CN103149931B - Fault diagnosing and tolerance control method for aircraft three-dimensional movement - Google Patents
Fault diagnosing and tolerance control method for aircraft three-dimensional movement Download PDFInfo
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Abstract
The invention provides a fault diagnosing and tolerance control method for aircraft three-dimensional movement. The method comprises the following steps of according to a model and a state combination index discriminant when an aircraft has no fault, carrying out the fault diagnosing; and when the aircraft has a fault, firstly, classifying the fault, analyzing the influence of the whole-aircraft parameters and the flight dynamic properties according to the determined fault type, and remodeling an abrupt-change system after the fault occurs, so as to obtain a new model description when the aircraft has the fault. The design method of a tolerance controller is characterized in that the output of a fault-free model can meet the requirements of military standards by the designed control amount; and when the aircraft has the fault, the new model output is stable by the control input which is formed by the control amount and the non-artificial additional control amount caused by the aircraft fault.
Description
Technical field
The present invention relates to aircraft fault diagnosis and Fault-tolerant Control Design method, particularly relate to a kind of aircraft three-dimensional motion fault diagnosis and fault tolerant control method.
Background technology
Military spacecraft is a kind of important operational weapon, civil aircraft is then a kind of important vehicles, in national defence and daily life and work, play more and more important role, the fault diagnosis research of aerocraft system is more and more subject to people's attention; Fault diagnosis is of crucial importance for the system that security requirement is high, detects the system failure in time, makes self-repairing system timely reconfigurable control rule, can avoid system crash and the material damage caused thus and casualties.
First the fault detection and diagnosis technology of taking as the leading factor with Analysis design grew up in the U.S. from early 1970s.1971, masschusetts, u.s.a Polytechnics Beard first proposed and replaces hardware redundancy with Analysis design in his PhD dissertation, and pass through systems self-organize, make system closed-loop stabilization, obtained the new thought of system failure information by the output of comparative observation device, indicate the birth of the fault diagnosis technology based on Analysis design.Research abroad for fault diagnosis mainly concentrates on industrial system and aerospace, at present based on the fault diagnosis of Study of Generalized system and nonlinear system, the method adopted is very various, specifically there is sliding mode observer method, neural net method, expert system method, small wave converting method, Unknown Input Observer method etc., and how several method is combined use.For the method for diagnosing faults comparative maturity of linear system, there is the more complete diagnostic method of a set of ratio, and had the application system of some reality.USAF have employed fault diagnosis technology on fighter experimental.The domestic paper to aircraft fault diagnosis and faults-tolerant control Study on Problems is also a lot, the method adopted also be intelligent or with the adaptive fusion method of state space (1, Zhang Junfeng, Hu Shousong be based on the fighter plane fault diagnosis of multi-kernel learning support vector machine, Southeast China University's journal (natural science edition), 2007,37th volume supplementary issue (I), pp1-5; Moral dragon Dong Chaoyang after the first Wang Qing of yellow happiness, based on the hypersonic aircraft faults-tolerant control of model reference adaptive, Nanjing Aero-Space University's journal, the 43rd volume supplementary issue in 2011).
When aircraft breaks down, complete machine aerodynamic force, moment, parameter and input quantity all may change, the all aerodynamic force of full machine, moment etc. will be caused all to suddenly change as aircraft wing penetrates an aperture, and different maneuvers can increase an equivalent inpnt; But, model before and after aircraft breaks down by current research method describes with identical structure and parameter, only the parameters such as input coefficient are adjusted, in design, do not consider air dynamic behaviour, cause the faults-tolerant control scheme provided lose contact with reality aircraft and be difficult to application.
Summary of the invention
Only aircraft input coefficient is adjusted to solve existing Fault Tolerance Control Technology, design in do not consider air dynamic behaviour, cause the faults-tolerant control scheme provided lose contact with reality aircraft and be difficult to apply technical matters, The present invention gives a kind of aircraft three-dimensional motion fault diagnosis and fault tolerant control method, the method carries out fault diagnosis according to model during aircraft non-fault and combinations of states index discriminant; First to failure modes when aircraft breaks down, then analyze the impact of the full machine parameter of aircraft and flight dynamics characteristic for the fault category determined, suddenly change after then again occurring fault system modelling, obtains the model description that aircraft breaks down stylish; The method for designing of fault-tolerant controller is that the controlled quentity controlled variable designed makes the output of non-fault model reach army's mark requirement, and the model stable output that the control inputs when aircraft that the non-artificial additional controlled quentity controlled variable that this controlled quentity controlled variable and aircraft fault cause forms jointly breaks down makes aircraft break down stylish.
The technical solution adopted for the present invention to solve the technical problems is: a kind of aircraft three-dimensional motion fault diagnosis and fault tolerant control method, be characterized in comprising the following steps:
1, the nonlinear model of three-dimensional motion during aircraft non-fault is:
（1）
In formula: state variable
represent flying height, flying speed, the air-flow angle of attack, yaw angle, angular velocity in roll, rate of pitch, yaw rate respectively;
refer to roll angle, the angle of pitch, crab angle respectively;
for acceleration of gravity;
represent dynamic pressure, wing area, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis respectively, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force respectively;
be respectively rolling moment, yawing, pitching moment;
for controlled quentity controlled variable; Usual yaw angle is less than
,
can not be zero;
2, fault method for detecting is: when combinations of states index
, then judge that aircraft has fault;
Wherein:
（2）
In formula:
be respectively according to (1) formula in the time
calculate
value;
be respectively in the time
actual measurement obtains
value,
for the detection threshold of setting;
for sign function;
represent the
individual sampled point, the corresponding sampling time is
;
for specification error adds up number,
for the sampling period;
3, first to failure modes when aircraft breaks down, the impact of the full machine parameter of aircraft and flight dynamics characteristic is analyzed again for the fault category determined, then suddenly change after again occurring fault system modelling, obtains the nonlinear model of three-dimensional motion when aircraft breaks down and be described as:
（3）
In formula:
not Biao Shi the wing area of aircraft when breaking down, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force when aircraft breaks down respectively;
,
with
rolling moment when aircraft breaks down, yawing, pitching moment;
control inputs when aircraft breaks down, the non-artificial additional controlled quentity controlled variable that aircraft fault causes
;
4, the method for designing of fault-tolerant controller is: the controlled quentity controlled variable designed
nonlinear model (1) formula stability margin is made to reach army's mark requirement, and this controlled quentity controlled variable
with the non-artificial additional controlled quentity controlled variable that aircraft fault causes
the control inputs when aircraft of common composition breaks down
make nonlinear model (3) formula stable output.
The invention has the beneficial effects as follows: carry out fault diagnosis by model during aircraft non-fault and combinations of states index discriminant, can effectively detect fault and failure message be classified, full machine parameter and the modeling again of flight dynamics characteristic is considered when aircraft breaks down, two model bases are designed fault-tolerant controller makes the output of non-fault model reach army's mark requirement, and the model stable output that the control inputs of the aircraft that the non-artificial additional controlled quentity controlled variable that this controlled quentity controlled variable and aircraft fault cause forms jointly when breaking down makes aircraft break down stylish, solve existing Fault Tolerance Control Technology only to adjust aircraft input coefficient, air dynamic behaviour is not considered in design, cause the faults-tolerant control scheme provided lose contact with reality aircraft and be difficult to apply technical matters.
Below in conjunction with embodiment, the present invention is elaborated.
Embodiment
1, the nonlinear model of three-dimensional motion during aircraft non-fault is:
（1）
In formula: state variable
represent flying height, flying speed, the air-flow angle of attack, yaw angle, angular velocity in roll, rate of pitch, yaw rate respectively;
refer to roll angle, the angle of pitch, crab angle respectively;
for acceleration of gravity;
represent dynamic pressure, wing area, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis respectively, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force respectively;
be respectively rolling moment, yawing, pitching moment;
for controlled quentity controlled variable; Usual yaw angle is less than usually
,
can not be zero;
2, fault method for detecting is: when combinations of states index
, then judge that aircraft has fault;
Wherein:
（2）
In formula:
be respectively according to (1) formula in the time
calculate
value;
be respectively in the time
actual measurement obtains
value,
for the detection threshold of setting;
for sign function;
represent the
individual sampled point, the corresponding sampling time is
;
for specification error adds up number,
for the sampling period;
3, first to failure modes when aircraft breaks down, the impact of the full machine parameter of aircraft and flight dynamics characteristic is analyzed again for the fault category determined, then suddenly change after again occurring fault system modelling, and the ternary nonlinear model obtaining three-dimensional motion when aircraft breaks down is described as:
（3）
In formula:
not Biao Shi the wing area of aircraft when breaking down, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force when aircraft breaks down respectively;
,
with
rolling moment when aircraft breaks down, yawing, pitching moment;
control inputs when aircraft breaks down, the non-artificial additional controlled quentity controlled variable that aircraft fault causes
;
4, the method for designing of fault-tolerant controller is: the controlled quentity controlled variable designed
nonlinear model (1) formula stability margin is made to reach army's mark requirement, and this controlled quentity controlled variable
with the non-artificial additional controlled quentity controlled variable that aircraft fault causes
the control inputs when aircraft of common composition breaks down
make nonlinear model (3) formula stable output.
Claims (1)
1. aircraft three-dimensional motion fault diagnosis and a fault tolerant control method, is characterized in comprising the following steps:
A during () aircraft non-fault, the nonlinear model of three-dimensional motion is:
（1）
In formula: state variable
represent flying height, flying speed, the air-flow angle of attack, yaw angle, angular velocity in roll, rate of pitch, yaw rate respectively;
refer to roll angle, the angle of pitch, crab angle respectively;
for acceleration of gravity;
represent dynamic pressure, wing area, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis respectively, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force respectively;
be respectively rolling moment, yawing, pitching moment;
for controlled quentity controlled variable; Usual yaw angle is less than
,
can not be zero;
B () fault method for detecting is: when combinations of states index
, then judge that aircraft has fault;
Wherein:
（2）
In formula:
be respectively according to (1) formula in the time
calculate
value;
be respectively in the time
actual measurement obtains
value,
for the detection threshold of setting;
for sign function;
represent the
individual sampled point, the corresponding sampling time is
;
for specification error adds up number,
for the sampling period;
First to failure modes when () aircraft breaks down c, the impact of the full machine parameter of aircraft and flight dynamics characteristic is analyzed again for the fault category determined, then suddenly change after again occurring fault system modelling, obtains the nonlinear model of three-dimensional motion when aircraft breaks down and be described as:
（3）
In formula:
not Biao Shi the wing area of aircraft when breaking down, vehicle mass, the moment of inertia around axis system X-axis, the moment of inertia around axis system Y-axis, amass around the moment of inertia of axis system Z axis, the moment of inertia of axis system XZ axle;
represent longitudinal force, side force and normal force when aircraft breaks down respectively;
,
with
rolling moment when aircraft breaks down, yawing, pitching moment;
control inputs when aircraft breaks down, the non-artificial additional controlled quentity controlled variable that aircraft fault causes
;
D the method for designing of () fault-tolerant controller is: the controlled quentity controlled variable designed
nonlinear model (1) formula stability margin is made to reach army's mark requirement, and this controlled quentity controlled variable
with the non-artificial additional controlled quentity controlled variable that aircraft fault causes
the control inputs when aircraft of common composition breaks down
make nonlinear model (3) formula stable output.
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| CN103616816B (en) * | 2013-11-15 | 2016-04-06 | 南京航空航天大学 | A kind of hypersonic aircraft elevator fault control method |
| CN104020670B (en) * | 2014-05-26 | 2017-02-22 | 南京航空航天大学 | Three-freedom helicopter fault tolerance control device based on support vector machine and method thereof |
| CN109696090B (en) * | 2019-01-16 | 2020-10-16 | 哈尔滨工业大学 | Online single-shot thrust identification method for carrier rocket/missile |
| CN113204193B (en) * | 2021-05-06 | 2022-10-25 | 北京航空航天大学 | Fault-tolerant control method and device for aircraft and electronic equipment |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0511730A2 (en) * | 1991-04-26 | 1992-11-04 | Litton Systems, Inc. | Fault-tolerant inertial navigation system |
| CN102200776A (en) * | 2010-03-25 | 2011-09-28 | 南京航空航天大学 | Fault diagnosis method for actuator of flight control system |
| CN102707616A (en) * | 2012-05-31 | 2012-10-03 | 西北工业大学 | Aircraft triangle model-based controller area design method |
| CN102854874A (en) * | 2012-06-18 | 2013-01-02 | 南京航空航天大学 | A plurality of united observer based fault diagnosis and fault-tolerant control device and method |
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2013
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Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0511730A2 (en) * | 1991-04-26 | 1992-11-04 | Litton Systems, Inc. | Fault-tolerant inertial navigation system |
| CN102200776A (en) * | 2010-03-25 | 2011-09-28 | 南京航空航天大学 | Fault diagnosis method for actuator of flight control system |
| CN102707616A (en) * | 2012-05-31 | 2012-10-03 | 西北工业大学 | Aircraft triangle model-based controller area design method |
| CN102854874A (en) * | 2012-06-18 | 2013-01-02 | 南京航空航天大学 | A plurality of united observer based fault diagnosis and fault-tolerant control device and method |
Non-Patent Citations (1)
| Title |
|---|
| 近空间飞行器故障诊断与容错控制的研究进展;姜斌 等;《南京航空航天大学学报》;20121015(第5期);第603-610页 * |
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