Aircraft guidance control device and method with landing angle constraint

Abstract

The present invention discloses a method for guiding and controlling an aircraft with a drop angle constraint, comprising the following steps: establishing a motion model, determining a sliding surface, setting a convergence law, and obtaining an overload control instruction. The method is based on sliding mode variable structure control, and by setting a fast non-singular terminal sliding surface, the guidance law can converge quickly, so that the miss distance of the aircraft and the angular rate of the missile-to-target line of sight converge quickly to zero, and then the terminal drop angle of the aircraft when hitting the target is the same as the expected value. The method for guiding and controlling an aircraft with a drop angle constraint disclosed by the present invention can realize a large drop angle constraint, and the drop angle can be greater than 90°, can realize vertical attack or reverse slope strike, and has many advantages such as high strike accuracy, high stability, and short convergence time.

Description

Technical Field

[0001] The invention relates to an aircraft guidance control method, in particular to an aircraft guidance control method with landing angle constraint, and belongs to the field of aircraft control. Background Art

[0002] For most tactical aircraft, the role of the guidance law is to generate appropriate guidance commands to guide the aircraft to hit the target with zero miss. However, for some aircraft, in addition to specific requirements for miss, it is also necessary to attack the target at a specific angle at the time of impact to enhance the damage effect.

[0003] For example, when attacking highly defensive targets such as aircraft carriers, it is necessary to select weak points for attack based on the location of their deck defense armor. This can effectively destroy the core facilities under the deck, which requires the aircraft to have the ability to strike at a specified angle.

[0004] For example, in a sea landing operation, a large number of important enemy targets are located in reverse slope caves facing the Pacific Ocean with their backs to the mainland. This requires the aircraft to have the ability to hit the target at a large angle of fall.

[0005] Existing aircraft guidance and control methods are difficult to achieve large landing angles, or have obvious vibration at large landing angles, a serious decrease in accuracy, and poor overload resistance.

[0006] Therefore, it is necessary to propose an aircraft guidance and control method with landing angle constraints to solve the above problems. Summary of the invention

[0007] In order to overcome the above problems, the inventors have conducted in-depth research and proposed an aircraft guidance control method with landing angle constraint, which includes the following steps:

[0008] S1. Establish motion model;

[0009] S2, determine the sliding surface;

[0010] S3, set the reaching law;

[0011] S4. Obtain an overload control instruction.

[0012] Furthermore, in step S1, the motion model is a relative motion model between a fixed target and a projectile, which is expressed as:

[0013]

[0014] Among them, q represents the sight angle of the missile, represents the angular velocity of the missile-target line of sight, r represents the distance between the aircraft and the target, θ M Indicates the ballistic inclination angle of the aircraft, VM Indicates the current speed of the aircraft, a M Indicates aircraft overload instruction;

[0015] According to the motion model, two system state variables are set, namely:

[0016]

[0017] Among them, q F Indicates the desired terminal landing angle.

[0018] Further, in step S2, the sliding surface is a fast non-singular terminal sliding surface.

[0019] In a preferred embodiment, the sliding surface s is:

[0020]

[0021] Among them, β1, β2, α1, α2 are constants, and β1, β2>0, 1<α2<2, α1>α2.

[0022] In a preferred embodiment, in step S3, a unique reaching law is set to improve the convergence speed of the sliding mode control.

[0023] In a preferred embodiment, in step S3, the reaching law is expressed as:

[0024]

[0025] Among them, k1, k2, k3, k4 are constants.

[0026] In a preferred embodiment, in step S4, the aircraft overload control instruction can be obtained by taking the first-order derivative of the sliding surface and combining it with the reaching law.

[0027] The aircraft overload control instruction can be expressed as:

[0028]

[0029] On the other hand, the present invention also provides an aircraft guidance and control system with landing angle constraint, preferably using the above control method, the system includes a satellite navigation module, an INS module, a microprocessor and an actuator, characterized in that:

[0030] The microprocessor obtains information transmitted by the satellite navigation module and the INS module, generates a rudder deviation instruction according to the aircraft overload control instruction, and transmits the rudder deviation instruction to the actuator.

[0031] The aircraft overload control instruction is expressed as:

[0032]

[0033] Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity Obtained from the satellite navigation module; β1, β2, α1, α2, k1, k2, k3, k4 are constants, s represents the sliding surface, and x1 and x2 can be obtained by solving the motion model.

[0034] In a preferred embodiment, the motion model is expressed as:

[0035]

[0036] Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity The current speed V of the aircraft is obtained from the satellite navigation module. M Obtained from the INS module.

[0037] In a preferred embodiment, the actuator comprises an electric steering gear, and the electric steering gear is a proportional electric steering gear.

[0038] The beneficial effects of the present invention include:

[0039] (1) It can achieve large drop angle constraints, and the drop angle can be greater than 90°, and can achieve vertical attack or reverse slope strike. Therefore, it can be used in more combat scenarios and has higher application value.

[0040] (2) It has a unique convergence law that can greatly reduce buffeting, improve the strike accuracy and stability of the aircraft, and shorten the convergence time;

[0041] (3) The system uses four synthetic antennas to receive satellite signals, which has stronger signal reception capability and high overload resistance, making the received signal range wider and more accurate. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Figure 1 A schematic flow chart of an aircraft guidance control method with landing angle constraint according to a preferred embodiment of the present invention is shown;

[0043] Figure 2 A schematic structural diagram of an aircraft guidance and control system with landing angle constraint according to a preferred embodiment of the present invention is shown;

[0044] Figure 3 The simulation result diagram in Example 1 of the present invention is shown;

[0045] Figure 4 The simulation result diagram in Example 1 of the present invention is shown;

[0046] Figure 5 The figure shows the simulation results in Example 1 according to the present invention. DETAILED DESCRIPTION

[0047] 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.

[0048] 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.

[0049] In one aspect, the present invention provides a method for aircraft guidance and control with landing angle constraint, comprising the following steps:

[0050] S1. Establish motion model;

[0051] S2, determine the sliding surface;

[0052] S3, set the reaching law;

[0053] S4. Obtain an overload control instruction.

[0054] In step S1, the motion model is a relative motion model between a fixed target and a projectile, which is expressed as:

[0055]

[0056] Among them, q represents the sight angle of the missile, represents the angular velocity of the missile-target line of sight, r represents the distance between the aircraft and the target, θ M Indicates the ballistic inclination angle of the aircraft, V M Indicates the current speed of the aircraft, a M Indicates an aircraft overload command.

[0057] Further, according to formula (1), the motion model can also be expressed as:

[0058]

[0059] Furthermore, in the present invention, according to the motion model, two system state variables are set, namely x1 and x2, where x1 represents the difference between the measured missile-target sight angle and the expected terminal landing angle, and x2 represents the measured missile-target sight angle rate:

[0060]

[0061] Among them, q F Indicates the desired terminal landing angle.

[0062] In the present invention, the guidance control method is based on sliding mode structure control. In the sliding mode structure control, the selection of the sliding surface has a great influence on the control effect.

[0063] Furthermore, in step S2, the sliding surface is a fast non-singular terminal sliding surface. The fast non-singular terminal sliding surface has the advantages of strong robustness, high precision, fast convergence speed and no singular problems, so that the obtained guidance law can converge quickly, so that the aircraft's miss distance and the missile-target line of sight angular rate converge quickly to zero, and then the terminal landing angle of the aircraft when it hits the target is the same as the expected value.

[0064] More preferably, the sliding surface s is:

[0065]

[0066] Among them, β1, β2, α1, α2 are constants, and β1, β2>0, 1<α2<2, α1>α2.

[0067] Furthermore, the applicant has conducted a large number of experiments and combined with design experience to determine the parameters of the sliding surface: β1=β2=1, α1=3, α2=9 / 7. Under this parameter, the aircraft's miss distance and the convergence speed of the missile-target line of sight angular rate are both fast, and have extremely strong robustness and accuracy.

[0068] The inventors found that chattering is common in sliding mode variable structure control. This is mainly because the control instructions contain discontinuous switching control laws. When the sliding surface reaches the zero point, the actuator’s speed is difficult to decay to the zero point due to inertia, causing the system state to continuously cross the sliding surface.

[0069] Chattering can reduce the stability accuracy and robustness of the system and may even stimulate unmodeled dynamics and destroy system stability.

[0070] In the present invention, in step S3, a unique reaching law is set to improve the convergence speed of the sliding mode variable control.

[0071] In step S3, the reaching law is expressed as:

[0072]

[0073] Among them, k1, k2, k3, k4 are constants.

[0074] In a preferred real-time mode, k1=800, k2=999, k3=800, k4=5.

[0075] In step S4, the overload control instruction is obtained according to the sliding surface and the reaching law.

[0076] Specifically, by taking the first-order derivative of the sliding surface and combining it with the reaching law, the aircraft overload control command can be obtained.

[0077] The aircraft overload control instruction can be expressed as:

[0078]

[0079] On the other hand, the present invention also provides an aircraft guidance and control system with landing angle constraint, which uses the above control method to achieve control of the aircraft.

[0080] The system includes a satellite navigation module, an INS module, a microprocessor and an actuator.

[0081] The satellite navigation module receives and resolves satellite signals to obtain the distance and angle information of the aircraft relative to the target, and transmits the information to the microprocessor;

[0082] Furthermore, the satellite navigation module transmits information including the distance r between the aircraft and the target, the trajectory inclination θ of the aircraft M and the missile-to-target line-of-sight angular velocity

[0083] The INS module includes a gyroscope and an accelerometer. The gyroscope measures the attitude of the aircraft, and the accelerometer measures the acceleration of the aircraft. The position and velocity information of the aircraft are obtained by numerically integrating the acceleration. The missile-target sight angle q is calculated by the position of the aircraft and the position of the target, and the information is transmitted to the microprocessor.

[0084] Furthermore, the information transmitted by the INS module includes the current speed V of the aircraft. M .

[0085] The microprocessor obtains information transmitted by the satellite navigation module and the INS module, obtains an aircraft overload control instruction according to the above control method, generates a rudder deviation instruction according to the aircraft overload control instruction, and transmits the rudder deviation instruction to the actuator;

[0086] The actuator includes an electric servo, which can generate corresponding torque according to the rudder deviation instruction, so as to change the attitude of the aircraft.

[0087] In a preferred embodiment, the electric servo is a proportional electric servo, which has better high overload resistance than traditional pneumatic servos, especially in control with landing angle constraints. When the aircraft is in a large maneuvering condition, precise control of the aircraft attitude can be achieved.

[0088] The inventors have found that at a large landing angle, the aircraft has a high overload, and at this time, the traditional satellite antenna has a poor ability to receive satellite signals.

[0089] In a preferred embodiment, the satellite navigation module uses a four-piece synthetic antenna to receive signals from satellites. The four-piece synthetic antenna is one of the antennas commonly used in the aviation field, and its specific model and structure are not particularly limited in the present invention.

[0090] More preferably, an anti-interference module is also provided in the satellite navigation module, and the satellite signal received by the antenna is filtered by the anti-interference module to weaken the influence of interference signals and noise. Any known navigation anti-interference module can be used, and there is no special limitation in the present invention.

[0091] According to the present invention, in the microprocessor, the aircraft overload control instruction is expressed as:

[0092]

[0093] Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity Obtained from the satellite navigation module; β1, β2, α1, α2, k1, k2, k3, k4 are constants, and s represents the sliding surface.

[0094] Preferably, the sliding surface is:

[0095]

[0096] Among them, x1 and x2 can be obtained by solving the motion model.

[0097]

[0098] The motion model is expressed as:

[0099]

[0100] Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity The current speed V of the aircraft is obtained from the satellite navigation module. M Obtained from the INS module.

[0101] Example

[0102] Example 1

[0103] Carry out simulation experiment and set the target coordinate as x t =10000m,y t=0, initial velocity of aircraft V m =250m / s, angle of incidence θ=55°, and the terminal angles are set to be θ f =80 ° ,θ f =90 ° and θ f =100 ° , perform simulation as follows:

[0104] S1. Establish motion model;

[0105] S2, determine the sliding surface;

[0106] S3, set the reaching law;

[0107] S4. Obtain an overload control instruction.

[0108] In step S1, the motion model is:

[0109]

[0110] In step S2, the sliding surface s is:

[0111]

[0112] Among them, β1=β2=1, α1=3, α2=9 / 7.

[0113] In step S3, the reaching law is:

[0114]

[0115] Among them, k1=800, k2=999, k3=800, k4=5.

[0116] In step S4, the aircraft overload control instruction is:

[0117]

[0118] The simulation results are as follows Figure 3 shown.

[0119] Example 2

[0120] The same simulation experiment as in Example 1 is performed, except that the target coordinate is set to x t =15000m,y t =0, initial velocity of aircraft V m =500m / s, angle of fire θ=55 ° .

[0121] The simulation results are as follows Figure 4 shown.

[0122] Example 3

[0123] The same simulation experiment as in Example 1 is performed, except that the target coordinate is set to x t =23000m,y t =0, initial velocity of aircraft V m =850m / s, angle of incidence θ=55°.

[0124] The simulation results are as follows Figure 5 shown.

[0125] It can be seen from the simulation result diagrams of Examples 1 to 3 that, under different simulation conditions, the methods in the embodiments can achieve a specified angle strike on a fixed target, and the landing angle can be greater than 90°, so that the aircraft can be used in combat scenarios that require large landing angle attacks, such as vertical strikes or reverse slope strikes.

[0126] In the description of the present invention, it should be noted that the terms "upper", "lower", "inner", "outer", "front", "rear", etc. indicate positions or positional relationships based on the working state of the present invention, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first", "second", "third", and "fourth" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0127] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected" and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0128] 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 method for guiding and controlling an aircraft with a landing angle constraint, comprising the following steps: S1. Establish motion model; S2, determine the sliding surface; S3, set the reaching law; S4, obtaining an overload control instruction; In step S1, the motion model is a relative motion model between a fixed target and a projectile, which is expressed as: Among them, q represents the sight angle of the missile, represents the angular velocity of the missile-target line of sight, r represents the distance between the aircraft and the target, θ M Indicates the ballistic inclination angle of the aircraft, V M Indicates the current speed of the aircraft, a M Indicates aircraft overload control instructions; According to the motion model, two system state variables are set, namely: Among them, q F Indicates the desired terminal landing angle; The sliding surface s is: Among them, β1, β2, α1, α2 are constants, and β1, β2>0, 1<α2<2, α1>α2; In step S3, the reaching law is expressed as: Among them, k1, k2, k3, k4 are constants; In step S4, the aircraft overload control instruction can be obtained by taking the first-order derivative of the sliding surface and combining it with the reaching law. The aircraft overload control instruction can be expressed as:

2. The method for aircraft guidance and control with landing angle constraint according to claim 1, characterized in that: In step S2, the sliding surface is a fast non-singular terminal sliding surface.

3. The method for aircraft guidance and control with landing angle constraint according to claim 1, characterized in that: In step S3, the convergence speed of the sliding mode control is improved by setting a unique reaching law.

4. An aircraft guidance and control system with landing angle constraint, adopting the control method as claimed in any one of claims 1 to 3, the system comprises a satellite navigation module, an INS module, a microprocessor and an actuator, characterized in that: The microprocessor obtains information transmitted by the satellite navigation module and the INS module, generates a rudder deviation instruction according to the aircraft overload control instruction, and transmits the rudder deviation instruction to the actuator. The aircraft overload control instruction is expressed as: Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity Obtained from the satellite navigation module; β1, β2, α1, α2, k1, k2, k3, k4 are constants, s represents the sliding surface, and x1 and x2 can be obtained by solving the motion model.

5. The aircraft guidance and control system with landing angle constraint according to claim 4, characterized in that: The motion model is expressed as: Among them, the distance between the aircraft and the target r, the sight angle q, and the trajectory inclination angle θ of the aircraft M and the missile-to-target line-of-sight angular velocity The current speed V of the aircraft is obtained from the satellite navigation module. M Obtained from the INS module.

6. The aircraft guidance and control system with landing angle constraint according to claim 4, characterized in that: The actuator comprises an electric steering gear, and the electric steering gear is a proportional electric steering gear.

Images