CN111813134B - Stability judgment method and judgment system of aircraft control system - Google Patents
Stability judgment method and judgment system of aircraft control system Download PDFInfo
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- CN111813134B CN111813134B CN202010549618.9A CN202010549618A CN111813134B CN 111813134 B CN111813134 B CN 111813134B CN 202010549618 A CN202010549618 A CN 202010549618A CN 111813134 B CN111813134 B CN 111813134B
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- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
- G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/10—Simultaneous control of position or course in three dimensions
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Abstract
本申请公开了一种飞行器控制系统的稳定性判别方法和判别系统，所述飞行器控制系统用于多个舵机的控制，所述稳定性判别方法包括以下步骤：监测每个所述舵机的控制角δi；根据监测到的所有所述舵机的控制角δi，换算所述舵机的控制角δi为所述飞行器控制系统的惯量中心坐标系下的相对控制角θi；根据所有所述舵机的相对控制角θi，合成欧几里得范数R，并求解所述欧几里得范数R的一阶导数；根据所述一阶导数判断所述飞行器控制系统是否稳定。本申请能够快速准确对多舵机运行场景下的飞行器控制系统的稳定性进行判断，以向飞行器控制系统后续的控制措施提供依据。
This application discloses a stability judgment method and judgment system for an aircraft control system. The aircraft control system is used to control multiple servos. The stability judgment method includes the following steps: monitoring the stability of each of the servos. Control angle δ i ; According to the monitored control angle δ i of all the servos, the control angle δ i of the servos is converted into the relative control angle θ i in the inertia center coordinate system of the aircraft control system; according to The relative control angles θ i of all the servos are synthesized into the Euclidean norm R, and the first-order derivative of the Euclidean norm R is solved; based on the first-order derivative, it is judged whether the aircraft control system is Stablize. This application can quickly and accurately judge the stability of the aircraft control system in a multi-server operation scenario, so as to provide a basis for subsequent control measures of the aircraft control system.
Description
Technical Field
The application relates to the technical field of electric power, in particular to a stability judging method and a stability judging system of an aircraft control system.
Background
The steering engine is widely applied in modern industry, in a plurality of application scenes, the control system comprises a plurality of steering engines, the steering engines are controlled by a central controller or a comprehensive control computer, real-time output angles (feedback angles for short) of the steering engines are collected, angle control (control angles for short) is carried out on the steering engines after upper-layer decision is carried out, and under normal conditions of the system, the feedback angles can follow the control angles to keep the balance of the system.
When the control system of the aircraft starts to lose stability due to certain disturbance or impact, the control angles of certain steering engines continuously increase, if the control system does not take measures to control at the moment, the control angles of certain steering engines tend to be infinite until the steering engines have no capability of outputting the control angles, and the control angles continue to increase due to the fact that the control system is not controlled, so that the control system is circulated until the control system is completely collapsed.
For example, in an aircraft control system, attitude control is performed by using a steering engine, the steering engine is controlled by means of steering engine output in the steering engine, the steering engine output angle is controlled by a central computer (control angle), and the actual rotating angle is fed back to the computer (feedback angle).
Therefore, timely and effective judgment of whether the control system of the aircraft is unstable is necessary for guaranteeing the control system which needs to control a plurality of steering engines to work normally.
Disclosure of Invention
The embodiment of the application provides a stability judging method and a judging system of an aircraft control system, which can rapidly and accurately judge the stability of the aircraft control system under a multi-steering engine operation scene so as to provide a basis for subsequent control measures of the aircraft control system.
In one aspect, an embodiment of the present application provides a method for determining stability of an aircraft control system, where the aircraft control system is used for controlling a plurality of steering engines, the method for determining stability includes the following steps:
monitoring the control angle delta of each steering engine i ；
According to the monitored control angle delta of all steering engines i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
According to the relative control angle theta of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
and judging whether the aircraft control system is stable or not according to the first derivative.
In this embodiment, preferably, the control angle δ of all the steering engines is monitored i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i The specific steps of (a) are as follows:
according to the inertia time constant M of all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI ；
Control angle delta of steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i 。
Preferably, the inertial time constant M according to all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI The calculation formula of (2) is as follows:
in delta COI Is the center angle of inertia; delta i The control angle of the ith steering engine; m is M i The inertia time constant of the ith steering engine;
preferably, the control angle theta is controlled according to the relative of all steering engines i The calculation formula of the synthesized euclidean norm R is:
wherein R is Euclidean norm, and n is the number of steering engines; θ i Is the relative control angle of the ith steering engine.
Preferably, after determining whether the aircraft control system is stable, the determining method further includes:
if the aircraft control system is judged to be stable, continuing to monitor the control angle delta of each steering engine i 。
Preferably, the specific step of judging whether the aircraft control system is stable according to the first derivative is as follows:
outputting a waveform of the first derivative according to the first derivative of the Euclidean norm R;
judging whether a minimum value exists when the first derivative is larger than zero according to the first derivative waveform, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
Preferably, the specific step of judging whether the aircraft control system is stable according to the first derivative is as follows:
solving a second derivative of the euclidean norm R;
and judging whether the first derivative and the second derivative are simultaneously larger than zero, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
Preferably, after solving the first derivative of the euclidean norm R, the stability discrimination method further includes:
solving a second derivative and a third derivative of the Euclidean norm R;
set at a known time t 0 When the first derivative, the second derivative and the third derivative are respectively b 1 、b 2 、b 3 ；
When said b 1 、b 2 、b 3 Satisfy b 1 ＜0，b 2 ＜0，b 3 ＞0，2b 1 b 3 -b 2 2 At > 0, the aircraft control system is predicted to destabilize at time t, wherein
On the other hand, the embodiment of the application also provides a stability judging system of an aircraft control system, wherein the aircraft control system is used for controlling a plurality of steering engines, and the stability judging system comprises:
monitoring means for monitoring the control angle delta of each steering engine i ；
Inertia center conversion means for converting the control angle delta of all the steering engines according to the detected control angle delta i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
The operation device is used for controlling the angle theta according to the relative of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
and the judging device is used for judging whether the aircraft control system is stable or not according to the first derivative.
In this embodiment, preferably, the inertia center conversion means includes:
a calculating unit for calculating an inertia time constant M according to all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI ；
A conversion unit for converting the control angle delta of the steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i 。
The beneficial effects that technical scheme that this application provided brought include:
the embodiment of the application provides a stability judging method and a stability judging system of an aircraft control system, wherein the control angles of a plurality of steering engines are subjected to nonlinear transformation of inertia centers to obtain relative control angles, all the relative control angles are subjected to Euclidean norms synthesis, the first derivative of the Euclidean norms is solved, whether the aircraft control system is unstable is judged according to the change condition of the first derivative, and a basis is provided for subsequent control measures of the aircraft control system.
Secondly, the embodiment of the application can also be used for predicting the time when the control system of the aircraft is unstable, and control measures can be taken in advance to prevent the control system of the aircraft from being unstable.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings that are needed in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
Fig. 1 is a flow chart of a method for determining stability of an aircraft control system according to an embodiment of the present application;
fig. 2 is a specific flow chart of a method for determining stability of an aircraft control system according to an embodiment of the present application;
fig. 3 is a block diagram of a stability determining system of an aircraft control system according to an embodiment of the present application;
fig. 4 is a block diagram of an inertia center conversion device according to an embodiment of the present application.
Detailed Description
For the purposes of making the objects, technical solutions and advantages of the embodiments of the present application more clear, the technical solutions of the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application, and it is apparent that the described embodiments are some embodiments of the present application, but not all embodiments. All other embodiments, which can be made by one of ordinary skill in the art without undue burden from the present disclosure, are within the scope of the present application based on the embodiments herein.
Referring to fig. 1, an embodiment of the present application provides a method for determining stability of an aircraft control system, where the aircraft control system is used for controlling a plurality of steering engines, the method includes the following steps:
step S1: monitoring the control angle delta of each steering engine i ；
Step S2: according to the monitored control angle delta of all steering engines i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
Step S3: according to the relative control angle theta of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
step S4: and judging whether the aircraft control system is stable or not according to the first derivative.
The embodiment of the application provides a stability judging method of an aircraft control system, which comprises the steps of carrying out nonlinear transformation of inertia centers on control angles of a plurality of steering engines to obtain relative control angles, synthesizing Euclidean norms on all the relative control angles, solving the first derivative of the Euclidean norms, judging whether the aircraft control system is unstable according to the change condition of the first derivative, and providing a basis for subsequent control measures of the aircraft control system.
Further, after determining whether the aircraft control system is stable, the determining method further includes:
if the aircraft control system is judged to be stable, continuing to monitor the control angle delta of each steering engine i 。
In this embodiment, if it is determined that the aircraft control system is stable, the output aircraft control system is stable, and the control angle δ of each steering engine is continuously monitored i . If the aircraft control system is judged to be unstable, outputting the instability of the aircraft control system so as to provide basis for subsequent control measures of the aircraft control system.
Specifically, the specific steps of the step S2 are as follows:
step S201: according to the inertia time constant M of all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI ；
Step S202: control angle delta of steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i I.e. θ i ＝δ i -δ COI ；
In the step S201, according to the inertia time constants M of all the steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI The calculation formula of (2) is as follows:
in delta COI Is the center angle of inertia; delta i The control angle of the ith steering engine; m is M i The inertia time constant of the ith steering engine;
in the step S3, according to the relative control angles theta of all steering engines i The calculation formula of the synthesized euclidean norm R is:
wherein R is Euclidean norm, and n is the number of steering engines; θ i Is the relative control angle of the ith steering engine.
In this embodiment, the control angle of the relative center of inertia (COI) of each steering engine varies with the aircraft control system during operation, and thus the euclidean norm R reflects the motion of the steering engine. When the aircraft control system is unstable, there must be at least one steering engine, and the control angle they are subjected to tends to infinity (taking additional limiting measures regardless of the control angle). If the attitude of the aircraft is out of control, in order to adjust the attitude of the aircraft, the control angle of one or more steering engines has a trend of increasing all the time, and then the difference of the control angle of one or more steering engines relative to the control angles of other steering engines also increases all the time.
Further, the specific steps of the step S4 are as follows:
step S401: outputting a waveform of the first derivative according to the first derivative of the Euclidean norm R;
step S402: judging whether a minimum value exists when the first derivative is larger than zero according to the first derivative waveform, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
In this embodiment, according to the waveform change curve of the first derivative, whether a minimum value exists or not can be intuitively observed in a section where the first derivative is greater than zero, and according to the change of the waveform of the first derivative of the synthesized euclidean norm R, it is simplified to determine that the aircraft control system cannot normally control the steering engine, that is, the aircraft control system is unstable.
Meanwhile, when the minimum value appears in the interval of which the first derivative is larger than zero, the state of the control system of the aircraft is changed, and the moment when the minimum value appears is the destabilizing moment.
The embodiment of the present application further provides a modification, in this modification, the specific steps of step S4 are:
step S401': solving a second derivative of the euclidean norm R;
step S402': and judging whether the first derivative and the second derivative are simultaneously larger than zero, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
In this modification, the time t is set to be a known time t 0 When the first derivative and the second derivative are b respectively 1 、b 2 The stability criteria of the aircraft control system are: if b 1 ＞0，b 2 < 0, after time t', becomes b 1 ＞0，b 2 And (3) judging the time t ' as the destabilizing time when the time t ' is more than 0, and slowly losing the stable state of the aircraft control system after the time t '.
Referring to fig. 2, an embodiment of the present application specifically provides a method for determining stability of an aircraft control system, where the method for determining stability specifically includes the following steps:
step A1: monitoring the control angle delta of each steering engine i And turning to step A2;
step A2: according to the monitored control angle delta of each steering engine i And an inertial time constant M corresponding to the steering engine i According to the calculation formulaCalculating an inertia center angle delta of the aircraft control system COI And turning to the step A3; wherein delta COI Is the center angle of inertia; delta i The control angle of the ith steering engine; m is M i The inertia time constant of the ith steering engine;
Step A3: control angle delta of each steering engine i Respectively subtracting the inertia center angles delta COI The relative control angle theta of each steering engine is obtained respectively i And turning to step A4;
step A4: according to the relative control angle theta of all steering engines i According to the calculation formulaSynthesizing Euclidean norm R, solving the first derivative and the second derivative of the Euclidean norm R, and turning to the step A5; wherein R is Euclidean norm, and n is the number of steering engines; θ i The relative control angle of the ith steering engine;
step A5: judging whether the first derivative is larger than zero, if so, turning to the step A6, otherwise, turning to the step A1;
step A6: and judging whether the second derivative is larger than zero, if so, turning to the step A7, otherwise, turning to the step A1.
Step A7: outputting the instability of the aircraft control system.
As a preferred embodiment of the embodiments of the present application, after solving the first derivative of the euclidean norm R, the stability discrimination method further includes:
solving a second derivative and a third derivative of the Euclidean norm R;
set at a known time t 0 When the first derivative, the second derivative and the third derivative are respectively b 1 、b 2 、b 3 ；
When said b 1 、b 2 、b 3 Satisfy b 1 ＜0，b 2 ＜0，b 3 ＞0，2b 1 b 3 -b 2 2 At > 0, the aircraft control system is predicted to destabilize at time t, wherein
In the preferred embodiment, the third derivative of the euclidean norm R is set at a known time t 0 When the first derivative, the second derivative and the third derivative are sequentially as follows：
R'(t 0 )＝b 1 ，R″(t 0 )＝b 2 ，R″'(t 0 )＝b 3 ；
Assuming that the third derivative of the euclidean norm R is a constant at time t, then
R″(t)＝b 3 (t-t 0 )+b 2 ，
Meanwhile, at time t, assuming R "(t) =0, then
Thus, the first and second heat exchangers are arranged,wherein the stability criterion of the aircraft control system is:
at time t 0 ，b 1 ＞0，b 2 < 0; at the time instant t of the time instant t,b 3 > 0; i.e. when said b 1 、b 2 、b 3 Satisfy b 1 ＜0，b 2 ＜0，b 3 ＞0，2b 1 b 3 -b 2 2 At > 0, at time->The stability of the aircraft control system is improved, and the stability criterion can be used for pre-judging the stability moment in advance so as to take measures in advance to prevent the instability of the aircraft control system.
As shown in fig. 3, the embodiment of the present application further provides a stability discriminating system of an aircraft control system, where the aircraft control system is used for controlling a plurality of steering engines, and the stability discriminating system includes:
monitoring device for monitoringMeasuring the control angle delta of each steering engine i ；
Inertia center conversion means for converting the control angle delta of all the steering engines according to the detected control angle delta i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
The operation device is used for controlling the angle theta according to the relative of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
and the judging device is used for judging whether the aircraft control system is stable or not according to the first derivative.
As shown in fig. 4, the inertia center conversion device includes a calculation unit and a conversion unit; the calculating unit is used for calculating the inertia time constant M of all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI The method comprises the steps of carrying out a first treatment on the surface of the The conversion unit is used for converting the control angle delta of the steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i The method comprises the steps of carrying out a first treatment on the surface of the Meanwhile, the calculation unit stores the inertia center angle delta COI The calculation formula of (2) is as follows:
in delta COI Is the center angle of inertia; delta i The control angle of the ith steering engine; m is M i The inertia time constant of the ith steering engine;
further, according to the relative control angle theta of all steering engines i The calculation formula of the synthesized euclidean norm R is:
wherein R is Euclidean norm, and n is the number of steering engines; θ i Is the relative control angle of the ith steering engine.
In this embodiment, the monitoring device is a sensor for detecting a steering engine control angle in the prior art, the inertia center conversion device, the operation device and the determination device are a terminal, the terminal includes a memory and a processor, the memory stores a computer program running on the processor, and the processor implements the method for determining the stability of the aircraft control system according to the above embodiment when executing the computer program.
In the description of the present application, it should be noted that the azimuth or positional relationship indicated by the terms "upper", "lower", etc. are based on the azimuth or positional relationship shown in the drawings, and are merely for convenience of description of the present application and simplification of the description, and are not indicative or implying that the apparatus or element in question must have a specific azimuth, be configured and operated in a specific azimuth, and thus should not be construed as limiting the present application. Unless specifically stated or limited otherwise, the terms "mounted," "connected," and "coupled" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the terms in this application will be understood by those of ordinary skill in the art as the case may be.
It should be noted that in this application, relational terms such as "first" and "second" and the like are used solely to distinguish one entity or action from another entity or action without necessarily requiring or implying any actual such relationship or order between such entities or actions. Moreover, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one … …" does not exclude the presence of other like elements in a process, method, article, or apparatus that comprises the element.
The foregoing is merely a specific embodiment of the application to enable one skilled in the art to understand or practice the application. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the application. Thus, the present application is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.
Claims (9)
1. A method for determining the stability of an aircraft control system, wherein the aircraft control system is used for controlling a plurality of steering engines, the method comprising the steps of:
monitoring the control angle delta of each steering engine i ；
According to the monitored control angle delta of all steering engines i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
According to the relative control angle theta of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
judging whether the aircraft control system is stable or not according to the first derivative;
after solving the first derivative of the euclidean norm R, the stability discrimination method further comprises:
solving a second derivative and a third derivative of the Euclidean norm R;
set at a known time t 0 When the first derivative, the second derivative and the third derivative are respectively b 1 、b 2 、b 3 ；
When said b 1 、b 2 、b 3 Satisfy b 1 ＜0，b 2 ＜0，b 3 ＞0，2b 1 b 3 -δ 2 2 At > 0, the aircraft control system is predicted to destabilize at time t, wherein
2. The method for determining the stability of an aircraft control system according to claim 1, wherein said control angle δ is determined based on all of said steering engines that are monitored i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i The specific steps of (a) are as follows:
according to the inertia time constant M of all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI ；
Control angle delta of steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i 。
3. The method for determining the stability of an aircraft control system according to claim 2, wherein said inertial time constant M is based on all of said steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI The calculation formula of (2) is as follows:
in delta COI Is the center angle of inertia; delta i The control angle of the ith steering engine; m is M i The inertia time constant of the ith steering engine;
4. a method for determining the stability of an aircraft control system according to claim 2 or 3, characterized in that said control angle θ is controlled according to the relative of all said steering engines i The calculation formula of the synthesized euclidean norm R is:
wherein R is Euclidean norm, and n is the number of steering engines; θ i Is the relative control angle of the ith steering engine.
5. The method of determining the stability of an aircraft control system according to claim 1, wherein after determining whether the aircraft control system is stable, the method further comprises:
if the aircraft control system is judged to be stable, continuing to monitor the control angle delta of each steering engine i 。
6. The method for determining the stability of an aircraft control system according to claim 1, wherein the specific step of determining whether the aircraft control system is stable according to the first derivative is:
outputting a waveform of the first derivative according to the first derivative of the Euclidean norm R;
judging whether a minimum value exists when the first derivative is larger than zero according to the first derivative waveform, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
7. The method for determining the stability of an aircraft control system according to claim 1, wherein the specific step of determining whether the aircraft control system is stable according to the first derivative is:
solving a second derivative of the euclidean norm R;
and judging whether the first derivative and the second derivative are simultaneously larger than zero, if so, destabilizing the aircraft control system, otherwise, stabilizing the aircraft control system.
8. A stability discrimination system of an aircraft control system that performs discrimination using the stability discrimination method of an aircraft control system according to claim 1, wherein the aircraft control system is used for control of a plurality of steering engines, the stability discrimination system comprising:
monitoring means for monitoring the control angle delta of each steering engine i ；
Inertia center conversion means for converting the control angle delta of all the steering engines according to the detected control angle delta i Converting the control angle delta of the steering engine i Is the relative control angle theta under the inertia center coordinate system of the aircraft control system i ；
The operation device is used for controlling the angle theta according to the relative of all steering engines i Synthesizing a Euclidean norm R, and solving a first derivative of the Euclidean norm R;
and the judging device is used for judging whether the aircraft control system is stable or not according to the first derivative.
9. The system for determining the stability of an aircraft control system according to claim 8, wherein the inertia center conversion means includes:
a calculating unit for calculating an inertia time constant M according to all steering engines i And control angle delta i Calculating the inertia center angle delta of the aircraft control system COI ；
A conversion unit for converting the control angle delta of the steering engine i Subtracting the inertia center angle delta COI Obtaining a relative control angle theta of the steering engine i 。
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