CN102346488B - Large aircraft roll channel control instruction computing method - Google Patents

Large aircraft roll channel control instruction computing method Download PDF

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CN102346488B
CN102346488B CN 201110294432 CN201110294432A CN102346488B CN 102346488 B CN102346488 B CN 102346488B CN 201110294432 CN201110294432 CN 201110294432 CN 201110294432 A CN201110294432 A CN 201110294432A CN 102346488 B CN102346488 B CN 102346488B
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pitch angle
instruction
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CN102346488A (en
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周海军
张翔伦
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No 618 Research Institute of China Aviation Industry
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No 618 Research Institute of China Aviation Industry
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Abstract

The invention belongs to a fly control technology and relates to improvement of a large aircraft roll channel control instruction computing method. The control instruction is computed through the following steps: a steering wheel angular displacement sensor measures a steering wheel angular displacement signal and inputs the steering wheel angular displacement signal into a fly control computer; the fly control computer receives a driver operation instruction, and the river operation instruction is resolved by an instruction forming filter to form a roll angular velocity instruction Pc; an inclination angle instruction gamma c is calculated; and an angular velocity instruction Wxc is determined. In the invention, the computing process does not need a feedback parameter of the aircraft, the designing and comprehensive difficulty is small, the evaluation process is simple, the posture limiting performance is ensured, and values in practical application are realized.

Description

A kind of large aircraft roll channel steering order computing method
Technical field
The invention belongs to flight control technology, relate to the improvement to large aircraft roll channel steering order computing method.
Background technology
This technology be full Flying by wire aircraft control law one of the emphasis that must solve, it is to concern flight safety, alleviate driver's burden and bring into play to greatest extent aeroplane performance, realizes the important means that the convenience of expectation is handled.The general design requirement of large aircraft roll channel is as follows:
(1) | γ |<γ hold, control wheel angular displacement ratio is in pitch angle speed, neutral spiral stability;
(2) γ hold<| γ |<γ max, control wheel angular displacement ratio is in γ, transport screw stability;
(3) | γ |<γ hold, loose bar keeps;
(4) γ hold<| γ |<γ max, loose bar keeps γ hold;
(5) | γ |<γ ' max, control wheel angular displacement ratio is in γ, hypervelocity pattern, transport screw stability.
In formula: the pitch angle that γ is aircraft, γ holdfor the pitch angle retention value of aircraft, γ maxfor aircraft allowable angle of inclination, γ ' maxfor the allowable angle of inclination of aircraft under the hypervelocity pattern.
This way to solve the problem mainly is divided into two classes at present: a class is the open loop control program, when the pitch angle of aircraft surpasses certain value, by the anti-pass actuator, to reverse control of control wheel, stop pitch angle further to increase, when the pine dish, opposite force makes within aircraft gets back to normal range.The drawback of the method is out of true control pitch angle, during practical flight, and after the driver unclamps control wheel, the pitch angle desired value that can't know for sure, this scheme also keeps function without the pine dish simultaneously.The another kind of close-loop control scheme that is based on the pitch angle feedback, when the pitch angle of aircraft reality surpasses certain value, reduce driver's forward direction steering order, reach the purpose of restriction, this scheme is due to the feedback parameters that comprises aircraft in command calculations, design and comprehensive difficulty are large, and the actual assessment process is comparatively complicated.
Summary of the invention
The objective of the invention is: propose a kind ofly not need feedback parameters, design and the comprehensive difficulty of aircraft little, the simple large aircraft roll channel of evaluation process steering order computing method.
Technical scheme of the present invention is: a kind of large aircraft roll channel steering order computing method, based on flight-control computer, air data system and inertial measurement system, it is characterized in that, and the calculation procedure of steering order is:
1, the control wheel angular displacement sensor is measured the control wheel angular displacement signal and is inputed to flight-control computer;
2, flight-control computer is received the pilot control instruction, forms roll angle speed instruction Pc after the instruction formed filter resolves;
3, calculate pitch angle instruction γ c: one of computing method are following method:
3.1, when air speed is in the normal flight envelope curve, hypervelocity sign overspeed sw=0;
3.1.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γn*γ offset………………………………………………………[1]
In formula: K γ nfor scale-up factor, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[2]
In formula: γ ' limfor the pitch angle of a upper computation period, the pitch angle deviation of the first computation period is zero, γ holdpitch angle retention value for aircraft;
3.1.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[3]
In formula: p climangular speed instruction through amplitude limit;
3.1.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[Pcmd-K fade*(γ′ lim-γ' int)]……………………………[4]
In formula: K fadefor the desalination factor, γ ' limfor the pitch angle of a upper computation period through amplitude limit, γ ' intintegration pitch angle for a upper computation period; K fade=50～100;
3.1.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxallowable angle of inclination for aircraft;
3.2, when aircraft during in overspeed condition, hypervelocity sign overspeed sw=1;
3.2.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γs*γ offset…………………………………………[5]
In formula: K γ sfor scale-up factor, K γ s=0～10, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[6]
In formula: γ holdit is the pitch angle retention value of aircraft;
Y ' limit is the pitch angle value of a upper computation of Period;
3.2.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[7]
In formula: p climangular speed instruction through amplitude limit;
3.2.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[Pcmd-K fade*(γ' lim-γ' int)]……………………………[8]
In formula: K fadefor the desalination factor;
In formula: γ ' intthe γ of a upper computation of Period intvalue;
3.2.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxit is aircraft allowable angle of inclination value;
4, determine angular speed instruction Wxc:
Wxc=Pcmd*Kp………………………………………[9]
In formula: Kp is scale factor, and value is 0～10.
Advantage of the present invention is: computation process does not need the feedback parameters of aircraft, and design and comprehensive difficulty are little, and evaluation process is simple, has guaranteed the attitude restrictions performance, has actual application value.
Embodiment
Below the present invention is described in further details.A kind of large aircraft roll channel steering order computing method, based on flight-control computer, air data system and inertial measurement system, is characterized in that, the calculation procedure of steering order is:
1, the control wheel angular displacement sensor is measured the control wheel angular displacement signal and is inputed to flight-control computer;
2, flight-control computer is received the pilot control instruction, forms roll angle speed instruction Pc after the instruction formed filter resolves;
3, calculate pitch angle instruction γ c: one of computing method are following method:
3.1, when air speed is in the normal flight envelope curve, hypervelocity sign overspeed sw=0;
3.1.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γn*γ offset…………………………………………[1]
In formula: K γ nfor scale-up factor, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[2]
In formula: γ ' limfor the pitch angle of a upper computation period, the pitch angle deviation of the first computation period is zero, γ holdpitch angle retention value for aircraft;
3.1.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[3]
In formula: p climangular speed instruction through amplitude limit;
3.1.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[Pcmd-K fade*(γ' lim-γ' int)]……………………………[4]
In formula: K fadefor the desalination factor, γ ' limfor the pitch angle of a upper computation period through amplitude limit, γ ' intintegration pitch angle for a upper computation period; K fade=50～100;
3.1.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxallowable angle of inclination for aircraft;
3.2, when aircraft during in overspeed condition, hypervelocity sign overspeed sw=1;
3.2.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γs*γ offset…………………………………………[5]
In formula: K γ sfor scale-up factor, K γ s=0～10, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[6]
In formula: γ holdit is the pitch angle retention value of aircraft;
γ ' limit is the pitch angle value of a upper computation of Period;
3.2.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[7]
In formula: p climangular speed instruction through amplitude limit;
3.2.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[Pcmd-K fade*(γ' lim-γ' int)]……………………………[8]
In formula: K fadefor the desalination factor;
In formula: γ ' intthe γ of a upper computation of Period intvalue;
3.2.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxit is aircraft allowable angle of inclination value;
4, determine angular speed instruction Wxc:
Wxc=Pcmd*Kp………………………………………[9]
In formula: Kp is scale factor, and value is 0～10.
Principle of work of the present invention is: by formula 8, carry out the integration pitch angle that tilt angle calculation must be expected, calculate the pitch angle deviation according to the pitch angle limits value, reach the purpose at restriction pitch angle, when unclamping control wheel, integration is input as zero, keeps current γ intthereby, reach the purpose of maintenance, avoid the synthtic price index of using the instruction of aircraft actual tilt angles to bring.
Embodiment 1
1, the control wheel angular displacement sensor is measured the control wheel angular displacement signal and is inputed to flight-control computer;
2, flight-control computer is received the pilot control instruction, forms roll angle speed instruction Pc after the instruction formed filter resolves;
3, calculate pitch angle instruction γ c:
3.1, when air speed is in the normal flight envelope curve, hypervelocity sign overspeed sw=0;
3.1.1, calculate feedback angular speed p according to the following formula equ:
3.1.2, calculate angular speed instruction p according to the following formula cmd:
3.1.3, calculate integration pitch angle γ according to the following formula int:
3.1.4, determine pitch angle instruction γ c:
Figure GDA00003169727700061
Figure GDA00003169727700071
Embodiment 2
1, the control wheel angular displacement sensor is measured the control wheel angular displacement signal and is inputed to flight-control computer;
2, flight-control computer is received the pilot control instruction, forms roll angle speed instruction Pc after the instruction formed filter resolves;
3, calculate pitch angle instruction γ c:
3.2, when aircraft during in overspeed condition, hypervelocity sign overspeed sw=1;
3.2.1, calculate feedback angular speed p according to the following formula equ:
3.2.2, calculate angular speed instruction p according to the following formula cmd:
3.2.3, calculate integration pitch angle γ according to the following formula int:
3.2.4, determine pitch angle instruction γ c:
4, determine angular speed instruction Wxc:
Figure GDA00003169727700072
Above-mentioned example meets the requirements fully.

Claims (1)

1. large aircraft roll channel steering order computing method, based on flight-control computer, air data system and inertial measurement system, is characterized in that, the calculation procedure of steering order is:
1.1, the control wheel angular displacement sensor measures the control wheel angular displacement signal and inputs to flight-control computer;
1.2, flight-control computer receives the pilot control instruction, forms roll angle speed instruction Pc after the instruction formed filter resolves;
1.3, calculate pitch angle instruction γ c: one of computing method are following method:
1.3.1, when air speed is in the normal flight envelope curve, hypervelocity sign overspeed sw=0;
1.3.1.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γn*γ offset………………………………………………………[1]
In formula: K γ nfor scale-up factor, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[2]
In formula: γ ' limfor the pitch angle of a upper computation period, the pitch angle deviation of the first computation period is zero, γ holdpitch angle retention value for aircraft;
1.3.1.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[3]
In formula: p climangular speed instruction through amplitude limit;
1.3.1.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[P cmd-K fade*(γ' lim-γ' int)]……………………………[4]
In formula: K fadefor the desalination factor, γ ' limfor the pitch angle of a upper computation period through amplitude limit, γ ' intintegration pitch angle for a upper computation period; K fade=50～100;
1.3.1.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxallowable angle of inclination for aircraft;
1.3.2, when aircraft during in overspeed condition, hypervelocity sign overspeed sw=1;
1.3.2.1, calculate feedback angular speed p according to the following formula equ:
p equ=K γs*γ offset…………………………………………[5]
In formula: K γ sfor scale-up factor, K γ s=0～10, γ offsetpitch angle deviation for current computation period;
When the craft inclination angle, absolute value is less than γ holdwhile spending, make γ offset=0;
When the craft inclination angle, absolute value is more than or equal to γ holdwhile spending,
γ offset=sign(γ' lim)*(|γ' lim|-γ hold)…………………………………[6]
In formula: γ holdit is the pitch angle retention value of aircraft;
γ ' limit is the pitch angle value of a upper computation of Period;
1.3.2.2, calculate angular speed instruction p according to the following formula cmd:
p cmd=p clim-p equ………………………………………………………[7]
In formula: p climangular speed instruction through amplitude limit;
1.3.2.3, calculate integration pitch angle γ according to the following formula int:
γ int=∫[P cmd-K fade*(γ' lim-γ' int)]……………………………[8]
In formula: K fadefor the desalination factor;
In formula: γ ' intthe γ of a upper computation of Period intvalue;
1.3.2.4, determine pitch angle instruction γ c:
Work as γ int<γ maxthe time, γ c=γ int;
Work as γ int>=γ maxthe time, γ c=γ max;
In formula: γ maxit is aircraft allowable angle of inclination value;
1.4, determine angular speed instruction Wxc:
Wxc=Pcmd*Kp………………………………………[9]
In formula: Kp is scale factor, and value is 0～10.
CN 201110294432 2011-09-26 2011-09-26 Large aircraft roll channel control instruction computing method Active CN102346488B (en)

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US9651948B2 (en) * 2015-09-14 2017-05-16 The Boeing Company Roll attitude-dependent roll rate limit
CN106697263B (en) * 2016-12-28 2019-03-01 中国航空工业集团公司西安飞机设计研究所 A kind of rolling aileron reversal control method
CN110161837B (en) * 2018-05-16 2021-12-10 北京机电工程研究所 Triple redundant integral signal equalization method

Citations (4)

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Publication number Priority date Publication date Assignee Title
WO2007042652A1 (en) * 2005-10-11 2007-04-19 Airbus France Method and device for attenuating on an aircraft the effects of a vertical turbulence
CN101804862A (en) * 2010-04-07 2010-08-18 南京航空航天大学 Thrust steering device of unmanned aerial vehicle and control method thereof
CN101807081A (en) * 2010-04-07 2010-08-18 南京航空航天大学 Autonomous navigation guidance method used for pilotless plane
CN201773322U (en) * 2010-08-04 2011-03-23 中国人民解放军第二炮兵工程学院 Dual-beam pseudo-monopulse tracking system for communication in moving

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Publication number Priority date Publication date Assignee Title
FR2924832B1 (en) * 2007-12-11 2010-11-19 Airbus France METHOD AND APPARATUS FOR GENERATING A CONTROLLED SPEED FOR AN AIRCRAFT RUNNING ON THE GROUND

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2007042652A1 (en) * 2005-10-11 2007-04-19 Airbus France Method and device for attenuating on an aircraft the effects of a vertical turbulence
CN101804862A (en) * 2010-04-07 2010-08-18 南京航空航天大学 Thrust steering device of unmanned aerial vehicle and control method thereof
CN101807081A (en) * 2010-04-07 2010-08-18 南京航空航天大学 Autonomous navigation guidance method used for pilotless plane
CN201773322U (en) * 2010-08-04 2011-03-23 中国人民解放军第二炮兵工程学院 Dual-beam pseudo-monopulse tracking system for communication in moving

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