CN106773782A - A kind of aeroelastic divergence hybrid modeling method - Google Patents
A kind of aeroelastic divergence hybrid modeling method Download PDFInfo
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- CN106773782A CN106773782A CN201611161879.3A CN201611161879A CN106773782A CN 106773782 A CN106773782 A CN 106773782A CN 201611161879 A CN201611161879 A CN 201611161879A CN 106773782 A CN106773782 A CN 106773782A
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- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
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Abstract
The invention belongs to aeroelastic divergence field, it is related to a kind of aeroelastic divergence modeling method.Structural model is set up by test data, calculating is set up aeroelasticity motion model and Controlling model, objectively reduces the model free degree, improves computational efficiency.
Description
Technical field
The invention belongs to aeroelastic divergence field, it is related to a kind of aeroelastic divergence modeling method.
Background technology
For the aircraft with servo-control system, aeroelastic divergence stability problem is one and inevitably asks
Topic.For aircraft first-fly and the great remodeling of aircraft, it is required for carrying out aeroelastic divergence stability analysis.
At present, aeroelastic divergence stability problem is mainly analyzed by Computer Simulation, and Computer Simulation is modeled
There is larger difference with aircraft actual conditions, thus simulation model is modified by test method mainly, but model
Amendment difficulty is than larger, and correction result also is difficult to fit like a glove.
The content of the invention
The purpose of the present invention:In order to solve, simulation model and actual airplane differ greatly and simulation model is difficult to amendment
Problem, is analyzed by test data, sets up test model, then enters promoting the circulation of qi by the mixed model tested and emulate
Dynamic servo flexibility analysis.
Technical scheme:A kind of aeroelastic divergence hybrid modeling method, it is characterised in that described method bag
Include following steps:
(1) n test point is chosen as the Degree of Structure Freedom, structural model is set up, full machine ground resonance test is carried out, and is measured
Modal frequency ω, Mode Shape Φh, modal damping Chh, modal mass Mhh;
(2) according to the modal mass M for measuringhhModal stiffness K is obtained with modal frequency ωhh;
Khh=ω2Mhh
(3) chain of command mode Φ is set up on its Degree of Structure Freedom according to test modelc;
(4) according to Mode Shape ΦhWith modal mass MhhMass M in the computation structure free degreeg;
(5) according to structural modal ΦhWith chain of command mode ΦcAnd mass MgSolve structural modal with
Coupling mass M between chain of command modehc;
(6) structure motion equation is set up:
ξ in formula, δ represent generalized structure displacement with control deflecting facet respectively;
(7) modal data obtained according to experiment, calculates unsteady using flow field calculation device or other numerical computation methods
Aerodynamic force, and identify broad sense aerodynamic force matrix Qh(s);
Q in formulah=[Qhh Qhc], An=[Ahhn Ahcn] (n=0,1,2), E=[Eh Ec].L is reference length, and V is air-flow
Speed,sIt is Laplace variable;
(8) using the broad sense aerodynamic force matrix Q of fittinghS () obtains broad sense aerodynamic force fa;
Q in formula∞Represent to flow pressure, q is generalized displacement, q=[ξ δ]T, including generalized structure displacement ξ and control surface deflection
δ;
(9) aerodynamic force state variable is taken:
It is transformed into time domain space:
Time domain broad sense aerodynamic force can be write as:
(10) the aeroelasticity equation of motion is set up:
(11) aeroelasticity equation is write as state space form:
In formula
(12) frequency response function of steering wheel is measured according to experiment, steering wheel state equation is obtained:
(13) due to xact=uae, the state equation of controlled device (plant) can represent with following formula:
In formula
Cp=[Cae Dae], Dp=0;
(14) consider control system state equation, can be obtained by simulation model:
(15) open-loop transfer function of controlled device and control system is set up:
In formulaCo=[DcCp Cc], Do=DcDp;
(16) state space equation is converted into frequency response function:
H (s)=Co(sI-Ao)-1Bo+Do
Bode figures and Nyquist figures are drawn, stability analysis can be carried out and analyzed with stability margin.
Beneficial effects of the present invention:Structural model is set up by test data, aeroelasticity motion mould is set up by calculating
Type and Controlling model, objectively reduce the model free degree, improve computational efficiency.
Specific embodiment
(1) n test point is chosen as the Degree of Structure Freedom, structural model is set up, full machine ground resonance test is carried out, and is measured
Modal frequency ω, Mode Shape Φh, modal damping Chh, modal mass Mhh。
(2) according to the modal mass M for measuringhhModal stiffness K is obtained with modal frequency ωhh。
Khh=ω2Mhh
(3) chain of command mode Φ is set up on its Degree of Structure Freedom according to test modelc。
(4) according to mass MgWith modal matrix ΦhWith modal mass matrix MhhBetween relation:
Can obtain:
Thus it is possible to obtain the coupling mass matrix between following structural modal and chain of command mode:
(5) modal data obtained according to experiment, solves broad sense aerodynamic force, is fitted broad sense aerodynamic force matrix:
Qh(p)=A0+A1p+A2p2+D(Ip-R)-1Ep
Q in formulah=[Qhh Qhc], An=[Ahhn Ahcn] (n=0,1,2), E=[Eh Ec].L is reference length, and V is air-flow
Speed, dimensionless Laplace variable p=sL/V, s are Laplace variable.Therefore, broad sense aerodynamic force matrix can be write as:
So, broad sense aerodynamic force can be write as:
Q in formula∞Represent to flow pressure, q is generalized displacement, including generalized structure displacement ξ and control surface deflection δ, q=[ξ δ
]T。
Take aerodynamic force state variable:
It is transformed into time domain space:
So, aerodynamic force can be write as:
(6) the aeroelastic divergence equation of motion can be write as:
Then, aeroelastic divergence equation can be write as state space form:
In formula
(7) frequency response function of steering wheel is measured according to experiment, steering wheel state equation is obtained:
(8) due to xact=uae, the state equation of controlled device (plant) can represent with following formula:
In formula
Cp=[Cae Dae], Dp=0.
(9) consider control system state equation, can be obtained being measured by experiment by simulation model:
(10) open-loop transfer function of controlled device and control system is set up
In formulaCo=[DcCp Cc], Do=DcDp
(11) state space equation is converted into frequency response function:
H (s)=Co(sI-Ao)-1Bo+Do
Bode figures are drawn with Nyquist figures.Stability analysis can be carried out to be analyzed with stability margin.
Claims (1)
1. a kind of aeroelastic divergence hybrid modeling method, it is characterised in that described method comprises the following steps:
(1) n test point is chosen as the Degree of Structure Freedom, structural model is set up, and carries out full machine ground resonance test, measurement mode
Frequencies omega, Mode Shape Φh, modal damping Chh, modal mass Mhh;
(2) according to the modal mass M for measuringhhModal stiffness K is obtained with modal frequency ωhh;
Khh=ω2Mhh
(3) chain of command mode Φ is set up on its Degree of Structure Freedom according to test modelc;
(4) according to Mode Shape ΦhWith modal mass MhhMass M in the computation structure free degreeg;
(5) according to structural modal ΦhWith chain of command mode ΦcAnd mass MgSolve between structural modal and chain of command mode
Coupling mass Mhc;
(6) structure motion equation is set up:
In formulaRepresent generalized structure displacement with control deflecting facet respectively;
(7) modal data obtained according to experiment, calculates unsteady pneumatic using flow field calculation device or other numerical computation methods
Power, and identify broad sense aerodynamic force matrix Qh(s);
Q in formulah=[Qhh Qhc], An=[Ahhn Ahcn] (n=0,1,2), E=[Eh Ec].L is reference length, and V is gas velocity
Degree, s is Laplace variable;
(8) using the broad sense aerodynamic force matrix Q of fittinghS () obtains broad sense aerodynamic force fa;
Q in formula∞Represent to flow pressure, q is generalized displacement, q=[ξ δ]T, including generalized structure displacement ξ and control surface deflection δ;
(9) aerodynamic force state variable is taken:
It is transformed into time domain space:
Time domain broad sense aerodynamic force can be write as:
(10) the aeroelasticity equation of motion is set up:
(11) aeroelasticity equation is write as state space form:
In formula
(12) frequency response function of steering wheel is measured according to experiment, steering wheel state equation is obtained:
(13) due to xact=uae, the state equation of controlled device (plant) can represent with following formula:
In formula
Cp=[Cae Dae], Dp=0;
(14) consider control system state equation, can be obtained by simulation model:
(15) open-loop transfer function of controlled device and control system is set up:
In formulaCo=[DcCp Cc], Do=DcDp;
(16) state space equation is converted into frequency response function:
H (s)=Co(sI-Ao)-1Bo+Do
Bode figures and Nyquist figures are drawn, stability analysis can be carried out and analyzed with stability margin.
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Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN108256264A (en) * | 2018-02-08 | 2018-07-06 | 北京航空航天大学 | A kind of aeroelastic divergence stability prediction method based on ground frequency response test |
CN109856989A (en) * | 2018-11-26 | 2019-06-07 | 广东工业大学 | A kind of pneumatic force servo system emulation modelling method |
CN110287505A (en) * | 2019-03-20 | 2019-09-27 | 北京机电工程研究所 | Stability of aircraft analysis method |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN102419238A (en) * | 2011-08-15 | 2012-04-18 | 中国航空工业集团公司西安飞机设计研究所 | Device for measuring mass moment of inertia of flutter model |
CN102566440A (en) * | 2011-12-29 | 2012-07-11 | 成都飞机工业(集团)有限责任公司 | Testing method for unmanned plane flight-control structure modal coupling |
US8393206B1 (en) * | 2010-02-09 | 2013-03-12 | Ping-Chih Chen | Dry wind tunnel system |
CN103530485A (en) * | 2013-11-05 | 2014-01-22 | 中国航空工业集团公司西安飞机设计研究所 | Full-aircraft beam frame type reduction stiffness combination modeling method |
CN103558842A (en) * | 2013-11-05 | 2014-02-05 | 中国航空工业集团公司西安飞机设计研究所 | Virtual open-loop frequency scanning method for airplane ground servo elasticity test |
CN105843073A (en) * | 2016-03-23 | 2016-08-10 | 北京航空航天大学 | Method for analyzing wing structure aero-elasticity stability based on aerodynamic force uncertain order reduction |
CN105929692A (en) * | 2016-05-16 | 2016-09-07 | 中国航空工业集团公司西安飞机设计研究所 | Elastic flying wing layout airplane gust alleviating controller parameter optimization method |
-
2016
- 2016-12-15 CN CN201611161879.3A patent/CN106773782B/en active Active
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8393206B1 (en) * | 2010-02-09 | 2013-03-12 | Ping-Chih Chen | Dry wind tunnel system |
CN102419238A (en) * | 2011-08-15 | 2012-04-18 | 中国航空工业集团公司西安飞机设计研究所 | Device for measuring mass moment of inertia of flutter model |
CN102566440A (en) * | 2011-12-29 | 2012-07-11 | 成都飞机工业(集团)有限责任公司 | Testing method for unmanned plane flight-control structure modal coupling |
CN103530485A (en) * | 2013-11-05 | 2014-01-22 | 中国航空工业集团公司西安飞机设计研究所 | Full-aircraft beam frame type reduction stiffness combination modeling method |
CN103558842A (en) * | 2013-11-05 | 2014-02-05 | 中国航空工业集团公司西安飞机设计研究所 | Virtual open-loop frequency scanning method for airplane ground servo elasticity test |
CN105843073A (en) * | 2016-03-23 | 2016-08-10 | 北京航空航天大学 | Method for analyzing wing structure aero-elasticity stability based on aerodynamic force uncertain order reduction |
CN105929692A (en) * | 2016-05-16 | 2016-09-07 | 中国航空工业集团公司西安飞机设计研究所 | Elastic flying wing layout airplane gust alleviating controller parameter optimization method |
Non-Patent Citations (2)
Title |
---|
乔洋 等: "弹性飞行器气动伺服弹性稳定性分析与控制", 《应用力学学报》 * |
惠俊鹏 等: "考虑参数摄动的飞行器气动伺服弹性鲁棒稳定性分析研究", 《导弹与航天运载技术》 * |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN108256264A (en) * | 2018-02-08 | 2018-07-06 | 北京航空航天大学 | A kind of aeroelastic divergence stability prediction method based on ground frequency response test |
CN109856989A (en) * | 2018-11-26 | 2019-06-07 | 广东工业大学 | A kind of pneumatic force servo system emulation modelling method |
CN110287505A (en) * | 2019-03-20 | 2019-09-27 | 北京机电工程研究所 | Stability of aircraft analysis method |
CN110287505B (en) * | 2019-03-20 | 2020-12-25 | 北京机电工程研究所 | Aircraft stability analysis method |
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