CN112486023A - Simulation design method of flight control system for cruise missile flight path and control system - Google Patents
Simulation design method of flight control system for cruise missile flight path and control system Download PDFInfo
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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
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—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
- 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
Abstract
The invention discloses a simulation design method of a flight control system of a patrol missile flight path and the control system, comprising the following steps: establishing a trajectory simulation mathematical model of the flying projectile; designing control parameters of a cruise missile track flight control system based on an L1 nonlinear track guidance law, an STT (spin transfer torque) stability control law and a roll stability control law; acquiring a six-degree-of-freedom controlled trajectory simulation program of the flying projectile; adding the interference factors and the deviation coefficient table into a six-degree-of-freedom controlled trajectory simulation program of the patrol missile; carrying out a six-degree-of-freedom controlled ballistic simulation test on the patrol missile according to the deviation coefficient table, and outputting a simulation result; judging whether the simulation result meets the expected requirement, if not, skipping to the step 2 and redesigning the control parameters; if yes, the design process is ended. The invention can solve the problems of pneumatic and control coupling, dynamic noise of projectile motion, seeker image rotation and the like caused by projectile rotation, and can effectively improve the flight fighting efficiency of the pneumatic layout of the patrol projectile track.
Description
Technical Field
The invention belongs to the field of flight control of a patrol missile flight path, and particularly relates to a simulation design method of a patrol missile flight path flight control system and the control system.
Background
Based on the requirement of improving and optimizing the flight fighting efficiency of the pneumatic layout of the patrol missile flight path, the patrol missile flight path control system needs to be optimally designed and subjected to simulation verification. Currently, the cruise missile trajectory flight control system is mainly designed based on the BTT (bank turn) control law, which has the following disadvantages:
1. the BTT control method is adopted to carry out turning control on the flight path flight of the flying bomb patrolling, and the flying bomb patrolling needs to rotate around a longitudinal axis at any time.
2. The projectile body rotation easily generates pneumatic and control coupling, and simultaneously introduces larger noise into the MEMS sensor, which is not beneficial to navigation calculation.
3. When the projectile body rotates, the image of the seeker is driven to rotate, and the aim identification and search by an operator in a 'person-in-loop' process are not facilitated.
4. The rotation of the projectile body brings difficulty to the application of the low-cost strapdown seeker, and is not beneficial to reducing the production cost of the flying projectile.
Therefore, the conventional fly-bomb flight path flight control system based on the BTT control law cannot fully exert the pneumatic advantages of the pneumatic layout fly-bomb patrol.
Disclosure of Invention
The invention aims to provide a simulation design method of a flight control system for a patrol missile flight path and the control system, aiming at the defects of the prior art, which can solve the problems of pneumatic and control coupling, dynamic noise of missile motion, image rotation of a seeker and the like caused by rotation of a missile and can effectively improve the flight fighting efficiency of the patrol missile flight path in pneumatic layout.
In order to solve the technical problems, the technical scheme adopted by the invention is as follows:
a simulation design method for a flight control system of a flying projectile flight path comprises the following steps:
it is characterized by also comprising:
step 2, designing control parameters of the cruise missile track flight control system based on an L1 nonlinear track guidance law, an STT stable control law and a rolling stable control law;
step 3, acquiring a six-degree-of-freedom controlled trajectory simulation program of the flying projectile based on the results of the step 1 and the step 2;
step 4, adding interference factors influencing the flight path flight segment of the patrol missile into a six-degree-of-freedom controlled trajectory simulation program of the patrol missile;
step 5, determining a deviation coefficient table of the flying projectile, and adding the deviation coefficient table into a six-degree-of-freedom controlled trajectory simulation program of the flying projectile;
step 6, developing a six-degree-of-freedom controlled ballistic simulation test of the patrol missile according to the deviation coefficient table, and outputting a simulation result;
step 7, judging whether the simulation result meets the expected requirement, if not, skipping to the step 2 and redesigning the control parameters; if yes, the design process is ended.
Further, the step 6 further includes obtaining robustness and anti-interference capability of the patrol missile six-degree-of-freedom controlled trajectory simulation program in the patrol missile track flight control process according to the simulation result.
As a preferable mode, the step 2 includes:
step 201, designing an L1 linearization model according to an L1 nonlinear track guidance law, designing an STT overload control block diagram according to an STT stability control law, and designing a rolling stability control block diagram according to a rolling stability control law;
step 202, resolving an L1 guidance control output equation according to an L1 linearized model, resolving an overload control output equation according to an STT overload control block diagram, and resolving a roll stability control output equation according to a roll stability control block diagram;
and step 203, designing control parameters of an L1 guidance law control output equation, an overload control output equation and a roll stability control output equation.
As a preferred mode, the design criteria of the control parameters include: the bandwidth of the inner loop is 1.5-3 times of the bandwidth of the outer loop.
As a preferred mode, the design criteria of the control parameters include: the bandwidth of the angular velocity loop is constrained by the bandwidth of the actuator.
As a preferred approach, the bandwidth design of all loops avoids the natural frequency of the cruise missile.
Based on the same invention concept, the invention also provides a flying round flight control system determined by the flying round flight control system simulation design method.
Compared with the prior art, the method can solve the problems of pneumatic and control coupling, projectile motion dynamic noise, seeker image rotation and the like caused by projectile rotation, and can effectively improve the flight fighting efficiency of the pneumatic layout of the patrol projectile flight path.
Drawings
Fig. 1 is a structure diagram of a linear model under the linear tracking of the L1 nonlinear track guidance law.
FIG. 2 is a block diagram of an exemplary acceleration control loop control.
FIG. 3 is a block diagram of an exemplary roll angle control loop control.
FIG. 4 is a diagram of the flight level trajectory of the flying projectile after the method of the present invention is adopted.
FIG. 5 is a vertical trace diagram of the flying projectile after the method of the invention is adopted.
Fig. 6 is a diagram of flying attitude angles of the flying round after the method of the invention is adopted.
Detailed Description
The following further describes the embodiments of the present invention with reference to the drawings.
The invention is based on the L1 nonlinear track guidance law + STT (sideslip turning) stable control method, and the simulation design method of the cruise missile track flight control system comprises the following steps:
And 2, designing control parameters of the flight path flight control system of the flying round based on an L1 nonlinear flight path guidance law, an STT stable control law and a rolling stable control law according to the overall design parameters, the pneumatic calculation parameters and the missile control principle of the flying round.
Step 3, acquiring a six-degree-of-freedom controlled trajectory simulation program of the flying projectile based on the results of the step 1 and the step 2; the mathematical model and the control algorithm are converted into a computer program.
And 4, analyzing interference factors influencing the flight path flight section of the patrol missile, and adding the interference factors influencing the flight path flight section of the patrol missile into a six-degree-of-freedom controlled trajectory simulation program of the patrol missile according to the influence mechanism of various interferences on the flight of the patrol missile.
And 5, designing a patrol missile quality parameter, a mass center parameter, a rotational inertia parameter, a thrust parameter, a pneumatic parameter, a single-item deviation and combined deviation coefficient table, and adding the deviation coefficient table into a patrol missile six-degree-of-freedom controlled trajectory simulation program.
And 6, checking a ballistic simulation mathematical model and a six-degree-of-freedom controlled ballistic simulation program of the patrol projectile to determine the credibility of the simulation program. And developing a six-degree-of-freedom controlled trajectory simulation test of the patrol missile according to the deviation coefficient table, and outputting a simulation result. And obtaining the robustness and the anti-interference capability of the patrol missile six-degree-of-freedom controlled trajectory simulation program in the patrol missile flight path control process according to the simulation result.
Step 7, judging whether the simulation result meets the expected requirement or not based on the simulation result, if not, skipping to the step 2 and redesigning the control parameters (screening out parameters which have more remarkable influence on the performance of the control system and repeatedly carrying out design optimization of the step 2); if so, finishing the design process, and controlling the overall performance of the system to meet the requirement of the design index.
The invention also provides a flying round track flight control system determined by the flying round track flight control system simulation design method.
Preferably, the step 2 includes:
(1) designing an L1 linearization model according to an L1 nonlinear track guidance law, designing an STT overload control block diagram according to an STT stability control law, and designing a rolling stability control block diagram according to a rolling stability control law. The structure of a linear model under the linear tracking of the L1 non-linear track guidance law is shown in fig. 1, fig. 2 is a control block diagram of a typical acceleration control loop, and fig. 3 is a control block diagram of a typical roll angle control loop.
(2) And solving an L1 guidance law control output equation according to an L1 linearized model.
When η is very small, the LI nonlinear guidance law output is:
wherein: a iszkIs a Z-direction acceleration instruction of a ballistic coordinate system of the flying projectile, V is the flying speed of the flying projectile, and L1The distance between the flying projectile and the reference point, d is the flight path error, eta is the included angle between the speed and the reference line, eta1Is the angle, η, between the velocity and the target path2Is the angle between the reference line and the target path,
is the track error differential.
The conversion relation between the ground coordinate system acceleration command and the ballistic coordinate system acceleration command is as follows:
the conversion relation between the projectile acceleration command and the ground coordinate system acceleration command is as follows:
when flight path flies, the vertical direction is fixed-height flight, then: a isyg＝0，ayk＝0，θ＝0。
Order:
the vertical type (2), (3) and (4) are combined to obtain:
order:
the vertical type (5) and (6) are combined:
wherein: a isxg、ayg、azgRespectively a three-dimensional acceleration instruction of a ground coordinate system of the flying projectile, axk、ayk、azkRespectively a three-way acceleration instruction of a coordinate system of a ballistic trajectory of the flying projectile, axb、ayb、azbRespectively a three-dimensional acceleration instruction G of a coordinate system of the missileKAs a transformation matrix from the ballistic coordinate system to the ground coordinate system, BGFor transforming matrix from ground coordinate system to missile coordinate system, theta, psiv、
Psi and gamma are trajectory inclination angle, trajectory deflection angle, pitch angle, yaw angle and roll angle respectively when the flying projectile flies, A11～A33、B11～B33Is an intermediate variable.
(3) And solving an overload control output equation according to the STT overload control block diagram.
Wherein: deltazc、δycRespectively elevator and rudder instruction, kωz、kα、kaThe proportional coefficient of the angular velocity loop, the proportional coefficient of the pseudo attack angle (sideslip angle) stability augmentation loop, the integral coefficient of the acceleration loop, ayb、aybf、azb、azbfRespectively as Y-direction and Z-direction acceleration commands and feedbacks, omega, of the flying projectile systemz、ωyRespectively as the pitch angle speed and the yaw angle speed of the flying projectile,
respectively, a pseudo attack angle and a pseudo sideslip angle of the flying projectile are observed, and dt is a resolving period of an integral algorithm.
(4) And resolving a roll stability control output equation according to the roll stability control block diagram.
δxc＝kωxωx+kωxkγ(γc-γ)+kωxkγi∫(γc-γ)dt (9)
Wherein: deltaxcRespectively, aileron rudder command, kωx、kγ、kγiThe roll angular velocity loop proportional coefficient, the roll angular loop integral coefficient, gammacGamma is the roll angle command and feedback of flying projectile, omegaxAnd calculating a period for the roll angular velocity and dt of the flying projectile as an integral algorithm.
(5) And designing control parameters of an L1 guidance law control output equation, an overload control output equation and a roll stability control output equation according to the design principle and the performance index of the control system.
The whole missile patrol control system is a three-channel stable control system, the pitching channel is a four-loop control system of height + acceleration + pseudo attack angle + pitching angle speed, the yawing channel is a four-loop control system of track + acceleration + pseudo sideslip angle + yaw angle speed, and the rolling channel is a two-loop control system of rolling angle + rolling angle speed.
Design criteria for the control parameters include:
principle 1: by the automatic control principle, the bandwidth of the inner loop is at least 1.5-3 times of that of the outer loop, so that the inner loop can well follow the output instruction of the outer loop.
Principle 2: the bandwidth of the angular velocity ring is constrained by the bandwidth of the executing steering engine according to the steering engine bandwidth constraint principle.
Principle 3: and the principle of avoiding the natural frequency of the patrol missile is adopted, and the bandwidth design of all loops avoids the natural frequency of the patrol missile.
And designing the control parameters of the whole control system of the flying projectile according to the principle.
Assuming that the natural frequency of the cruise missile is xHz and the bandwidth of the steering engine is 15xHz, the bandwidths of the track loop, the acceleration loop, the pseudo-sideslip angle loop and the yaw rate loop can be respectively designed to be 0.5xHz, 2xHz, 4xHz and 7 xHz.
(6) And adding the designed control parameters into a six-degree-of-freedom controlled trajectory simulation program of the flying projectile for simulation verification.
Preferably, the table of the coefficient of bias in step 5 is shown in table 1 below.
TABLE 1 table of coefficient of pull bias
The six-degree-of-freedom controlled trajectory simulation program for the patrol missile comprises a patrol missile flight dynamics module, a target motion dynamics module, a missile-target relative motion module, a guidance law resolving module, a control law resolving module, an interference input module and the like. Control parameters of the control system are designed according to the step 2, and after the robustness and the anti-interference capability of the control system are verified through batch simulation of the parameter deviation table in the step 5, trajectory simulation curves shown in the figures 4, 5 and 6 are obtained, and simulation results meet design requirements.
In conclusion, the method of the invention can keep the roll angle stable to be 0 when the patrol missile flight path flies, can well solve the problems of pneumatic coupling and control coupling, dynamic noise of missile motion, seeker image rotation and the like caused by the rotation of the patrol missile body, and meets the requirement of improving the optimization design of the pneumatic layout patrol missile flight path flight fighting efficiency.
While the present invention has been described with reference to the embodiments shown in the drawings, the present invention is not limited to the embodiments, which are illustrative and not restrictive, and it will be apparent to those skilled in the art that various changes and modifications can be made therein without departing from the spirit and scope of the invention as defined in the appended claims.
Claims (7)
1. A simulation design method for a flight control system of a flying projectile flight path comprises the following steps:
step 1, establishing a ballistic simulation mathematical model of a flying projectile;
it is characterized by also comprising:
step 2, designing control parameters of the cruise missile track flight control system based on an L1 nonlinear track guidance law, an STT stable control law and a rolling stable control law;
step 3, acquiring a six-degree-of-freedom controlled trajectory simulation program of the flying projectile based on the results of the step 1 and the step 2;
step 4, adding interference factors influencing the flight path flight segment of the patrol missile into a six-degree-of-freedom controlled trajectory simulation program of the patrol missile;
step 5, determining a deviation coefficient table of the flying projectile, and adding the deviation coefficient table into a six-degree-of-freedom controlled trajectory simulation program of the flying projectile;
step 6, developing a six-degree-of-freedom controlled ballistic simulation test of the patrol missile according to the deviation coefficient table, and outputting a simulation result;
step 7, judging whether the simulation result meets the expected requirement, if not, skipping to the step 2 and redesigning the control parameters; if yes, the design process is ended.
2. The method according to claim 1, wherein the step 6 further includes obtaining, according to the simulation result, robustness and anti-interference capability of a patrol missile six-degree-of-freedom controlled trajectory simulation program in the patrol missile trajectory flight control process.
3. The method for the simulation design of the cruise missile flight path flight control system according to the claim 1, wherein the step 2 comprises the following steps:
step 201, designing an L1 linearization model according to an L1 nonlinear track guidance law, designing an STT overload control block diagram according to an STT stability control law, and designing a rolling stability control block diagram according to a rolling stability control law;
step 202, resolving an L1 guidance control output equation according to an L1 linearized model, resolving an overload control output equation according to an STT overload control block diagram, and resolving a roll stability control output equation according to a roll stability control block diagram;
and step 203, designing control parameters of an L1 guidance law control output equation, an overload control output equation and a roll stability control output equation.
4. The method for the simulation design of the cruise missile trajectory flight control system according to claim 3, wherein the design criteria of the control parameters comprise: the bandwidth of the inner loop is 1.5-3 times of the bandwidth of the outer loop.
5. The method for the simulation design of the cruise missile trajectory flight control system according to claim 3, wherein the design criteria of the control parameters comprise: the bandwidth of the angular velocity loop is constrained by the bandwidth of the actuator.
6. The method according to claim 3, wherein the bandwidth design of all loops avoids the natural frequency of the cruise missile.
7. A patrolling missile trajectory flight control system determined by the patrolling missile trajectory flight control system simulation design method according to any one of claims 1 to 6.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113739635A (en) * | 2021-08-26 | 2021-12-03 | 北京航天飞腾装备技术有限责任公司 | Guidance method for realizing missile large-sector-angle launching |
| CN115328191A (en) * | 2022-07-15 | 2022-11-11 | 北京星途探索科技有限公司 | Method, system, equipment and storage medium for controlling cruise of grazing cruise target |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106054612A (en) * | 2016-06-29 | 2016-10-26 | 河南科技大学 | BTT missile flight trajectory automatic control method |
| CN111176315A (en) * | 2020-01-19 | 2020-05-19 | 西北工业大学 | Variant patrol missile control method based on L1 adaptive control |
-
2020
- 2020-12-18 CN CN202011508865.0A patent/CN112486023B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106054612A (en) * | 2016-06-29 | 2016-10-26 | 河南科技大学 | BTT missile flight trajectory automatic control method |
| CN111176315A (en) * | 2020-01-19 | 2020-05-19 | 西北工业大学 | Variant patrol missile control method based on L1 adaptive control |
Non-Patent Citations (4)
| Title |
|---|
| 孙宝彩等: "巡飞导弹BTT自动驾驶仪设计及仿真", 《弹箭与制导学报》 * |
| 李怀涛等: "一种基于导航坐标系的巡飞弹航迹跟踪方法", 《弹箭与制导学报》 * |
| 杨黔龙 等: "引入侧向加速度反馈的侧滑转弯控制方法", 《探测与控制学报》 * |
| 陈伟 等: "针对移动目标的简控弹制导控制方法研究", 《计算机技术与发展》 * |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113739635A (en) * | 2021-08-26 | 2021-12-03 | 北京航天飞腾装备技术有限责任公司 | Guidance method for realizing missile large-sector-angle launching |
| CN113739635B (en) * | 2021-08-26 | 2023-01-24 | 北京航天飞腾装备技术有限责任公司 | Guidance method for realizing missile large-sector-angle launching |
| CN115328191A (en) * | 2022-07-15 | 2022-11-11 | 北京星途探索科技有限公司 | Method, system, equipment and storage medium for controlling cruise of grazing cruise target |
| CN115328191B (en) * | 2022-07-15 | 2023-09-12 | 北京星途探索科技有限公司 | Method, system, equipment and storage medium for controlling cruising of ground sweeping cruising target |
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