CN110162822A  The quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode  Google Patents
The quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode Download PDFInfo
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 CN110162822A CN110162822A CN201910206453.2A CN201910206453A CN110162822A CN 110162822 A CN110162822 A CN 110162822A CN 201910206453 A CN201910206453 A CN 201910206453A CN 110162822 A CN110162822 A CN 110162822A
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Abstract
The present invention provides a kind of quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode, this method comprises: one, grid dividing is carried out to pneumatic face is deformed；Two, it obtains structural modal information and is interpolated on the aerodynamic grid for deforming pneumatic face, determine the structure motion rule for deforming pneumatic face；Three, determine timedomain analysis material calculation；Four, it is based on full potential flow theory, arranges collar vortex on each aerodynamic grid, is calculated using unsteady Bernoulli equation and obtains the size and distribution that deform the aerodynamic force on pneumatic surface grids；Five, rear trailing vortex is moved into the displacement of a time step with local airflow speed to form downstream trailing vortex flow field；Six, the structural modal information under future time steplength is interpolated on the pneumatic surface grids of deformation, step 4 and step 5 are repeated, until completing the solution for promoting unsteady aerodynamic force in time domain.It applies the technical scheme of the present invention, to solve the technical problem that the modeling of time domain aerodynamics evaluation is complicated in the prior art and computational efficiency is low.
Description
Technical field
The present invention relates to flight vehicle aerodynamic elasticity technical field more particularly to a kind of time domain of coupled structure mode are quickly non
Unsteady Flow calculation method.
Background technique
The calculating of unsteady aerodynamic force is the key link of Flight Vehicle Design and analysis, directly affects the motordriven of aircraft
Property, flight stability and the assessment of safety.Therefore unsteady aerodynamic force calculating is aeroelasticity point in Flight Vehicle Design link
The pith of analysis, servo stabilization analysis, flight mechanics emulation, flying quality assessment, the reasonability and standard of Modeling Calculation
True property is to Flight Vehicle Design and analysis important in inhibiting.
In aeroelasticity field, unsteady aerodynamic force is aircraft flutter analysis, aeroelastic divergence analysis and the sound of something astir
The important input that should be analyzed.But due to the complexity that unsteady aerodynamic force calculates, aeroelasticity field is unsteady for a long time
Aerodynamic force is passed through frequently with resonance oscillation it is assumed that being simplified in frequency domain and being calculated, this makes based on unsteady aerodynamic force
The flutter analysis of calculation and stability analysis are also confined in frequency domain mostly, are needed if timedomain analysis to be carried out relatively complicated
Conversion, also to undertake timefrequency domain conversion modeling bring error.With contemporary aircraft high speed, lightweight, flexible etc.
Growth requirement, the timedomain analysis of multidisciplinary synthesis become the inevitable approach of aircraft analysis and design, this is just to unsteady pneumatic
Power proposes fast and accurately timedomain calculation demand.But current time domain aerodynamic force mostly uses CFD to calculate, and computation modeling is complicated,
Computational efficiency is low, is not appropriate for the analysis application of engineering phase.
Summary of the invention
The present invention provides a kind of quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode, it is able to solve existing
There is the technical problem that the modeling of time domain aerodynamics evaluation is complicated in technology and computational efficiency is low.
The present invention provides a kind of quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode, time domain is quickly non
Unsteady Flow calculation method includes: step 1, carries out grid dividing to the pneumatic face of the deformation of aircraft；Step 2 obtains and flies
Structural modal information is simultaneously interpolated on the aerodynamic grid for deforming pneumatic face by the structural modal information of row device, is determined and is deformed pneumatic face
Structure motion rule；Step 3, according to aerodynamic grid size, speed of incoming flow and the knot in the pneumatic face of deformation for deforming pneumatic face
The structure characteristics of motion determines timedomain analysis material calculation；Step 4 is based on full potential flow theory, is deforming each pneumatic of pneumatic face
Collar vortex is arranged on grid, according to the aerodynamic influence matrix of pacesetting equations aircraft, according to aerodynamic influence
Coefficient matrix and normal direction Solid boundary condition calculate the collar vortex circular rector size for obtaining current time, according to the collar vortex at current time
Circular rector size simultaneously calculates the size and distribution for obtaining the aerodynamic force on the pneumatic surface grids of deformation using unsteady Bernoulli equation；Step
Rapid five, keep rear trailing vortex circular rector it is constant, by rear trailing vortex with local airflow speed move a time step displacement with
Form downstream trailing vortex flow field；Step 6, on the basis of the trailing vortex flow field of downstream, by the structural modal information under future time steplength
It is interpolated on the pneumatic surface grids of deformation, repeats step 4 and step 5, until completing to promote asking for unsteady aerodynamic force in time domain
Solution.
Further, the size of the collar vortex circular rector Γ at current time can be according to aerodynamic influence matrix A and normal direction
Solid boundary conditionTo obtain, wherein A is wing aerodynamic power influence coefficient matrix, and Γ is wing
Adhere to the column vector of collar vortex intensity composition, V (t)=[U (t), V (t), W (t)], V (t) are to include wing motion speed and Fei Ding
The speed of normal speed of incoming flow, V_{w}For trailing vortex at the wing caused by induced velocity,For the normal vector of wing locality.
Further, the aerodynamic force deformed on pneumatic surface grids can be according to the collar vortex circular rector size and benefit according to current time
With unsteady Bernoulli equationCome
It obtains, wherein ρ is to carry out current density, τ_{i}For the tangent vector for deforming the direction pneumatic surface grids i, τ_{j}To deform the direction pneumatic surface grids j
Tangent vector, Δ c_{ij}The chord length to jth ith, Zhan Xiang of grid, Δ b are arranged to deform pneumatic face_{ij}For deform pneumatic face arrange to
Ith, the length of jth of grid of Zhan Xiang, Γ_{ij}The collar vortex circular rector to jth ith, Zhan Xiang of grid is arranged to deform pneumatic face,
Γ_{i1,j}The collar vortex circular rector to jth (i1)th, Zhan Xiang of grid, Γ are arranged to deform pneumatic face_{i,j1}It arranges to deform pneumatic face to the
I, collar vortex circular rector to 1 grid of jth, Δ p are opened up_{ij}For deform pneumatic face arrange to ith, jth of grid of Zhan Xiang it is pneumatic
Power.
Further, in step 1, for the tangential grid in pneumatic face at least more than 5, opening up can be according to net to grid after deformation
Lattice unit slenderness ratio is determined.
Further, in step 2, the structural modal information of aircraft includes rigid motion mode or elastic movement mould
State.
Further, the structure motion rule for deforming pneumatic face includes simple harmonic motion.
It applies the technical scheme of the present invention, the quick unsteady aerodynamic force of time domain for providing a kind of coupled structure mode calculates
Method, this method by space to threedimensional pneumatic surface grids divide can the bending of accurate description pneumatic face, torsional deflection it is several
What feature, the Modeling Calculation suitable for aerodynamic force under the various movements of aircraft and deformation state；By being arranged on aerodynamic grid
Collar vortex is solved using full Patential Flow Equation, and eliminate microvariations small deformation it is assumed that be suitable for the wider change of lifting surface
The unsteady aerodynamic force of shape and motion state calculates；By analyzing coupled structure mode of motion, it is applicable to any true
The rigid body and elastic movement mode of setting formula, compared with the existing technology frequency domain unsteady aerodynamic force calculating are only applicable to simple harmonic quantity
The structural elasticity mode of vibration greatly expands the scope of application of the unsteady aerodynamic force in conjunction with structure motion mode.Therefore,
The quick unsteady aerodynamic force calculation method of time domain provided by the present invention is calculated with time domain CFD unsteady aerodynamic force in the prior art
Compare, modeling is simple, and time domain promotes computational efficiency high, can adapt to contemporary aircraft design multimode, more rounds it is quick when
The demand of domain simulation analysis.
Detailed description of the invention
Included attached drawing is used to provide to be further understood from the embodiment of the present invention, and which constitute one of specification
Point, for illustrating the embodiment of the present invention, and come together to illustrate the principle of the present invention with verbal description.It should be evident that below
Attached drawing in description is only some embodiments of the present invention, for those of ordinary skill in the art, is not paying creation
Property labour under the premise of, be also possible to obtain other drawings based on these drawings.
The time domain that Fig. 1 shows the coupled structure mode provided according to a particular embodiment of the invention is quickly unsteady pneumatic
The flow diagram of power calculation method；
Fig. 2 shows the structural representations that the pneumatic surface grids of the threedimension curved surface provided according to a particular embodiment of the invention divide
Figure；
Fig. 3 shows the schematic diagram of the collar vortex time stepping method provided according to a particular embodiment of the invention；
Fig. 4 shows the plane before the pneumatic face of deformed curved surface provided according to a particular embodiment of the invention and deformation
Pneumatic face contrast schematic diagram；
Fig. 5 shows the schematic diagram in the pneumatic face of the torsion mode provided according to a particular embodiment of the invention；
Fig. 6 shows the flow field schematic diagram of the torsion mode provided according to a particular embodiment of the invention；
Fig. 7 shows the curve of the wing aerodynamic power under the sin characteristics of motion provided according to a particular embodiment of the invention
Schematic diagram；
Fig. 8 shows the song of the wing aerodynamic power under the 1cos characteristics of motion provided according to a particular embodiment of the invention
Line schematic diagram.
Specific embodiment
It should be noted that in the absence of conflict, the features in the embodiments and the embodiments of the present application can phase
Mutually combination.Following will be combined with the drawings in the embodiments of the present invention, and technical solution in the embodiment of the present invention carries out clear, complete
Ground description, it is clear that described embodiments are only a part of the embodiments of the present invention, instead of all the embodiments.It is right below
The description only actually of at least one exemplary embodiment be it is illustrative, never as to the present invention and its application or use
Any restrictions.Based on the embodiments of the present invention, those of ordinary skill in the art are without creative efforts
Every other embodiment obtained, shall fall within the protection scope of the present invention.
It should be noted that term used herein above is merely to describe specific embodiment, and be not intended to restricted root
According to the illustrative embodiments of the application.As used herein, unless the context clearly indicates otherwise, otherwise singular
Also it is intended to include plural form, additionally, it should be understood that, when in the present specification using term "comprising" and/or " packet
Include " when, indicate existing characteristics, step, operation, device, component and/or their combination.
Unless specifically stated otherwise, positioned opposite, the digital table of the component and step that otherwise illustrate in these embodiments
It is not limited the scope of the invention up to formula and numerical value.Simultaneously, it should be appreciated that for ease of description, each portion shown in attached drawing
The size divided not is to draw according to actual proportionate relationship.For technology, side known to person of ordinary skill in the relevant
Method and equipment may be not discussed in detail, but in the appropriate case, and the technology, method and apparatus should be considered as authorizing explanation
A part of book.In shown here and discussion all examples, any occurrence should be construed as merely illustratively, and
Not by way of limitation.Therefore, the other examples of exemplary embodiment can have different values.It should also be noted that similar label
Similar terms are indicated in following attached drawing with letter, therefore, once it is defined in a certain Xiang Yi attached drawing, then subsequent attached
It does not need that it is further discussed in figure.
As shown in Figures 1 to 8, the time domain for providing a kind of coupled structure mode according to a particular embodiment of the invention is quick
Unsteady aerodynamic force calculation method, the quick unsteady aerodynamic force calculation method of the time domain includes: step 1, the deformation to aircraft
Pneumatic face carries out grid dividing；Step 2 obtains the structural modal information of aircraft and structural modal information is interpolated into deformation
On the aerodynamic grid in pneumatic face, the structure motion rule for deforming pneumatic face is determined；Step 3, according to the pneumatic net for deforming pneumatic face
The structure motion rule of lattice size, speed of incoming flow and the pneumatic face of deformation determines timedomain analysis material calculation；Step 4, based on complete
Potential flow theory arranges collar vortex on each aerodynamic grid for deforming pneumatic face, according to pacesetting equations aircraft
Aerodynamic influence matrix calculates according to aerodynamic influence matrix and normal direction Solid boundary condition and obtains current time
Collar vortex circular rector size, according to the collar vortex circular rector size at current time and using unsteady Bernoulli equation calculate obtain deformation gas
The size and distribution of aerodynamic force on dynamic surface grids；Step 5 keeps the circular rector of rear trailing vortex constant, by rear trailing vortex with locality
Flow field velocity moves the displacement of a time step to form downstream trailing vortex flow field；Step 6, on the basis in downstream trailing vortex flow field
On, the structural modal information under future time steplength is interpolated on the pneumatic surface grids of deformation, repeats step 4 and step 5, directly
To the solution for completing to promote unsteady aerodynamic force in time domain.
Using such configuration mode, a kind of quick unsteady aerodynamic force calculating side of time domain of coupled structure mode is provided
Method, this method by space to threedimensional pneumatic surface grids divide can the bending of accurate description pneumatic face, torsional deflection geometry
Feature, the Modeling Calculation suitable for aerodynamic force under the various movements of aircraft and deformation state；By arranging whirlpool on aerodynamic grid
Ring is solved using full Patential Flow Equation, and eliminate microvariations small deformation it is assumed that be suitable for the wider deformation of lifting surface
And the unsteady aerodynamic force of motion state calculates；By analyzing coupled structure mode of motion, it is applicable to arbitrarily determine
The rigid body and elastic movement mode of form, the calculating of frequency domain unsteady aerodynamic force is only applicable to simple harmonic quantity vibration compared with the existing technology
Dynamic structural elasticity mode greatly expands the scope of application of the unsteady aerodynamic force in conjunction with structure motion mode.Therefore, originally
The quick unsteady aerodynamic force calculation method of time domain provided by inventing calculates phase with time domain CFD unsteady aerodynamic force in the prior art
Than modeling is simple, and time domain promotes computational efficiency high, can adapt to contemporary aircraft design multimode, more round fast timedomains
The demand of simulation analysis.
In the present invention, for the calculating of the quick unsteady aerodynamic force of the time domain for realizing coupled structure mode, it is necessary first to
Grid dividing is carried out to the pneumatic face of the deformation of aircraft.Specifically, in the present invention, to the flight vehicle aerodynamic face analyzed of needs into
When row grid dividing, grid dividing standard can refer to the grid dividing requirement of panel method.Panel method is unsteady in aeroelasticity
The conventional means of aerodynamics evaluation generally requires grid downstream direction not interlock unanimously, and grid slenderness ratio is suitable.Pneumatic face
Grid should not it is too close can not be too thick.The number of tangential grid at least more than 5, open up to grid can according to grid cell slenderness ratio into
Row determines.In the present invention, pneumatic surface grids division is not limited to plane lifting surface situation, can be bent according to lifting surface, torsion
And aerofoil profile feature carries out the division of space threedimensional grid.
After completing aircraft and deforming the grid dividing in pneumatic face, need to obtain the structural modal information of aircraft simultaneously
Structural modal information is interpolated on the aerodynamic grid for deforming pneumatic face, determines the structure motion rule for deforming pneumatic face.Specifically
Ground, in the present invention, the structural modal meaning of aircraft can be aircraft overall structure mode, is also possible to than broad
The mode of the components such as wing or the mode oneself write.Structural modal source is unlimited, as long as a kind of movement in pneumatic face
Form, can be calculated by software or oneself is write.After the structural modal information of aircraft has been determined
Calculate the unsteady aerodynamic force under the movement situation with structure Coupling.
As a specific embodiment of the invention, the structural modal information of aircraft either elastic movement mode,
It is also possible to rigid motion mode.Structure motion mode is not limited to small deformation it is assumed that but super large amplitude can not occur herein
Movement, i.e. modal amplitudes cannot be too big, in order to avoid cause the complicated nonlinear effects such as airflow separation.Aircraft has been determined
Structural modal information is simultaneously interpolated into after aerodynamic grid by structural modal information, it is thus necessary to determine that deforms the structure motion rule in pneumatic face
Rule.In the present embodiment, it is regular as the structure motion for deforming pneumatic face that simple simple harmonic motion can be used, it is alternative
The characteristics of motion of other forms can also be used as the structure motion rule for deforming pneumatic face, herein with no restrictions in ground.
Further, structural modal information is being interpolated into the structure motion that aerodynamic grid and having determined deforms pneumatic face
After rule, the structure motion rule according to analysis object, calculating situation and the pneumatic face of deformation is needed, determines timedomain analysis meter
Steplength is calculated, that is, determines that unsteady time domain promotes the time span calculated.In the present invention, analysis object refers mainly to pneumatic face division
Sizing grid, calculating situation refers mainly to speed of incoming flow.In order to guarantee the accuracy of unsteady aerodynamic force calculating, when needing to guarantee
Between steplength cannot be differed too much with pneumatic surface grids size with the product of speed of incoming flow, i.e., domain analytical calculation steplength and calculate
The product of speed of incoming flow should be identical as pneumatic surface grids order of magnitude.
After timedomain analysis material calculation has been determined, the calculating of the unsteady aerodynamic force at current time can be carried out.Specifically
Collar vortex is arranged on each pneumatic surface grids in the present invention in ground, according to BiotSavart law (i.e. pacesetting formula)
Aerodynamic influence matrix is solved, is calculated and is obtained currently according to aerodynamic influence matrix and normal direction Solid boundary condition
The collar vortex circular rector size at moment is finally obtained according to the collar vortex circular rector size at current time and using the calculating of unsteady Bernoulli equation
Take the size and distribution for deforming the aerodynamic force on pneumatic surface grids.
Specifically, in the present invention, normal direction Solid boundary condition refers to the normal direction movement velocity of pneumatic surface grids, according to
The boundary condition that normal velocity is zero limits, and the movement of the induced velocity and curved surface itself washed under aerodynamic force and speed are zero, i.e.,Therefore, the size of the collar vortex circular rector Γ at current time can be according to aerodynamic influence matrix A
And normal direction Solid boundary conditionTo obtain, wherein A is wing aerodynamic power influence coefficient matrix,
Γ is the column vector that wing adheres to collar vortex intensity composition, and V (t)=[U (t), V (t), W (t)], V (t) are comprising wing motion speed
The speed of degree and unsteady speed of incoming flow, V_{w}For trailing vortex at the wing caused by induced velocity,For the normal vector of wing locality.
Further, after obtaining the collar vortex circular rector size at current time, the aerodynamic force on pneumatic surface grids is deformed
It can be according to the collar vortex circular rector size and the unsteady Bernoulli equation of utilization according to current timeTo obtain, wherein ρ is incoming flow
Density, τ_{i}For the tangent vector for deforming the direction pneumatic surface grids i, τ_{j}For the tangent vector for deforming the direction pneumatic surface grids j, Δ c_{ij}To become
The pneumatic face of shape arranges the chord length to jth ith, Zhan Xiang of grid, Δ b_{ij}It arranges to deform pneumatic face to jth ith, Zhan Xiang of grid
Length, Γ_{ij}The collar vortex circular rector to jth ith, Zhan Xiang of grid, Γ are arranged to deform pneumatic face_{i1,j}To deform pneumatic face column
To (i1)th, the collar vortex circular rector of jth of grid of Zhan Xiang, Γ_{i,j1}For deform pneumatic face arrange to ith, open up to 1 grid of jth
Collar vortex circular rector, Δ p_{ij}The aerodynamic force to jth ith, Zhan Xiang of grid is arranged to deform pneumatic face.It can be completed as a result, current
The calculating of moment unsteady aerodynamic force.
After the calculating for completing current time unsteady aerodynamic force, keep the circular rector of rear trailing vortex constant, by rear
Trailing vortex moves the displacement of a time step with local airflow speed to form downstream trailing vortex flow field.The step 5 is mainly used for shape
At downstream trailing vortex collar vortex, exist due to during the timedomain calculation of unsteady aerodynamic force, needing to acquire each moment trailing vortex respectively
Induced velocity caused by wing, it is therefore desirable to obtain the trailing vortex collar vortex at each moment.Specifically, in the present invention, entire stream
The collar vortex of field is divided into two parts, as shown in figure 3, a part of collar vortex is arranged on wing, which is attachment collar vortex, attached
The least significant end of collar vortex be rear trailing vortex, another part collar vortex is arranged in wing catchment, which is downstream trailing vortex collar vortex,
In, downstream trailing vortex collar vortex is gradually formed in time stepping method by the abjection of rear trailing vortex.
Further, on the basis of the trailing vortex flow field of downstream, the structural modal information under future time steplength is interpolated into
It deforms on pneumatic surface grids, keeps trailing vortex collar vortex in downstream constant, repeat step 4 and step 5, until completing to promote in time domain non
The solution of Unsteady Flow.
Further understand to have to the present invention, below with reference to Fig. 1 to Fig. 8 to coupled structure mode of the invention when
The quick unsteady aerodynamic force calculation method in domain is described in detail.
As shown in Figures 1 to 8, the time domain for providing a kind of coupled structure mode according to a particular embodiment of the invention is quick
Unsteady aerodynamic force calculation method, in the present embodiment, by taking a high aspect ratio rectangular wing as an example, to coupling knot of the invention
The time domain unsteady aerodynamic force calculation method of structure mode is introduced.Airfoil chord a length of 60mm, length 480mm calculate incoming flow
30m/s, 2 degree of angles of attack.Since example is high aspect ratio wing, using symmetrical airfoil, do not consider that camber influences, wing is
It is deformed into space curved surface under load effect, as shown in Figure 4.Only method applicable cases are illustrated herein, therefore select example
Relatively simple, pneumatic face divides more coarse.Practical application be should be carried out according to the requirement of technical solution it is more careful pneumatic
Face divides and modeling.The quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode provided by the present embodiment is specifically wrapped
Include following steps.
Step 1 carries out grid dividing, grid of the grid dividing standard referring to panel method to the pneumatic face of the deformation of aircraft
It divides and requires, it is desirable that grid downstream direction does not interlock unanimously, and grid slenderness ratio is suitable.Pneumatic surface grids should not it is too close can not
It is too thick.At least more than 5, opening up can be determined the number of tangential grid to grid according to grid cell slenderness ratio.
Step 2 obtains the structural modal information of aircraft and structural modal information is interpolated into the pneumatic of the pneumatic face of deformation
On grid, the structure motion rule for deforming pneumatic face is determined.In the present embodiment, as shown in figure 5, a kind of torsion of wing is transported
Movement is structure Coupling mode of motion, and structure Coupling mode of motion is interpolated on the aerodynamic grid for deforming pneumatic face.Deformation
The structure motion rule in pneumatic face is simple simple harmonic motion.
Step 3, according to aerodynamic grid size, speed of incoming flow and the structure motion in the pneumatic face of deformation for deforming pneumatic face
Rule determines timedomain analysis material calculation.In the present embodiment, timedomain analysis material calculation is 0.001s.
Step 4 is based on full potential flow theory, collar vortex is arranged on each aerodynamic grid for deforming pneumatic face, according to speed
Induced formulas solves the aerodynamic influence matrix of aircraft, according to aerodynamic influence matrix and normal direction object plane boundary
Condition calculates the collar vortex circular rector size for obtaining current time, is exerted according to the collar vortex circular rector size at current time and using unsteady uncle
Sharp equation calculation obtains the size and distribution for deforming the aerodynamic force on pneumatic surface grids.
Step 5 keeps the circular rector of rear trailing vortex constant, and rear trailing vortex is moved a time step with local airflow speed
Long displacement is to form downstream trailing vortex flow field.
Step 6, on the basis of the trailing vortex flow field of downstream, one time step of every propulsion, by the knot under future time steplength
Structure modal information is interpolated on wing, is updated wing and is deformed pneumatic surface grids, solves the aerodynamic influence coefficient after updating configuration
Matrix calculates the collar vortex circular rector size for obtaining current time using normal direction Solid boundary condition, according to the collar vortex ring at current time
It measures size and is calculated using unsteady Bernoulli equation and obtain the size and distribution that deform the aerodynamic force on pneumatic surface grids.Later
It keeps its circular rector constant rear trailing vortex, the displacement of a time step is moved with local airflow speed, form downstream vortex wake
?.It is similar, step 4 and step 5 are constantly repeated, until completing the solution for promoting unsteady aerodynamic force in time domain.
In conclusion the present invention provides a kind of quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode,
This method is coupled with mode of motion known to structure, is quickly analyzed the unsteady aerodynamic force under moving known to aircraft, and
Numerical simulation calculated result is intuitively provided in time domain scale, convenient for designer to the visual understanding of unsteady phenomenon and non
The accurate understanding of Unsteady Flow lays the foundation for the response analysis of time domain aeroelasticity, timedomain stability analysis.In addition, this hair
The bright division by threedimensional pneumatic surface grids not only can be with the geometrical feature of the pneumatic face bending of accurate description, torsional deflection, base
The limitation for breaching traditional unsteady aerodynamic force microvariations small deformation and assuming is solved in the whole flow field domain of full Patential Flow Equation, can be fitted
The solution of unsteady aerodynamic force for arbitrary motion form (except the strong nonlinearities situation such as big angle of attack separation, stall).Quickly
Time domain method for solving convenient for intuitively recognize and examine unsteady aerodynamic force calculating process and as a result, be conducive to it is multidisciplinary,
Multiduty aircraft unsteady regime timedomainsimulation analytical calculation.
Therefore, the quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode provided by the present invention has been broken existing
Stage aeroelasticity analysis project unsteady aerodynamic force calculates the limitation for being often confined to frequency domain, and structure motion mode is also no longer limited to
In small deformation range, the scope of application of unsteady aerodynamic force calculating is expanded.It is calculated with conventional Timedomain CFD unsteady aerodynamic force
It compares, unsteady aerodynamic force calculation method modeling proposed by the invention is simple, and time domain promotes computational efficiency high, has adapted to the modern times
The demand of Flight Vehicle Design multimode, more round fast timedomain simulation analysis.
For ease of description, spatially relative term can be used herein, as " ... on ", " ... top ",
" ... upper surface ", " above " etc., for describing such as a device shown in the figure or feature and other devices or spy
The spatial relation of sign.It should be understood that spatially relative term is intended to comprising the orientation in addition to device described in figure
Except different direction in use or operation.For example, being described as if the device in attached drawing is squeezed " in other devices
It will be positioned as " under other devices or construction after part or construction top " or the device of " on other devices or construction "
Side " or " under other devices or construction ".Thus, exemplary term " ... top " may include " ... top " and
" in ... lower section " two kinds of orientation.The device can also be positioned with other different modes and (is rotated by 90 ° or in other orientation), and
And respective explanations are made to the opposite description in space used herein above.
In addition, it should be noted that, limiting components using the words such as " first ", " second ", it is only for be convenient for
Corresponding components are distinguished, do not have Stated otherwise such as, there is no particular meanings for abovementioned word, therefore should not be understood as to this
The limitation of invention protection scope.
The foregoing is only a preferred embodiment of the present invention, is not intended to restrict the invention, for the skill of this field
For art personnel, the invention may be variously modified and varied.All within the spirits and principles of the present invention, made any to repair
Change, equivalent replacement, improvement etc., should all be included in the protection scope of the present invention.
Claims (6)
1. a kind of quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode, which is characterized in that the time domain is quick
Unsteady aerodynamic force calculation method includes:
Step 1 carries out grid dividing to the pneumatic face of the deformation of aircraft；
Step 2 obtains the structural modal information of aircraft and the structural modal information is interpolated into the pneumatic of the pneumatic face of deformation
On grid, the structure motion rule for deforming pneumatic face is determined；
Step 3, according to the structure of the aerodynamic grid size in the pneumatic face of deformation, speed of incoming flow and the pneumatic face of deformation
The characteristics of motion determines timedomain analysis material calculation；
Step 4 is based on full potential flow theory, collar vortex is arranged on each aerodynamic grid in the pneumatic face of deformation, according to speed
Induced formulas solves the aerodynamic influence matrix of aircraft, according to the aerodynamic influence matrix and normal direction object plane
Boundary condition calculates the collar vortex circular rector size for obtaining current time, according to the collar vortex circular rector size at current time and utilizes unsteady
Bernoulli equation calculates the size and distribution for obtaining and deforming the aerodynamic force on pneumatic surface grids；
Step 5 keeps the circular rector of rear trailing vortex constant, and the rear trailing vortex is moved a time step with local airflow speed
Long displacement is to form downstream trailing vortex flow field；
Structural modal information under future time steplength is interpolated into change on the basis of the downstream trailing vortex flow field by step 6
On the pneumatic surface grids of shape, step 4 and step 5 are repeated, until completing the solution for promoting unsteady aerodynamic force in time domain.
2. the quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode according to claim 1, feature exist
In the size of the collar vortex circular rector Γ at current time can be according to the aerodynamic influence matrix A and normal direction object plane side
Boundary's conditionTo obtain, wherein A is wing aerodynamic power influence coefficient matrix, and Γ is that wing adheres to whirlpool
The column vector of ring intensity composition, V (t)=[U (t), V (t), W (t)], V (t) are to include wing motion speed and unsteady incoming flow
The speed of speed, V_{w}For trailing vortex at the wing caused by induced velocity,For the normal vector of wing locality.
3. the quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode according to claim 2, feature exist
In the aerodynamic force deformed on pneumatic surface grids can be according to the collar vortex circular rector size according to current time and the unsteady Bernoulli Jacob of utilization
EquationTo obtain, wherein ρ is
Come current density, τ_{i}For the tangent vector for deforming the direction pneumatic surface grids i, τ_{j}For the tangent vector for deforming the direction pneumatic surface grids j, Δ c_{ij}
The chord length to jth ith, Zhan Xiang of grid, Δ b are arranged to deform pneumatic face_{ij}It arranges to deform pneumatic face to ith, jth of Zhan Xiang
The length of grid, Γ_{ij}The collar vortex circular rector to jth ith, Zhan Xiang of grid, Γ are arranged to deform pneumatic face_{i1,j}It is pneumatic to deform
Face arranges the collar vortex circular rector to jth (i1)th, Zhan Xiang of grid, Γ_{i,j1}For deform pneumatic face arrange to ith, open up to jth 1
The collar vortex circular rector of grid, Δ p_{ij}The aerodynamic force to jth ith, Zhan Xiang of grid is arranged to deform pneumatic face.
4. the quick unsteady aerodynamic force calculating side of time domain of coupled structure mode according to any one of claim 1 to 3
Method, which is characterized in that in said step 1, for the tangential grid in pneumatic face at least more than 5, opening up can root to grid after the deformation
It is determined according to grid cell slenderness ratio.
5. the quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode according to claim 4, feature exist
In in the step 2, the structural modal information of the aircraft includes rigid motion mode or elastic movement mode.
6. the quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode according to claim 5, feature exist
In the structure motion rule in the pneumatic face of deformation includes simple harmonic motion.
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Citations (7)
Publication number  Priority date  Publication date  Assignee  Title 

CN102012953A (en) *  20101104  20110413  西北工业大学  CFD (computational fluid dynamics)/CSD (circuit switch data) coupled solving nonlinear aeroelasticity simulation method 
CN102866637A (en) *  20121007  20130109  西北工业大学  Quadratic orderreduction based method for simulating unsteady aerodynamic force of aerofoil with operation surface 
CN106508028B (en) *  20100930  20140702  上海机电工程研究所  A kind of determination complex appearance aircraft supersonic speed, the hypersonic method for having angle of attack tremor secure border 
CN105046021A (en) *  20150825  20151111  西北工业大学  Nonlinear optimization algorithm for rational approximation of unsteady aerodynamic minimum state 
CN105843073A (en) *  20160323  20160810  北京航空航天大学  Method for analyzing wing structure aeroelasticity stability based on aerodynamic force uncertain order reduction 
US20160312765A1 (en) *  20150423  20161027  Continuum Dynamics, Inc.  Vertical axis liftdriven wind turbine with force canceling blade configuration 
CN109459206A (en) *  20181217  20190312  西北工业大学  Ground experiment unsteady aerodynamic force loading method 

2019
 20190319 CN CN201910206453.2A patent/CN110162822B/en active Active
Patent Citations (7)
Publication number  Priority date  Publication date  Assignee  Title 

CN106508028B (en) *  20100930  20140702  上海机电工程研究所  A kind of determination complex appearance aircraft supersonic speed, the hypersonic method for having angle of attack tremor secure border 
CN102012953A (en) *  20101104  20110413  西北工业大学  CFD (computational fluid dynamics)/CSD (circuit switch data) coupled solving nonlinear aeroelasticity simulation method 
CN102866637A (en) *  20121007  20130109  西北工业大学  Quadratic orderreduction based method for simulating unsteady aerodynamic force of aerofoil with operation surface 
US20160312765A1 (en) *  20150423  20161027  Continuum Dynamics, Inc.  Vertical axis liftdriven wind turbine with force canceling blade configuration 
CN105046021A (en) *  20150825  20151111  西北工业大学  Nonlinear optimization algorithm for rational approximation of unsteady aerodynamic minimum state 
CN105843073A (en) *  20160323  20160810  北京航空航天大学  Method for analyzing wing structure aeroelasticity stability based on aerodynamic force uncertain order reduction 
CN109459206A (en) *  20181217  20190312  西北工业大学  Ground experiment unsteady aerodynamic force loading method 
NonPatent Citations (4)
Title 

USMAN ASGHAR 等: "Modelling and simulation of flow induced vibrations in vertical axis wind turbine blade", 《2017 14TH INTERNATIONAL BHURBAN CONFERENCE ON APPLIED SCIENCES AND TECHNOLOGY (IBCAST)》 * 
XIE CHANGCHUAN等: "Static aeroelastic analysis including geometric nonlinearities based on reduced order model", 《CHINESE JOURNAL OF AERONAUTICS》 * 
姚伟刚 等: "基于特征正交分解的非定常气动力建模技术", 《力学学报》 * 
张红波 等: "弹性飞机尾涡遭遇动响应分析方法", 《航空科学技术》 * 
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