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 PDF

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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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pneumatic
grid
unsteady
aerodynamic force
time domain
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CN110162822B (en
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刘燚
刘凯
操小龙
何海波
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Beijing Research Institute of Mechanical and Electrical Technology
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Beijing Research Institute of Mechanical and Electrical Technology
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    • G06F17/5009
    • G06F17/5095
    • G06F2217/78
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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 time-domain 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 step-length 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

The quick unsteady aerodynamic force calculation method of the time domain of coupled structure mode
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 motor-driven 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 time-domain analysis to be carried out relatively complicated Conversion, also to undertake time-frequency domain conversion modeling bring error.With contemporary aircraft high speed, lightweight, flexible etc. Growth requirement, the time-domain analysis of multidisciplinary synthesis become the inevitable approach of aircraft analysis and design, this is just to unsteady pneumatic Power proposes fast and accurately time-domain 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 time-domain 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 step-length 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, VwFor 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, τiFor the tangent vector for deforming the direction pneumatic surface grids i, τjTo deform the direction pneumatic surface grids j Tangent vector, Δ cijThe chord length to j-th i-th, Zhan Xiang of grid, Δ b are arranged to deform pneumatic faceijFor deform pneumatic face arrange to I-th, the length of j-th of grid of Zhan Xiang, ΓijThe collar vortex circular rector to j-th i-th, Zhan Xiang of grid is arranged to deform pneumatic face, Γi-1,jThe collar vortex circular rector to j-th (i-1)-th, Zhan Xiang of grid, Γ are arranged to deform pneumatic facei,j-1It arranges to deform pneumatic face to the I, collar vortex circular rector to -1 grid of jth, Δ p are opened upijFor deform pneumatic face arrange to i-th, j-th 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 three-dimensional 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 three-dimension 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 1-cos 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 time-domain 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 step-length 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 three-dimensional 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 time-domains 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 three-dimensional 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 air-flow 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 time-domain analysis meter Step-length 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 step-length cannot be differed too much with pneumatic surface grids size with the product of speed of incoming flow, i.e., domain analytical calculation step-length and calculate The product of speed of incoming flow should be identical as pneumatic surface grids order of magnitude.
After time-domain 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 Biot-Savart 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, VwFor 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, τiFor the tangent vector for deforming the direction pneumatic surface grids i, τjFor the tangent vector for deforming the direction pneumatic surface grids j, Δ cijTo become The pneumatic face of shape arranges the chord length to j-th i-th, Zhan Xiang of grid, Δ bijIt arranges to deform pneumatic face to j-th i-th, Zhan Xiang of grid Length, ΓijThe collar vortex circular rector to j-th i-th, Zhan Xiang of grid, Γ are arranged to deform pneumatic facei-1,jTo deform pneumatic face column To (i-1)-th, the collar vortex circular rector of j-th of grid of Zhan Xiang, Γi,j-1For deform pneumatic face arrange to i-th, open up to -1 grid of jth Collar vortex circular rector, Δ pijThe aerodynamic force to j-th i-th, 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 time-domain 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 step-length 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 time-domain analysis material calculation.In the present embodiment, time-domain 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 step-length 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, time-domain stability analysis.In addition, this hair The bright division by three-dimensional 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 time-domain-simulation 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 Time-domain 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 time-domain 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 above-mentioned 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 time-domain 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 step-length 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, VwFor 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, τiFor the tangent vector for deforming the direction pneumatic surface grids i, τjFor the tangent vector for deforming the direction pneumatic surface grids j, Δ cij The chord length to j-th i-th, Zhan Xiang of grid, Δ b are arranged to deform pneumatic faceijIt arranges to deform pneumatic face to i-th, j-th of Zhan Xiang The length of grid, ΓijThe collar vortex circular rector to j-th i-th, Zhan Xiang of grid, Γ are arranged to deform pneumatic facei-1,jIt is pneumatic to deform Face arranges the collar vortex circular rector to j-th (i-1)-th, Zhan Xiang of grid, Γi,j-1For deform pneumatic face arrange to i-th, open up to jth -1 The collar vortex circular rector of grid, Δ pijThe aerodynamic force to j-th i-th, 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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