CN116203849A - Falling angle constraint control system applied to remote composite guidance aircraft - Google Patents

Falling angle constraint control system applied to remote composite guidance aircraft Download PDF

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CN116203849A
CN116203849A CN202310499107.4A CN202310499107A CN116203849A CN 116203849 A CN116203849 A CN 116203849A CN 202310499107 A CN202310499107 A CN 202310499107A CN 116203849 A CN116203849 A CN 116203849A
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aircraft
guidance
bullet
angle
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CN116203849B (en
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王伟
于之晨
林时尧
王少龙
王小康
林德福
王江
王辉
王雨辰
陈仕伟
张宏岩
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
Ordnance Science and Research Academy of China
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
Ordnance Science and Research Academy of China
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    • 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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    • G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
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Abstract

The invention discloses a falling angle constraint control system applied to a remote composite guidance aircraft, wherein the system adopts a satellite laser composite guidance aircraft to conduct guidance control; the system adopts a satellite guidance strategy to perform middle guidance gliding so as to achieve the purpose of range extension; in the terminal guidance section, a novel sliding mode guidance law is adopted, and on the basis of guaranteeing hit accuracy, the accurate control of the falling angle is realized.

Description

Falling angle constraint control system applied to remote composite guidance aircraft
Technical Field
The invention relates to a range-extending and falling angle constraint control system of an aircraft, in particular to a falling angle constraint control system applied to a remote composite guidance aircraft.
Background
With advances in equipment technology and changes in actual demand, accurate striking of targets has become an increasingly important technical requirement. The device has the capability of precisely striking the key target of the opposite party, can quickly establish advantages of the opposite party, and has the advantages of difficult generation of accidental injury, low cost and the like; the precision striking technique has become increasingly critical to combat.
The precision guidance aircraft is used as important equipment for precisely striking the target, and with the progress of technology, the requirements on the precision guidance aircraft are gradually increased; in the hardware level, single guidance detection equipment can cause information error measurement due to self limitation and deficiency, hit accuracy is reduced, and application requirements of modern environments can not be met gradually.
While at the software level, guidance systems are critical parts of aircraft, improvements are needed to cope with new needs. On the one hand, increasing the range is one direction that aircraft are continually pursuing in order to allow the aircraft to launch outside the range of detection of the other party, in order to increase the difficulty of interception of the other party. On the other hand, in order that the aircraft may hit the weak part of the counterpart's important facility without affecting the surrounding innocent masses, the falling angle constraint is also of great concern as a key ring for accurate hit. The method has the advantages of large damage to key targets and small damage to irrelevant things, but also has the defect of difficult control, a guidance system with stronger robustness is needed, and accordingly, the guidance law is used as the key of the guidance system, and new improvement is needed.
Based on the above problems, the present inventors have made intensive studies on a guidance control system for an aircraft, in hopes of designing a new guidance control system capable of solving the above problems.
Disclosure of Invention
In order to overcome the problems, the inventor has conducted intensive researches and designs a falling angle constraint control system applied to a remote composite guidance aircraft, wherein the system adopts a satellite laser composite guidance aircraft to conduct guidance control; the system adopts a satellite guidance strategy to perform middle guidance gliding so as to achieve the purpose of range extension; in the terminal guidance section, a novel sliding mode guidance law is adopted, and on the basis of guaranteeing hit accuracy, the accurate control of the falling angle is realized, so that the invention is completed.
In particular, it is an object of the present invention to provide a roll angle constraint control system for a remotely compound guided vehicle, the system comprising,
the inertial component module 1 is used for obtaining the roll angle of the aircraft in real time through the inertial component module 1;
the satellite guidance module 2 is used for acquiring the position and speed information of the aircraft in real time through the satellite guidance module 2;
the laser guidance module 3 is used for acquiring the missile vision angular rate in real time after the aircraft enters the terminal guidance section;
the central processing module 4 is used for obtaining the overload required by the aircraft in real time according to the received information, decomposing the overload required by the aircraft according to the rolling angle information, generating a rudder deflection instruction, transmitting the rudder deflection instruction to the actuating mechanism, and controlling the actuating mechanism to perform rudder operation.
Wherein, when the aircraft is in the middle guidance section, the central processing module 4 obtains the overload required by the guidance law of the overweight proportion,
the central processing module 4 obtains the required overload through a novel sliding mode guidance law when the aircraft is in the terminal guidance section.
Wherein, the overload proportional guidance law is obtained by the following formula (I):
Figure SMS_1
(one)
wherein ,
Figure SMS_2
indicating a need for overload;
Figure SMS_3
representing the proportional guide coefficient;
Figure SMS_4
representing the relative movement speed of the bullet mesh;
Figure SMS_5
representing the bullet eye line of sight angular rate;
Figure SMS_6
representing the gravity compensation coefficient;
Figure SMS_7
indicating the gravitational acceleration.
Wherein the relative movement speed of the bullet mesh
Figure SMS_8
By relative distance to the bullet eye->
Figure SMS_9
Obtaining a derivative;
the bullet eye sight angular rate
Figure SMS_10
By visual observation of the angle ∈>
Figure SMS_11
Obtaining a derivative;
wherein, the relative distance between the bullet and the eye
Figure SMS_12
Obtained by the following formula (II):
Figure SMS_13
(II)
Angle of view of bullet
Figure SMS_14
Obtained by the following formula (III):
Figure SMS_15
(III)
Figure SMS_16
Representing the position coordinates of the aircraft obtained by the satellite guidance module 2;
Figure SMS_17
representing the coordinates of the target location pre-filled in the aircraft.
The novel sliding mode guidance law obtains the overload required by the following formula (IV):
Figure SMS_18
(IV)
wherein ,
Figure SMS_19
indicating a need for overload;
Figure SMS_20
representing variable gain, wherein the value is positive;
Figure SMS_21
representing the slide face;
Figure SMS_22
and />
Figure SMS_23
Each independently representing a design parameter, all positive values;
Figure SMS_24
indicating a desired falling angle; />
Figure SMS_25
and />
Figure SMS_26
Each independently represents a control variable;
Figure SMS_27
and />
Figure SMS_28
Each independently represents a function with respect to a control variable.
Wherein the variable gain
Figure SMS_29
Obtained by the following formula (five):
Figure SMS_30
(V)
wherein ,
Figure SMS_31
representing design parameters;
Figure SMS_32
representing the gain factor;
Figure SMS_33
representing the slide face.
Wherein, the slip form surface
Figure SMS_34
The time-dependent change is represented by the following formula (six):
Figure SMS_35
(six)
wherein ,
Figure SMS_36
representing tracking errors;
Time
Figure SMS_37
is +.>
Figure SMS_38
Starting and controlling time for the laser guidance module (3);
Figure SMS_39
the derivative representing the tracking error by p +.>
Figure SMS_40
Obtaining a derivative;
Figure SMS_41
indicating the initial moment +.>
Figure SMS_42
Is a tracking error of (2);
Figure SMS_43
indicating the initial moment +.>
Figure SMS_44
Derivative of tracking error;
Figure SMS_45
and />
Figure SMS_46
Each independently represents a design parameter, all positive values.
Wherein the tracking error
Figure SMS_47
Obtained by the following formula (seven):
Figure SMS_48
(seven)
wherein ,
Figure SMS_49
indicating a desired falling angle;
Figure SMS_50
representation->
Figure SMS_51
Bullet view angle at moment.
Wherein the control variable
Figure SMS_52
Selected as->
Figure SMS_53
；
Controlling variables
Figure SMS_54
Selected as->
Figure SMS_55
；
Figure SMS_56
The bullet visual angle is shown;
Figure SMS_57
indicating the bullet eye gaze angular rate. />
Wherein the function is related to the control variable
Figure SMS_58
Obtained by the following formula (eight):
Figure SMS_59
(eight)
Function of control variables
Figure SMS_60
Obtained by the following formula (nine):
Figure SMS_61
(nine)
wherein ,
Figure SMS_62
the bullet visual angle is shown;
Figure SMS_63
representing the relative movement speed of the bullet mesh;
Figure SMS_64
representing the relative distance of the bullet meshes;
Figure SMS_65
representing the ballistic dip angle of the aircraft;
Figure SMS_66
representing the control variable.
The invention has the beneficial effects that:
(1) According to the falling angle constraint control system applied to the remote composite guidance aircraft, disclosed by the invention, the satellite laser composite guidance is adopted to improve the target detection and information collection capacity of the guidance equipment, so that the accurate hit capacity of the aircraft is improved; the satellite laser composite guidance is adopted to make up for the problem that single guidance equipment is insufficient in information acquisition capacity; the middle guidance section adopts a satellite guidance system to provide information for the aircraft to glide, so that the aircraft cannot deviate from the direction during range extension; the satellite laser composite guidance is adopted in the terminal guidance section, the relative distance between the missile and the target is obtained through a satellite guidance system, and the laser guidance system obtains the visual angle of the missile so as to cope with the modern complex actual situation;
(2) According to the falling angle constraint control system applied to the remote composite guidance aircraft, the satellite guidance system is adopted in the middle guidance section, and the guidance law is guided by the overweight compensation proportion to control the aircraft to perform gliding range increase; because the guidance law adds a gravity compensation signal in the proportional guidance loop, the path of the aircraft can be lifted upwards in the initial stage of proportional guidance so as to increase the flight envelope of the aircraft;
(3) According to the invention, in the falling angle constraint control system applied to the remote composite guidance aircraft, the terminal guidance section adopts a novel sliding mode guidance law, so that the robustness of the aircraft is improved, and the expected falling angle can be realized with smaller error; the PID sliding mode surface can realize smaller steady-state error and bring the advantage of stronger integral control robustness; the novel approach law can solve the buffeting problem in the approach process, so that the terminal guidance process is more stable.
Drawings
FIG. 1 is a logic diagram of the overall architecture of a roll angle constraint control system applied to a remotely compound guided vehicle in accordance with a preferred embodiment of the present invention;
FIG. 2 illustrates an aircraft trajectory curve in an embodiment;
fig. 3 shows the aircraft trajectory curve in the comparative example.
Reference numerals
1-inertial component module
2-satellite guidance module
3-laser guidance module
4-central processing module
Detailed Description
The invention is further described in detail below by means of the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used 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 illustrated in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the invention, a falling angle constraint control system applied to a remote composite guidance aircraft is provided, and as shown in fig. 1, the system comprises an inertial component module 1, a satellite guidance module 2, a laser guidance module 3 and a central processing module 4. Preferably, the system also comprises a power supply master control device which is connected with a thermal power supply on the aircraft and is responsible for supplying power to the whole system, and each module can normally work under rated power according to the voltage requirement.
The roll angle of the aircraft is obtained in real time through the inertial component module 1, and the roll angle information is transmitted to the central processing module as input quantity. The inertial component module 1 comprises INS inertial elements such as a triaxial MEMS gyroscope and an accelerometer, zero alignment is required after the inertial component module is started, and attitude information measurement can be converged to a true value only after a certain time; the inertial component module 1 can also be replaced by a mechanical gyroscope, the mechanical gyroscope does not need to consume time to determine a zero reference, and the triaxial angular velocity of the projectile body can be sensitively measured immediately after the inertial component module is started.
Acquiring the position and speed information of the aircraft in real time through the satellite guidance module 2; the satellite guidance module mainly comprises: GPS receiver, big dipper receiver, GLONASS receiver, antenna, treater and anti-interference module.
The antenna is designed into four synthetic antennas, and the cross distribution is adopted, so that the capability of capturing satellite signals when an aircraft rolls at a high speed can be improved, the captured satellite signals are transmitted to an anti-interference module for modulating and filtering the signals, and then the signals are transmitted to corresponding satellite receivers according to the corresponding satellite signals. The satellite receiver is composed of a GPS receiver, a Beidou receiver and a GLONASS receiver and is responsible for preprocessing satellite signals and transmitting navigation messages after signal conversion to the processor. The processor calculates and analyzes the received navigation message to obtain the position information of the aircraft and the speed information of the aircraft, and can further obtain information such as the relative distance of the bullet based on the target position. And meanwhile, the relative distance between the missile and the target is obtained through the satellite guidance module 2. In the application, the target position information is pre-stored in the aircraft before the aircraft takes off, the target is captured through the laser guide head in the terminal guidance section, and the target position information is corrected after the target is captured.
Acquiring the missile vision angular velocity in real time after the aircraft enters the terminal guidance section through the laser guidance module 3; the laser guidance module 3 is formed by an external laser target indicator and a laser platform guidance head on the aircraft. The laser target indicator outside the aircraft irradiates the target, the laser platform seeker receives the diffuse reflection laser echo signal of the target through the four-quadrant photoelectric detector assembly, calculates the bullet visual line angle deviation information and the target sight line angle rate, calculates the bullet visual line angle rate, and sends the bullet visual line angle rate to the central processing module as an input quantity.
The central processing module 4 is used for obtaining the overload required by the aircraft in real time according to the received information, decomposing the overload required by the aircraft according to the rolling angle information, generating a rudder deflection instruction, transmitting the rudder deflection instruction to the actuating mechanism, and controlling the actuating mechanism to perform rudder operation. The received information comprises the roll angle of the aircraft provided by the inertial component module 1 in real time, and also comprises the position and speed information of the aircraft provided by the satellite guidance module 2 in real time in the middle guidance section, and the angular rate information of the missile vision provided by the laser guidance module 3 in real time in the terminal guidance section.
In a preferred embodiment, the aircraft is started after the end of the uncontrolled section, and enters the middle guided section; when the aircraft is in the middle guidance section, the central processing module 4 generates the overload required by the guidance law of the overgravity proportion, the aircraft enters a Cheng Huaxiang increasing stage at the moment, when the relative distance of the missile is smaller than 8km, the aircraft enters the terminal guidance section, the laser guidance head captures a target, the central processing module 4 generates the overload required by the guidance law of the novel sliding mode, and the aircraft finally hits the target at a desired falling angle.
In a preferred embodiment, the overgrowth proportional guidance law obtains the required overload by the following formula (one):
Figure SMS_67
(one)
wherein ,
Figure SMS_68
indicating a need for overload;
Figure SMS_69
representing the proportional guide coefficient; typically 2-6, preferably 4;
Figure SMS_70
representing the relative movement speed of the bullet mesh;
Figure SMS_71
representing the bullet eye line of sight angular rate;
Figure SMS_72
representing the gravity compensation coefficient; preferably the value is +.>
Figure SMS_73
；
Figure SMS_74
Representing gravitational acceleration; preferably the value is +.>
Figure SMS_75
。
Preferably, the relative movement speed of the bullet mesh
Figure SMS_76
By relative distance to the bullet eye->
Figure SMS_77
Obtaining a derivative;
the bullet eye sight angular rate
Figure SMS_78
By visual observation of the angle ∈>
Figure SMS_79
Obtaining a derivative;
wherein, the relative distance between the bullet and the eye
Figure SMS_80
Obtained by the following formula (II):
Figure SMS_81
(II)
Angle of view of bullet
Figure SMS_82
Obtained by the following formula (III):
Figure SMS_83
(III)
Figure SMS_84
Representing the position coordinates of the aircraft obtained by the satellite guidance module 2;
Figure SMS_85
representing the coordinates of the target location pre-filled in the aircraft.
In a preferred embodiment, when the mesh is relatively distant
Figure SMS_86
When the laser guidance module is started, the laser guidance module enters a terminal guidance section, guidance control is performed by adopting a novel sliding mode guidance law, and the novel sliding mode guidance law obtains overload required by the following formula (IV):
Figure SMS_87
(IV)
wherein ,
Figure SMS_88
indicating a need for overload;
Figure SMS_89
representing variable gain, wherein the value is positive;
Figure SMS_90
representing the slide face;
Figure SMS_91
and />
Figure SMS_92
Each independently representing a design parameter, all positive values; preferably the value is +.>
Figure SMS_93
,
Figure SMS_94
；
Figure SMS_95
Indicating the expected falling angle, and pre-filling the aircraft in the aircraft before the aircraft takes off;
Figure SMS_96
and />
Figure SMS_97
Each independently represents a control variable;
Figure SMS_98
and />
Figure SMS_99
Each independently represents a function with respect to a control variable.
Preferably, the variable gain
Figure SMS_100
Obtained by the following formula (five):
Figure SMS_101
(V)
wherein ,
Figure SMS_102
representing design parameters; preferably take the value +.>
Figure SMS_103
The method comprises the steps of carrying out a first treatment on the surface of the In this application by setting the design parameters +.>
Figure SMS_104
Can control->
Figure SMS_105
The upper limit of (2) is not too large to influence the control effect;
Figure SMS_106
representing the gain factor; preferably take the value +.>
Figure SMS_107
The method comprises the steps of carrying out a first treatment on the surface of the In this application +_ is determined by setting the gain factor>
Figure SMS_108
Approach->
Figure SMS_109
Is a speed of (2);
Figure SMS_110
representing the slide face.
By setting the variable gain in the present application
Figure SMS_111
Replacing the constant velocity approach law in the traditional scheme, namely eliminating the occurrence of buffeting phenomenon, can lead the constant velocity approach law to have satisfactory convergence rate.
Preferably, the slip form surface
Figure SMS_112
Is a PID slip plane, which varies with time as shown in the following formula (six):
Figure SMS_113
(six)
wherein ,
Figure SMS_114
representing tracking errors;
Time
Figure SMS_115
is +.>
Figure SMS_116
Starting and controlling the moment for the laser guidance module 3;
Figure SMS_117
the derivative representing the tracking error by p +.>
Figure SMS_118
Obtaining a derivative;
Figure SMS_119
indicating the initial moment +.>
Figure SMS_120
Is a tracking error of (2);
Figure SMS_121
indicating the initial moment +.>
Figure SMS_122
Derivative of tracking error;/>
Figure SMS_123
and />
Figure SMS_124
Each independently represents a design parameter, all positive values.
The upper integral limit in formula (six)
Figure SMS_125
Time of presentation->
Figure SMS_126
Is->
Figure SMS_127
and />
Figure SMS_128
Is->
Figure SMS_129
All represent the integrand, and the integrand in equation (six) is still time in particular, so all use +.>
Figure SMS_130
And (3) representing.
Preferably, tracking error
Figure SMS_131
Obtained by the following formula (seven):
Figure SMS_132
(seven)
wherein ,
Figure SMS_133
indicating a desired falling angle;
Figure SMS_134
representation->
Figure SMS_135
Bullet view angle at moment.
By arranging the sliding mode surface
Figure SMS_136
The method can generate smaller steady-state errors in the process of realizing the sliding mode control, and can also increase the robustness of the sliding mode control.
Preferably, the control variable
Figure SMS_137
Selected as->
Figure SMS_138
；
Controlling variables
Figure SMS_139
Selected as->
Figure SMS_140
；
wherein ,
Figure SMS_141
the bullet visual angle is shown;
Figure SMS_142
indicating the angular velocity of the view line of sight of the bullet, which angular velocity of view line of sight of the bullet is +.>
Figure SMS_143
Is obtained by a laser guidance module.
Preferably, the function is a function of the control variable
Figure SMS_144
Obtained by the following formula (eight):
Figure SMS_145
(eight)
Function of control variables
Figure SMS_146
Obtained by the following formula (nine):
Figure SMS_147
(nine)
wherein ,
Figure SMS_148
the bullet visual angle is shown;
Figure SMS_149
representing the relative movement speed of the bullet mesh;
Figure SMS_150
representing the relative distance of the bullet meshes;
Figure SMS_151
representing a ballistic tilt angle of the aircraft, the ballistic tilt angle being measured by an inertial measurement assembly on the aircraft;
Figure SMS_152
representing the control variable.
The invention also provides a falling angle constraint control method applied to the remote composite guidance aircraft, which is realized through the control system;
preferably, the method comprises the steps of:
step 1, acquiring the overload needed by the aircraft in real time through the overweight proportion guidance law in the guidance section of the aircraft,
step 2, when the relative distance between the bullets and the eyes
Figure SMS_153
And when the laser guidance module is started, the laser guidance module enters the terminal guidance section, and the required overload of the aircraft is obtained in real time by adopting a novel sliding mode guidance law.
In the method, the central processing module 4 receives the roll angle information, decomposes the overload to be used based on the roll angle information, generates a rudder deflection instruction, and transmits the rudder deflection instruction to the actuating mechanism to control the actuating mechanism to perform rudder operation.
Preferably, the overload proportional guidance law is obtained by the following formula (one):
Figure SMS_154
(one)
wherein ,
Figure SMS_155
indicating a need for overload;
Figure SMS_156
representing the proportional guide coefficient; typically 2-6, preferably 4;
Figure SMS_157
representing the relative movement speed of the bullet mesh;
Figure SMS_158
representing the bullet eye line of sight angular rate;
Figure SMS_159
representing the gravity compensation coefficient; preferably the value is +.>
Figure SMS_160
；
Figure SMS_161
Representing gravitational acceleration; preferably the value is +.>
Figure SMS_162
。
Preferably, the relative movement speed of the bullet mesh
Figure SMS_163
By relative distance to the bullet eye->
Figure SMS_164
Obtaining a derivative;
the bullet eye sight angular rate
Figure SMS_165
By visual observation of the angle ∈>
Figure SMS_166
Obtaining a derivative;
wherein, the relative distance between the bullet and the eye
Figure SMS_167
Obtained by the following formula (II):
Figure SMS_168
(II)
Angle of view of bullet
Figure SMS_169
Obtained by the following formula (III):
Figure SMS_170
(III)
Figure SMS_171
Representing the position coordinates of the aircraft obtained by the satellite guidance module 2;
Figure SMS_172
representing the coordinates of the target location pre-filled in the aircraft.
The novel sliding mode guidance law obtains the overload needed by the user through the following formula (IV):
Figure SMS_173
(IV)
wherein ,
Figure SMS_174
indicating a need for overload; />
Figure SMS_175
Representing variable gain, wherein the value is positive;
Figure SMS_176
representing the slide face;
Figure SMS_177
and />
Figure SMS_178
Each independently representing a design parameter, all positive values; preferably the value is +.>
Figure SMS_179
,
Figure SMS_180
；
Figure SMS_181
Indicating a desired falling angle;
Figure SMS_182
and />
Figure SMS_183
Each independently represents a control variable;
Figure SMS_184
and />
Figure SMS_185
Each independently represents a function with respect to a control variable.
Preferably, the variable gain
Figure SMS_186
Obtained by the following formula (five):
Figure SMS_187
(V)
wherein ,
Figure SMS_188
representing design parameters; preferably take the value +.>
Figure SMS_189
；
Figure SMS_190
Representing the gain factor; preferably take the value +.>
Figure SMS_191
；
Figure SMS_192
Representing the slide face.
Preferably, the slip form surface
Figure SMS_193
Is a PID slip plane, which varies with time as shown in the following formula (six):
Figure SMS_194
(six)
wherein ,
Figure SMS_195
representing tracking errors;
Time
Figure SMS_196
is +.>
Figure SMS_197
Starting and controlling the moment for the laser guidance module 3;
Figure SMS_198
the derivative representing the tracking error by p +.>
Figure SMS_199
Obtaining a derivative;
Figure SMS_200
indicating the initial moment +.>
Figure SMS_201
Is a tracking error of (2);
Figure SMS_202
indicating the initial moment +.>
Figure SMS_203
Derivative of tracking error;
Figure SMS_204
and />
Figure SMS_205
Each independently represents a design parameter, all positive values.
Preferably, tracking error
Figure SMS_206
Obtained by the following formula (seven):
Figure SMS_207
(seven)
wherein ,
Figure SMS_208
indicating a desired falling angle; />
Figure SMS_209
Representation->
Figure SMS_210
Bullet view angle at moment.
Preferably, the control variable
Figure SMS_211
Selected as->
Figure SMS_212
；
Controlling variables
Figure SMS_213
Selected as->
Figure SMS_214
；
wherein ,
Figure SMS_215
the bullet visual angle is shown; />
Figure SMS_216
Indicating the angular velocity of the view line of sight of the bullet, which angular velocity of view line of sight of the bullet is +.>
Figure SMS_217
Is obtained by a laser guidance module.
Preferably, the function is a function of the control variable
Figure SMS_218
Obtained by the following formula (eight):
Figure SMS_219
(eight)
Function of control variables
Figure SMS_220
Obtained by the following formula (nine):
Figure SMS_221
(nine)
wherein ,
Figure SMS_222
the bullet visual angle is shown;
Figure SMS_223
representing the relative movement speed of the bullet mesh;
Figure SMS_224
representing the relative distance of the bullet meshes;
Figure SMS_225
representing a ballistic tilt angle of the aircraft, the ballistic tilt angle being measured by an inertial measurement assembly on the aircraft;
Figure SMS_226
representing the control variable.
Examples
Setting the initial position of the aircraft under an inertial coordinate system as the origin of the coordinate system, and setting the initial relative distance of the missile at the terminal guidance stage:
Figure SMS_227
the method comprises the steps of carrying out a first treatment on the surface of the Aircraft speed: />
Figure SMS_228
The method comprises the steps of carrying out a first treatment on the surface of the Initial ballistic tilt angle of aircraft: />
Figure SMS_229
The method comprises the steps of carrying out a first treatment on the surface of the Initial bullet visual angle: />
Figure SMS_230
。
The target is a stationary object on the ground.
And the middle guidance section adopts an overweight supplementary proportion guidance law to control the aircraft to glide, namely the required overload of the aircraft is obtained through the following formula (I):
Figure SMS_231
(one)
wherein ：
Figure SMS_232
，/>
Figure SMS_233
，/>
Figure SMS_234
。
and a novel sliding mode guidance law is adopted in the terminal guidance section to control the aircraft to fly to a target, namely the required overload of the aircraft is obtained through the following formula (IV):
Figure SMS_235
(IV)
Wherein the variable gain
Figure SMS_236
Obtained by the following formula (five):
Figure SMS_237
(V)
Sliding die surface
Figure SMS_238
The time-dependent change is represented by the following formula (six):
Figure SMS_239
(six)
Tracking error
Figure SMS_240
Obtained by the following formula (seven):
Figure SMS_241
(seven)
Controlling variables
Figure SMS_242
Selected as->
Figure SMS_243
；
Controlling variables
Figure SMS_244
Selected as->
Figure SMS_245
；
Function of control variables
Figure SMS_246
Obtained by the following formula (eight):
Figure SMS_247
(eight)
Function of control variables
Figure SMS_248
Obtained by the following formula (nine):
Figure SMS_249
(nine)
The design is as follows: gain coefficient
Figure SMS_250
The method comprises the steps of carrying out a first treatment on the surface of the Design parameters->
Figure SMS_251
、/>
Figure SMS_252
、/>
Figure SMS_253
。
Desired falling angle
Figure SMS_254
I.e. a total of 4 values of the desired landing angle, 4 flight trajectories are obtained correspondingly, as shown in fig. 2.
According to the embodiment, the falling angle constraint control system applied to the remote composite guidance aircraft can control the aircraft to accurately hit a target at a preset falling angle.
Comparative example
The same aircraft, target and test environment as in the examples were selected, the desired landing angle of the aircraft was set to 70 degrees, and the aircraft was guided using substantially the same control scheme as in the examples, except for the variable gain
Figure SMS_255
Taking a fixed value, and specifically taking the value as 2; the flight path obtained by controlling the aircraft to fly to the target according to the method is shown in a traditional sliding mode control method curve in fig. 3, and the flight path when the angle of the hope drop is 70 degrees in the embodiment is shown in fig. 3 at the same time, as shown in a novel sliding mode control method curve.
As can be seen from fig. 3, the novel sliding mode guidance law adopted in the terminal guidance section in the present application can improve the hit accuracy of the aircraft, and the aircraft adopting the conventional sliding mode guidance rate can cause the miss result.
The invention has been described above in connection with preferred embodiments, which are, however, exemplary only and for illustrative purposes. On this basis, the invention can be subjected to various substitutions and improvements, and all fall within the protection scope of the invention.

Claims (10)

1. A falling angle constraint control system applied to a remote compound guided vehicle is characterized in that the system comprises,
the inertial component module (1) is used for acquiring the roll angle of the aircraft in real time through the inertial component module (1);
the satellite guidance module (2) is used for acquiring the position and speed information of the aircraft in real time through the satellite guidance module (2);
the laser guidance module (3) is used for acquiring the view line angular rate of the bullet in real time after the aircraft enters the terminal guidance section;
the central processing module (4) is used for obtaining the overload required by the aircraft in real time according to the received information, decomposing the overload required by the aircraft according to the rolling angle information, generating a rudder deflection instruction, transmitting the rudder deflection instruction to the actuating mechanism, and controlling the actuating mechanism to perform rudder operation.
2. The system of claim 1, wherein the system comprises a control system for controlling the landing angle of a remotely compound guided vehicle,
when the aircraft is in the middle guidance section, the central processing module (4) obtains overload through the overweight proportion guidance law,
when the aircraft is in the terminal guidance section, the central processing module (4) obtains the overload through a novel sliding mode guidance law.
3. The system of claim 2, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
the overload proportional guidance law is obtained through the following formula (I):
Figure QLYQS_1
(one)
wherein ,
Figure QLYQS_2
indicating a need for overload;
Figure QLYQS_3
representing the proportional guide coefficient;
Figure QLYQS_4
representing the relative movement speed of the bullet mesh;
Figure QLYQS_5
representing the bullet eye line of sight angular rate;
Figure QLYQS_6
representing the gravity compensation coefficient;
Figure QLYQS_7
indicating the gravitational acceleration.
4. The system of claim 3, wherein the system comprises a control unit for controlling the landing angle of the remotely compound guided vehicle,
the relative movement speed of the bullet mesh
Figure QLYQS_8
By relative distance to the bullet eye->
Figure QLYQS_9
Obtaining a derivative;
the bullet eye sight angular rate
Figure QLYQS_10
By visual observation of the angle ∈>
Figure QLYQS_11
Obtaining a derivative;
wherein, the relative distance between the bullet and the eye
Figure QLYQS_12
Obtained by the following formula (II):
Figure QLYQS_13
(II)
Angle of view of bullet
Figure QLYQS_14
Obtained by the following formula (III):
Figure QLYQS_15
(III) Supports>
Figure QLYQS_16
Representing the position coordinates of the aircraft obtained by the satellite guidance module (2);
Figure QLYQS_17
representing the coordinates of the target location pre-filled in the aircraft.
5. The system of claim 2, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
the novel sliding mode guidance law obtains the overload needed by the user through the following formula (IV):
Figure QLYQS_18
(IV)
wherein ,
Figure QLYQS_19
indicating a need for overload;
Figure QLYQS_20
representing variable gain, wherein the value is positive;
Figure QLYQS_21
representing the slide face;
Figure QLYQS_22
and />
Figure QLYQS_23
Each independently representing a design parameter, all positive values;
Figure QLYQS_24
indicating the desired falling angle；
Figure QLYQS_25
and />
Figure QLYQS_26
Each independently represents a control variable;
Figure QLYQS_27
and />
Figure QLYQS_28
Each independently represents a function with respect to a control variable.
6. The system of claim 5, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
variable gain
Figure QLYQS_29
Obtained by the following formula (five):
Figure QLYQS_30
(V)
wherein ,
Figure QLYQS_31
representing design parameters;
Figure QLYQS_32
representing the gain factor;
Figure QLYQS_33
representing the slide face.
7. The system for controlling the attitude constraint of a remotely compound guided vehicle according to claim 5 or 6,
sliding die surface
Figure QLYQS_34
The time-dependent change is represented by the following formula (six):
Figure QLYQS_35
(six)
wherein ,
Figure QLYQS_36
representing tracking errors;
Time
Figure QLYQS_37
is +.>
Figure QLYQS_38
Starting and controlling time for the laser guidance module (3); />
Figure QLYQS_39
The derivative representing the tracking error by p +.>
Figure QLYQS_40
Obtaining a derivative;
Figure QLYQS_41
indicating the initial moment +.>
Figure QLYQS_42
Is a tracking error of (2);
Figure QLYQS_43
indicating the initial moment +.>
Figure QLYQS_44
Tracking errorIs a derivative of (2);
Figure QLYQS_45
and />
Figure QLYQS_46
Each independently represents a design parameter, all positive values.
8. The system of claim 7, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
tracking error
Figure QLYQS_47
Obtained by the following formula (seven):
Figure QLYQS_48
(seven)
wherein ,
Figure QLYQS_49
indicating a desired falling angle;
Figure QLYQS_50
representation->
Figure QLYQS_51
Bullet view angle at moment.
9. The system of claim 5, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
controlling variables
Figure QLYQS_52
Selected as->
Figure QLYQS_53
；
Controlling variables
Figure QLYQS_54
Selected as->
Figure QLYQS_55
；
Figure QLYQS_56
The bullet visual angle is shown;
Figure QLYQS_57
indicating the bullet eye gaze angular rate.
10. The system of claim 5, wherein the system comprises a control system for controlling the landing angle of the remotely compound guided vehicle,
function of control variables
Figure QLYQS_58
Obtained by the following formula (eight):
Figure QLYQS_59
(eight)
Function of control variables
Figure QLYQS_60
Obtained by the following formula (nine):
Figure QLYQS_61
(nine)
wherein ,
Figure QLYQS_62
the bullet visual angle is shown;
Figure QLYQS_63
representing the relative movement speed of the bullet mesh; />
Figure QLYQS_64
Representing the relative distance of the bullet meshes;
Figure QLYQS_65
representing the ballistic dip angle of the aircraft;
Figure QLYQS_66
representing the control variable. />
CN202310499107.4A 2023-05-06 2023-05-06 Falling angle constraint control system applied to remote composite guidance aircraft Active CN116203849B (en)

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Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111220031A (en) * 2018-11-26 2020-06-02 北京理工大学 Remote guidance aircraft with full range coverage
CN111377064A (en) * 2018-12-27 2020-07-07 北京理工大学 Satellite-loss-preventing remote guidance aircraft with full range coverage
CN111397441A (en) * 2019-01-03 2020-07-10 北京理工大学 Full range coverage guidance system for remotely guided vehicles with strapdown seeker
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