CN112486023B - Simulation design method for flight control system of cruise missile flight path and control system - Google Patents
Simulation design method for flight control system of cruise missile flight path and control system Download PDFInfo
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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 torque transfer) stability control law and a rolling stability control law; acquiring a six-degree-of-freedom controlled trajectory simulation program of a flying round; 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 battle efficiency of the pneumatic layout of the patrol projectile track flight.
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. The projectile body drives the seeker image to rotate when rotating, and the aim of identifying and searching by an operator is not facilitated when a person is in a loop.
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 patrol missile flight path flight control system 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, seeker image rotation and the like caused by rotation of a missile and can effectively improve the battle efficiency of pneumatic layout patrol missile flight path flight.
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:
the method is characterized by also comprising the following steps:
step 2, designing control parameters of the cruise missile track flight control system based on an L1 nonlinear track guidance law, an STT (spin torque transfer) stability control law and a rolling stability 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 law control output equation according to an L1 linearization 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 203, designing control parameters of an L1 guidance law control output equation, an overload control output equation and a rolling 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 to 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 mode, the bandwidth of all loops is designed to avoid the natural frequency of the missile.
Based on the same inventive concept, the invention also provides a patrol missile flight path flight control system determined by the patrol missile flight path 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 patrol projectile flight path.
Drawings
FIG. 1 is a diagram of a linear model structure under L1 nonlinear track guidance law linear tracking.
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 a flight level trajectory of a flying round after the method of the 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 the flying attitude angle 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 relates to a method for stably controlling a cruise missile track flight control system based on an L1 nonlinear track guidance law and an STT (spin transfer rate), wherein the method for the simulation design 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 round 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 single-item deviation and combined deviation coefficient table of the mass parameter, the mass center parameter, the rotational inertia parameter, the thrust parameter and the pneumatic parameter of the inspection missile, and adding the deviation coefficient table into a six-degree-of-freedom controlled trajectory simulation program of the inspection missile.
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 carrying out 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 nonlinear 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 resolving an L1 guidance law control output equation according to the L1 linearized model.
When η is very small, the LI nonlinear guidance law output is:
wherein: a is a zk Is a Z-direction acceleration instruction of a ballistic coordinate system of the flying round, V is the flying speed of the flying round, L 1 The 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, eta 1 As a nip between speed and target pathAngle η 2 Is 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:
the vertical direction is the level flight when flight path flight, then: a is a yg ＝0，a yk ＝0，θ＝0。
Order:
the vertical type (2), (3) and (4) are combined to obtain:
order:
the combined type (5) and (6) are as follows:
wherein: a is a xg 、a yg 、a zg Respectively as flying roundThree-directional acceleration command of ground coordinate system, a xk 、a yk 、a zk Respectively a three-way acceleration instruction of a coordinate system of a ballistic trajectory of the flying projectile, a xb 、a yb 、a zb Respectively a three-dimensional acceleration instruction G of a coordinate system of the missile K As a transformation matrix from the ballistic coordinate system to the ground coordinate system, B G For transforming matrix from ground coordinate system to missile coordinate system, theta, psi v 、
Psi and gamma are trajectory inclination angle, trajectory deflection angle, pitch angle, yaw angle and roll angle respectively when the flying projectile flies, A 11 ～A 33 、B 11 ～B 33 Is an intermediate variable.
(3) And solving an overload control output equation according to the STT overload control block diagram.
Wherein: delta. For the preparation of a coating zc 、δ yc Respectively elevator and rudder instruction, k ωz 、k α 、k a The 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, a yb 、a ybf 、a zb 、a zbf Respectively as Y-direction and Z-direction acceleration commands and feedbacks, omega, of the flying projectile system z 、ω y Respectively as the pitch angle speed and the yaw angle speed of the flying projectile,
the false attack angle and the false sideslip angle of the flying projectile are respectively represented, and dt is an integral algorithm resolving period.
(4) And resolving a roll stability control output equation according to the roll stability control block diagram.
δ xc ＝k ωx ω x +k ωx k γ (γ c -γ)+k ωx k γi ∫(γ c -γ)dt (9)
Wherein: delta xc Respectively, aileron rudder command, k ωx 、k γ 、k γi The roll angular velocity loop proportional coefficient, the roll angular loop integral coefficient, gamma c Gamma is the roll angle command and feedback of flying projectile, omega x And 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 rolling 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 7xHz.
(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, which is designed by the invention, 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 (6)
1. A simulation design method for a flight control system of a cruise missile 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 a cruise missile track flight control system based on an L1 nonlinear track guidance law, an STT (spin torque transfer) stability control law and a roll stability control law; 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 law control output equation according to an L1 linearization 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;
the L1 guidance law control output equation is as follows:
wherein: a is xg 、a yg 、a zg Respectively a three-dimensional acceleration instruction of a ground coordinate system of the flying projectile, a xk 、a yk 、a zk Respectively a three-way acceleration instruction of a coordinate system of a ballistic trajectory of the flying projectile, a xb 、a yb 、a zb Respectively a three-way acceleration instruction G of a coordinate system of the missile body K As a transformation matrix from the ballistic coordinate system to the ground coordinate system, B G For transforming matrix from ground coordinate system to missile coordinate system, theta, psi v 、
Psi and gamma are trajectory inclination angle, trajectory deflection angle, pitch angle, yaw angle and roll angle of the flying projectile during flying，A 11 ～A 33 、B 11 ～B 33 Is an intermediate variable;
the overload control output equation is as follows:
wherein: delta zc 、δ yc Respectively elevator and rudder instruction, k ωz 、k α 、k a The proportional coefficient of the angular velocity loop, the proportional coefficient of the pseudo attack angle and the stability augmentation loop of the sideslip angle, the integral coefficient of the acceleration loop, a yb 、a ybf 、a zb 、a zbf Respectively as Y-direction and Z-direction acceleration commands and feedbacks, omega, of the flying projectile system z 、ω y Respectively 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, and dt is an integral algorithm resolving period;
the roll stability control output equation is as follows:
δ xc ＝k ωx ω x +k ωx k γ (γ c -γ)+k ωx k γi ∫(γ c -γ)dt
wherein: delta xc Respectively, aileron rudder command, k ωx 、k γ 、k γi The roll angular velocity loop proportional coefficient, the roll angular loop integral coefficient, gamma c Gamma is the roll angle command and feedback of flying projectile, omega x Calculating a period for the roll angular velocity and dt of the flying projectile as an integral algorithm;
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;
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 round, and adding the deviation coefficient table into a six-degree-of-freedom controlled trajectory simulation program of the flying round;
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 trajectory flight control system according to claim 1, wherein the design criteria of the control parameters comprise: the bandwidth of the inner loop is 1.5 to 3 times of the bandwidth of the outer loop.
4. The method for the simulation design of the cruise missile trajectory flight control system according to claim 1, wherein the design criteria of the control parameters comprise: the bandwidth of the angular velocity loop is constrained by the bandwidth of the actuator.
5. The method according to claim 1, wherein the bandwidth design of all loops avoids the natural frequency of the cruise missile.
6. A flying round trajectory flight control system determined via the flying round trajectory flight control system simulation design method according to any one of claims 1 to 5.
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