CN108132134A - Aerodynamic derivative discrimination method and system based on wind tunnel free flight test - Google Patents
Aerodynamic derivative discrimination method and system based on wind tunnel free flight test Download PDFInfo
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- CN108132134A CN108132134A CN201711127037.0A CN201711127037A CN108132134A CN 108132134 A CN108132134 A CN 108132134A CN 201711127037 A CN201711127037 A CN 201711127037A CN 108132134 A CN108132134 A CN 108132134A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/08—Aerodynamic models
Abstract
The invention discloses aerodynamic derivative discrimination methods and system based on wind tunnel free flight test, are related to model aircraft identification technique field, can real simulation flight status, obtain accurate aerodynamic derivative.The present invention includes：Wind tunnel free flight test system, the wind tunnel free flight test system carry out degree of freedom release test to model aircraft, obtain test data.It is parsed to obtaining test data, the posture and angular velocity information of aircraft are obtained, establishes corresponding kinetic model, and determine unknown parameter therein, then the output-error method based on maximum-likelihood method is used, gained quantity of state will be calculated, i.e., corresponding attitude angle and angular speed and experiment measure attitude angle and angular speed substitutes into object function and its derivative, pass through iterative calculation, object function is continued to optimize, and its derivative is made to reach minimum value, determines that unknown parameters ' value at this time is exactly identification result.Identification process of the present invention is simple, and model accuracy is high, improves the convenience of model parameter acquisition.
Description
Technical field
The present invention relates to model aircraft identification technique more particularly to the aerodynamic derivative identification side based on wind tunnel free flight test
Method and system.
Background technology
In the aerodynamic derivative Research on Identification of aircraft, currently used test method is that forced oscillation tries in traditional wind-tunnel
It tests and Flight Test.
The experimental condition of Flight Test is very dependent on weather conditions, because the limitation of weather condition, and have in air
There is a large amount of disturbance, the experimental condition is difficult to control, it is also difficult to make the experiment of repeatability, affect the accurate of test data
Property.Moreover, in Flight Test, the identification action that aircraft is made is also relatively fierce, is easy to cause aviation accident, has peace
The not high risk of full property.
In order to avoid the above problem, the prior art forces foam stability high come the high danger coefficient of alternative cost using wind-tunnel
Flight Test.Wind-tunnel is forced in foam stability, in order to simulate the airflow state in reality, the ten of wind-tunnel and support device design
Point complexity, however model aircraft but very simple.Model aircraft and support device by the way of being connected or akinesia or
Model aircraft can only be driven to move by complicated movements design by support device, this so that result of the test is pneumatic by strut
Elasticity effect is larger, and motion design post-processes extremely complex with test data.However, although the design of support device is answered
Miscellaneous cost is higher, but can not simulate aircraft in reality and flexibly be acted in flight, and rudder face can not adjust at any time, need
Wind-tunnel shutdown could be adjusted into rudder face, there is limitation and stiffness.In addition to this, wind-tunnel forces foam stability to need higher
Wind speed could obtain enough signal-to-noise ratio aerodynamic datas with forced oscillation frequency from balance, and forced oscillation uses sine wave conduct
Input, sine wave further comprise the high and low frequency wave band of a part other than the frequency of oscillation that experiment needs.High frequency band meeting
Lead to the movement being coupled with structural modal, low-frequency band can lead to some plunging motions and linear movement, and experiment is caused to input
Not purely, result of the test accuracy is affected.
To sum up, for the aerodynamic derivative Research on Identification of aircraft, lack at present can real simulation flight status, obtain
The experimental rig and method of accurate aerodynamic derivative.
Invention content
The present invention provides aerodynamic derivative discrimination methods and system based on wind tunnel free flight test, can be in wind tunnel test
In really simulate the state of flight of aircraft, more accurately aerodynamic derivative is obtained by experiment.
In order to achieve the above objectives, the present invention adopts the following technical scheme that：
Aerodynamic derivative discrimination method based on wind tunnel free flight test, including：
S1, ground monitoring center are loaded into kinetic model；
S2, wind-tunnel is opened, the ground monitoring center sends closed loop instruction to MEMS controllers, and mould is compared in the contracting of trim aircraft
Type；
S3, the ground monitoring center send pumping signal respectively to the MEMS controllers, and micro- steering gear system drives institute
It states aircraft scale model and carries out motion of rudder, the aircraft scale model makes exciter response；
S4, flight control system acquire and resolve the exciter response, obtain the posture and angular speed of the aircraft scale model
Information, and send back ground monitoring center；
S5, by the pumping signal, i.e., the rudder number of believing one side only is loaded into the kinetic model, assigns in the kinetic model
Unknown parameter initial value, initial value are rule of thumb provided with static force measurement result of the test, resolve the kinetic model, kinetic model
For differential equation of first order, posture and angular speed calculated value is calculated；
S6, the posture measured and the angular speed tested is loaded into majorized function with calculating gained posture and magnitude of angular velocity,
Using the output-error method based on maximum-likelihood method, by iterative calculation, object function is continued to optimize, when the derivative of object function
During for minimum value, the unknown parameter in kinetic model at this time is labeled as identification result.
Further, the kinetic model includes:Longitudinal model, lateral model and course model,
Longitudinal model is：
Wherein, θ is pitch angle, and q is pitch rate, δeFor elevator, MαFor static-stability pitching moment, MqIt is hindered for pitching
Buddhist nun's torque,For pitch control efficiency, unknown parameter includes
Lateral model is：
Wherein, p be rolling angular speed, δaFor aileron, LpFor roll damping,For roll guidance efficiency, unknown parameter packet
It includes
Course model is：
Wherein, ψ is yaw angle, and r is yawrate, NβStatic-stability yawing, NrFor yaw damping torque, unknown ginseng
Number includes Θ=[Nβ, Nr]。
Further, the pumping signal includes：Pitching times pulse excitation signal, rolling small echo pumping signal, yaw times
Pulse excitation signal.J (Θ)=det (R)
Further, the rolling small echo pumping signal is Meyer small echos.Small echo signal has the dual of time domain and frequency domain
Locality：When axis on, pumping signal, which is only concentrated, to be appeared in several seconds；And on frequency domain, signal energy concentrate on it is given in
In frequency range.Since low frequency component is zero, after small echo excitation-off, the rolling angular response of aircraft is finally returned in initial value.
Further, the expression formula of majorized function is：
The expression formula of the object function is：
J (Θ)=det (R) (5)
The derivative of the object function is
The present invention also provides the aerodynamic derivative identification systems based on wind tunnel free flight test, contract including wind-tunnel (1), aircraft
It is freer than model (2), MEMS controllers (3), degree of freedom release device (4) and ground monitoring center (5), the middle installation of wind-tunnel (1)
Release device (4) is spent, aircraft scale model (2) is installed in degree of freedom release device (4), aircraft scale model (2) is beaten using 3D
Print technology completes, and is realized according to similarity criterion similar to geometric similarity, the quality phase Sihe inertia of entity aircraft.Aircraft contracts
Than MEMS controllers (3) embedded in model (2), ground monitoring center (5) are communicated by wireless network and MEMS controllers (3),
MEMS controllers (3) include micro- steering gear system, flight control system and magnetic coder, and micro- steering gear system is received from ground monitoring
The control instruction at center (5) drives motion of rudder；Flight control system solves the posture of aircraft scale model (2) in real time
It calculates and records the information such as the posture of aircraft scale model (2), angular speed, magnetic coder measures and records the actual bit of rudder face in real time
It puts.
Further, degree of freedom release device (4) can realize the Three Degree Of Freedom release of aircraft scale model (2).
The beneficial effects of the invention are as follows：The present invention provides the aerodynamic derivative discrimination method based on wind tunnel free flight test and
System can really simulate the state of flight of aircraft in wind tunnel test, and more accurately aerodynamic derivative is obtained by experiment.Its
In, degree of freedom release device releases rotational freedom, has been effectively isolated the aeroelasticity of support system to aircraft scale model
Influence so that aircraft scale model can be acted flexibly；Wind speed is relatively low needed for experiment, between 15m/s-25m/s, micro- rudder
Machine system can drive aircraft scale model to carry out motion of rudder, realize the automatic trim of aircraft scale model, simulate aircraft
Trim maneuver in flight course, the situation of the data of return closer to aircraft in live flying so that final measuring and calculating
As a result more precisely.
Description of the drawings
It to describe the technical solutions in the embodiments of the present invention more clearly, below will be to needed in the embodiment
Attached drawing is briefly described, it should be apparent that, the accompanying drawings in the following description is only some embodiments of the present invention, for ability
For the those of ordinary skill of domain, without creative efforts, it can also be obtained according to these attached drawings other attached
Figure.
Fig. 1 is the structure diagram of the aerodynamic derivative identification system based on wind tunnel free flight test；
Fig. 2 is MEMS controllers and the information exchange schematic diagram at ground monitoring center；
Fig. 3 is identification flow chart schematic diagram；
Fig. 4 is pitching identification result schematic diagram；
Fig. 5 is rolling identification result schematic diagram；
Fig. 6 is yaw identification result schematic diagram.
Wherein, 1- wind-tunnel, 2- aircraft scale models, 3-MEMS controllers, 4- degree of freedom release devices, in 5- ground monitorings
The heart.
Specific embodiment
For those skilled in the art is made to more fully understand technical scheme of the present invention, With reference to embodiment to this
Invention is described in further detail.
An embodiment of the present invention provides the aerodynamic derivative identification system based on wind tunnel free flight test, as shown in Figure 1, packet
It includes：Including wind-tunnel 1, aircraft scale model 2, MEMS controllers 3, degree of freedom release device 4 and ground monitoring center 5.In wind-tunnel 1
Degree of freedom release device 4 is installed, degree of freedom release device 4 is coupled aircraft scale model 2, and aircraft scale model 2 is embedded in MEMS
Controller 3.MEMS controllers 3 include micro- steering gear system, flight control system and magnetic coder.
As shown in Fig. 2, MEMS controllers 3 and ground monitoring center 1, by wireless network connection, MEMS controllers 3 receive
The command information at ground monitoring center 1 makes aircraft scale model 2 make corresponding response, to the posture of aircraft scale model 2 and
Angular speed record and resolve in real time.Magnetic coder measures and records the physical location of 2 rudder face of aircraft scale model in real time,
These information are sent to ground monitoring center 1, the state of flight of 1 real-time observation airplane scale model 2 of ground monitoring center.
Aerodynamic derivative discrimination method based on wind tunnel free flight test, including：
S1, ground monitoring center be loaded into kinetic model, including longitudinally, laterally with course kinetic model, wherein, longitudinal direction
Model is：
θ is pitch angle, and q is pitch rate, δeFor elevator, MαFor static-stability pitching moment, MqFor damping in pitch power
Square,For pitch control efficiency, unknown parameter includes
Lateral model is：
P be rolling angular speed, δaFor aileron, LpFor roll damping,For roll guidance efficiency, unknown parameter includes
Course model is：
ψ is yaw angle, and r is yawrate, NβStatic-stability yawing, NrFor yaw damping torque, unknown parameter includes
Θ=[Nβ, Nr]。
S2, wind-tunnel is opened, the wind speed of wind-tunnel is 20m/s, and the ground monitoring is centrally generated and is sent to MEMS controllers
Closed loop instructs, trim aircraft scale model.
It is small that S3, the ground monitoring center to the MEMS controllers send pitching times pulse excitation signal, rolling respectively
Wave excitation signal and yaw times pulse excitation signal, micro- steering gear system drive the aircraft scale model to carry out motion of rudder, institute
It states aircraft scale model and makes pitching excitation response, rolling exciter response and yaw exciter response respectively.
S4, flight control system acquire and resolve pitching excitation response, rolling exciter response and yaw exciter response, obtain aircraft
The posture and angular velocity information under each exciter response of scale model, obtain posture and angular velocity measurement value, and send back
Ground monitoring center.
S5, by the pumping signal, i.e., the rudder number of believing one side only is loaded into the kinetic model, assigns in the kinetic model
Unknown parameter initial value, initial value are rule of thumb provided with static force measurement result of the test, for longitudinal model, unknown parameterBy times pulse excitation signal δeIt substitutes into, to unknown parameterIn
It is every rule of thumb to assign initial value with static test result, it thus can solve differential equation of first order(1), the attitude angle θ and rate of pitch q calculated, i.e. y=[θ, q]；For transverse direction
Model, unknown parameterBy meyer small echo pumping signals δaIt substitutes into, to unknown parameterIn
It is every rule of thumb to assign initial value with static test result, it thus can solve differential equation of first order(2), it obtains
The angular velocity in roll p of calculating；For course model, unknown parameter Θ=[Nβ, Nr], by times pulse excitation signal δrIt substitutes into, to not
Know parameter Θ=[Nβ, Nr] in it is every rule of thumb assign initial value with static test result, thus can solve differential equation of first order(3), the attitude angle ψ and yaw rate r calculated；
S6, the posture measured and the angular speed tested is loaded into majorized function with calculating gained posture and magnitude of angular velocity,
Using the output-error method based on maximum-likelihood method, by iterative calculation, object function J (Θ) is continued to optimize, works as object function
Derivative G when being minimum value, it is identification result to determine unknown parameters ' value at this time.
The expression formula of the object function and majorized function is respectively：
J (Θ)=det (R) (4)
The derivative of the object function is
In formula, vector y is the pitch angle and rate of pitch of computation model output, and vector z is the analog value that experiment obtains,
tkFor sampling instant, multiple mobile process have sampled n times altogether.
To longitudinal identification, result of the test is as shown in figure 4, real imaginary curve is respectively that left and right elevon rudder is partially defeated in figure
Enter；Lower two width subgraphs are respectively then pitch angle and pitch rate model response (dotted line) and test value (solid line).It is defeated to become pulse
Most of response phase after entering, model coincide preferably with test value.Prove the validity of this discrimination method.
To lateral identification, result of the test is as shown in figure 5, model response is closer to experimental result.It rubs due to ignoring
Wiping acts on, and there are certain differences for response under small angular velocity in roll.
To course recognize, result of the test as shown in fig. 6, modeling and simulating obtain magnetic heading angle and yaw rate when
Between course and experimental result it is very close
The beneficial effects of the invention are as follows：
Rotational freedom is released, being effectively isolated the aeroelasticity of support system influences；
Only need relatively low the test wind that can recognize dynamic derivative information, experimental cost is low；
Test period is short, and risk is small, is protected from weather influences, and identification process is simple, improves the convenient of model parameter acquisition
Property.
The above description is merely a specific embodiment, but protection scope of the present invention is not limited thereto, any
Those familiar with the art in the technical scope disclosed by the present invention, all should by the change or replacement that can be readily occurred in
It is included within the scope of the present invention.Therefore, protection scope of the present invention should be subject to the protection scope in claims.
Claims (7)
1. the aerodynamic derivative discrimination method based on wind tunnel free flight test, which is characterized in that including：
S1, ground monitoring center are loaded into kinetic model；
S2, wind-tunnel is opened, the ground monitoring is centrally generated and sends closed loop instruction to MEMS controllers, and mould is compared in the contracting of trim aircraft
Type；
S3, the ground monitoring center send pumping signal respectively to the MEMS controllers, and micro- steering gear system driving is described to fly
Machine scale model carries out motion of rudder, and the aircraft scale model makes exciter response；
S4, flight control system acquire and resolve the exciter response, and the aircraft scale model is obtained according to the exciter response
Posture and angular velocity measurement value, and send back ground monitoring center；
S5, the pumping signal is loaded into the kinetic model, resolves the kinetic model, obtain posture and angular speed meter
Calculation value；
S6, by the posture and angular velocity measurement value, be loaded into majorized function with the posture and angular speed calculated value, using based on
The output-error method of maximum-likelihood method by iterative calculation, continues to optimize object function, when the derivative of object function is minimum value
When, the unknown parameter in kinetic model at this time is labeled as identification result.
2. the aerodynamic derivative discrimination method based on wind tunnel free flight test according to claims 1, it is characterised in that, institute
Kinetic model is stated to include:Longitudinal model, lateral model and course model,
Longitudinal model is：
Wherein, θ is pitch angle, and q is pitch rate, δeFor elevator, MαFor static-stability pitching moment, MqFor damping in pitch power
Square,For pitch control efficiency, the unknown parameter includesLateral model is：
Wherein, p be rolling angular speed, δaFor aileron, LpFor roll damping,For roll guidance efficiency, the unknown parameter includes
Course model is：
Wherein, ψ is yaw angle, and r is yawrate, NβStatic-stability yawing, NrFor yaw damping torque, the unknown ginseng
Number includes Θ=[Nβ, Nr]。
3. the aerodynamic derivative discrimination method based on wind tunnel free flight test according to claims 1, which is characterized in that institute
Pumping signal is stated to include：Pitching times pulse excitation signal, rolling small echo pumping signal, yaw times pulse excitation signal.
4. the aerodynamic derivative discrimination method based on wind tunnel free flight test according to claims 3, which is characterized in that institute
Rolling small echo pumping signal is stated as Meyer small echos.
5. the aerodynamic derivative discrimination method based on wind tunnel free flight test according to claims 1, it is characterised in that institute
Stating the expression formula of majorized function is：
The expression formula of the object function is:
J (Θ)=det (R) (5)
The derivative of the object function is
6. the aerodynamic derivative identification system based on wind tunnel free flight test, which is characterized in that compare mould including wind-tunnel (1), aircraft contracting
Type (2), MEMS controllers (3), degree of freedom release device (4) and ground monitoring center (5), the middle installation degree of freedom of wind-tunnel (1) are released
Device (4) is put, degree of freedom release device (4) connects aircraft scale model (2), and embedded MEMS is controlled in aircraft scale model (2)
Device (3), ground monitoring center (5) are communicated by wireless network and MEMS controllers (3), and MEMS controllers (3) include micro- steering engine
System, flight control system and magnetic coder.
7. the aerodynamic derivative identification system based on wind tunnel free flight test according to claims 1, which is characterized in that from
The Three Degree Of Freedom that can realize aircraft scale model (2) by degree release device (4) discharges.
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| CN110940481B (en) * | 2019-11-13 | 2021-09-07 | 中国航天空气动力技术研究院 | Dynamic derivative test model of high-speed wind tunnel of flying wing layout aircraft |
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| CN111220345A (en) * | 2020-01-19 | 2020-06-02 | 南京航空航天大学 | Formation flight aerodynamic interference and attitude control experiment system and experiment method |
| CN112014062A (en) * | 2020-08-19 | 2020-12-01 | 中国航天空气动力技术研究院 | Pose measurement system and measurement method for wind tunnel free flight test model |
| CN112577709A (en) * | 2020-12-29 | 2021-03-30 | 中国航天空气动力技术研究院 | System and method for combining non-contact measurement and free flight in laser wind tunnel |
| CN113252288A (en) * | 2021-07-19 | 2021-08-13 | 中国空气动力研究与发展中心高速空气动力研究所 | Wind tunnel test method based on jet flow and control surface pitch angle closed-loop control |
| CN113252288B (en) * | 2021-07-19 | 2022-02-11 | 中国空气动力研究与发展中心高速空气动力研究所 | Wind tunnel test method based on jet flow and control surface pitch angle closed-loop control |
| CN114001906A (en) * | 2021-11-02 | 2022-02-01 | 中国空气动力研究与发展中心超高速空气动力研究所 | Automatic rudder deflection angle changing device for hypersonic wind tunnel hinge moment measurement test and use method thereof |
| CN115186376A (en) * | 2022-06-21 | 2022-10-14 | 中国航空工业集团公司哈尔滨空气动力研究所 | Small-aspect-ratio flying wing aircraft rolling torque model and parameter identification method thereof |
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| CN116933400B (en) * | 2023-09-13 | 2023-11-21 | 中国航空工业集团公司哈尔滨空气动力研究所 | Method for constructing aerodynamic moment model of coupling uncertainty |
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