CN101694364B - Method for quickly converting perturbation guidance and iteration guidance - Google Patents

Abstract

The invention relates to a method for quickly converting perturbation guidance and iteration guidance. The method comprises the following steps: (1) at the time of a first period t0 after entering iteration guidance, determining a program angle output by perturbation guidance according to a ground theoretical trajectory and meanwhile calculating a program angle output by the iteration guidance; (2) calculating a program angle required to be rotated and the maximum angular acceleration amax allowed to be rotated according to the program angle output by the perturbation guidance and the program angle output by the iteration guidance and determining time deltat required by a rotating process, wherein the formula of the program angle required to be rotated is disclosed in the specification; and (3) calculating the program angle within the time of t-(t+deltat) according to a result of the step (2) and utilizing the program angle to control so as to realize the quick conversion of the perturbation guidance and the iteration guidance. The invention overcomes the defects of the prior art and can ensure that attitudes in the process of guidance rule switching have quick and steady transition.

Description

Perturbation guidance and iteration guidance quick conversion method

Technical Field

The invention relates to a rapid conversion method of perturbation guidance and iterative guidance, which is mainly applied to the technical field of carrier rocket guidance.

Background

Perturbation guidance and iterative guidance are guidance methods applied to a carrier rocket, and are real-time algorithms which are used as guidance laws, run in an rocket-borne computer and guide the rocket to a target orbit according to a flight state in the rocket flight process. The guiding function is mainly realized by program corners and other auxiliary information.

The iterative guidance is an optimal guidance method developed on the basis of a modern optimal control principle and a computer application technology, a carrier rocket is generally used after flying out of the atmosphere when the iterative guidance is adopted, perturbation guidance is adopted in the atmosphere, the perturbation guidance is based on a standard trajectory, the iterative guidance is used for calculating the optimal flight trajectory in real time, when the deviation of control instructions (program angles) output by the carrier rocket and the control instructions (program angles) is large, direct and rapid switching or linear switching in a short time can cause the program angles to jump in the iterative guidance switching process of perturbation guidance, the attitude control system is not favorable for stably controlling the rocket body, and the slow switching can cause error accumulation.

The tactical missile control technology 2006 No.3, named 'novel ballistic missile closed-circuit guidance research', analyzes the basic principle and the main defects of perturbation guidance of a missile with a latent projectile channel, summarizes the advantages of closed-circuit guidance, but does not introduce a conversion method between perturbation guidance and iterative guidance.

Foreign relevant introduction on the technology is not retrieved.

Disclosure of Invention

The technical problem to be solved by the invention is as follows: the method overcomes the defects of the prior art, and provides a rapid switching method of perturbation guidance and iterative guidance, which can make the posture in the guidance law switching process transition rapidly and stably.

The technical solution of the invention is as follows: the perturbation guidance and iteration guidance fast conversion method comprises the following steps:

(1) the first period t after entering the iterative guidance₀Determining a program angle of perturbation guidance output according to a ground theoretical trajectory at moment

And simultaneously calculating the program angle of the iterative guidance output

(2) According to the program angle output by the perturbation guidance

And program angle of iterative guidance outputCalculating the programmed angle of rotationAnd maximum angular acceleration a of the allowed rotation_maxDetermining the time delta t required by the rotation process; the program angle required to rotate

(3) Calculating a program angle in the time of t-t + delta t according to the result of the step (2)

Using the angle of the program

Controlling to realize the rapid conversion between perturbation guidance and iterative guidance;the calculation formula of (2) is as follows:

wherein t is the current control time.

Maximum angular acceleration a of the allowable rotation in the step (2)_maxThe calculation formula of (2):

wherein, J_z1Is the moment of inertia about the normal or transverse axis of the launch vehicle;

M_z1the maximum transverse normal control moment of the carrier rocket is obtained;

sign () denotes a symbol that takes data within parentheses.

The calculation formula of the time delta t required by the rotation process in the step (2) is as follows:

compared with the prior art, the invention has the beneficial effects that:

(1) the invention controls the program angle

The calculation formula of (1) divides the conversion process into two sections, each section adopts a time quadratic curve form, the angle of a connecting point between the two sections of curves and the angle change rate are continuous and have no jump, and the transition time is calculated according to the maximum available rotation angular acceleration in the conversion process, thereby ensuring the conversion process to be rapid and stable.

(2) The maximum angular acceleration allowed to rotate is obtained by calculation according to the moment of inertia around the normal axis or the transverse axis of the carrier rocket and the maximum transverse normal control moment of the carrier rocket, and specific data of the conversion moment can be adopted, so that the calculation result is the maximum angular acceleration at the conversion moment instead of the maximum angular acceleration allowed in the whole process of the rocket, and the shortest conversion time is further ensured.

Drawings

FIG. 1 is a flow chart of the method of the present invention;

FIG. 2 is a tracking curve of attitude angle in the direct switching of program angle;

FIG. 3 is a plot of attitude angle tracking using a program angle linear transition mode;

fig. 4 is an attitude angle tracking curve in the procedure angle quadratic curve transition mode according to the present invention.

Detailed Description

Because the existing rocket mostly adopts a three-axis stable control method, wherein the rolling axis is generally kept at 0 degree or a fixed angle, no change occurs in the flight, and the problem of conversion does not exist. The invention can be applied to the program angle conversion of two channels of pitch and yaw, which can be independently performed, and only the pitch channel is described here. The following describes the implementation process of the present invention in detail with reference to fig. 1, and the specific steps are as follows:

(1) in the first period (t) after entering the iterative guidance₀Time of day), determining a program angle of perturbation guidance output according to a ground theory trajectoryAnd simultaneously calculating the program angle of the iterative guidance output

Program angle for perturbation guidance

The program angle in the ground-theoretical trajectory is an array that varies with time, determined from a standard trajectory calculated on the ground (i.e., the ground-theoretical trajectory), during the rocket's flight

According to time t₀Obtained by interpolation calculation.

Program angle for iterative guidance output

The calculation result is obtained according to the iterative guidance equation, which can be referred to the iterative guidance equation and the pitch program angle described in pages 283 to 293 of the control system of aerospace publishers

Determined by equation (3-112):

wherein,

is the mean procedure angle;

is the program angle adjustment;

is a timing from the current time, and is obtained for real-time calculation $\tilde{t}=0,$ Then:

the calculation formulas of the parameters in the formulas can refer to the relevant contents from page 283 to page 293 in the control system (up), because the timing t started by the current time in the formulas in the control system (up) conflicts with the current time t in the method of the present invention, so that the t of the formulas in the control system (up) is used as the t

Instead.

(2) According to the program angle output by the perturbation guidance

And iterationProgrammed angle of guidance output

Calculating the programmed angle of rotationAnd maximum angular acceleration a of the allowed rotation_maxDetermining the time delta t required by the rotation process;

angle of program required to be rotated

Maximum angular acceleration a of the allowed rotation_maxCalculating the formula:

wherein J_z1For moment of inertia about the normal or transverse axis of the launch vehicle, which is generally equal, and which varies with the rocket structure and the total mass, it can be obtained from the ground theoretical ballistic data, if necessary with some margin, e.g. by increasing the moment of inertia of the theoretical trajectory by 20% as J_z1And the method is used for calculation, so that the engineering application process is safer.

M_z1And calculating the maximum transverse normal control moment of the carrier rocket according to the configuration of the engine and the maximum available swing angle. For example, four engines in a "+" configuration, each engine thrust P, and maximum pivot angle δ_maxAnd the distance from the thrust action point of the engine to the mass center is L, then:

M_z1＝2×P×sin(δ_max)×L

the above-mentioned ground theoretical trajectory and engineConfiguration, thrust P, maximum pivot angle delta_maxThe engine thrust action point to centroid distance L data is obtained by a rocket ensemble scheme, which must be known and not a point of innovation for those skilled in the art to utilize the present invention, and therefore, will not be described in detail herein.

The time delta t required by the rotation process is calculated by the formula:

(3) calculating a program angle in the time of t-t + delta t according to the result of the step (2)

Using the angle of the program

Controlling to realize the rapid conversion between perturbation guidance and iterative guidance;

the calculation formula of (2) divides the conversion process into two sections, each section adopts a quadratic curve form of time, the angle of a joint point between the two sections of curves and the change rate of the angle are continuous and have no jump, and therefore, the curve is ensured to be smooth. The formula according to the requirements is as follows:

wherein t is the current control time.

By comparing fig. 2, fig. 3, and fig. 4, it can be seen that: the attitude angle tracking condition under the program angle quadratic curve transition mode is best, the overshoot is minimum, and the transition time is shortest, so that the quick and smooth transition effect is achieved, and the attitude control system is facilitated.

When the method is used for controlling the yaw channel, only the corresponding pitching channel information in the introduction is needed to be replaced by the yaw channel information, and a person skilled in the art can convert the yaw channel by using the method according to the introduction without creative labor, so that the detailed description is omitted.

The invention is not described in detail and is within the knowledge of a person skilled in the art.

Claims (1)

1. The perturbation guidance and iteration guidance fast conversion method is characterized by comprising the following steps:

(1) the first period t after entering the iterative guidance₀Determining a program angle of perturbation guidance output according to a ground theoretical trajectory at moment

And simultaneously calculating the program angle of the iterative guidance output

(2) According to the program angle output by the perturbation guidanceAnd program angle of iterative guidance outputCalculating the programmed angle of rotation

And maximum angular acceleration a of the allowed rotation_maxDetermining the time delta t required by the rotation process; the program angle required to rotate

Said maximum angular acceleration a of the allowed rotation_maxThe calculation formula of (2):

wherein, J_z1Is the moment of inertia about the normal or transverse axis of the launch vehicle;

M _z1the maximum transverse normal control moment of the carrier rocket is obtained;

sign () represents a symbol to take data in parentheses;

the calculation formula of the time delta t required by the rotation process is as follows:

(3) calculating a program angle in the time of t-t + delta t according to the result of the step (2)

Using the angle of the programControlling to realize the rapid conversion between perturbation guidance and iterative guidance;the calculation formula of (2) is as follows:

wherein t is the current control time.

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