CN103482061A - Harmonic wave recognition correction method of self-adaption helicopter structure response control - Google Patents
Harmonic wave recognition correction method of self-adaption helicopter structure response control Download PDFInfo
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- CN103482061A CN103482061A CN201310406632.3A CN201310406632A CN103482061A CN 103482061 A CN103482061 A CN 103482061A CN 201310406632 A CN201310406632 A CN 201310406632A CN 103482061 A CN103482061 A CN 103482061A
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Abstract
The invention discloses a harmonic wave synchronous-recognition correction method of self-adaption helicopter structure response control, which belongs to the field of active helicopter vibration control. The harmonic wave synchronous-recognition correction method of self-adaption helicopter structure response control is established according to the vibration characteristic that the frequency vibration and the steady-state harmonic wave vibration of high order harmonic waves of a rotor wing are taken as main vibration of a helicopter body, control error response harmonic wave coefficient recognition and control input harmonic wave coefficient synchronous-correction in self-adaption helicopter structure response control are realized, the aim of reducing the helicopter body control point vibration response is achieved, and the self-adaption vibration control capability for tracking the vibration environment variation is provided.
Description
Technical field
The present invention relates to the harmonic synchronous identification revised law of Structural Response for Helicopters adaptive control, belong to the technical field of helicopter vibration ACTIVE CONTROL.
Background technology
Helicopter is in flight course, and rotor blade changes with azimuth the asymmetric aerodynamic environment that the experience cycle changes, need feathering by blade to the asymmetric aerodynamic environment with compensation.The propeller hub six power elements that the Airflow Environment that cycle changes and feathering produce rotor are except containing the 0 rank load such as pulling force and operating torque, also have the humorous wave exciting force relevant with rotating speed to rotor blade sheet number, frequency is that (N is rotor blade sheet number to N Ω, Ω is gyroplane rotate speed) and the vibrational load of high-order harmonic wave by the actuated structure of rotor shaft, make airframe structure all the time in severe exciting environment.High level of vibration has a strong impact on the work efficiency of chaufeur, the reliability of airborne equipment and crew member's traveling comfort etc., and the fuselage shaking level has become an important indicator of evaluation helicopter performance.The high level of vibration, the raising control system performance that reduce helicopter are the Focal point and difficult points in Helicopter Technology field always.Active Control of Structural Response for Helicopters due to its strong adaptability, control effective, energy consumption is low, become effective ways and the important development direction of helicopter vibration ACTIVE CONTROL without advantages such as navigability problems.
It is principal character that the stable state harmonic vibration of rotor by frequency and high-order harmonic wave thereof take in helicopter body vibration, based on discrete Fourier transformation (Discrete Fourier Transform, DFT) and the frequency domain control method of inverse transformation be applied to the helicopter vibration ACTIVE CONTROL, piece treatment characteristic due to DFT, make the control inputs correction asynchronous with the departure signal sampling, but, after the sampling error response signal in the cycle is carried out the DFT conversion by certain hour, extract the multiple amplitude of signal and carry out the Correction and Control input harmonics; In addition because the resolution of DFT is limited, to effectively extract the multiple amplitude of sampled signal, sampling frequency need to be set to rotor usually by the integral multiple of frequency, otherwise just need a large amount of sampled datas could more accurately realize extracting at certain discrete frequency place the multiple amplitude of sampled signal, thereby cause the delay of control inputs correction, reduce controller performance.Except the frequency domain control method, the time domain feedback control strategy as
control, the controlled reset based on internal model principle, and the Linear-Quadratic Problem controlled reset of frequency domain cost function shaping be applied to the research of helicopter vibration ACTIVE CONTROL, but the design of time domain feedback controller needs the dynamic (dynamical) state-space model of helicopter structure usually, the controller exponent number obtained is generally very high, and the vibration suppression performance at discrete point in frequency place and the dynamic property of controlled reset need balance; In addition, the time domain feed forward control method based on auto-adaptive filtering technique Fx-LMS is also for the helicopter vibration ACTIVE CONTROL, but repeatedly occurs the situation of dispersing after the control system convergence in the flight experiment process, and robustness is poor.
Summary of the invention
The invention provides a kind of harmonic synchronous identification revised law of Structural Response for Helicopters adaptive control, realize that the control inputs harmonic wave synchronizes and revise with the departure signal sampling in frequency domain.
The present invention adopts following technical scheme: a kind of harmonic synchronous identification revised law of Structural Response for Helicopters adaptive control, it comprises the steps
Step 1, according to the lifting airscrew feature, extract the rotor excitation frequency, determines and need to control harmonic order number, initialization control inputs harmonic signal;
Step 3, the departure response signal sampled value Correction and Control error responses harmonic coefficient that utilizes step 2 to obtain;
Step 4, utilize the departure response harmonic coefficient Correction and Control input harmonics coefficient obtained in step 3, by revised control inputs harmonic coefficient, determines next time domain control inputs constantly, returns to step 2.
In described step 3, by departure response signal sampled value and this harmonic wave basic function sampled value constantly, adopt Recursive Least Squares at each sampling instant Real time identification departure response harmonic coefficient.
Described step 4 comprises steps A and step B, wherein
Steps A: the departure that obtains according to step 3 response harmonic coefficient, adopt the steepest gradient algorithm to synchronize real-time Correction and Control input harmonics coefficient with sampling frequency,
Step B: the harmonic wave basic function sampled value in the control inputs harmonic coefficient obtained by steps A and this moment obtains the time domain control inputs signal of next sampling instant.
The harmonic wave basic function sampled value that in described step 3, in the identification of departure response harmonic coefficient and step B, the real-time output of control inputs signal needs is produced in real time by the function of the digital signal processor inside of control module.
The present invention has following beneficial effect:
(1) realized that the control inputs harmonic wave revises in real time with departure response signal sample-synchronous in frequency domain, the control inputs correction time lag that the restriction of having avoided sampling frequency need to be set to the forcing frequency integral multiple causes while reaching not for integral multiple;
(2) utilize the harmonic synchronous identification revised law of Structural Response for Helicopters adaptive control to realize the lower harmonic vibration of multistage rotor excitation is controlled, and the method to helicopter vibration, environmental change has stronger Adaptive vibration control ability.
The accompanying drawing explanation
Fig. 1 is the harmonic synchronous identification revised law block diagram of Structural Response for Helicopters adaptive control of the present invention.
Fig. 2 is the departure response of double frequency stable state harmonic vibration.
Fig. 3 is actual measurement single-frequency vibration control error responses on certain helicopter.
Fig. 4 is actual measurement double frequency vibrating departure response on certain helicopter.
The specific embodiment
Below in conjunction with accompanying drawing, technical scheme of the present invention is elaborated.
Please refer to shown in Fig. 1, the harmonic synchronous identification revised law of Structural Response for Helicopters adaptive control of the present invention, comprise the steps:
Step 1, according to the lifting airscrew feature, extract the rotor excitation frequency
ω i , determine and control harmonic order number
r, obtain the control inputs harmonic signal of current time
,
In formula (1)
with
be
lof individual control inputs
ithe cosine of order harmonics and sinusoidal coefficient are in the value of initial time, and 0 means initial time.
Step 3, the departure response signal sampled value Correction and Control error responses harmonic coefficient that utilizes step 2 to obtain, it adopts Recursive Least Squares at each sampling instant Real time identification departure response harmonic coefficient by departure response signal sampled value and this harmonic wave basic function sampled value constantly
Step 4, wherein step 4 can be divided into again two steps, wherein
Steps A: the departure response harmonic coefficient that utilizes identification in step 3
, adopt the steepest gradient adaptive algorithm to synchronize real-time Correction and Control input harmonics coefficient with sampling frequency, update equation is
In formula,
ifor identity matrix, its exponent number with
it is identical,
e i (
k)=[
a 1
i (
k)
b 1
i (
k)
a 2
i (
k)
b 2
i (
k)
a mi (
k)
b mi (
k)]
tfor
mindividual measuring point place respective frequencies
the estimated valve of harmonic coefficient (come from
),
u i (
k)=[
x 1
i (
k)
y 1
i (
k)
x 2
i (
k)
y 2
i (
k)
x li (
k)
y li (
k)]
tfor
lindividual measuring point place respective frequencies
harmonic coefficient,
with
be respectively the weighting matrix of departure response and control inputs,
be
ithe correction step-length of rank control inputs harmonic coefficient,
for frequency
the transfer matrix at place, definition
, wherein,
,
be
lindividual control inputs to the
mthe transfer function at place, individual response controlling point is in frequency
the value at place, Re means to get real part, and Im means to get imaginary part.
Step B: the revised control inputs harmonic coefficient obtained by steps A
and
(
the vector that each control inputs i order harmonics coefficient forms, the third line pair as lower as formula (3)
expression to make an amendment, obtain
and
) and the harmonic wave basic function sampled value in this moment time domain control inputs signal that obtains next sampling instant be
Return to step 2.
The harmonic wave basic function sampled value that in step 3, in the identification of departure response harmonic coefficient and step B, the real-time output of control inputs signal needs is produced in real time by the function of the digital signal processor inside of control module.
Fig. 2 has provided the departure response that under double frequency harmonic excitation, the stable state harmonic vibration is controlled, Fig. 3 and Fig. 4 have provided respectively on certain helicopter the departure response of surveying single-frequency, double frequency vibrating control, the harmonic synchronous identification revised law that shows the Structural Response for Helicopters adaptive control can effectively reduce the structural vibration level, and environmental change has stronger Adaptive vibration control ability to helicopter vibration.
The harmonic wave identification revised law of Structural Response for Helicopters adaptive control of the present invention has realized that the control inputs harmonic wave revises in real time with departure response signal sample-synchronous in frequency domain, the control inputs correction time lag that the restriction of having avoided sampling frequency need to be set to the forcing frequency integral multiple causes while reaching not for integral multiple.Realized the lower harmonic vibration of multistage rotor excitation is controlled, to helicopter vibration, environmental change has stronger Adaptive vibration control ability simultaneously.
The above is only the preferred embodiment of the present invention, it should be pointed out that for those skilled in the art, can also make some improvement under the premise without departing from the principles of the invention, and these improvement also should be considered as protection scope of the present invention.
Claims (4)
1. the harmonic synchronous of Structural Response for Helicopters adaptive control identification revised law, is characterized in that: comprise the steps
Step 1, according to the lifting airscrew feature, extract the rotor excitation frequency, determines and need to control harmonic order number, initialization control inputs harmonic signal;
Step 2, using current time domain control inputs harmonic signal as actuator control inputs drive configuration, gathers the departure response signal at the place, controlling point simultaneously;
Step 3, the departure response signal sampled value Correction and Control error responses harmonic coefficient that utilizes step 2 to obtain;
Step 4, utilize the departure response harmonic coefficient Correction and Control input harmonics coefficient obtained in step 3, by revised control inputs harmonic coefficient, determines next time domain control inputs constantly, returns to step 2.
2. the harmonic synchronous of Structural Response for Helicopters adaptive control as claimed in claim 1 identification revised law, is characterized in that: in described step 3, by departure response signal sampled value and the harmonic wave basic function sampled value in this moment, adopt Recursive Least Squares at each sampling instant Real time identification departure response harmonic coefficient.
3. the harmonic synchronous of Structural Response for Helicopters adaptive control as claimed in claim 1 is identified revised law, and it is characterized in that: described step 4 comprises steps A and step B, wherein
Steps A: the departure that obtains according to step 3 response harmonic coefficient, adopt the steepest gradient algorithm to synchronize real-time Correction and Control input harmonics coefficient with sampling frequency,
Step B: the harmonic wave basic function sampled value in the control inputs harmonic coefficient obtained by steps A and this moment obtains the time domain control inputs signal of next sampling instant.
4. the harmonic synchronous of Structural Response for Helicopters adaptive control as claimed in claim 3 identification revised law is characterized in that: in described step 3 in the identification of departure response harmonic coefficient and step B the harmonic wave basic function sampled value of the real-time output needs of control inputs signal by the function of the digital signal processor inside of control module, produced in real time.
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Cited By (10)
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|---|---|---|---|---|
| CN103955239A (en) * | 2014-05-05 | 2014-07-30 | 南昌华梦达航空科技发展有限公司 | Self-adaption shock resistance control method of unmanned helicopter |
| CN104978450A (en) * | 2015-04-27 | 2015-10-14 | 中国直升机设计研究所 | Position optimal selection method for active vibration control of helicopter |
| CN105843270A (en) * | 2016-03-31 | 2016-08-10 | 南京航空航天大学 | Helicopter multi-frequency vibration active control method |
| CN106945831A (en) * | 2017-03-29 | 2017-07-14 | 南京航空航天大学 | Helicopter body vibrates multiple-harmonic multiple-input and multiple-output adaptive feedforward control method |
| CN108945405A (en) * | 2018-04-23 | 2018-12-07 | 南京航空航天大学 | Helicopter body vibrates adaptive harmonic wave feedforward-sliding formwork and feeds back mixing control method |
| CN109240087A (en) * | 2018-10-23 | 2019-01-18 | 固高科技(深圳)有限公司 | Change the method and system that instruction plan frequency inhibits vibration in real time |
| CN109376449A (en) * | 2018-11-09 | 2019-02-22 | 中国直升机设计研究所 | Helicopter body level of vibration evaluation method and device |
| CN110531624A (en) * | 2019-09-09 | 2019-12-03 | 南京航空航天大学 | A kind of helicopter vibration absorber and its control method |
| CN111624582A (en) * | 2020-07-07 | 2020-09-04 | Oppo广东移动通信有限公司 | Periodic error calibration method, device and system |
| CN113110021A (en) * | 2021-03-17 | 2021-07-13 | 华南理工大学 | Method for identifying servo system and designing controller |
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Cited By (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103955239B (en) * | 2014-05-05 | 2015-12-30 | 南昌华梦达航空科技发展有限公司 | A kind of depopulated helicopter Adaptive vibration-resistant control method |
| CN103955239A (en) * | 2014-05-05 | 2014-07-30 | 南昌华梦达航空科技发展有限公司 | Self-adaption shock resistance control method of unmanned helicopter |
| CN104978450B (en) * | 2015-04-27 | 2019-03-29 | 中国直升机设计研究所 | A kind of helicopter vibration active control position preferred method |
| CN104978450A (en) * | 2015-04-27 | 2015-10-14 | 中国直升机设计研究所 | Position optimal selection method for active vibration control of helicopter |
| CN105843270A (en) * | 2016-03-31 | 2016-08-10 | 南京航空航天大学 | Helicopter multi-frequency vibration active control method |
| CN106945831A (en) * | 2017-03-29 | 2017-07-14 | 南京航空航天大学 | Helicopter body vibrates multiple-harmonic multiple-input and multiple-output adaptive feedforward control method |
| CN106945831B (en) * | 2017-03-29 | 2020-04-24 | 南京航空航天大学 | Helicopter body vibration multi-harmonic multi-input multi-output feedforward self-adaptive control method |
| CN108945405A (en) * | 2018-04-23 | 2018-12-07 | 南京航空航天大学 | Helicopter body vibrates adaptive harmonic wave feedforward-sliding formwork and feeds back mixing control method |
| CN108945405B (en) * | 2018-04-23 | 2021-08-06 | 南京航空航天大学 | Helicopter body vibration self-adaptive harmonic feedforward-sliding mode feedback hybrid control method |
| CN109240087A (en) * | 2018-10-23 | 2019-01-18 | 固高科技(深圳)有限公司 | Change the method and system that instruction plan frequency inhibits vibration in real time |
| CN109240087B (en) * | 2018-10-23 | 2022-03-01 | 固高科技股份有限公司 | Method and system for inhibiting vibration by changing command planning frequency in real time |
| CN109376449A (en) * | 2018-11-09 | 2019-02-22 | 中国直升机设计研究所 | Helicopter body level of vibration evaluation method and device |
| CN109376449B (en) * | 2018-11-09 | 2022-07-01 | 中国直升机设计研究所 | Helicopter body vibration level evaluation method and device |
| CN110531624A (en) * | 2019-09-09 | 2019-12-03 | 南京航空航天大学 | A kind of helicopter vibration absorber and its control method |
| CN111624582A (en) * | 2020-07-07 | 2020-09-04 | Oppo广东移动通信有限公司 | Periodic error calibration method, device and system |
| CN113110021A (en) * | 2021-03-17 | 2021-07-13 | 华南理工大学 | Method for identifying servo system and designing controller |
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