CN104483090B - Dynamitic control and measurement method - Google Patents
Dynamitic control and measurement method Download PDFInfo
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- CN104483090B CN104483090B CN201410808133.1A CN201410808133A CN104483090B CN 104483090 B CN104483090 B CN 104483090B CN 201410808133 A CN201410808133 A CN 201410808133A CN 104483090 B CN104483090 B CN 104483090B
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Abstract
The invention provides a dynamitic control and measurement method which is used for a wind tunnel test. An attack angle and a control plane deflection angle of an airplane model supported in a wind tunnel by a controllable support system are measured from flat flight to pull-up to overload G. According to the dynamitic control and measurement method, a real-time dynamitic control and measurement manner is adopted, the angle of attack and the control plane deflection angle of the airplane model are dynamically controlled and adjusted in real time by means of the support system and a steering engine, so that the wind tunnel blowing test can be completed once from beginning to end with the method, time for the wind tunnel test is far shorter than that of a static test in the prior art, the test cost is greatly saved, the available time of wind tunnel resources is greatly prolonged, and the equipment utilization efficiency is improved.
Description
Technical field
The present invention relates to a kind of aviation aerodynamic measurement method, especially one kind are used for model aircraft wind tunnel test
Cheng Zhong, dynamic control and the angle of attack of survey aircraft model and the method for rudder face drift angle.
Background technology
Wind tunnel test is that by model aircraft or its part, such as fuselage, wing etc. are fixed on according to aerodynamic principle
In wind-tunnel, flow through model aircraft or its part by applying artificial airflow, various complicated state of flights in the air are simulated with this, obtains
Take test data.Wind-tunnel is by the aerodynamic studies testing equipment most basic with aircraft development, each type aircraft
Development be required for carrying out substantial amounts of test in wind-tunnel.The main purpose of wind tunnel test is intended to obtain the various skies of model aircraft
The Changing Pattern of aerodynamic parameter.Evaluate the flying quality of each aircraft, except such as speed, highly, aircraft weight and starting
Outside the key elements such as machine thrust, one of most important standard is the aerodynamic quality of aircraft.
The Chinese patent application that the relative theory of model aircraft wind tunnel test and process were submitted on October 12nd, 2012
Clearly described in 201210387483.6, the Chinese patent application in addition submitted to for 2013 01 month 14
Also provide dependency structure and the operation principle of the support system that can be used for wind tunnel test in 201310011601.8, quote herein
As reference, understood in order to those skilled in the art.
During existing wind tunnel test, for the dynamic parameter of survey aircraft certain maneuver of model, such as aircraft mould
The angle of attack of type and rudder face drift angle, often carry out wind tunnel test using static mode, then using the various aerodynamic of measurement
Parameter calculates and obtains the angle of attack of model aircraft and rudder face drift angle, then pass through wind-tunnel flyoff checking calculate the angle of attack obtaining with
Rudder face drift angle whether there is deviation with the actual angle of attack and rudder face drift angle, is then added by way of linear interpolation if there is deviation
To revise, optimal state is to carry out wind-tunnel flyoff again to the corrected angle of attack and rudder face drift angle to verify, examination
Test that process troublesome calculation workload is very big, and the result obtaining be also approximate interpolation, not very accurate although repeatedly drying
Checking can gradually approaching to reality value, but whole process is also static.For example, in order to survey aircraft flies to pull-up and reaches from flat
To the maneuver setting overload, the angle of attack of real aircraft is dynamic change, and rudder face drift angle is also dynamic change.And show
There is wind tunnel test to be that model aircraft is fixed in wind-tunnel, model aircraft is fixed an angle of attack, one angle of rudder face fixed deflection
Degree blows, and obtains the overload of aircraft according to the aerodynamic force of model aircraft that measurement obtains, if the overload that obtains of measurement with
Want the overload reaching to there is deviation, then row interpolation is entered according to the size of deviation, model aircraft is repeatedly adjusted according to interpolation solid
It is scheduled on the fixing degree of bias of the angle of attack in wind-tunnel and rudder face, until the overload that measurement obtains is less than error with wanting the overload reaching
Till scope.The wind tunnel test cycle of this static measurement method of prior art is very very long, needs repeatedly to dry, precision test
Cost substantially unacceptable, often can only test a limited number of time after pass through interpolation calculation and obtain one approximately
Value.
Content of the invention
The technical problem to be solved in the present invention is to provide a kind of dynamic control and measuring method, so that above institute is reduced or avoided
The problem mentioned.
For solving above-mentioned technical problem, the present invention proposes a kind of dynamic control and measuring method, in wind tunnel test,
Measurement flies to described aircraft when pull-up reaches overload g using the model aircraft that controllable support system is supported in wind-tunnel from flat
The angle of attack of model and rudder face drift angle, wherein, are provided with the control surface deflection controlling described model aircraft inside described model aircraft
Steering wheel, described steering wheel has a minimum angle adjustable m, and methods described comprises the steps:
Step 1: described model aircraft is supported in described wind-tunnel by described support system, using described support system
Adjust described model aircraft and be in flat winged state, be zero using rudder face drift angle described in described servos control, start wind-tunnel and blown
Wind.
Step 2: increase by one step-length angle a of deflection on the basis of original using rudder face described in described servos control, utilize
Described support system adjusts the angle of attack of described model aircraft, and measurement makes the pitching of described model aircraft by adjusting the described angle of attack
Moment is overload g when zero1.
Step 3: if g1Less than g, then repeat n step 2, until gnMore than or equal to g；If gnEqual to g, then measurement obtains
Rudder face drift angle is n*a, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；If gnMore than g, then enter next step；
Wherein n is positive integer.
Step 4: reduce the angle of deflection a/2 on the basis of original using rudder face described in described servos control, using described
Support system controls the angle of attack of described model aircraft to be adjusted, and measurement makes described model aircraft by adjusting the described angle of attack
Pitching moment is overload g when zeron+1.
Step 5: if gn+1More than g, then Repeated m time step 4, until gn+mLess than or equal to g；If gn+mEqual to g, then measurement obtains
The described rudder face drift angle obtaining is n*a-m*a/2, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein m is just
Integer.
Step 6: if gn+mLess than g, then increase the rudder face described in angular deflection of deflection a/4, weight on the basis of step 5 again
Multiple k step 2 arrives step 3 until gn+m+kMore than or equal to g, if gn+m+kEqual to g, then the described rudder face drift angle measuring acquisition is n*
A-m*a/2+k*a/4, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein k is positive integer.
Step 7: if gn+m+kMore than g, then reduce the rudder face described in angular deflection of deflection a/8, weight on the basis of step 6 again
Multiple p step 4 arrives step 5 until gn+m+k+pLess than or equal to g, if gn+m+k+pEqual to g, then the described rudder face drift angle measuring acquisition is
N*a-m*a/2+k*a/4-p*a/8, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein p is positive integer.
Step 8: as long as the overload that measurement obtains is not equal to g, then the circulation of repeat step 6 to 7, control and increase or reduce
The angle of control surface deflection halves every time on the basis of original, until the final deflection angle of described rudder face is less than described steering wheel
Minimum angle adjustable m, that is, described steering wheel adjusts described control surface deflection angle and is less than the scope of its mechanical adjustable and cannot continue
Till adjusting described control surface deflection, now the described rudder face drift angle obtaining with measurement is controlled to can be considered required according to abovementioned steps
The overload of measurement is the described angle of attack and described rudder face drift angle during g.
Preferably, shown support system real-time dynamicly controls the angle of attack of described model aircraft using electronic joint.
Preferably, described steering wheel real-time dynamicly controls described rudder face to enter horizontal deflection during described wind tunnel test.
Preferably, described m is 0.01 degree, and described a is 0.1 degree.
In above-mentioned dynamic control provided by the present invention and measuring method, the complete mould of the angle of attack of model aircraft and rudder face drift angle
Intend real flight progress, belonged to the method real-time dynamicly controlling with measurement, methods described is from startup wind-tunnel flyoff
Start to finish can disposably complete, and the time of wind tunnel test is far less than the time of the static test of prior art, significantly
Save experimentation cost, the pot life of wind-tunnel resource is greatly improved, and improves the utilization ratio of equipment.
Brief description
The following drawings is only intended to, in doing schematic illustration and explanation to the present invention, not delimit the scope of the invention.Wherein,
Fig. 1 is shown that dynamic control and survey aircraft mould in the wind tunnel test according to a specific embodiment of the present invention
The model demonstration schematic diagram of the angle of attack of type and rudder face drift angle.
Specific embodiment
In order to be more clearly understood to the technical characteristic of the present invention, purpose and effect, now comparison brief description this
Bright specific embodiment.Wherein, identical part adopts identical label.
Fig. 1 is shown that dynamic control and survey aircraft mould in the wind tunnel test according to a specific embodiment of the present invention
The model demonstration schematic diagram of the angle of attack of type and rudder face drift angle.As illustrated, the model aircraft 3 of the present invention utilizes controllable support
System 1 is supported in wind-tunnel 2, need measurement be model aircraft 3 from flat fly to overload g that pull-up reaches setting when, aircraft mould
The angle of attack of type 3 and rudder face drift angle.
Due to present invention needs measurement is the angle of attack during flat winged pull-up for the model aircraft 3 and rudder face drift angle, then
Described rudder face 31 can be the elevator of model aircraft, flying tail etc., or other is during operating aircraft model pull-up
Need the auxiliary lifting device used, such as front wing etc..It will be appreciated by those skilled in the art that the present invention is only schematically
Using specific embodiment, one general test method is described, those skilled in the art can expand according to the scheme that the present invention introduces
Open up in the measurement process of yaw angle, roll angle and other any rudder face drift angle.
Support system 1 shown in Fig. 1 can using background section refer to 201310011601.8 in similar
Support system, described support system can provide the angle of attack to model aircraft, yaw angle, roll angle flexibly to be controlled, Er Qieke
Real-time dynamicly to control the angle of attack of model aircraft 3 using electronic joint etc., in addition, being also equipped with inside described model aircraft 3
Control the steering wheel 4 of rudder face 31 deflection of model aircraft 3, for real-time dynamicly controlling rudder face 31 to carry out during wind tunnel test
Deflection.Support system 1 shown in Fig. 1 of the present invention and steering wheel 4, for realizing dynamic control of the present invention and survey aircraft mould
The angle of attack of type and rudder face drift angle provide preliminary experimental basis.
Just it has been observed that being a difference in that with prior art maximum, the present invention is to provide a kind of dynamic control and measurement
Method, methods described is disposably to complete it is not necessary to wind-tunnel shutdown is adjusted repeatedly from starting wind-tunnel flyoff start to finish
The position of whole model aircraft, and the angle of attack of model aircraft 3 and rudder face drift angle are also to simulate real flight progress to carry out in real time
Dynamically control.
Specifically, dynamic control and the angle of attack of survey aircraft model and the method for rudder face drift angle in the wind tunnel test of the present invention
Comprise the steps:
Step 1: model aircraft 3 is supported in wind-tunnel 2 by support system 1, adjusts model aircraft using support system 1
3 are in flat winged state, and controlling rudder face drift angle using steering wheel 4 is zero, starts wind-tunnel 2 and blows.Certainly, those skilled in the art
It should be appreciated that step 1 is in order to full text describes a reduced condition of easy setting, actually it is contemplated that the present invention may be use with from flying
Any one state of angle of attack of machine model 3 and rudder face drift angle start wind tunnel test, and this test objective to be seen is in order to which kind of obtains
Parameter under state of flight.
Step 2: control rudder face 31 to increase by one step-length angle a of deflection, wherein said rudder on the basis of original using steering wheel 4
Machine 4 has a minimum angle adjustable m, that is, the minimum angles value that steering wheel 4 drives rudder face 31 deflection for machinery is m.Example
As the precision according to Machine Design and cost consideration, this minimum angle adjustable m may be 0.01 degree, that is, adjusts rudder every time
Face drift angle minimum can only achieve 0.01 degree of precision, less angle is exactly that frame for movement cannot be accomplished.In addition, step-length angle
Degree a can be set as a numerical value according to the experience of this area, and this numerical value is preferably able to need measurement to obtain much smaller than final
Transship for rudder face drift angle during g, and this numerical value is also much larger than minimum angle adjustable m, for example, if final need measurement to obtain
Transship and may be 10 degree about for rudder face drift angle during g, then step-length angle a can select 0.1 degree.Certainly, step-length angle a is permissible
It is controlled adjusting according to practical situation, to meet the balance of experimental precision and economy.
After steering wheel 4 controls rudder face 31 to deflect step-length angle a, the aerodynamic force distribution of whole model aircraft 3 there occurs
Change, with regard to amount of unbalance, this is consistent with real state of flight to the pitching moment leading to model aircraft 3, therefore truly
Aircraft can be pulled upwardly.Likewise, in order to simulate real state of flight, it is possible to use support system 1 controls model aircraft 3
The angle of attack be adjusted, when the pitching moment of model aircraft 3 is zero, again reach poised state.Now namely control
After rudder face 31 has deflected step-length angle a, the statokinetic that model aircraft should be at, certainly, this attitude is got off the plane model 3
Aerodynamic force size can also measure by balance, obtain this attitude and get off the plane model 3 through calculating also just can measure
Overload g1.
Due to selecting before when step-length angle a it has to be considered that its numerical value is much smaller than final deflection angle, therefore warp
Cross once obtained overload g of deflection1Certainly it is less than g's it is therefore desirable to accumulative be repeated several times the same step-length angle of deflection
A, the final deflection angle of Step wise approximation.Namely back to back step is:
Step 3: if g1Less than g, then repeat n step 2, until gnMore than or equal to g；If gnEqual to g, then measurement obtains
Rudder face drift angle is n*a, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；If gnMore than g, then enter next step；
Wherein n is positive integer.
In this step, through accumulative repeat for n time to deflect same step-length angle a and then adjustment angle of attack rebalancing moment it
Afterwards, overload g that measurement obtainsnIt is possible to exactly equal to (certainly this probability is very little) of g, in this case just not
Need further to operate, n time the deflection angle of adjustment i.e. n times of step-length angle a are exactly final deflection angle altogether
, the angle of attack of the model aircraft of the as required measurement of the angle of attack now.Certainly, most possible situation is that rudder face 31 adjusted
Head, that is, the angle that adjusts is too big, and (this is that the measuring method of the present invention wishes the situation that occurs in fact, shows finally to deflect
Angle is certainly between (n-1) * a and n*a) so that the g of measurement acquisitionnMore than overload g setting, then remaining thing is just
It is the deflection back controlling rudder face, gradually approach final rudder face drift angle.
Step 4: control rudder face 31 to reduce the angle of deflection a/2 on the basis of original using steering wheel 4, support system using described
The angle of attack of system 1 control model aircraft 3 is adjusted, and measurement makes the pitching power of described model aircraft 3 by adjusting the described angle of attack
Square is overload g when zeron+1.
This step back adjusts the deflection of rudder face, and the step-length of selection is a/2, is the mistake that same measurement of comparison obtains afterwards
Carry the size with g.
Step 5: if gn+1More than g, then Repeated m time step 4, until gn+mLess than or equal to g；If gn+mEqual to g, then measurement obtains
The described rudder face drift angle obtaining is n*a-m*a/2, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein m is just
Integer.
In this step, repeat for m time back to deflect same step-length angle a/2 and then adjust angle of attack rebalancing through accumulative
After moment, overload g that measurement obtainsn+mEqually it is possible to exactly equal to (this probability is also very little) of g, this
In the case of again without further operating, final deflection angle is exactly n*a-m*a/2, and the angle of attack now is required
The angle of attack of the model aircraft of measurement.Certainly, it is desirable also to the situation occurring is to have adjusted again anti-excessive so that measurement obtains
gn+mLess than overload g setting, then remaining thing is exactly and in turn further toward the big deflection adjusting rudder face, gradually forces
Closely final rudder face drift angle.
Actually in most cases may only need to adjustment and will adjust anti-excessive several times, perhaps once just recall to
, that is, m be possible to equal to 1 it is also possible to be equal to 10, depending on the numerical values recited of step-length angle a, therefore step-length angle a
With the number of times of adjustment and the time correlation of wind tunnel test.Certainly, because measurement, calculating, adjustment are all by background computer control
System is processed, and the time hinge structure repeatedly approaching operation is almost flashy thing, under prior art conditions, no matter
The selection of step-length angle a is much, and the time of wind tunnel test has all been far less than the time of the static test of prior art, substantially
On only need in 1 minute just permissible, greatly save experimentation cost, reduce energy loss, the pot life of wind-tunnel resource
Greatly improved, improved the utilization ratio of equipment.
Step 6: if gn+mLess than g, then increase the angular deflection rudder face 31 of deflection a/4 on the basis of step 5 again, repeat k
Secondary step 2 arrives step 3 until gn+m+kMore than or equal to g, if gn+m+kEqual to g, then the described rudder face drift angle measuring acquisition is n*a-m*
A/2+k*a/4, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein k is positive integer.
This step is actually step 2 and 3 and repeats to approach the mistake of final rudder face drift angle toward the deflection of big iteration adjustment rudder face
Journey, simply the step-length of iteration be reduced to the 1/4 of step-length angle a.
Step 7: if gn+m+kMore than g, then reduce the rudder face 31 described in angular deflection of deflection a/8 on the basis of step 6 again,
Repeat p step 4 and arrive step 5 until gn+m+k+pLess than or equal to g, if gn+m+k+pEqual to g, then measure the described rudder face drift angle of acquisition
For n*a-m*a/2+k*a/4-p*a/8, the angle of attack of the model aircraft of the as required measurement of the angle of attack now；Wherein p is positive integer.
Likewise, this step is actually step 4 and 5 repeating to approach final deflection angle toward the deflection of little iteration adjustment rudder face
Degree process, simply the step-length of iteration be reduced to the 1/8 of step-length angle a.
Step 8: as long as the overload that measurement obtains is not equal to g, then the circulation of repeat step 6 to 7, control and increase or reduce
The angle of control surface deflection halves every time on the basis of original, until the final deflection angle of described rudder face 31 is less than described steering wheel
4 minimum angle adjustable m, that is, described steering wheel 4 adjusts described rudder face 31 deflection angle less than the scope of its mechanical adjustable no
Till method continues adjustment described rudder face 31 deflection, now can be considered required survey according to the rudder face drift angle that abovementioned steps measurement obtains
The overload of amount is the described angle of attack and described rudder face drift angle during g.
This step is that the process of repeatedly approaching adjustment is explained further, and actually carries out with regard to it in step before
Detailed explanation, the situation herein summarized is, in most cases can not control so that measuring the overload obtaining just etc.
In g, then as long as being not equal to g it is necessary to ceaselessly repeatedly approach, being finally stopped the condition approached is, when every time in original base
Deflection angle is halved on plinth, always reaches the mechanical adjustable limit of steering wheel 4, continue below the mechanical adjustable limit of steering wheel 4
Adjustment cannot be carried out, simultaneously economically also unnecessary, has now controlled the angle of attack obtaining with measurement and rudder face drift angle
Can be regarded as the final angle of the required measurement of infinite approach, you can be considered as final result.
In sum, in above-mentioned dynamic control provided by the present invention and measuring method, the angle of attack of model aircraft and rudder face
Drift angle simulates real flight progress completely, belongs to the method real-time dynamicly controlling with measurement, and methods described is from startup wind
Hole blowing test start to finish can disposably complete, and the time of wind tunnel test is far less than the static test of prior art
Time, greatly save experimentation cost, the pot life of wind-tunnel resource is greatly improved, improve the utilization effect of equipment
Rate.
It will be appreciated by those skilled in the art that although the present invention is to be described according to the mode of multiple embodiments,
It is that not each embodiment only comprises an independent technical scheme.For the sake of in description, so narration is just for the sake of understanding,
Description should be understood by those skilled in the art as an entirety, and by involved technical scheme in each embodiment
Regard as and can be mutually combined into the mode of different embodiments to understand protection scope of the present invention.
The foregoing is only the schematic specific embodiment of the present invention, be not limited to the scope of the present invention.Any
Those skilled in the art, the equivalent variations made on the premise of the design without departing from the present invention and principle, modification and combination,
The scope of protection of the invention all should be belonged to.
Claims (4)
1. a kind of dynamic control and measuring method, for, in wind tunnel test, measurement is supported on using controllable support system (1)
Model aircraft (3) in wind-tunnel (2) flies to the angle of attack of described model aircraft (3) and rudder face drift angle when pull-up reaches overload g from flat,
Wherein, the steering wheel (4) controlling the rudder face (31) of described model aircraft (3) to deflect is installed inside described model aircraft (3), described
Steering wheel (4) has a minimum angle adjustable m, and methods described comprises the steps:
Step 1: described model aircraft (3) is supported in described wind-tunnel (2) by described support system (1), using described
Support system (1) adjusts described model aircraft (3) and is in flat winged state, and controlling described rudder face drift angle using described steering wheel (4) is zero,
Start wind-tunnel (2) to blow；
Step 2: control described rudder face (31) to increase by one step-length angle a of deflection, profit on the basis of original using described steering wheel (4)
Adjust the angle of attack of described model aircraft (3) with described support system (1), measurement makes described aircraft mould by adjusting the described angle of attack
The pitching moment of type (3) is overload g when zero1；
Step 3: if g1Less than g, then repeat n step 2, until gnMore than or equal to g；If gnEqual to g, then measure the rudder face of acquisition
Drift angle is n*a, the angle of attack of the model aircraft (3) of the as required measurement of the angle of attack now；If gnMore than g, then enter next step；Its
Middle n is positive integer；
Step 4: control described rudder face (31) to reduce the angle of deflection a/2 on the basis of original using described steering wheel (4), using institute
Stating support system (1) controls the angle of attack of described model aircraft (3) to be adjusted, and measurement makes described flying by adjusting the described angle of attack
The pitching moment of machine model (3) is overload g when zeron+1；
Step 5: if gn+1More than g, then Repeated m time step 4, until gn+mLess than or equal to g；If gn+mEqual to g, then measurement obtains
Described rudder face drift angle is n*a-m*a/2, the angle of attack of the model aircraft (3) of the as required measurement of the angle of attack now；Wherein m is just whole
Number；
Step 6: if gn+mLess than g, then increase rudder face (31) described in angular deflection of deflection a/4, weight on the basis of step 5 again
Multiple k step 2 arrives step 3 until gn+m+kMore than or equal to g, if gn+m+kEqual to g, then the described rudder face drift angle measuring acquisition is n*
a-m*a/2+k*a/4；Wherein k is positive integer；
Step 7: if gn+m+kMore than g, then reduce rudder face (31) described in angular deflection of deflection a/8, weight on the basis of step 6 again
Multiple p step 4 arrives step 5 until gn+m+k+pLess than or equal to g, if gn+m+k+pEqual to g, then the described rudder face drift angle measuring acquisition is
N*a-m*a/2+k*a/4-p*a/8, the angle of attack of the model aircraft (3) of the as required measurement of the angle of attack now；Wherein p is just whole
Number；
Step 8: as long as the overload that measurement obtains is not equal to g, then the circulation of repeat step 6 to 7, control and increase or reduce rudder face
The angle of deflection halves every time on the basis of original, until the final deflection angle of described rudder face (31) is less than described steering wheel
(4) minimum angle adjustable m, that is, described steering wheel (4) adjusts the model that described rudder face (31) deflection angle is less than its mechanical adjustable
Enclose and till cannot continuing to adjust described rudder face (31) deflection, now control, according to abovementioned steps, the described aircraft obtaining with measurement
The angle of attack of model (3) and rudder face drift angle can be considered that the overload of required measurement is the described angle of attack and described rudder face drift angle during g.
2. the method for claim 1 is it is characterised in that shown support system (1) adopts electronic joint real-time dynamicly
Control the angle of attack of described model aircraft (3).
3. method as claimed in claim 1 or 2 is it is characterised in that described steering wheel (4) is real-time during described wind tunnel test
Described rudder face (31) is dynamically controlled to enter horizontal deflection.
4. it is characterised in that described m is 0.01 degree, described a is 0.1 degree to method as claimed in claim 3.
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| CN201410808133.1A CN104483090B (en) | 2014-12-22 | 2014-12-22 | Dynamitic control and measurement method |
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| CN201410808133.1A CN104483090B (en) | 2014-12-22 | 2014-12-22 | Dynamitic control and measurement method |
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| CN113252285B (en) * | 2021-07-15 | 2021-10-08 | 中国空气动力研究与发展中心低速空气动力研究所 | Vertical wind tunnel model pitching-rolling test device and use method |
| CN113772115B (en) * | 2021-11-12 | 2022-02-11 | 中国空气动力研究与发展中心低速空气动力研究所 | Design method for controlling deflection angle of rear-mounted horizontal tail control surface of helicopter |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| JPH1144607A (en) * | 1997-07-25 | 1999-02-16 | Mitsubishi Heavy Ind Ltd | Wind tunnel |
| JP4310440B2 (en) * | 2006-03-22 | 2009-08-12 | 防衛省技術研究本部長 | Dynamic wind tunnel test method |
| CN101699237A (en) * | 2009-11-20 | 2010-04-28 | 中国航空工业空气动力研究院 | Three-dimensional model attitude angle video measuring system for wind tunnel model test |
| CN102305699A (en) * | 2011-05-19 | 2012-01-04 | 北京航空航天大学 | Wind tunnel experiment system for free flight model |
| CN102941925B (en) * | 2012-11-05 | 2015-05-13 | 北京航空航天大学 | Elastic correction method of airplane lifting resistance aerodynamic characteristics |
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