculated programming time to the aircraft's onboard computer. When the aircraft captures the illuminator signal, the aircraft enters the terminal guidance phase.

[0072] In a preferred embodiment, in the terminal guidance stage, the guidance law with landing angle constraint obtains the required overload through the following formula (III):

[0073]

[0074] Among them, a M Indicates the required overload, through which the flight attitude of the aircraft in the terminal guidance phase is controlled;

[0075] Represents the relative velocity between the projectile and the target, which is obtained in real time through formula (IV).

[0076]

[0077] V a Indicates the aircraft speed;

[0078] represents the angular velocity of the missile-target line of sight, which is obtained in real time through formula (5),

[0079]

[0080] R represents the relative distance between the missile and the target, which is obtained by real-time calculation by the computer on board the aircraft. Before the aircraft is launched, the target position information and the position information of the aircraft launch point are loaded into the computer. During the flight of the aircraft, the position information of the aircraft itself is obtained in real time through the satellite signal receiving system and the inertial navigation system, that is, P a , so that the relative distance can be solved in real time;

[0081] K represents an adaptive parameter; its value is -100 to 100, preferably -20 to 30, and more preferably 5;

[0082] s represents the non-singular terminal sliding surface; obtained by formula (VI),

[0083] s=e 2 +β|e 1 |ρsign(e 1 ) (six)

[0084] sign(e 1 ) represents a symbolic function;

[0085] β represents the design parameter, and its value is β = 0.7;

[0086] ρ represents the design parameter, and its value is ρ=1.4;

[0087] e 1 Represents the angle deviation, e 2 represents the angular velocity deviation,

[0088] Among them, e 1 =λ-λ d ,

[0089] λ d It represents the expected terminal angle, that is, the angle at which the aircraft is expected to hit the target.

[0090] σ represents the aircraft lead angle, which is obtained by σ = θ-λ;

[0091] θ represents the ballistic inclination angle, which is obtained by real-time calculation. During the calculation process, the derivative of the ballistic inclination angle is first calculated. Then, the value of θ is obtained by solving the differential equation; further, in represents the derivative of the ballistic inclination angle, a M ′ represents the required overload calculated at the previous moment, that is, the trajectory inclination angle θ is obtained through real-time iteration. At the initial moment, the initial value of the required overload is selected as 0;

[0092] λ represents the sight angle between the missile and the target, which is obtained by real-time calculation by the computer on board the aircraft. Specifically, before launching, the target coordinates (x T ,y T ) is bound to the missile-borne computer. Since it is aimed at a fixed target, the coordinates of the target will not change. The coordinates of the aircraft are obtained in real time according to the GPS\INS integrated navigation system, and the coordinates are (P σx ,P σy ),according to Calculate the missile-target sight angle.

[0093] Further preferably, the aircraft is equipped with a GPS / INS combined navigation system, wherein the GPS part includes four synthetic antennas for receiving satellite signals, and also includes an anti-interference module for filtering the received satellite signals. The speed information and position information of the aircraft can be obtained through the GPS part in real time, but in actual working process, the information may be inaccurate due to interference; compared with the traditional conical antenna and the improved ring antenna, the four synthetic antennas not only have a stronger satellite signal receiving capability, but also have the characteristic of resisting high overload, and are more suitable for installation on high-range and high-overload aircraft.

[0094] The INS part can also obtain the speed and position information of the aircraft in real time, but the INS will produce error accumulation during the flight.

[0095] Therefore, the present application adopts the weighted output of GPS+INS to obtain the speed and position of the aircraft.

[0096] Specifically, accurate aircraft position and velocity information is obtained through the following formula:

[0097]

[0098] Among them, P a Indicates the position of the aircraft obtained by real-time solution;

[0099] V a Indicates the aircraft speed calculated in real time;

[0100] P G Indicates the aircraft position information obtained by GPS measurement;

[0101] P I Indicates the aircraft position information measured by INS;

[0102] V G Indicates the aircraft speed information measured by GPS;

[0103] V I Indicates the aircraft speed information measured by INS;

[0104] σ 1 Indicates the weight of the information obtained by GPS measurement;

[0105] σ 2 Indicates the weight of the information measured by INS;

[0106] σ 1 Obtained by the following formula:

[0107] σ 1 =1-σ 2

[0108] σ 2 Obtained by the following formula:

[0109]

[0110] Where R 0 It indicates the initial relative distance between the projectile and the target before launch and is bound in the aircraft's computer before the aircraft is launched.

[0111] R′ represents the relative distance between the missile and the target obtained by coupling the aircraft position information measured by GPS and the target position information.

[0112] Embodiment 1:

[0113] Select two identical aircraft and set the same launch velocity V m =750m / s, launch angle 53 degrees, launch the two aircrafts successively in the same direction with the same initial conditions;

[0114] Among them, for the first aircraft, after it passes the highest point, the angle of attack is adjusted by deflecting the pitch rudder by 5° so that its angle of attack is between 20° and 21°;

[0115] Its movement trajectory is Figure 1 As shown by the solid line in the figure, its angle of fall is 40° and its range is 82km.

[0116] For the second aircraft, no special control is performed after it passes the highest point, and its motion trajectory is as follows Figure 1 As shown by the dotted line in the figure, its angle of fall is 70° and its range is 26km.

[0117] It can be seen from the above embodiments that the extended-range aircraft guidance control method can significantly increase the range of the aircraft while reducing the landing angle.

[0118] Embodiment 2,

[0119] Set the distance between the target and the aircraft launch point to 40km, select 6 identical aircraft, and the initial launch velocity of the aircraft is V m =750m / s, the launch angle is 53 degrees, and the six aircraft are launched one after another with the same initial conditions.

[0120] Among them, the expected terminal landing angles of the six aircraft are λ d =30°, λ d =40°,λ d =50°,λ d =60°,λ d =70°,λ d =80°.

[0121] After all aircraft pass the highest point and before entering the terminal guidance phase, the angle of attack of the aircraft is adjusted by pitching the rudder 5 degrees to a value above 20 degrees.

[0122] When the six aircraft enter the terminal guidance phase, the required overload is calculated in real time by the following formula (III), and the aircraft is steered and controlled according to the required overload:

[0123]

[0124] in, The adaptive parameter K is set to 5, s = e2 +β|e 1 |ρsign(e 1 ); design parameters β = 0.7, ρ = 1.4; e 1 =λ-λ d ,

[0125] The final flight trajectories of the six aircraft are as follows: Figure 2 As shown in the figure, it can be seen from the trajectory diagram that all six aircraft hit the target 40 km away, and

[0126] The expected terminal landing angle is λ d =30°, the actual terminal landing angle is 29.5°;

[0127] The expected terminal landing angle is λ d =40°, the actual terminal landing angle is 40.3°;

[0128] The expected terminal landing angle is λ d =50°, the actual terminal landing angle is 50°;

[0129] The expected terminal landing angle is λ d = 60°, the actual terminal landing angle is 60.2°;

[0130] The expected terminal landing angle is λ d =70°, the actual terminal landing angle is 69.1°;

[0131] The expected terminal landing angle is λ d =80°, its actual terminal angle of attack is 78.3°. It can be seen from the above embodiment 2 that the extended-range aircraft guidance control method with angle of attack constraint can not only enable the aircraft to hit the target after increasing the range, but also select the desired terminal angle of attack as needed, so as to select a more appropriate terminal angle of attack according to the nature of the target.

[0132] The present invention has been described above in conjunction with preferred embodiments, but these embodiments are only exemplary and serve only as an illustration. On this basis, the present invention may be subjected to a variety of substitutions and improvements, all of which fall within the scope of protection of the present invention.

Claims

1. A guidance and control method for an extended-range aircraft with landing angle constraint, characterized in that: In this method, After passing the highest point, the aircraft enters the gliding stage. In the gliding stage, the aircraft adjusts the angle of attack by deflecting the pitch rudder, so that the aircraft glides at a balanced angle of attack. After the aircraft enters the terminal guidance phase, it is guided and controlled by the guidance law with landing angle constraint. In the gliding stage, the required pitch rudder deflection angle is obtained in real time by the following formula (1): Among them, α represents the angle of attack, δ represents the pitch rudder angle, and C g represents the center of mass of the aircraft, C ym Indicates the aircraft pressure center, C yd represents the pitch rudder pressure center; τ represents the lift-to-drag ratio; The lift-to-drag ratio τ is obtained by the following formula (II): Among them, F y Indicates the total lift, F x Indicates total elastic resistance; In the terminal guidance stage, the guidance law with landing angle constraint obtains the required overload through the following formula (III): Among them, a M Indicates that overload is required. Indicates the relative speed of the projectile and target; It represents the angular velocity of the missile-target line of sight; R represents the relative distance between the projectile and the target; K represents the adaptive parameter; s represents the non-singular terminal sliding surface; β represents the design gain; ρ represents the design gain; e1 represents the angle deviation; e2 represents the angular velocity deviation; σ represents the aircraft lead angle; The non-singular terminal sliding surface s is obtained by the following formula (VI): s=e2+β|e1| ρ sign(e1) (vi) sign(e1) represents the sign function.

2. The guidance and control method for an extended-range aircraft with landing angle constraint according to claim 1, characterized in that: In the gliding section, the balance angle of attack is to control the angle of attack of the aircraft to be above a set value, and the set value is 18 to 25 degrees.

3. The guidance and control method for an extended-range aircraft with landing angle constraint according to claim 1, characterized in that: The relative speed of the projectile Obtained in real time through the following formula (IV): The projectile-to-target line-of-sight angular velocity Obtained in real time through the following formula (V): Among them, V a Indicates the aircraft speed.

4. The guidance and control method for an extended-range aircraft with landing angle constraint according to claim 1, characterized in that: The angle deviation e1=λ-λ d , The angular velocity deviation Among them, λ represents the sight angle of the missile and the target, λ d Indicates the desired terminal angle.

5. The guidance and control method for an extended-range aircraft with landing angle constraint according to claim 3, characterized in that: The aircraft speed and position are obtained by the following formula (VII): Among them, P a Indicates the position of the aircraft; V a Indicates the aircraft speed; P G Indicates the aircraft position information obtained by GPS measurement; P I Indicates the aircraft position information measured by INS; V G Indicates the aircraft speed information measured by GPS; V I Indicates the aircraft speed information measured by INS; σ1 represents the weight of information obtained by GPS measurement; σ2 represents the weight of the information obtained by INS measurement; σ1 is obtained by the following formula: σ1=1-σ2 σ2 is obtained by the following formula: Where R0 represents the initial relative distance between the projectile and the target at the time of launch; R′ represents the relative distance between the missile and the target obtained by coupling the aircraft position information measured by GPS and the target position information.

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