CN103466102B - Flying vehicles control effect redistribution method in cross configuration actuator failure situation - Google Patents
Flying vehicles control effect redistribution method in cross configuration actuator failure situation Download PDFInfo
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- CN103466102B CN103466102B CN201310438997.4A CN201310438997A CN103466102B CN 103466102 B CN103466102 B CN 103466102B CN 201310438997 A CN201310438997 A CN 201310438997A CN 103466102 B CN103466102 B CN 103466102B
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Abstract
Flying vehicles control effect redistribution method in cross configuration actuator failure situation, step is as follows: (1) to position actuator failure according to the aircraft state information of Real-time Collection and judges; (2) according to the control law that aircraft has designed, the control system actuating unit pivot angle size of cross configuration actuating unit in non-faulting situation is calculated; (3) according to cross actuating unit configuring condition, according to fault localization and result of determination, distribution is recalculated to control system actuating unit pivot angle size; (4) actuating unit swings according to the pivot angle size recalculating distribution, completes the effective control to aircraft.
Description
Technical field
The present invention relates to aerospacecraft Reconfigurable Control technical field.
Background technology
Along with the development of space technology and the variation of task, the emission density of aerospacecraft increases, mission requirements are complicated, and requirement aircraft security reliably being completed to aerial mission is also more and more higher.Because aerospacecraft manufacture, launch cost are high, guarantee that each reliable flight is particularly important.One of normal measure adopted at present is the quality control strengthening development, emission process, improves flight reliability; Another measure is redundant configuration control system equipment, when certain device fails, utilizes alternate device to replace.The enforcement of a rear measure not only increases development cost, and is often subject to the restriction of aircraft space and weight etc. and more difficult enforcement.When therefore there is non-lethal fault in aerospacecraft flight course, if by fully taping the latent power to control system ability, realize the reasonable disposition of flying vehicles control ability when aircraft generation non-lethal fault, complete the effective control to aircraft, significant to the reliability promoting aerospacecraft.
Summary of the invention
Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, a kind of cross is provided to configure flying vehicles control effect redistribution method in actuator failure situation, when there is non-lethal fault in the actuating unit that the method configures for " ten " font, control system ability is fully taped the latent power, require to recalculate distribution to the control system actuating unit pivot angle calculated in non-faulting situation, realize the reasonable disposition of flying vehicles control ability when actuating unit generation non-lethal fault, complete the effective control to aircraft.
Technical solution of the present invention is: flying vehicles control effect redistribution method in cross configuration actuator failure situation, described cross configuration actuating unit refers to that flight control system has four actuating units, and these four actuating units are by " ten " font Install and configure; Described actuator failure situation refers to actuating unit generation non-lethal fault, namely, the fault that one stuck or pendulum angle occurs and diminishes is only had in four actuating units, and compensated by the reasonable swing of other three non-faulting actuating units, control system has ability aircraft being implemented to effectively control; Step is as follows:
(1) according to the aircraft state information of Real-time Collection actuator failure positioned and judge;
(2) according to the control law that aircraft has designed, the expectation pivot angle of cross configuration actuating unit in non-faulting situation is calculated;
(3) according to cross actuating unit configuring condition, according to fault localization and result of determination, distribution is recalculated to control system actuating unit pivot angle size;
(4) actuating unit swings according to the pivot angle size recalculating distribution, completes the effective control to aircraft.
Being implemented as follows of described step (3):
(3.1) according to the fault localization in step (1) and result of determination, the angle that the actuating unit of et out of order can swing is determined;
(3.2) the expectation pivot angle δ of four actuating units of cross configuration in non-faulting situation is calculated
1d, δ
2d, δ
3dand δ
4d, determine the angle of the actual swing of actuating unit of et out of order in step (3.1), and calculate the actual pivot angle of this actuating unit and the deviation expecting pivot angle;
(3.3) according to the deviation that step (3.2) calculates, calculate the actual pivot angle of other three actuating units of non-et out of order, be specially:
When actuating unit 1 et out of order, then the actual pivot angle of actuating unit 2, actuating unit 3 and actuating unit 4 calculates as follows:
δ
2a＝δ
2d+de
1
δ
3a＝δ
3d+de
1
δ
4a＝δ
4d-de
1
When actuating unit 2 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 3 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
2
δ
3a＝δ
3d-de
2
δ
4a＝δ
4d+de
2
When actuating unit 3 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
3
δ
2a＝δ
2d-de
3
δ
4a＝δ
4d+de
3
When actuating unit 4 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 3 calculates as follows:
δ
1a＝δ
1d-de
4
δ
2a＝δ
2d+de
4
δ
3a＝δ
3d+de
4
Above-mentioned, de
1, de
2, de
3, de
4be respectively the deviation that corresponding fault actuating unit calculates according to step (3.2).
The present invention compared with prior art beneficial effect is:
(1) the present invention is directed to the actuating unit of " ten " font configuration when there is non-lethal fault, control system ability is fully taped the latent power, require to recalculate distribution to the control system actuating unit pivot angle calculated in non-faulting situation, realize the reasonable disposition of flying vehicles control ability when actuating unit generation non-lethal fault, complete the effective control to aircraft.The method that the present invention proposes can promote the ability that the anti-driving engine of aerospacecraft stops, controls the actuator failures such as rudder face is stuck, therefore to aircraft development and aerospacecraft, highly reliable to complete aerial mission significant, has broad application prospects in aerospacecraft development.
(2) in the present invention control action to recalculate assigning process calculated amount little, obviously can not increase the calculated amount of aircraft computer, the real-time of flying vehicles control process can be ensured.
Accompanying drawing explanation
Fig. 1 is cross of the present invention configuration actuating unit schematic rear view.
Fig. 2 is diagram of circuit of the present invention.
When Fig. 3 is the generation of actuating unit trouble free, the deviation curve at pitch attitude angle in aircraft flight process.
Fig. 4 is actuating unit generation non-lethal fault, when not adopting control action redistribution method of the present invention, and the deviation curve at pitch attitude angle in aircraft flight process.
Fig. 5 is actuating unit generation non-lethal fault, when adopting control action redistribution method of the present invention, and the deviation curve at pitch attitude angle in aircraft flight process.
Detailed description of the invention
For the situation of cross configuration actuating unit generation non-lethal fault, namely, aerospacecraft control system in current flight process has four actuating units, these four actuating units are by " ten " font Install and configure, as shown in Figure 1, the fault stuck or pendulum angle occurring diminish when only having one in four actuating units, and compensated by reasonable swings of other three non-faulting actuating units, control system has ability aircraft being implemented to effectively control.In this case, first the control system actuating unit pivot angle size requirements in non-faulting situation is calculated, then according to cross actuating unit configuring condition, according to fault localization and result of determination, distribution is recalculated to control system actuating unit pivot angle size, completes the effective control to aircraft.Flying vehicles control effect redistribution method in cross configuration actuator failure situation, as shown in Figure 2, step is as follows:
(1) according to the aircraft state information of Real-time Collection actuator failure positioned and judge;
Actuator failure is positioned and judges to refer to that determining is the fault which actuating unit there occurs stuck or pendulum angle and diminishes in this step.This is by being that each actuating unit installs pendulum angle survey sensor, and each actuating unit pivot angle data analyzing Real-time Collection with expect that the difference of pivot angle angle value completes and actuator failure positioned and judges.
(2) according to the control law that aircraft has designed, the control system actuating unit pivot angle size of cross configuration actuating unit in non-faulting situation is calculated;
According to the control law that aircraft has designed in this step, first try to achieve the equivalent pivot angle of pitching, driftage and rolling three passages
, δ
ψ, and δ
γ(this step is known); Then according to " ten " font actuating unit layout, the expectation pivot angle δ trying to achieve four actuating units in non-faulting situation is decomposed
1d, δ
2d, δ
3dand δ
4d.Computation process is as follows:
δ
1d＝δ
ψ-δ
γ
δ
3d＝δ
ψ+δ
γ
(3) according to cross actuating unit configuring condition, according to fault localization and result of determination, distribution is recalculated to control system actuating unit pivot angle size;
For the situation of cross configuration actuating unit generation non-lethal fault in this step, namely, aerospacecraft control system in current flight process has four actuating units, these four actuating units are by " ten " font Install and configure, as shown in Figure 1, the fault stuck or pendulum angle occurring diminish when only having one in four actuating units, and compensated by reasonable swings of other three non-faulting actuating units, control system has ability aircraft being implemented to effectively control.In this case, the process of distribution is recalculated to actuating unit pivot angle size as follows:
When actuating unit 1 et out of order, according to failure condition, if the angle that actuating unit 1 reality swings is δ
1a, with expectation pivot angle δ
1ddeviation be de
1=δ
1d-δ
1a, then the actual pivot angle of actuating unit 2, actuating unit 3 and actuating unit 4 calculates as follows:
δ
2a＝δ
2d+de
1
δ
3a＝δ
3d+de
1
δ
4a＝δ
4d-de
1
When actuating unit 2 et out of order, according to failure condition, if the angle that actuating unit 2 reality swings is δ
2a, with expectation pivot angle δ
2ddeviation be de
2=δ
2d-δ
2a, then the actual pivot angle of actuating unit 1, actuating unit 3 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
2
δ
3a＝δ
3d-de
2
δ
4a＝δ
4d+de
2
When actuating unit 3 et out of order, according to failure condition, if the angle that actuating unit 3 reality swings is δ
3a, with expectation pivot angle δ
3ddeviation be de
3=δ
3d-δ
3a, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
3
δ
2a＝δ
2d-de
3
δ
4a＝δ
4d+de
3
When actuating unit 4 et out of order, according to failure condition, if the angle that actuating unit 4 reality swings is δ
4a, with expectation pivot angle δ
4ddeviation be de
4=δ
4d-δ
4a, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 3 calculates as follows:
δ
1a＝δ
1d-de
4
δ
2a＝δ
2d+de
4。
δ
3a＝δ
3d+de
4
(4) actuating unit swings according to the pivot angle size recalculating distribution, completes the effective control to aircraft.
Namely according to current actuator failure situation, the δ distributing and obtain will be recalculated
1a, δ
2a, δ
3a, and δ
4aoutput to corresponding actuating unit, actuating unit is swung according to the pivot angle size recalculating distribution, completes the effective control to aircraft.
Embodiment
For certain aircraft vertical passage pitch attitude angle control process, when actuating unit trouble free occurs, under the control law effect designed, in aircraft flight process, the deviation at pitch attitude angle as shown in Figure 3, pitch attitude angular deviation controls in less scope, Fig. 3 shows that control system is functional, and aircraft can omnidistance stabilized flight; If in aircraft flight to there occurs the actuating unit 2 stuck non-lethal fault at positive 10 degree when 10 seconds, do not adopt control action redistribution method of the present invention, with under described identical control law effect above, in aircraft flight process, the deviation at pitch attitude angle as shown in Figure 4, pitch attitude angular deviation increases rapidly, Fig. 4 shows aircraft vertical passage pitch attitude angle rapid divergence, and aircraft can not complete aerial mission; If in aircraft flight to there occurs the actuating unit 2 stuck non-lethal fault at positive 10 degree when 10 seconds, and adopt control action redistribution method of the present invention to recalculate distribution to actuating unit pivot angle size, with under described identical control law effect above, in aircraft flight process, the deviation at pitch attitude angle as shown in Figure 5, pitch attitude angular deviation control with trouble free a situation arises lower identical numerically, Fig. 5 shows that control system reaches and trouble free a situation arises lower identical performance, ensure that aircraft can omnidistance stabilized flight.
The unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.
Claims (2)
1. flying vehicles control effect redistribution method in cross configuration actuator failure situation, described cross configuration actuating unit refers to that flight control system has four actuating units, and these four actuating units are by " ten " font Install and configure; Described actuator failure situation refers to actuating unit generation non-lethal fault, namely, the fault that one stuck or pendulum angle occurs and diminishes is only had in four actuating units, and compensated by the reasonable swing of other three non-faulting actuating units, control system has ability aircraft being implemented to effectively control; It is characterized in that step is as follows:
(1) according to the aircraft state information of Real-time Collection actuator failure positioned and judge;
(2) according to the control law that aircraft has designed, the expectation pivot angle of cross configuration actuating unit in non-faulting situation is calculated;
(3) according to cross actuating unit configuring condition, according to fault localization and result of determination, distribution is recalculated to control system actuating unit pivot angle size;
(4) other three non-faulting actuating units swing according to the pivot angle size recalculating distribution, complete the effective control to aircraft.
2. flying vehicles control effect redistribution method in cross configuration actuator failure situation according to claim 1, is characterized in that: being implemented as follows of described step (3):
(3.1) according to the fault localization in step (1) and result of determination, the angle that the actuating unit of et out of order can swing is determined;
(3.2) the expectation pivot angle δ of four actuating units of cross configuration in non-faulting situation is calculated
1d, δ
2d, δ
3dand δ
4d, determine the angle of the actual swing of actuating unit of et out of order in step (3.1), and calculate the actual pivot angle of this actuating unit and the deviation expecting pivot angle;
(3.3) according to the deviation that step (3.2) calculates, the actual pivot angle of other three actuating units of non-et out of order is calculated; Be specially:
When actuating unit 1 et out of order, then the actual pivot angle of actuating unit 2, actuating unit 3 and actuating unit 4 calculates as follows:
δ
2a＝δ
2d+de
1
δ
3a＝δ
3d+de
1
δ
4a＝δ
4d-de
1
When actuating unit 2 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 3 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
2
δ
3a＝δ
3d-de
2
δ
4a＝δ
4d+de
2
When actuating unit 3 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 4 calculates as follows:
δ
1a＝δ
1d+de
3
δ
2a＝δ
2d-de
3
δ
4a＝δ
4d+de
3
When actuating unit 4 et out of order, then the actual pivot angle of actuating unit 1, actuating unit 2 and actuating unit 3 calculates as follows:
δ
1a＝δ
1d-de
4
δ
2a＝δ
2d+de
4
δ
3a＝δ
3d+de
4
Above-mentioned, de
1, de
2, de
3, de
4be respectively the deviation that corresponding fault actuating unit calculates according to step (3.2).
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| CN104200016B (en) * | 2014-08-20 | 2017-05-10 | 中国运载火箭技术研究院 | Multi-control-surface aircraft modal calculation and verification method |
| CN104898635B (en) * | 2014-10-27 | 2017-11-07 | 中国运载火箭技术研究院 | A kind of high thrust liquid rocket failure reconfiguration control method |
| CN104743100B (en) * | 2015-03-03 | 2017-01-25 | 北京航天自动控制研究所 | Redistribution method of control action for aircraft under fault condition of executing mechanisms for X-type configuration |
| CN105157487B (en) * | 2015-09-01 | 2017-08-29 | 四川航天系统工程研究所 | Missile Actuator failure tolerant control method based on Analysis design |
| CN108454884B (en) * | 2018-02-27 | 2020-09-18 | 北京控制工程研究所 | Power rise safety guidance method and system |
| CN110989559B (en) * | 2019-12-20 | 2021-07-06 | 北京中科宇航探索技术有限公司 | Fault redundancy attitude control device and attitude control method of actuating mechanism |
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