CN112526872B - Method for processing guidance and terminal guidance handover and guidance information in constraint with large falling angle - Google Patents

Method for processing guidance and terminal guidance handover and guidance information in constraint with large falling angle Download PDF

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CN112526872B
CN112526872B CN202011407568.7A CN202011407568A CN112526872B CN 112526872 B CN112526872 B CN 112526872B CN 202011407568 A CN202011407568 A CN 202011407568A CN 112526872 B CN112526872 B CN 112526872B
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CN112526872A (en
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贾庆忠
刘俊辉
单家元
刘永善
丁艳
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Beijing Institute of Technology BIT
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
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Abstract

The invention relates to a method for processing guidance and terminal guidance handover and guidance information in large-falling-angle constraint, and belongs to the technical field of guidance aircrafts. The invention aims to provide a method for processing guidance and terminal guidance handover and guidance information in large-falling-angle constraint, which is used for processing and obtaining guidance information required by middle guidance and terminal guidance according to the position, speed, attitude information and the like of a projectile body measured by inertial navigation; the method comprises the steps of starting a guidance system, starting guidance, and stopping guidance; the aircraft can fly farther in the middle guidance section, a large drop angle is formed in the final guidance section, and simultaneously, the target can be hit accurately.

Description

Method for processing guidance and terminal guidance handover and guidance information in constraint with large falling angle
Technical Field
The invention relates to a method for processing guidance and terminal guidance handover and guidance information in large-falling-angle constraint, and belongs to the technical field of guidance aircrafts.
Background
In order to improve the hitting capability of a deep-buried reinforcement target, a target under a bridge opening and a target in a building in the air and the throwing capability outside a vehicle defense area, the development of a guidance aircraft with a long range, a large falling angle, a large falling speed and good maneuvering performance is urgently required. The method enriches the hitting means of the open space continuously, enhances the hitting capability, and particularly develops the capability of accurately hitting firm and deep targets of enemies.
In order to meet the requirements of range, falling angle and falling speed of a guided aircraft, a new aerodynamic layout, a new trajectory scheme and a new guidance control strategy need to be designed. A guidance section in a conventional guidance aircraft adopts an uncontrolled trajectory, so that the flight distance is short and the launching area is small. In order to realize that the guidance aircraft has certain high-altitude gliding capability and meet the maneuvering capability of low-altitude gliding, a duck-type pneumatic layout scheme, namely a folding empennage and a normal-type pneumatic layout of a tailless type rudder with a trailing edge, is adopted. The method adopts a ballistic scheme combining an initial stable section, a gliding middle-made guide section and a large-falling-angle final guide section. And the middle guidance section adopts a trajectory inclination angle tracking scheme which takes trajectory inclination angles, throwing angles, gravity compensation and the like as polynomials. The final guidance adopts a sliding mode variable structure guidance law with large falling angle constraint, which is combined by proportional guidance, line-of-sight angle constraint, gravity compensation and sliding mode control items.
The control targets of the middle guidance and the terminal guidance are different from the required measurement information, and the middle guidance and terminal guidance junction has a large influence on the trajectory track and the terminal guidance effect, so that a reasonable middle guidance and terminal guidance junction strategy and a guidance information processing method need to be designed.
Disclosure of Invention
The invention aims to provide a method for processing guidance and terminal guidance handover and guidance information in large-falling-angle constraint, which is used for processing and obtaining guidance information required by middle guidance and terminal guidance according to the position, speed, attitude information and the like of a projectile body measured by inertial navigation; the method comprises the steps of starting a guidance system, starting guidance, and stopping guidance; the aircraft can fly farther in the middle guidance section, a large drop angle is formed in the final guidance section, and simultaneously, the target can be hit accurately.
The purpose of the invention is realized by the following technical scheme.
The method for processing guidance and terminal guidance handover and guidance information in constraint with a large falling angle specifically comprises the following steps:
step one, position information X of guided missile on ground coordinate system by using navigation measurementm,Ym,ZmVelocity information Vx,Vy,VzAnd a known target position X in the ground coordinate systemt,Yt,ZtAnd calculating to obtain the high and low sight line angle q under the ground coordinate systemesAzimuth line of sight angle qbsAnd high and low line of sight angular velocities
Figure BDA0002818014890000011
And azimuth line-of-sight angular velocity
Figure BDA0002818014890000012
Xg＝Xt-Xm,Yg＝Yt-Ym,Zg＝Zt-Zm (1)
Figure BDA0002818014890000021
Figure BDA0002818014890000022
Figure BDA0002818014890000023
Wherein
Figure BDA0002818014890000024
Tanx＝(Yg/Sqrtx). Relative position coordinate (X) of target relative to missileg,Yg,Zg)。
The ground coordinate system takes the emitting point as the origin of coordinates, takes the connecting line of the projection of the emitting point on the horizontal plane and the target point as the X axis, and takes the direction facing the target point as the positive direction; taking the direction vertical to the horizontal plane as the Y axis and the direction towards the sky as the positive direction; the coordinate axis Z forms a right-hand coordinate system with the X and Y axes.
Step two, measuring information by using navigation equipment and calculating the relative distance R of the bullet eyesgThe synthesis velocity VcBallistic declination psiVAnd ballistic inclination angle thetadAnd (4) information.
Figure BDA0002818014890000025
Figure BDA0002818014890000026
θd＝atan(Vy/Vx) (6)
ψV＝atan(-Vz/Vx) (7)
And step three, before the guided aircraft passes through the maneuvering point, all real-time measurement information is obtained from a navigation system, and gliding flight is carried out according to an overload instruction generated by a guidance law in the following process.
Uf1＝a-H/Hr-b×Vr/Vc+57.3c×Q0-57.3θd+d×cos(θd) (8)
Figure BDA0002818014890000027
Wherein U isf1And Uf2Respectively pitch and yaw direction overload commands, and the relative height H is Yg，Q0For ballistic inclination at the moment of delivery, VrIs the nominal flying speed of the aircraft, HrAnd a, b, c, d, l and m are coefficients.
And step four, judging whether the current state information meets the shift switching condition of the maneuvering point.
Judging the current high and low sight line angle qesAngle q with desired tip falldWhether the difference is greater than or equal to qε(ii) a Judging the current flying height YgY or morezh；qεAnd YzhIs a parameter preset according to requirements; if the current flight state information simultaneously meets the two conditions, the aircraft reaches a maneuvering point, and then enters a final guidance control section; otherwise, entering a middle guidance control section, namely performing gliding flight according to the method in the third step;
step five, entering a final guidance control section; and finishing the switching from the middle guidance section to the final guidance section.
And 5.1, switching guidance instructions from the middle guidance section to the final guidance section in an exponential fade-in mode, and ensuring stable switching of the guidance instructions.
Rksa＝e(-(t1-t1a)/1.0)
Uf1b＝Uf1×Rksa+Acc_y×(1-Rksa) (10)
Uf2b＝Uf2×Rksa+Acc_z×(1-Rksa)
Wherein t1 is the current flight time given by the missile-borne computer, t1a is the time of just entering the maneuvering point, the Rksa instruction cross-linking index coefficient, Uf1、Uf2The pitch and yaw direction overload instructions of the middle guidance section are Acc _ y and Acc _ z, the pitch and yaw direction overload instructions of the final guidance section are Uf1b and Uf2b, and the pitch and yaw direction overload instructions are merged by instructions output by the guidance system.
And 5.2, in the final guidance stage, all real-time measurement information is obtained and processed from the navigation system, and the vertical direction and the lateral direction fly according to overload instructions generated by the following final guidance law respectively.
Acc_y＝(Acc1+Acc2+Acc3+Ug)×Rkk (11)
Figure BDA0002818014890000031
Wherein l1、m1Rkk is a proportionality coefficient, Acc1, Acc2, Acc3, UgThe overload instruction components are calculated in the steps 5.3-5.6.
Step 5.3, calculating a proportion guidance item overload command Acc 1:
Figure BDA0002818014890000032
wherein Xk _ bili, Clamda are coefficients.
Rb_l1＝Rg/(Rg+d1) (14)
Wherein d is1Are coefficients.
Step 5.4, calculating an overload command Acc2 of the corner constraint item:
Acc2＝Vc×(Xk_bili+1)×Clamda×Rb_l2×(qes-qd)/(Rg+d2)/57.3 (15)
Rb_l2＝Rg/(Rg+d2) (16)
where Rb _ l2 is an intermediate variable, d2Are coefficients.
Step 5.5, calculating an overload instruction Acc3 of the sliding mode control item:
Acc3＝Epsl×Sgns×Rb_l1 (17)
where Epsl is a coefficient.
The calculation formula of the saturation function term Sgns is as follows:
Figure BDA0002818014890000041
Figure BDA0002818014890000042
step 5.6, calculating an overload instruction U of the gravity compensation itemg：
Ug＝9.8×cos(θd)＝9.8×cos(atan(Vy/Vx)) (20)
Wherein theta isdIs the ballistic inclination angle.
Advantageous effects
The invention provides a middle guidance and terminal guidance handover strategy based on navigation measurement output information and a guidance information processing method. The guidance law in the large-drop-angle guidance aircraft based on the trajectory inclination angle, the launch angle, the altitude and the speed is provided, the switching logic of the middle guidance law and the end guidance law is designed, and the guidance information processing method and the stable switching method of the middle guidance section and the end guidance section are designed, so that the aircraft can fly farther in the middle guidance section, the large drop angle is formed in the end guidance section, and the target can be accurately hit.
Drawings
FIG. 1 is a diagram of a method for implementing a ballistic implementation of a guided vehicle in accordance with an embodiment of the present invention;
FIG. 2 shows the switching judgment conditions of the middle guidance law and the last guidance law;
FIG. 3 is a three-dimensional ballistic curve;
FIG. 4 is a longitudinal ballistic curve;
FIG. 5 is a lateral ballistic curve;
FIG. 6 is a ballistic dip curve;
fig. 7 is a velocity profile.
Detailed Description
The invention will be further explained with reference to the following drawings and examples
Example 1
In the embodiment, taking an air-ground percussion guidance aircraft as an example, the altitude of a target point is 1000m, the throwing height is 5000m, and the throwing speed is Vr250m/s, throw angle Q 00 degree, off-axis emission angle of 5.3 degrees, target position 10000m from horizontal distance guidance aircraft, and expected falling angle qd-70 degrees. The trajectory scheme of the guided aircraft is shown in fig. 1 and is divided into a stable section, a middle guidance section and a final guidance section, and the invention aims at the last two guidance stages.
The method for handing over guidance and terminal guidance in constraint with a large drop angle and processing guidance information comprises the following specific implementation steps:
the method comprises the following steps that firstly, a stable section, a middle guidance section and a final guidance section are adopted in a ballistic scheme, three channels of the stable section adopt attitude drivers fed back by angular velocity and angular position, attitude drivers are adopted in a rolling channel of the middle guidance section and a rolling channel of the final guidance section, overload drivers are adopted in a pitching channel and a yawing channel, a ballistic inclination angle tracking guidance law is adopted in the middle guidance law, and a sliding mode with a falling angle constraint is adopted in the final guidance law to control the guidance law.
Step two, utilizing the position information X of the guided missile in the ground coordinate system measured by navigationm,Ym,ZmVelocity information Vx,Vy,VzAnd a known target position X in the ground coordinate systemt,Yt,ZtAnd calculating to obtain the high and low sight line angle q under the ground coordinate systemesAzimuth line of sight angle qbsAnd high and low line of sight angular velocities
Figure BDA0002818014890000051
And azimuth line-of-sight angular velocity
Figure BDA0002818014890000052
Xg＝Xt-Xm,Yg＝Yt-Ym,Zg＝Zt-Zm (1)
Figure BDA0002818014890000053
Figure BDA0002818014890000054
Figure BDA0002818014890000055
Wherein
Figure BDA0002818014890000056
Tanx＝(Yg/Sqrtx). Relative position coordinate (X) of target relative to missileg,Yg,Zg)。
The ground coordinate system takes the emitting point as the origin of coordinates, takes the connecting line of the projection of the emitting point on the horizontal plane and the target point as the X axis, and takes the direction facing the target point as the positive direction; taking the direction vertical to the horizontal plane as the Y axis and the direction towards the sky as the positive direction; the coordinate axis Z forms a right-hand coordinate system with the X and Y axes.
Thirdly, measuring information by using navigation equipment and calculating the relative distance R of the bullet eyesgThe synthesis velocity VcBallistic declination psiVAnd ballistic inclination angle thetadAnd (4) information.
Figure BDA0002818014890000061
Figure BDA0002818014890000062
θd＝atan(Vy/Vx) (6)
ψV＝atan(-Vz/Vx) (7)
Step four, before the guided aircraft passes the maneuvering point, all real-time measurement information is obtained from a navigation system, a reference instruction a is taken to be 9, and the reference altitude H is takenr4000, 20 for b, 0.3 for c, 1.5 for d, 0.2 for l, and 4 for m, and the gliding flight is performed according to the overload command generated by the guidance law as follows.
Uf1＝a-H/Hr-b×Vr/Vc+57.3c×Q0-57.3θd+d×cos(θd) (8)
Figure BDA0002818014890000063
Wherein U isf1And Uf2Respectively pitch and yaw direction overload commands, and the relative height H is Yg，Q0For ballistic inclination at the moment of delivery, VrAnd the flying speed of the aircraft at the moment of launching.
And step five, judging whether the current state information meets the shift switching condition of the maneuvering point. Judging the current high and low sight line angle qesAngle q from desired end line of sightdWhether the difference is greater than or equal to q ε30 °; judging the current flying height YgWhether or not Y is greater than or equal tozh4000 m; if the front flight state information simultaneously meets the two conditions, the aircraft reaches a maneuvering point, and then enters a final guidance control section; otherwise, entering a middle guidance control section, namely performing gliding flight according to the method of the step four; fig. 2 shows an implementation of the switching determination between the middle guidance law control segment and the last guidance law control segment.
Step six, entering a final guidance control section; and finishing the switching from the middle guidance section to the final guidance section.
And 6.1, switching guidance instructions from the middle guidance section to the final guidance section in an exponential fade-in mode, and ensuring stable switching of the guidance instructions.
Rksa＝e(-(t1-t1a)/1.0)
Uf1b＝Uf1×Rksa+Acc_y×(1-Rksa) (10)
Uf2b＝Uf2×Rksa+Acc_z×(1-Rksa)
Wherein t1 is the current flight time given by the missile-borne computer, t1a is the time of just entering the maneuvering point, the Rksa instruction cross-linking index coefficient, Uf1、Uf2The pitch and yaw direction overload instructions of the middle guidance section are Acc _ y and Acc _ z, the pitch and yaw direction overload instructions of the final guidance section are Uf1b and Uf2b, and the pitch and yaw direction overload instructions are merged by instructions output by the guidance system.
And 6.2, in the final guidance stage, all real-time measurement information is obtained and processed from a navigation system, and the vertical direction and the lateral direction fly according to overload instructions generated by the following final guidance law respectively.
Acc_y＝(Acc1+Acc2+Acc3+Ug)×Rkk (11)
Figure BDA0002818014890000071
Wherein Rkk is 0.4, l1＝0.4、m 15 is a proportionality coefficient Acc1, Acc2, Acc3 and UgThe overload instruction components are calculated in the steps 6.3-6.6.
Step 6.3, calculating a proportion guidance item overload command Acc 1:
Figure BDA0002818014890000072
wherein Xk _ bili ═ 1.7 and Clamda ═ 1.5 are coefficients.
Rb_l1＝Rg/(Rg+d1) (14)
Wherein d is1200 is a coefficient.
Step 6.4, calculating an overload command Acc2 of the corner constraint item:
Acc2＝Vc×(Xk_bili+1)×Clamda×Rb_l2×(qes-qd)/(Rg+d2)/57.3 (15)
Rb_l2＝Rg/(Rg+d2) (16)
where Rb _ l2 is an intermediate variable, d2350 is a coefficient.
Step 6.5, calculating an overload instruction Acc3 of the sliding mode control item:
Acc3＝Epsl×Sgns×Rb_l1 (17)
wherein, Epsl is 0.4 as a coefficient.
The calculation formula of the saturation function term Sgns is as follows:
Figure BDA0002818014890000081
Figure BDA0002818014890000082
6.6, calculating an overload instruction U of the gravity compensation itemg：
Ug＝9.8×cos(θd)＝9.8×cos(atan(Vy/Vx)) (20)
Wherein theta isdIs the ballistic inclination angle.
Through the steps, medium and terminal guidance handover can be realized, the requirements of long range, large falling angle constraint and accurate guidance can be met simultaneously, and large falling angle attack is carried out on a long-distance ground target. The specific implementation effect is shown in figures 3-7. Fig. 3 to 5 show that it is possible to achieve long-distance glide with a range of 10000m and hit the target with high accuracy and a miss distance of 0.13m by the embodiment 1. Figure 5 shows that the proposed guidance scheme can correct lateral deviations due to off-axis emission with high accuracy. Figure 6 shows that the terminal ballistic dip reaches the desired drop angle with a drop angle deviation of 0.12 degrees. FIG. 7 shows that the final flight speed is high and can reach over 320 m/s.
The above detailed description is intended to illustrate the objects, aspects and advantages of the present invention, and it should be understood that the above detailed description is only exemplary of the present invention and is not intended to limit the scope of the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the scope of the present invention.

Claims (1)

1. The method for processing guidance and terminal guidance handover and guidance information in constraint with a large falling angle is characterized by comprising the following steps of: the method comprises the following concrete steps:
step one, position information X of guided missile on ground coordinate system by using navigation measurementm,Ym,ZmVelocity information Vx,Vy,VzAnd a known target position X in the ground coordinate systemt,Yt,ZtAnd calculating to obtain the high and low sight line angle q under the ground coordinate systemesAzimuth line of sight angle qbsAnd high and low line of sight angular velocities
Figure FDA0002818014880000011
And azimuth line-of-sight angular velocity
Figure FDA0002818014880000012
Xg＝Xt-Xm,Yg＝Yt-Ym,Zg＝Zt-Zm (1)
Figure FDA0002818014880000013
Figure FDA0002818014880000014
Wherein
Figure FDA0002818014880000015
Tanx＝(Yg/Sqrtx); target relative toRelative position coordinates (X) of the missileg,Yg,Zg)；
The ground coordinate system takes the emitting point as the origin of coordinates, takes the connecting line of the projection of the emitting point on the horizontal plane and the target point as the X axis, and takes the direction facing the target point as the positive direction; taking the direction vertical to the horizontal plane as the Y axis and the direction towards the sky as the positive direction; the coordinate axis Z and the X and Y axes form a right-hand coordinate system;
step two, measuring information by using navigation equipment and calculating the relative distance R of the bullet eyesgThe synthesis velocity VcBallistic declination psiVAnd ballistic inclination angle thetadInformation;
Figure FDA0002818014880000016
Figure FDA0002818014880000017
θd＝atan(Vy/Vx) (6)
ψV＝atan(-Vz/Vx) (7)
before the guided aircraft passes through a maneuvering point, all real-time measurement information is obtained from a navigation system, and gliding flight is carried out according to an overload instruction generated by a guidance law in the following steps;
Uf1＝a-H/Hr-b×Vr/Vc+57.3c×Q0-57.3θd+d×cos(θd) (8)
Figure FDA0002818014880000021
wherein U isf1And Uf2Respectively pitch and yaw direction overload commands, and the relative height H is Yg，Q0For ballistic inclination at the moment of delivery, VrIs the nominal flying speed of the aircraft, HrA, b, c, d, l, m are coefficients;
judging whether the current state information meets the shift switching condition of the maneuvering point;
judging the current high and low sight line angle qesAngle q from desired end line of sightdWhether the difference is greater than or equal to qε(ii) a Judging the current flying height YgY or morezh；qεAnd YzhIs a parameter preset according to requirements; if the current flight state information simultaneously meets the two conditions, the aircraft reaches a maneuvering point, and then enters a final guidance control section; otherwise, entering a middle guidance control section, namely performing gliding flight according to the method in the third step;
step five, entering a final guidance control section; completing the switching from the middle-made guide section to the last-made guide section;
5.1, switching guidance instructions from the middle guidance section to the final guidance section in an exponential fade-in mode to ensure stable switching of the guidance instructions;
Figure FDA0002818014880000022
wherein t1 is the current flight time given by the missile-borne computer, t1a is the time of just entering the maneuvering point, Rksa is the command cross-linking index coefficient, Uf1、Uf2The command is a pitch and yaw direction overload command of the middle guidance section, Acc _ y and Acc _ z are final guidance section pitch and yaw direction overload commands, and Uf1b and Uf2b are pitch and yaw direction overload commands after command fusion output by the guidance system;
step 5.2, entering a terminal guidance control section; all real-time measurement information is obtained and processed from a navigation system, and the vertical direction and the lateral direction fly according to overload instructions generated by the following last guidance law respectively;
Acc_y＝(Acc1+Acc2+Acc3+Ug)×Rkk (11)
Figure FDA0002818014880000031
wherein l1、m1Rkk is a proportionality coefficient, Acc1, Acc2, Acc3, UgRespectively, each overload instruction component is calculated in the step 5.3-5.6;
step 5.3, calculating a proportion guidance item overload command Acc 1:
Figure FDA0002818014880000034
wherein Xk _ bili, Clamda are coefficients;
Rb_l1＝Rg/(Rg+d1) (14)
wherein d is1Is a coefficient;
step 5.4, calculating an overload command Acc2 of the corner constraint item:
Acc2＝Vc×(Xk_bili+1)×Clamda×Rb_l2×(qes-qd)/(Rg+d2)/57.3 (15)
Rb_l2＝Rg/(Rg+d2) (16)
wherein q isdFor desired landing angle information, Rb _ l2 is an intermediate variable, d2Is a coefficient;
step 5.5, calculating an overload instruction Acc3 of the sliding mode control item:
Acc3＝Epsl×Sgns×Rb_l1 (17)
wherein Epsl is a coefficient;
the calculation formula of the saturation function term Sgns is as follows:
Figure FDA0002818014880000032
Figure FDA0002818014880000033
step 5.6, calculating an overload instruction U of the gravity compensation itemg：
Ug＝9.8×cos(θd)＝9.8×cos(atan(Vy/Vx)) (20)
Wherein theta isdIs a ballistic inclination angle; at the moment, the medium and terminal guidance handover can be realized, the requirements of long range, large falling angle restriction and accurate guidance are met, and large falling angle attack is carried out on a long-distance ground target.
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