CN115113640A - Extended range aircraft guidance control method with falling angle constraint - Google Patents
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Abstract
The invention discloses a range-extended aircraft guidance control method with a landing angle constraint function.
Description
Technical Field
The invention relates to the field of aircraft guidance control, in particular to a range-extending aircraft guidance control method with a falling angle constraint function.
Background
With the development of technology, guidance aircrafts are receiving much attention due to the high precision characteristic thereof. The traditional guidance aircraft adopts a proportional guidance law, and can complete the task of colliding and striking targets. Considering the range problem of the aircraft, the traditional gun-launched aircraft experiences a high overload state when launched, the element has poor high overload resistance, so that in the inner trajectory, the aircraft element is damaged and fails, and the range of the aircraft is limited.
Due to the contradictory relationship between the range and the large landing angle, the inventor of the present invention has made intensive studies on the existing aircraft guidance control scheme in order to wait for the design of a new guidance control method capable of solving the above-mentioned problems.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention makes a keen study and designs a range-extending aircraft guidance control method with a landing angle constraint, in the method, in a middle guidance section, a pitching steering engine is controlled to steer, the attack angle of an aircraft is improved to enable the aircraft to be capable of gliding farther, when the aircraft enters a final guidance section, a guidance law based on a sliding mode surface is set, the required overload of the aircraft is solved in real time under the condition of setting a desired landing angle, the aircraft is controlled to fly according to the overload, and finally the aircraft is controlled to land according to the desired landing angle and hit a target under the condition of ever increasing the range of the aircraft, so that the invention is completed.
In particular, the invention aims to provide a range-extended aircraft guidance control method with a landing angle constraint, wherein in the method,
the aircraft enters the gliding section after passing through the highest point, the aircraft adjusts the attack angle in the gliding section through the deflection of the pitching rudder, so that the aircraft glides and flies under the condition of balancing the attack angle,
and after the aircraft enters the final guide section, the aircraft is guided and controlled through a guidance law with a falling angle constraint.
In the glide section, the balance attack angle is used for controlling the attack angle of the aircraft to be more than a set value, the set value is 18-25 degrees, and the selected angle is 20 degrees.
Wherein, in the glide section, a required pitch rudder deflection angle is obtained in real time by the following formula (one):
wherein alpha represents an angle of attack, delta represents a pitch rudder deflection angle, C g Representing the aircraft center of mass, C ym Indicating aircraft core pressure, C yd Representing pitch rudder pressure center; τ represents a lift-to-drag ratio;
preferably, the lift-to-drag ratio τ is obtained by the following formula (two):
wherein, F y Indicating full spring lift, F x Indicating full spring resistance.
Wherein, in the final guidance stage, the guidance law with the falling angle constraint obtains the overload demand through the following formula (three):
wherein, a M It is indicated that an overload is required,
r represents the relative distance of the bullet eyes;
k represents an adaptive parameter;
s represents a nonsingular terminal sliding mode surface;
β represents a design gain;
ρ represents a design gain;
e 1 indicating an angular deviation;
e 2 indicating an angular velocity deviation;
σ denotes the aircraft nose angle.
Wherein the relative speed of the eyes
Obtained in real time by the following formula (iv):
the angular velocity of the line of sight of the bullet
Obtained in real time by the following formula (five):
wherein, V a Representing the aircraft speed.
Wherein the nonsingular terminal sliding mode surface s is obtained by the following formula (six):
s＝e 2 +β|e 1 |ρsign(e 1 ) And (VI).
sign(e 1 ) Representing a symbolic function;
wherein the angular deviation e 1 ＝λ-λ d
Deviation of angular velocity
Wherein λ represents the viewing angle of the bullet eye, λ d Indicating the desired terminal angle.
Wherein the aircraft speed and position are obtained by the following equation (seven):
wherein, P a Representing an aircraft position;
V a representing the aircraft speed;
P G representing aircraft position information measured by a GPS;
P I representing aircraft position information measured by the INS;
V G representing aircraft speed information measured by a GPS;
V I representing aircraft speed information measured by the INS;
σ 1 representing the weight occupied by the information obtained by GPS measurement;
σ 2 representing the weight occupied by the information measured by the INS;
σ 1 obtained by the following formula:
σ 1 ＝1-σ 2
σ 2 obtained by the following formula:
in the formula R 0 Representing the initial relative distance of the missile eyes when the missile eyes are launched;
r' represents the relative distance between the bullet and the target by coupling the aircraft position information measured by GPS and the target position information of the filling.
The invention has the advantages that:
(1) according to the control method for the extended range aircraft with the falling angle constraint, provided by the invention, the gliding distance is increased by controlling the gliding posture in the middle guidance section, so that the range of the aircraft is greatly increased;
(2) according to the extended range aircraft guidance control method with the falling angle constraint, the guidance control is carried out on the final guidance section by setting a unique guidance law, so that the aircraft can hit a target according to an expected falling angle on the basis of increasing the range.
Drawings
FIG. 1 illustrates a trajectory diagram of an aircraft in accordance with embodiment 1 of the present invention;
fig. 2 shows a trajectory diagram of an aircraft according to embodiment 2 of the present invention.
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
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. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the invention, the extended range aircraft guidance control method with the falling angle constraint is provided, the aircraft is a non-rolling aircraft, in the method,
the aircraft enters the gliding section after passing through the highest point, the aircraft adjusts the attack angle in the gliding section through the deflection of the pitching rudder, so that the aircraft glides and flies under the condition of balancing the attack angle,
and after the aircraft enters the final guide section, the aircraft is guided and controlled through a guidance law with a falling angle constraint.
Preferably, in the gliding section, the balance attack angle is a value which controls the attack angle of the aircraft to be more than a set value, and the set value is 18-25 degrees, more preferably 20 degrees; the inventor finds that by setting the critical value of the angle of attack, the range of the aircraft can be improved to the maximum extent, and the angle of view limitation of the seeker can be ensured.
In the glide section, the required pitching rudder deflection angle is obtained in real time through the following formula (one):
where α represents the angle of attack, δ represents the pitch rudder angle, C g Representing the aircraft center of mass, C ym Indicating aircraft core pressure, C yd Representing pitch rudder pressure center; τ represents a lift-to-drag ratio;
preferably, the lift-to-drag ratio τ is obtained by the following formula (two):
wherein, F y The total elastic lift force is expressed by the formula
Wherein α represents an angle of attack, δ z And the detected rudder deflection angle is represented, the attack angle is obtained by an inertial navigation system, and the detected rudder deflection angle is obtained by a steering engine sensor.
The method is characterized in that attack angle lift coefficient and rudder deflection angle lift coefficient are respectively represented, the two lift coefficients are obtained by a blowing experiment during design and are bound in an missile-borne computer before launching. F x Indicating the total resistance
Representing the drag coefficient, obtained from the design-time blow-through experiment, already bound in the missile-borne computer before launch. W represents the missile wing area, which is a known quantity at design time.
In a preferred embodiment, the aircraft is operated according to a pre-launch journey time, and prior to launch, the experimenter is programmed to obtain a rough time for the aircraft to peak and to order the calculated journey time to the aircraft's missile computer. When the illuminator signal is captured by the aircraft, the aircraft enters the terminal guidance segment.
In a preferred embodiment, in the final guidance period, the guidance law with the falling angle constraint obtains the overload demand by the following formula (III)
Wherein, a M Indicating the requirement overload, and controlling the flight attitude of the aircraft at the tail guide section through the requirement overload;
V a representing the aircraft speed;
r represents the order of the bulletThe relative distance is obtained by real-time calculation of a computer carried on the aircraft; wherein, before the aircraft is launched, the target position information and the position information of the aircraft launching point are both filled into the computer, and in the flight process of the aircraft, the position information of the aircraft, namely P, is obtained in real time through the satellite signal receiving system and the inertial navigation system a So that the relative distance can be calculated in real time;
k represents an adaptive parameter; the value is-100, preferably-20-30, and more preferably 5;
s represents a nonsingular terminal sliding mode surface; obtained by the formula (VI) and the formula,
s＝e 2 +β|e 1 |ρsign(e 1 ) (VI)
sign(e 1 ) Representing a symbolic function;
beta represents a design parameter, and the value of beta is 0.7;
ρ represents a design parameter, and the value of ρ is 1.4;
e 1 indicates the angular deviation, e 2 The deviation of the angular velocity is represented by,
wherein e is 1 ＝λ-λ d ，
λ d Indicating a desired terminal angle at which the aircraft is expected to hit the target.
σ represents the aircraft nose angle, obtained by θ - λ;
theta represents the ballistic inclination angle, which is calculated in real time, and in the calculation process, the derivative of the ballistic inclination angle is firstly calculated
Then solving a differential equation to obtain a value theta; further, the air conditioner is provided with a fan,
wherein
Representing the derivative of the ballistic inclination angle, a M ' indicating the overload to be used which is calculated at the last moment, namely the trajectory inclination angle theta is obtained through real-time iteration, and selecting the initial value of the overload to be used as 0 at the initial moment;
lambda denotes the viewing angle of the bullet eye, obtained by real-time calculation by a computer onboard the aircraft, and specifically the coordinates (x) of the target are taken before launch T ,y T ) When bound to a pop-up computer, the coordinates of the object do not change because the object is fixed. The coordinate of the aircraft is obtained in real time according to the GPS/INS integrated navigation system, and the coordinate is (P) σx ,P σy ) According to
And calculating to obtain the visual line angle of the bullet.
Further preferably, a GPS/INS integrated navigation system is mounted on the aircraft, wherein the GPS section includes four synthetic antennas for receiving satellite signals, and an anti-jamming module for performing filtering processing on the received satellite signals, and obtaining speed information and position information of the aircraft through the GPS section, but in an actual working process, the information may be misaligned due to jamming; compared with the traditional conical antenna and the improved loop antenna, the four-piece combined antenna has stronger satellite signal receiving capability and high overload resistance, and is more suitable for being installed on an aircraft with a large range and a high overload.
The INS part can also obtain speed information and position information of the aircraft in real time, but the INS generates error accumulation during the flight process.
Therefore, the speed and the position of the aircraft are obtained by weighting the output of the GPS + INS.
Specifically, accurate aircraft position and velocity information is obtained by:
wherein, P a Representing the aircraft position obtained by real-time solution;
V a representing the real-time calculated aircraft speed;
P G representing aircraft position information measured by a GPS;
P I representing aircraft position information measured by the INS;
V G representing aircraft speed information measured by a GPS;
V I representing aircraft speed information measured by the INS;
σ 1 representing the weight occupied by the information obtained by GPS measurement;
σ 2 representing the weight occupied by the information measured by the INS;
σ 1 obtained by the following formula:
σ 1 ＝1-σ 2
σ 2 obtained by the following formula:
in the formula R 0 Representing the initial relative distance of the projectile prior to launch, is mounted in the aircraft's computer prior to launch.
R' represents the relative distance between the bullet and the target by coupling the aircraft position information measured by GPS and the target position information of the filling.
Example 1:
selecting two same aircrafts and setting the same initial launching speed V m 750m/s, launch angle 53 degrees, launch these two aircrafts one after the other towards the same direction with the same initial conditions;
after the first aircraft passes through the highest point, the first aircraft deflects by 5 degrees through the pitching rudder to adjust the attack angle, so that the attack angle is 20-21 degrees;
the motion locus is shown by a solid line in fig. 1, the falling angle is 40 degrees, and the range is 82 km.
The second aircraft, which was not specifically controlled after passing through the highest point, had a trajectory of motion as indicated by the dashed line in fig. 1, a landing angle of 70 ° and a range of 26 km.
According to the embodiment, the range-extended aircraft guidance control method can obviously improve the range of the aircraft under the condition of reducing the falling angle.
In the case of the example 2, the following examples are given,
setting the distance between the target and the launching point of the aircraft to be 40km, selecting 6 same aircraft, and setting the launching initial speed of the aircraft to be V m The 6 aircraft were launched in succession under the same initial conditions, with an angle of launch of 53 degrees, 750 m/s.
Wherein the desired terminal falling angles of the 6 aircrafts are respectively lambda d ＝30°、 λ d ＝40°、λ d ＝50°、λ d ＝60°、λ d ＝70°、λ d ＝80°。
After the aircrafts pass through the highest point and before entering the final guide section, the attack angle of the aircrafts is adjusted by deflecting the pitching rudder by 5 degrees, so that the attack angle of the aircrafts is more than 20 degrees;
when 6 aircrafts enter a final pilot segment, calculating the required overload in real time through the following formula (III), and controlling the aircrafts by steering according to the required overload:
wherein,
the adaptive parameter K takes the value 5, s-e 2 +β|e 1 |ρsign(e 1 ) (ii) a The design parameter beta is 0.7, and rho is 1.4; e.g. of a cylinder 1 ＝λ-λ d ， 
The flight trajectories of the 6 aircraft are shown in fig. 2, and it can be seen from the trajectory graph that the 6 aircraft hit targets outside 40km, and
desired terminal fall angle of λ d 30 ° for an aircraft with a true terminal landing angle of 29.5 °;
desired terminal fall angle of λ d A 40 ° aircraft with a true terminal landing angle of 40.3 °;
desired terminal fall angle of λ d A 50 ° aircraft with a true terminal landing angle of 50 °;
desired terminal fall angle of λ d 60 ° for an aircraft, its true terminal landing angle is 60.2 °;
desired terminal fall angle of λ d 70 ° aircraft with a true terminal landing angle of 69.1 °;
desired terminal fall angle of λ d An 80 ° aircraft has a true terminal landing angle of 78.3 °. It can be known from the above embodiment 2 that the extended range aircraft guidance control method with the falling angle constraint not only enables the aircraft to hit the target after increasing the range, but also enables the aircraft to select the desired terminal falling angle as required, so that a more appropriate terminal falling angle can be selected according to the property of the target.
The present invention has been described above in connection with preferred embodiments, which are intended to be exemplary only and illustrative only. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.
Claims (8)
1. A range-extended aircraft guidance control method with a landing angle constraint is characterized in that in the method,
the aircraft enters the gliding section after passing through the highest point, the aircraft adjusts the attack angle in the gliding section through the deflection of the pitching rudder so that the aircraft glides and flies under the condition of balancing the attack angle,
and after the aircraft enters the final guidance section, the aircraft performs guidance control through a guidance law with a falling angle constraint.
2. The extended range aircraft guidance control method with the landing angle constraint of claim 1,
in the gliding section, the balance attack angle is used for controlling the attack angle of the aircraft to be more than a set value, and the set value is 18-25 degrees, and is preferably 20 degrees.
3. The extended range aircraft guidance control method with the landing angle constraint of claim 2,
in the gliding section, the required pitching rudder deflection angle is obtained in real time through the following formula (one):
wherein alpha represents an angle of attack, delta represents a pitch rudder deflection angle, C g Representing the aircraft center of mass, C ym Indicating aircraft core pressure, C yd Representing pitch rudder pressure center; τ represents a lift-to-drag ratio;
preferably, the lift-to-drag ratio τ is obtained by the following formula (two):
wherein, F y Indicating full spring lift, F x Indicating full spring resistance.
4. The extended range aircraft guidance control method with the landing angle constraint according to claim 1,
in the final guidance segment, the guidance law with the falling angle constraint obtains the required overload through the following formula (three):
wherein, a M It is indicated that an overload is required,
r represents the relative distance of the bullet;
k represents an adaptive parameter;
s represents a nonsingular terminal sliding mode surface;
β represents a design gain;
ρ represents a design gain;
e 1 indicating an angular deviation;
e 2 indicating an angular velocity deviation;
σ denotes the aircraft nose angle.
5. The extended range aircraft guidance control method with the landing angle constraint of claim 4,
relative speed of the bullet
Obtained in real time by the following formula (iv):
angular velocity of the line of sight of the bullet
Obtained in real time by the following formula (five):
wherein, V a Representing the aircraft speed.
6. The extended range aircraft guidance control method with the landing angle constraint of claim 4,
the nonsingular terminal sliding mode surface s is obtained by the following formula (six):
s＝e 2 +β|e 1 | ρ sign(e 1 ) (VI)
sign(e 1 ) Representing a symbolic function.
7. The extended range aircraft guidance control method with the falling angle constraint according to claim 4 or 6,
the angular deviation e 1 ＝λ-λ d ，
Deviation of angular velocity
Wherein λ represents the viewing angle of the bullet eye, λ d Indicating the desired terminal angle.
8. The extended range aircraft guidance control method with the landing angle constraint of claim 4,
the aircraft speed and position are obtained by the following equation (seven):
wherein, P a Representing an aircraft position;
V a representing the aircraft speed;
P G representing aircraft position information measured by a GPS;
P I representing aircraft position information measured by the INS;
V G representing aircraft speed information measured by a GPS;
V I representing aircraft speed information measured by the INS;
σ 1 to representThe weight occupied by the information obtained by GPS measurement;
σ 2 representing the weight occupied by the information measured by the INS;
σ 1 obtained by the following formula:
σ 1 ＝1-σ 2
σ 2 obtained by the following formula:
in the formula R 0 Representing the initial relative distance of the missile eyes when the missile eyes are launched;
r' represents the relative distance between the bullet and the target, coupled by the GPS measurement of the aircraft position and the target position of the filling.
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| CN110032206A (en) * | 2019-05-06 | 2019-07-19 | 北京理工大学 | Top control method and control system are attacked in the big angle of fall of long-range guidance aircraft |
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| CN111679687A (en) * | 2020-05-22 | 2020-09-18 | 北京空天技术研究所 | Guide control integration method with falling angle constraint |
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