CN106650095A - Method for correcting unmanned aerial vehicle control matrix based on wind tunnel test data and CFD calculation - Google Patents
Method for correcting unmanned aerial vehicle control matrix based on wind tunnel test data and CFD calculation Download PDFInfo
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- CN106650095A CN106650095A CN201611194028.9A CN201611194028A CN106650095A CN 106650095 A CN106650095 A CN 106650095A CN 201611194028 A CN201611194028 A CN 201611194028A CN 106650095 A CN106650095 A CN 106650095A
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Abstract
Provided is a method for correcting an unmanned aerial vehicle control matrix based on wind tunnel test data and CFD calculation. Due to the fact that flight control needs pneumatic control, the reliable control matrix is provided. The wind tunnel test data of a planar model with an unconventional pneumatic layout is extracted, the dynamic derivative of a design point is subjected to local correction, and the influence of the unconventional layout on a traditional layout empirical formula is effectively eliminated. Compared with an existing control matrix providing method, an original method refers to ab control matrix research researched through a traditional unmanned aerial vehicle engineering design method, is obtained mainly based on experience summarized through many years of conventional appearance pneumatic layout design, has limitation to the unconventional layout and is disconnected from a wind tunnel test link. The conventional force measurement data of a wind tunnel test of the planar model is utilized creatively, the control matrix is corrected, and a more precise control matrix is provided for PID parameter control.
Description
Technical field
The present invention relates to a kind of modification method of the unmanned aerial vehicle (UAV) control matrix calculated based on wind tunnel test data and CFD, together
When take into account Wind Tunnel Data and CFD result of calculations are referred to, normal arrangement and unconventional can be quickly realized in engineering
The calculating of layout flight control matrix, mainly used in flight vehicle aerodynamic design process, belongs to aviation aircraft pneumatic design
Technical field.
Background technology
During Flight Vehicle Design, the design of control rate generally needs pneumatic cooperation department to provide, once flight control square
Battle array offer error is larger, can cause the prolongation of design cycle and the decline of the precision of design.And fly control matrix, mainly by
What pneumatic department provided, the utilization of traditional empirical equation is confined to conventional aerodynamic arrangement, for unconventional layout and approximately
The profile of normal arrangement is calculated has certain error, and as following aerodynamic arrangement (is obtained based on pneumatic design optimization
Profile) development error can be increasing.It is therefore desirable on the basis of project designing accuracy and time cycle requirement is met,
Necessary amendment is carried out to Conventional wisdom formula.
The content of the invention
The present invention technology solve problem be：Overcome the deficiencies in the prior art, there is provided one kind is based on wind tunnel test data
The modification method of the unmanned aerial vehicle (UAV) control matrix calculated with CFD, realizes unmanned aerial vehicle (UAV) control matrix engineering method application, reliable quick
Realize wind tunnel test data and CFD and calculate data and Conventional wisdom formula, at utmost meet flight control matrix engineering
Precision and cycle request.
The present invention technical solution be：The unmanned aerial vehicle (UAV) control matrix that calculated with CFD based on wind tunnel test data is repaiied
Correction method, step is as follows：
(1) Mathematical Modeling is built according to the formal parameter of unmanned plane, and carries out CFD calculating, obtain the pneumatic ginseng of unmanned plane
Number；
(2) judge whether Mathematical Modeling belongs to Conventional pneumatic layout, if into step (3), if not into step
(4)；
(3) according to the Mathematical Modeling of above-mentioned determination, using traditional empirical equation control matrix is calculated, and with gas in (1)
C in dynamic parameterL,CDAnd CmWith reference to traditional empirical equation, the respective element in control matrix is modified, obtains new
Control matrix, using the control adjustment of matrix unmanned plane pid control parameter；The CLFor aircraft lift coefficient, CDAircraft
Resistance coefficient, CmFor aircraft pitching moment coefficient；
(4) according to the Mathematical Modeling of above-mentioned determination, the traditional empirical equation of Selection utilization or datcom calculate control matrix；
Model to being obtained based on Mathematical Modeling processing in step (1) carries out wind tunnel test, using the C in wind tunnel test dataLα,Cmα
AndRespective element in controlling matrix is modified, and new control matrix is obtained, using the control adjustment of matrix
Unmanned plane pid control parameter；
The CLαThe full machine lift coefficient of aircraft is to angle of attack derivative, CmαThe full machine pitching moment coefficient of aircraft is led to the angle of attack
Number,Because of the airplane ascensional force derivative coefficient that change in angle of attack causes,Because the aircraft pitch moment coefficient that change in angle of attack causes is led
Number.
When the outfield experiments data of unmanned plane can be obtained, (3) or (4) are obtained control using outfield experiments data
Matrix is modified, and is finally controlled matrix.
Mathematical Modeling to setting up in step (1) simplifies, and obtains the simplified mathematical modulo of blended wing-body plus empennage
Type.
Cfd calculating is carried out to simplified Mathematical Modeling, whether lift-drag ratio, torque factor, lift coefficient, resistance coefficient is judged
Meet design objective, if meeting, step (2) is continued executing with to simplified Mathematical Modeling；Otherwise, Mathematical Modeling carried out again
Simplify.
Described simplification is processed as the protruding part by unmanned plane middle fuselage length direction less than the 10% of equivalent chord length and enters
Row smooth treatment, either horizontal tail plus double fins or horizontal tail add single vertical tail either vee tail to be modified as into T-shaped empennage；Or will
Tail supporting rod is removed；Or by inlet lip and fuselage fusion treatment；Or by nozzle and fuselage fusion treatment；When wingtip it is little
When wing area is less than wing area 10%, winglet is directly removed, or when winglet area is more than or equal to airfoil
During product 10%, after winglet is removed, in the case where leading edge of a wing line and trailing edge line are constant, increase length, wing area
Incrementss are winglet area；Or store Combinations or screw are ignored into process.
Air flue lip is continuous with the second dervative that fuselage fusion treatment is the guide line for merging surface.
Nozzle is that nozzle and fuselage trailing edge are enclosed construction with fuselage fusion treatment, it is to avoid La Vier jet pipe effect.
Vee tail is modified as into T-shaped empennage using equivalent perspective plane area method, and either horizontal tail plus double fins or horizontal tail add list
Vertical tail.
Equivalent perspective plane area method is that vee tail is projected to into horizontal plane and vertical face, and two after being projected area makes
Projected area of the modification back segment aerofoil profile in horizontal plane and vertical face is equal thereto.
When T-shaped empennage is revised as using vee tail, the use environment of unmanned plane is necessary for high and medium environment.
Compared with the prior art, the invention has the advantages that：
(1) modification method of the unmanned aerial vehicle (UAV) control matrix by being calculated with CFD based on wind tunnel test data, effectively by gas
In combination with dynamic design is with control parameter design (pid parameter), it is to avoid in conventional Flight Vehicle Design development process pneumatic design and
Control parameter design contact discrepancy.Constantly weed out the old and bring forth the new for aerodynamic configuration of aircraft, new layout is continued to bring out, Yuan Youdan
The increasingly lower present situation of pure conventional method estimation precision, based on the unmanned aerial vehicle (UAV) control matrix that wind tunnel test data and CFD are calculated
Modification method, effectively utilize new tool entering to precision by means of wind tunnel test, CFD data, and datcom various dimensions
Row is greatly lifted.
(2) the simplified means of this paper simplify means compared to other, the parameter being effectively guaranteed in CFD calculating process
Precision, and using the precision in datcom software calculating process.
(3) because vee tail design is easy to control system to be designed, the frequency occurred in the design of following aerodynamic arrangement
Can dramatically increase.The present invention according to use environment (flying height), engine location, and empennage and main wing relative position, with
And whether Flying-wing to determine whether to be modified as T-shaped empennage, either horizontal tail plus double fins or horizontal tail add single vertical tail.
(4) the control matrix computations that the data in existing wind tunnel data storehouse are introduced unconventional layout of the invention
In, original method precision is improved.
(5) according to the flight test in outfield, and free flight experimental data, the aerodynamic parameter of acquisition is to flight control matrix
Further corrected, and then obtained control matrix.Future can be used for the modified to the model aircraft, unmanned plane PID controls
Parameter designing processed is used.
Description of the drawings
Fig. 1 is flow chart of the present invention.
Specific embodiment
The specific embodiment of the present invention is further described in detail below in conjunction with the accompanying drawings.
Flight control matrix in flight vehicle aerodynamic design process, it usually needs accurately provide for aircraft PID designs, be
Ensure the progress of design cycle and engineering design, the means of generally conventional upper employing are simple empirical equation estimations.Using
During, the invention introduces wind tunnel data and CFD data wherein, used as correcting unconventional layout and Conventional pneumatic
The foundation of profile.
As shown in figure 1, a kind of modification method of the unmanned aerial vehicle (UAV) control matrix calculated based on wind tunnel test data and CFD, real
Apply step as follows：
(1) Mathematical Modeling is built according to the formal parameter of unmanned plane, and carries out CFD calculating, obtain the pneumatic ginseng of unmanned plane
Number, such as CL,CDAnd Cm；The CLFor aircraft lift coefficient, CDAircraft resistance coefficient, CmFor aircraft pitching moment system
Number；
Under the requirement for ensureing Flight Vehicle Design index parameter, for the ease of solving to unmanned aerial vehicle (UAV) control matrix, it is
PID control design proposes control matrix.The present invention simplifies to the Mathematical Modeling set up, and obtains blended wing-body plus empennage
Simplify Mathematical Modeling, and then CFD calculating is carried out to simplified Mathematical Modeling, judge lift-drag ratio, torque factor, lift coefficient, resistance
Whether force coefficient meets design objective, if meeting, step (2) is continued executing with to simplified Mathematical Modeling；Otherwise, logarithm again
Learn model to be simplified.
The above-mentioned simplification for referring to can adopt following manner：
Protruding parts of the A by unmanned plane middle fuselage length direction less than the 10% of equivalent chord length carries out smooth treatment,
Vee tail is modified as T-shaped empennage by B, and either horizontal tail plus double fins or horizontal tail add single vertical tail；
C removes tail supporting rod；
D is by inlet lip and fuselage fusion treatment；
E is by nozzle and fuselage fusion treatment；
F directly removes winglet when winglet area is less than wing area 10%, or when winglet face
When product is more than or equal to wing area 10%, after winglet is removed, in the case where leading edge of a wing line and trailing edge line are constant, increase
Plus length, wing area incrementss are winglet area；
Store Combinations or screw are ignored process by G.
Wherein, air flue lip is continuous with the second dervative that fuselage fusion treatment is the guide line for merging surface.Nozzle with
Fuselage fusion treatment is nozzle and fuselage trailing edge is enclosed construction, it is to avoid La Vier jet pipe effect.
Vee tail is modified as T-shaped empennage either horizontal tail plus double fins or flat by the present invention using equivalent perspective plane area method
Tail adds single vertical tail.Equivalent perspective plane area method is that vee tail is projected to into horizontal plane and vertical face, two after being projected face
Product, makes projected area of the modification back segment aerofoil profile in horizontal plane and vertical face equal thereto.The empennage obtained after to simplifying, is protecting
Under demonstrate,proving equivalent perspective plane area method, particular/special requirement is not done in the relative position present invention, complete to simplify according to custom requirements.
When T-shaped empennage is revised as using vee tail, the use environment of unmanned plane is necessary for high and medium environment.
(2) judge whether Mathematical Modeling belongs to Conventional pneumatic layout, if into step (3), if not into step
(4)；Above-mentioned Conventional pneumatic layout be main wing in front tailplane rear, have one or two vertical fin, for example, ARJ21.
(3) according to the Mathematical Modeling of above-mentioned determination, counted using traditional empirical equation (referring specifically to airplane design handbook)
Control matrix is calculated, and with the C in aerodynamic parameter in (1)L,CDAnd CmRespective element in controlling matrix is modified, right
CL,CDAnd CmParameter item substituted, new control matrix is obtained, using control adjustment of matrix unmanned plane PID control ginseng
Number, the CLFor aircraft lift coefficient, CDAircraft resistance coefficient, CmFor aircraft pitching moment coefficient；
(4) according to the Mathematical Modeling of above-mentioned determination, the traditional empirical equation of Selection utilization or datcom calculate control matrix；
Model to being obtained based on simplified Mathematical Modeling processing carries out wind tunnel test, using the C in wind tunnel test dataLα,CmαAndRespective element in controlling matrix is modified, to CLα,CmαAndParameter item substituted, obtain
New control matrix, by the use of the control matrix as the design for being originally inputted parameter and being applied to unmanned plane pid control parameter；It is described
CLαThe full machine lift coefficient of aircraft is to angle of attack derivative, CmαThe full machine pitching moment coefficient of aircraft to angle of attack derivative,Because of the angle of attack
The airplane ascensional force derivative coefficient that change causes,Because of the aircraft pitch moment derivative coefficient that change in angle of attack causes.
When the outfield experiments data of unmanned plane can be obtained, (3) or (4) are obtained control using outfield experiments data
Matrix is modified, and is finally controlled matrix.Future can be used for the modified to the model aircraft, unmanned plane PID control
Parameter designing is used.
The content not being described in detail in description of the invention belongs to the known technology of professional and technical personnel in the field.
Claims (10)
1. the modification method of the unmanned aerial vehicle (UAV) control matrix for being calculated with CFD based on wind tunnel test data, it is characterised in that step is as follows：
(1) Mathematical Modeling is built according to the formal parameter of unmanned plane, and carries out CFD calculating, obtain the aerodynamic parameter of unmanned plane；
(2) judge whether Mathematical Modeling belongs to Conventional pneumatic layout, if into step (3), if not into step (4)；
(3) according to the Mathematical Modeling of above-mentioned determination, using traditional empirical equation control matrix is calculated, and with pneumatic ginseng in (1)
C in numberL,CDAnd CmWith reference to traditional empirical equation, the respective element in control matrix is modified, obtains new control
Matrix processed, using the control adjustment of matrix unmanned plane pid control parameter；The CLFor aircraft lift coefficient, CDAircraft resistance
Coefficient, CmFor aircraft pitching moment coefficient；
(4) according to the Mathematical Modeling of above-mentioned determination, the traditional empirical equation of Selection utilization or datcom calculate control matrix；To base
The model that Mathematical Modeling processing is obtained in step (1) carries out wind tunnel test, using the C in wind tunnel test dataLα,CmαAndTo control matrix in respective element be modified, obtain new control matrix, using the control adjustment of matrix nobody
Machine pid control parameter；
The CLαThe full machine lift coefficient of aircraft is to angle of attack derivative, CmαThe full machine pitching moment coefficient of aircraft to angle of attack derivative,
Because of the airplane ascensional force derivative coefficient that change in angle of attack causes,Because of the aircraft pitch moment derivative coefficient that change in angle of attack causes.
2. method according to claim 1, it is characterised in that：When the outfield experiments data of unmanned plane can be obtained, profit
The control matrix that (3) or (4) obtain is modified with outfield experiments data, is finally controlled matrix.
3. method according to claim 1, it is characterised in that：Mathematical Modeling to setting up in step (1) simplifies,
Obtain the simplified Mathematical Modeling of blended wing-body plus empennage.
4. method according to claim 3, it is characterised in that：Cfd calculating is carried out to simplified Mathematical Modeling, judges to rise resistance
Than, torque factor, lift coefficient, resistance coefficient whether meet design objective, if meet, simplified Mathematical Modeling is continued to hold
Row step (2)；Otherwise, Mathematical Modeling simplified again.
5. the method according to claim 3 or 4, it is characterised in that：Described simplification is processed as unmanned plane middle fuselage is long
Degree direction carries out smooth treatment less than 10% protruding part of equivalent chord length, either by vee tail be modified as T-shaped empennage or
Horizontal tail adds double fins or horizontal tail to add single vertical tail；Or remove tail supporting rod；Or by inlet lip and fuselage fusion treatment；
Or by nozzle and fuselage fusion treatment；When winglet area is less than wing area 10%, winglet is directly gone
Remove, or when winglet area is more than or equal to wing area 10%, after winglet is removed, in leading edge of a wing line with after
In the case that edge line is constant, increase length, wing area incrementss are winglet area；Or neglect store Combinations or screw
Slightly process.
6. method according to claim 5, it is characterised in that：Air flue lip is to merge drawing for surface with fuselage fusion treatment
The second dervative of wire is continuous.
7. method according to claim 5, it is characterised in that：Nozzle is after nozzle and fuselage with fuselage fusion treatment
Edge is enclosed construction, it is to avoid La Vier jet pipe effect.
8. method according to claim 5, it is characterised in that：Vee tail is modified as using equivalent perspective plane area method T-shaped
Either horizontal tail adds double fins or horizontal tail to add single vertical tail to empennage.
9. method according to claim 8, it is characterised in that：Equivalent perspective plane area method is that vee tail is projected to into level
Face and vertical face, two after being projected area, make modification back segment aerofoil profile the projected area in horizontal plane and vertical face and its
It is equal.
10. method according to claim 5, it is characterised in that：When T-shaped empennage is revised as using vee tail, unmanned plane
Use environment be necessary for high and medium environment.
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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 |
| CN111563292A (en) * | 2020-04-15 | 2020-08-21 | 成都飞机工业（集团）有限责任公司 | Laminar flow airfoil type Re number effect correction method based on flow transition |
| CN112362290A (en) * | 2020-09-30 | 2021-02-12 | 成都飞机工业（集团）有限责任公司 | Method for rapidly analyzing influence of thickness tolerance of wing on resistance 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 |
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| CN111563292A (en) * | 2020-04-15 | 2020-08-21 | 成都飞机工业（集团）有限责任公司 | Laminar flow airfoil type Re number effect correction method based on flow transition |
| CN112362290A (en) * | 2020-09-30 | 2021-02-12 | 成都飞机工业（集团）有限责任公司 | Method for rapidly analyzing influence of thickness tolerance of wing on resistance coefficient |
| CN112362290B (en) * | 2020-09-30 | 2021-08-03 | 成都飞机工业（集团）有限责任公司 | Method for rapidly analyzing influence of thickness tolerance of wing on resistance coefficient |
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