CN106840572A - A kind of near space high aspect ratio flexible flier wind tunnel test data correcting method - Google Patents
A kind of near space high aspect ratio flexible flier wind tunnel test data correcting method Download PDFInfo
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Abstract
The invention discloses a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method, including the amendment of wind tunnel test essential state data, the adjustment of horizontal tail established angle and wind tunnel test steerage data correction.The present invention combines domestic existing wind tunnel test experimental condition, by aeroelasticity and flow dynamics analysis means, completes the amendment to rigid body profile wind tunnel test data.The inventive method can be that the design of near space high aspect ratio flexible flier control rate and flight simulation provide accurate aerodynamic data, it is ensured that the aircraft being capable of the stabilized flight in the case of near space large deformation.
Description
Technical field
The present invention relates to a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method, for closing on
Space high aspect ratio flexible flier control rate design and flight simulation, belong to aerodynamic scope.
Background technology
With continuing to develop for scientific theory, engineering design and manufacturing technology, the stagnant empty ability of overlength turn into it is modern nobody
One important directions of machine development, near space high aspect ratio flexible flier with solar energy unmanned plane as Typical Representative, such as Europe
Zephyr, Heliplat, Solar Impulse in continent, Vulture, the SolarEagle in the U.S., domestic CH-T3, CH-T4
It is the cruising altitude for improving aircraft and the stagnant empty ability of overlength for realizing aircraft Deng, its primary design object on the whole,
And based on this two performances, reach including many military, civilian purposes including scouting, communication, scientific research, early warning etc..
At present, single shoe or double shoe cloth are added using high aspect ratio wing more than domestic and international near space solar energy unmanned plane
Office.Such unmanned plane aspect ratio is generally higher than 25, and due to the features such as its flexible structure, the low wing are carried, in the work of aerodynamic loading
Under, it may occur that obvious deformation, the deformation of wherein wing is especially pronounced, and wing tip maximum flexibility is up to semispan
25%.With the increase of deformation, there is larger change in the structural dynamic characteristics and aerodynamic characteristic of unmanned plane.
At present, domestic near space solar energy unmanned plane wind tunnel test is still using conventional low-speed wind tunnel scale model
Test method.However, conventional low-speed wind tunnel test method cannot accurate simulation solar energy unmanned plane gas after by aerodynamic loading
Dynamic characteristic, results of wind tunnel differs larger with actual result, it is impossible to be that control rate design and flight simulation provide accurate gas
Dynamic data.Near space solar energy unmanned plane aerodynamic characteristic how is obtained, is technical problem urgently to be resolved hurrily.
The content of the invention
Technology solve problem of the invention is：Overcome the deficiencies in the prior art, there is provided a kind of near space high aspect ratio is soft
Property aircraft wind tunnel test data correcting method, be near space high aspect ratio flexible flier control rate design and flight simulation
Accurate aerodynamic data is provided.
Technical solution of the invention is：A kind of near space high aspect ratio flexible flier wind tunnel test data correction
Method, comprises the following steps：
Step one：Wind tunnel test essential state data is modified using the analysis of gas bullet and method for numerical simulation, is obtained
The revised horizontal course data of aircraft；
Step 2：The horizontal tail established angle obtained to wind tunnel test using gas bullet analysis method is adjusted；
Step 3：Wind tunnel test steerage data are modified by testing gentle bullet analysis, obtain revised steerage
Data.
The implementation method of the step one is：
(2.1) according to the architectural characteristic of near space high aspect ratio flexible flier, flown using gas bullet analysis software
Row device deformation data；
(2.2) the gas bullet profile after the initial rigid body profile of aircraft and deformation is entered respectively by flow dynamics analysis software
Row CFD numerical simulations, obtain the gentle aerodynamic force and aerodynamic moment for playing profile of initial rigid body profile under every group of angle of attack and yaw angle；
(2.3) according to aerodynamic force and aerodynamic moment, the initial gentle side force for playing profile of rigid body profile under each angle of attack is calculated
Coefficient sideslip angular derivative, rolling moment coefficient sideslip angular derivative, yawing moment coefficient sideslip angular derivative, and further calculate accordingly
Lateral force coefficient sideslip angular derivative rate of change K (c before and after deformationz β), rolling moment coefficient sideslip angular derivative rate of change K (mx β), partially
Boat moment coefficient sideslip angular derivative rate of change K (my β)；
(2.4) using the K (c under the different angles of attackz β)、K(mx β) and K (my β) under the correspondence angle of attack in wind tunnel test data
Horizontal course data cz, mx, myIt is modified, correction formula is as follows：
cz'=cz·K(cz β)
mx'=mx·K(mx β)
my'=my·K(my β)
Wherein cz', mx', my' it is the horizontal course data of revised aircraft.
The implementation method of the step 2 is：
(3.1) according to the results of wind tunnel of near space high aspect ratio flexible flier, established angle, the rudder of horizontal tail are obtained
System of Pianguan County and elevator steerage mz δz；
(3.2) flexural deformation of the tail supporting rod under aerodynamic force and self gravitation effect, root are obtained using gas bullet analysis software
According to the flexural deformation and elevator steerage mz δz, it is compensated angle of rudder reflection δ1, to ensure that aircraft is in trim condition, enter
Step (3.3)；
(3.3) compensation angle of rudder reflection δ is judged1Whether it is 0, if compensation angle of rudder reflection δ1It is not 0, then according to system of rudder Pianguan County and benefit
Repay angle of rudder reflection δ1Horizontal tail established angle is adjusted, the horizontal tail established angle after being adjusted, is then back to step (3.2), updated
Compensation angle of rudder reflection δ1；Otherwise, horizontal tail established angle adjustment terminates.
The implementation method of the step 3 is：
(4.1) according to the results of wind tunnel of near space high aspect ratio flexible flier, the manipulation effect of elevator is obtained
Rate, i.e. steerage mz δz；
(4.2) power of 1kg is loaded in aircraft horizontal tail, measurement tail supporting rod end deflection angle Δ α；
(4.3) the power f that 1 ° of horizontal tail of control surface deflection is produced is obtained according to steerage0, 1 ° of change for causing of control surface deflection is obtained accordingly
Shape angle f0Δ α, and then obtain actual deflection angle (1-f0Δ α) and horizontal tail produce power f1=(1-f0·Δα)·
f0；
(4.4) according to the method for (4.3), the power produced according to horizontal tail continues to calculate actual deflection angle and the deflection angle
The power that lower horizontal tail is produced, constantly circulation, until the power f that horizontal tail is producednUntill constant；
(4.5) elevator steerage is modified, correction formula is as follows：
mz δz'=mz δz·fn/f0
Wherein fn/f0It is steerage attenuation coefficient.
Compared with the prior art, the invention has the advantages that：
The invention provides a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method, comprehensively examine
Consider domestic existing wind tunnel test condition, with reference to aeroelasticity and flow dynamics analysis means, complete and rigid body profile wind-tunnel is tried
Test the amendment of data.It is the design of near space high aspect ratio flexible flier control rate and flight simulation by the inventive method
There is provided accurate aerodynamic data, it is ensured that the aircraft being capable of the stabilized flight in the case of near space large deformation.
Brief description of the drawings
Fig. 1 is flow chart of the present invention；
Fig. 2 is wind tunnel test essential state data modification method flow chart of the present invention；
Fig. 3 is horizontal tail established angle method of adjustment flow chart of the present invention；
Fig. 4 is steerage data correcting method flow chart of the present invention.
Specific embodiment
Domestic wind tunnel test at present is the conventional survey that rigid model is carried out by aircraft scale model reasonable in design
Power is tested.Pilot project mainly includes three：
1) tail setting angle experiment：Suitable tail setting angle is found by test method.
2) basic status experiment：Measure the aerodynamic force c under different angle of attack and yaw angle βx, cy, czWith aerodynamic force away from mx, my,
mz.Wherein cxIt is resistance coefficient, cyIt is lift coefficient, czIt is lateral force coefficient, mxIt is rolling moment coefficient, myIt is yawing system
Number, mzIt is pitching moment coefficient.
3) rudder effectiveness test：When measuring primary control surface deflection angle, the c under different α and βx, cy, cz, mx, my, mzVariable quantity.
The present invention according to existing rigid model wind tunnel test data, with reference to aeroelasticity and flow dynamics analysis means,
Wind tunnel test data are modified.Overall flow of the present invention is as shown in figure 1, comprise the following steps that：
(1) wind tunnel test essential state data is modified, flow chart is as shown in Fig. 2 specifically amendment step is as follows：
(1.1) according to the composite structure characteristic of near space high aspect ratio flexible flier, analyze soft using gas bullet
Part, provides full machine deformation data；
(1.2) the full machine deformation data obtained according to gas bullet analysis software, is flown by three-dimensional digital-to-analogue software modeling
Gas bullet profile after the initial rigid body profile of device and deformation, then by flow dynamics analysis software respectively to the gentle bullet of rigid body profile
Profile carries out CFD numerical simulations, obtains under every group of α and β, and the six component aerodynamic force and aerodynamic force of rigid body profile are away from cx1, cy1, cz1,
mx1, my1, mz1, the gentle six component aerodynamic force for playing profile are with aerodynamic force away from cx2, cy2, cz2, mx2, my2, mz2.Wherein, current fluid
Mechanical analysis software category is more, and wind tunnel test mark need to be carried out to selected flow dynamics analysis software before this step is carried out
Mould is verified, to ensure the accuracy and reliability of the analysis software result of calculation.
(1.3) aerodynamic force and aerodynamic moment for being obtained according to flow dynamics analysis software, are obtained by the method for linear fit
The lateral force coefficient sideslip angular derivative c of initial rigid body profile under to each angle of attackz1 β, rolling moment coefficient sideslip angular derivative mx1 β, partially
Boat moment coefficient sideslip angular derivative my1 β, and gas bullet profile lateral force coefficient sideslip angular derivative cz2 β, rolling moment coefficient break away
Angular derivative mx2 β, yawing moment coefficient sideslip angular derivative my2 β, aircraft is bullied before bullet deformation under being then calculated each angle of attack
Lateral force coefficient sideslip angular derivative rate of change K (c afterwardsz β), rolling moment coefficient sideslip angular derivative rate of change K (mx β), yawing
Coefficient sideslip angular derivative rate of change K (my β)；K(cz β)=cz2 β/cz1 β,
K(mx β)=mx2 β/mx1 β,K(my β)=my2 β/my1 β。
(1.4) using the K (c under the different angles of attackz β)、K(mx β) and K (my β) under the correspondence angle of attack in wind tunnel test data
Horizontal course data cz, mx, myIt is modified, correction formula is as follows：
cz'=cz·K(cz β)
mx'=mx·K(mx β)
my'=my·K(my β)
Wherein cz', mx', my' it is the horizontal course data of revised aircraft.
Because the deformation of near space high aspect ratio flexible flier gas bullet influences smaller to longitudinal aerodynamic data, therefore longitudinal direction
cx, cy, mzDo not correct.
(2) horizontal tail established angle is adjusted, flow chart is as shown in figure 3, detailed process is：
(2.1) according to the rigid body results of wind tunnel of near space high aspect ratio flexible flier, obtain aircraft and patrolling
The horizontal tail established angle φ needed from trim is realized under boat state0, elevator steerage mz δzAnd horizontal tail established angle and lifting rudder kick
Equivalent relation (the note at angle：Generally, wind tunnel test can obtain deflecting the pitching moment m produced by 1 ° of horizontal tail established anglezVariable quantity
It is equivalent to how many degree elevator angles).
(2.2) according to horizontal tail established angle and elevator steerage, using the deformation data of gas bullet analysis software calculating aircraft,
Flexural deformation of the tail supporting rod under aerodynamic force and self gravitation effect, is obtained, so as to obtain the compensation rudder that full machine is also needed to from trim
Drift angle δ1, into step (2.3).
(2.3) according to the compensation angle of rudder reflection δ for obtaining1Adjustment horizontal tail established angle, is then repeated using gas bullet analysis software
(2.2), loop iteration is until compensation angle of rudder reflection δ1It is 0 °.
After iteration terminates, final horizontal tail established angle φ is obtainedn, as cruising condition horizontal tail established angle.
(3) steerage data are modified, flow is as shown in figure 4, detailed process is；
(3.1) elevator is obtained according to the rigid body results of wind tunnel of near space high aspect ratio flexible flier first
Driving efficiency, i.e. steerage mz δz。
(3.2) by loading the power of 1kg in horizontal tail, the deflection angle Δ α that tail supporting rod end causes by deformation is obtained.Mesh
Before can realize this step by two methods, the first is：Realize loading 1kg in tail supporting rod end using gas bullet analysis software
Power, be calculated because of the deflection angle Δ α that deformation causes；Second is the loading 1kg on the structure exemplar produced
Counterweight, actual measurement deflection angle Δ α.Second method is closer to actual value.
(3.3) according to steerage, using formula f0=mz δz·ρV2/ 2Sc/L can be obtained produced by 1 ° of rudder kick of lifting
Power f0, wherein, ρ V2/ 2 represent dynamic pressures, and S represents area of reference, and c represents reference length, L represent horizontal tail press the heart to center of gravity away from
From.
Two steps obtain the f of the inclined 1 ° of generation of rudder according to more than0The deformation angle Δ α caused with loading 1kg power, so that available
The deformation angle f that 1 ° of control surface deflection causes0Δ α, then obtaining actual deflection angle is
(1-f0Δ α), the corresponding power for producing is f1=(1-f0·Δα)·f0；
According to f1Obtain deflection angle (1-f1Δ α), and then power is obtained for f2=(1-f1·Δα)·f0；According to f2
To deflection angle (1-f2Δ α), and then power is obtained for f3=(1-f2·Δα)·f0... ...
Loop iteration, until fnIt is constant.Wherein f0,f1......fnRepresent the power produced by control surface deflection.
(3.4) steerage attenuation coefficient f is calculatedn/f0, then to elevator steerage m in wind tunnel test dataz δzIt is modified,
Obtain new steerage mz δz′, mz δz′=mz δz·fn/f0。
The present invention considers domestic existing wind tunnel test condition, complete with reference to aeroelasticity and flow dynamics analysis means
Into the amendment to rigid body profile wind tunnel test data.The method passes through flight test checking, can be big near space
Aspect ratio flexible flier control rate is designed and flight simulation provides accurate aerodynamic data, it is ensured that aircraft can close on sky
Between stabilized flight in the case of large deformation.
The undisclosed technology of the present invention belongs to general knowledge as well known to those skilled in the art.
Claims (4)
1. a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method, it is characterised in that including following step
Suddenly：
Step one：Wind tunnel test essential state data is modified using the analysis of gas bullet and method for numerical simulation, is corrected
The horizontal course data of aircraft afterwards；
Step 2：The horizontal tail established angle obtained to wind tunnel test using gas bullet analysis method is adjusted；
Step 3：Wind tunnel test steerage data are modified by testing gentle bullet analysis, obtain revised steerage data.
2. a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method according to claim 1,
It is characterized in that：The implementation method of the step one is：
(2.1) according to the architectural characteristic of near space high aspect ratio flexible flier, aircraft is obtained using gas bullet analysis software
Deformation data；
(2.2) the gas bullet profile after the initial rigid body profile of aircraft and deformation is carried out respectively by flow dynamics analysis software
CFD numerical simulations, obtain the gentle aerodynamic force and aerodynamic moment for playing profile of initial rigid body profile under every group of angle of attack and yaw angle；
(2.3) according to aerodynamic force and aerodynamic moment, the initial gentle lateral force coefficient for playing profile of rigid body profile under each angle of attack is calculated
Sideslip angular derivative, rolling moment coefficient sideslip angular derivative, yawing moment coefficient sideslip angular derivative, and deformation is further calculated accordingly
Front and rear lateral force coefficient sideslip angular derivative rate of change K (cz β), rolling moment coefficient sideslip angular derivative rate of change K (mx β), yaw forces
Moment coefficient sideslip angular derivative rate of change K (my β)；
(2.4) using the K (c under the different angles of attackz β)、K(mx β) and K (my β) to the horizontal boat under the correspondence angle of attack in wind tunnel test data
To data cz, mx, myIt is modified, correction formula is as follows：
cz'=cz·K(cz β)
mx'=mx·K(mx β)
my'=my·K(my β)
Wherein cz', mx', my' it is the horizontal course data of revised aircraft.
3. a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method according to claim 1,
It is characterized in that：The implementation method of the step 2 is：
(3.1) according to the results of wind tunnel of near space high aspect ratio flexible flier, established angle, rudder Pianguan County of horizontal tail are obtained
System and elevator steerage mz δz；
(3.2) flexural deformation of the tail supporting rod under aerodynamic force and self gravitation effect, is obtained using gas bullet analysis software, according to institute
State flexural deformation and elevator steerage mz δz, it is compensated angle of rudder reflection δ1, to ensure that aircraft is in trim condition, into step
(3.3)；
(3.3) compensation angle of rudder reflection δ is judged1Whether it is 0, if compensation angle of rudder reflection δ1It is not 0, then according to system of rudder Pianguan County and compensation rudder
Drift angle δ1Horizontal tail established angle is adjusted, the horizontal tail established angle after being adjusted, is then back to step (3.2), update compensation
Angle of rudder reflection δ1；Otherwise, horizontal tail established angle adjustment terminates.
4. a kind of near space high aspect ratio flexible flier wind tunnel test data correcting method according to claim 1,
It is characterized in that：The implementation method of the step 3 is：
(4.1) according to the results of wind tunnel of near space high aspect ratio flexible flier, the driving efficiency of elevator is obtained, i.e.,
Steerage mz δz；
(4.2) power of 1kg is loaded in aircraft horizontal tail, measurement tail supporting rod end deflection angle Δ α；
(4.3) the power f that 1 ° of horizontal tail of control surface deflection is produced is obtained according to steerage0, 1 ° of deformation angle for causing of control surface deflection is obtained accordingly
f0Δ α, and then obtain actual deflection angle (1-f0Δ α) and horizontal tail produce power f1=(1-f0·Δα)·f0；
(4.4) according to the method for (4.3), the power produced according to horizontal tail continues to be put down under calculating actual deflection angle and the deflection angle
The power that tail is produced, constantly circulation, until the power f that horizontal tail is producednUntill constant；
(4.5) elevator steerage is modified, correction formula is as follows：
mz δz'=mz δz·fn/f0
Wherein fn/f0It is steerage attenuation coefficient.
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| CN111006835A (en) * | 2019-11-19 | 2020-04-14 | 蓝箭航天空间科技股份有限公司 | Rocket projectile pitching moment coefficient and pressure center coefficient correction method and storage medium |
| CN111220347B (en) * | 2020-03-03 | 2021-06-22 | 四川腾盾科技有限公司 | Aircraft pneumatic coordination correction method |
| CN111220347A (en) * | 2020-03-03 | 2020-06-02 | 四川腾盾科技有限公司 | Aircraft pneumatic coordination correction method |
| CN113987794A (en) * | 2021-10-26 | 2022-01-28 | 成都飞机工业（集团）有限责任公司 | Nonlinear rigid pneumatic data correction method, device, equipment and storage medium for airplane |
| CN113987794B (en) * | 2021-10-26 | 2024-06-07 | 成都飞机工业（集团）有限责任公司 | Nonlinear rigid pneumatic data correction method, device and equipment for airplane and storage medium |
| CN114001913A (en) * | 2021-10-27 | 2022-02-01 | 江西洪都航空工业股份有限公司 | Component pneumatic load correction method based on wind tunnel test |
| CN114001913B (en) * | 2021-10-27 | 2024-04-02 | 江西洪都航空工业股份有限公司 | Component pneumatic load correction method based on wind tunnel test |
| CN114383801B (en) * | 2021-12-20 | 2024-03-19 | 中国航天空气动力技术研究院 | Wind tunnel virtual flight test system and method for variant aircraft |
| CN114383801A (en) * | 2021-12-20 | 2022-04-22 | 中国航天空气动力技术研究院 | Wind tunnel virtual flight test system and method for morphing aircraft |
| CN114608794B (en) * | 2022-05-11 | 2022-07-19 | 中国航空工业集团公司哈尔滨空气动力研究所 | Method for measuring aerodynamic coefficient of model wind tunnel virtual flight test |
| CN114608794A (en) * | 2022-05-11 | 2022-06-10 | 中国航空工业集团公司哈尔滨空气动力研究所 | Method for measuring aerodynamic coefficient of model wind tunnel virtual flight test |
| CN116183154B (en) * | 2023-04-24 | 2023-06-27 | 中国空气动力研究与发展中心高速空气动力研究所 | Ultra-large attack angle wind tunnel test method for slender body model |
| CN116183154A (en) * | 2023-04-24 | 2023-05-30 | 中国空气动力研究与发展中心高速空气动力研究所 | Ultra-large attack angle wind tunnel test method for slender body model |
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