CN103077259A - Hypersonic speed guided missile multi-field coupling dynamics integrated simulation analysis method - Google Patents

Abstract

The invention provides a hypersonic speed guided missile multi-field coupling dynamics integrated simulation analysis method, which comprises the following steps of: establishing a coupled structural motion equation; improving an aerodynamic heating calculation method; improving an unsteady aerodynamics calculation method; establishing an integrated simulation analysis technique process; designing a relationship between an integrated simulation analysis operation model and data; and realizing an integrated simulation analysis system. Compared with the prior art, the aerodynamic heating calculation method and the unsteady aerodynamics calculation method are improved; various techniques of flutter analysis, thermal flutter analysis, pneumatic servo elastic analysis and thermal pneumatic servo elastic analysis are integrated; and an effective integration method for integrally solving and optimally designing a high-speed winged missile all-missile combination body coupling dynamics problem is provided.

Description

The integrated simulating analysis of many Coupled Dynamics of hypersonic missile

Technical field

The present invention relates to the Missile Preliminary technology, specifically belong to a kind of hypersonic missile and consider aeroelasticity and the pneumatic servo Analyse dlasto method of Aerodynamic Heating impact, realize the complex appearance guided missile is carried out Aerodynamic Heating, three-dimensional temperature field, hot-die attitude, non-Unsteady Flow, heat flutter and the integrated simulation analysis of pneumatic servo elastic stability.

Background technology

Guided missile is when high speed and high maneuvering flight, can bring serious Aerodynamic Heating and aeroelasticity coupling, and servo-control system and the coupling of body elastic deformation, have a strong impact on the aeroelastic divergence stability of guided missile, cause sudden and catastrophic body fracture accident, be exist in the high-speed missile integrated design, must add the problem of giving solution.

The design analysis of aeroelasticity and pneumatic servo elastic stability mainly adopts numerical value emulation method.In traditional analytical approach, the general Coupled Interaction Analysis that only carries out the simple aeroelasticity of parts or assembly or aeroelastic divergence Aerodynamic-structure considers that the aerothermoelastic analysis of Aerodynamic Heating also just carries out respectively Aerodynamic Heating, hot-die attitude and heat flutter analysis.Although some method and software can be finished some factor coupling analysis, such as Aerodynamic Heating and structure temperature field, thermal stress, hot strength coupling analysis, temperature field and hot-die attitude, heat flutter coupling analysis etc. consider Aerodynamic Heating under the hypersonic and complex appearance condition, structure temperature field, thermal stress, hot-die attitude, aeroelasticity and and many multifactor coupling integrated simulation analysis problems considering the servo-control system coupling but fail to solve.

Summary of the invention

The object of the present invention is to provide the integrated simulating analysis of many Coupled Dynamics of hypersonic missile, comprehensively solve aeroelasticity and pneumatic servo elastic stability problem.

Achieve the above object, the present invention realizes by following technical scheme.The integrated simulating analysis of many Coupled Dynamics of hypersonic missile comprises the steps:

1) sets up the structure motion equation that is coupled

Utilize mode superposition method or branch mode method to set up the full quality battle array of assembly and the differential equation of motion of Stiffness Matrix decoupling zero or accurate decoupling zero of playing, the state space equation of domain space carries out simulation analysis in the time domain upper integral and by the flow process of closed-loop path when using state-space method that the transport function of control system is changed into.

2) improve the calculation of aerodynamic heating method

From solving the demand of the wing-body-tail assembly complex appearance calculation of aerodynamic heating, developed the method that the CFD numerical simulation combines with engineering calculation, finding the solution the Euler equation without the viscosity flow Flow Field Numerical and determine boundary layer outer rim parameter the outside, then adopt reference enthalpy method gauging surface hot-fluid in the active inside boundary of viscosity, both overcome the difficulty that Engineering Algorithm is difficult to find the solution complex appearance boundary layer outer rim parameter, and avoided again numerical solution Navier-Stokes equation to need the problem of huge calculated amount and computational resource.

3) improve non-Unsteady Flow computing method

The technology solution that employing CFD numerical simulation and local stream piston theory combine by no means permanent power is calculated a difficult problem.For the Arbitrary 3 D object, suppose that air-flow is along the either direction of object plane tangent plane, set up the lower expression formula of washing that causes perpendicular to the object plane vibration at tangent plane, utilize the geometric relationship between practical structures direction of vibration and the object plane outer normal direction, set up the expression formula that local stream piston theory calculates the non-Unsteady Flow of random appearance, and the locality of complex appearance stream obtains by the CFD method of finding the solution the Euler equation, and the large angle of attack, assembly flow field are disturbed, washed down etc. and all find the solution acquisition in the stream of locality.

4) set up integrated simulation analysis techniqueflow

For integrated hypersonic aeroelasticity mathematical model and each single technology, form the technical Analysis flow process of many integrated simulation analysis systems of Coupled Dynamics, provided realization approach and the mutual relationship of each individual event functional module and whole integrated analysis platform from the angle of integrated many Coupled Dynamics simulation analysis, comprised that FLUTTER CALCULATION (analysis of structure aeroelastic stability), heat flutter, aeroelastic divergence analysis and hot gas move servo flexibility analysis four partial contents.

5) design integration simulation analysis job model and data relationship

For specific implementation step 4) shown in four functions, job model and the data relationship of integrated simulation analysis system have been designed, comprise modal calculation operation module, non-Unsteady Flow computational tasks module, local airflow computational tasks module, calculation of aerodynamic heating operation module, temperature field, thermal stress and hot-die attitude computational tasks module, heat shield ablation Calculation operation module, control computational tasks module and Coupled Dynamics emulation and critical parameters identification operation module, and major function and the data relationship of each operation are described.

6) realize integrated simulation analysis system

According to step 4) techniqueflow and the step 5 of design) the operation module of design, set up data structure diagram and the database structure figure of simulation analysis system; Business software technology and self-compiling program that integrated whole integral method is required are finished separate part or are entirely played assembly in different Mach number and the moving servo elastokinetics emulation of flutter, heat flutter analysis, aeroelastic divergence dynamics simulation and hot gas under the different angle of attack conditions.

The inventive method compared with prior art, its advantage and beneficial effect are: improve Aerodynamic Heating and non-Unsteady Flow computing method, integrated flutter analysis, heat flutter analysis, aeroelastic divergence analysis and hot gas move the every technology of servo flexibility analysis, provide integrated and have found the solution and optimal design high speed aerodynamic missile plays the effective integration method of the Coupled Dynamics problem of assembly entirely.Compare result of calculation with wind tunnel experiment correctly reliable.Compare with pure CFD numerical method, in the situation that satisfy accuracy requirement, counting yield improves more than 100 times.

Description of drawings

Fig. 1 is the integrated simulation analysis flow chart of steps of many Coupled Dynamics of the present invention;

Fig. 2 is the integrated simulation analysis techniqueflow charts of many Coupled Dynamics of full bullet of the present invention;

Fig. 3 is full bullet of the present invention many Coupled Dynamics integrated simulation analysis system operation module and entity relationship diagram;

Fig. 4 is many Coupled Dynamics integrated simulation systems of full bullet of the present invention data structure diagram;

Fig. 5 is the integrated simulation data base structural drawing of many Coupled Dynamics of full bullet of the present invention;

Fig. 6 (a) is the full bullet assembly of the present invention critical dynamic pressure of flutter comparison diagram as a result under a=0 ° of condition;

Fig. 6 (b) is the full bullet assembly of the present invention critical dynamic pressure of flutter comparison diagram as a result under the Ma=4 condition;

Fig. 7 (a) is the bomb body displacement response diagram of full bullet assembly of the present invention flutter;

Fig. 7 (b) be full bullet assembly of the present invention flutter for missile wing displacement response diagram;

Fig. 7 (c) is the empennage displacement response diagram of full bullet assembly of the present invention flutter;

Fig. 8 (a) is as a result figure of full bullet assembly aluminum structure of the present invention heat flutter;

Fig. 8 (b) is the heat flutter figure as a result of full bullet assembly titanium alloy structure of the present invention.

Embodiment

The invention will be further described below with reference to drawings and Examples.

As shown in Figure 1, the integrated simulation analysis step of many Coupled Dynamics of the present invention is as follows:

1) sets up the structure motion equation that is coupled

The high speed aerodynamic missile physical model that limited consideration is comprised of bomb body, missile wing and empennage when carrying out structural dynamic Epidemiological Analysis or emulation, utilizes mode superposition method or the branch mode method can be to the differential equation of motion of quality battle array and Stiffness Matrix decoupling zero or accurate decoupling zero usually

$\left[M\right]\left\{\stackrel{\·\·}{\mathrm{\ξ}}\right\}+\left[C\right]\left\{\stackrel{\·}{\mathrm{\ξ}}\right\}+\left[K\right]\left\{\mathrm{\ξ}\right\}=\left\{Q_A\right\}+\left\{Q_C\right\}---\left(1\right)$

[M], [C] and [K] are respectively modal mass, modal damping and the modal stiffness matrix of structure in the formula;

{ Q _AThe broad sense aerodynamic force array that produces for elastic vibration:

$\left\{Q_A\right\}=\underset{\mathrm{\σ}}{\∫\∫}{\left[\mathrm{\Φ}\right]}^T\mathrm{\Δ}{p}_A(x,y,t)\mathrm{d\σ}---\left(2\right)$

{ Q _CBe the Generalized Control power array of rudder oblique presentation life:

$\left\{Q_C\right\}=\underset{r}{\∫\∫}{\left[\mathrm{\Φ}\right]}^T\mathrm{\Δ}{p}_C(x,y,t)\mathrm{d\σ}---\left(3\right)$

Limit of integration σ represents the whole surperficial integration to aircraft in the formula; R represents the rudder face integration.

Generally speaking, { Q _AAnd { Q _CCan be directly or write as the form of identical with equation (1) left side Second Order Differential Operator about modal coordinate through depression of order and conversion, namely

$\left\{Q\right\}=\left[A\right]\left\{\stackrel{\·\·}{\mathrm{\ξ}}\right\}+\left[B\right]\left\{\stackrel{\·}{\mathrm{\ξ}}\right\}+\left[C\right]\left\{\mathrm{\ξ}\right\}---\left(4\right)$

[A], [B] and [C] are the aerodynamic coefficient matrix in the formula, generally are full battle array, independently are exactly mutually by the coupling of aerodynamic coefficient matrix cross term between each mode degree of freedom, realize inspiring each other through aerodynamic force.In addition, control system is also experienced structural vibration by sensitive element, and to export rudder inclined to one side, and the relation table between its input and output is shown

$\mathrm{\α}\left(t\right)=f\left(\stackrel{\·}{\mathrm{\φ}}\right),\stackrel{\·}{\mathrm{\φ}}=[{{\mathrm{\Φ}}_b}^{\′}\left(x_m\right),\mathrm{0,0}]\left\{\stackrel{\·}{\mathrm{\ξ}}\right\}---\left(5\right)$

Function f is the control law of control system, and the transport function by frequency domain represents usually.Formula (5) has formed the closed-loop path relation of inputting each other, exporting between control system and the structure motion equation.The transport function of equation (1) and formula (4) all can be converted to the state space equation that is become by first-order ordinary differential equation system, carry out simulation analysis in the time domain upper integral and by the flow process of closed-loop path.

The structure motion equation that represents in order to set up modal coordinate needs the modal parameter of Flight Vehicle Structure.For the low mach situation of not considering Aerodynamic Heating, modal parameter [M], [C] and [K] directly obtain by FEM (finite element) calculation or by the modal test of routine.For hypersonic situation, for obtaining the hot-die attitude under the Aerodynamic Heating environment, must carry out calculation of aerodynamic heating along trajectory, take each constantly the heat flow density of Aerodynamic Heating as initial conditions, from initial time, find the solution next transient state temperature field constantly every a time step, then calculate thermal stress and hot-die attitude according to the temperature field.Next is broad sense aerodynamic force and the Generalized Control power on equation the right, is the integration of the product of time-dependent pressure distribution and Mode Shape.Analyze above-mentioned needed each single technology of the equation of motion of setting up, except finite element modal analysis, Analysis of Three-Dimensional Temperature, thermal stress and the analysis of hot-die attitude all are the technology of comparative maturity, remaining calculation of aerodynamic heating and non-Unsteady Flow calculate needs self-developing.

2) improve the calculation of aerodynamic heating method

The method that improved heating Aerodynamic Heating computing method have adopted the CFD numerical simulation to combine with engineering calculation, finding the solution the Euler equation without the viscosity flow Flow Field Numerical and determine boundary layer outer rim parameter the outside, then adopt reference enthalpy method gauging surface hot-fluid in the active inside boundary of viscosity, both overcome the difficulty that Engineering Algorithm is difficult to find the solution complex appearance boundary layer outer rim parameter, and avoided again numerical solution Navier-Stokes equation to need the problem of huge calculated amount and computational resource.When hypersonic three-dimensional flow field outer rim calculation of parameter, the Eulerian equation numerical evaluation adopts fluid mechanics business software FLUENT to calculate, the spatial spreading form adopts Second-order Up-wind Roe and AUSM form, adopt in time the alternating direction implicit solving method, after calculating convergence, from FLUENT, derive respectively surface coordinate and the aerodynamic parameter of bomb body and lifting surface.In inside boundary, utilize reference enthalpy Aerodynamic Heating method gauging surface hot-fluid, and with the link of computer programs of finding the solution One-dimensional Heat Conduction Equation, calculate body surface heat flux and temperature.

3) improve non-Unsteady Flow computing method

The technology that improved non-Unsteady Flow computing method have adopted the CFD numerical simulation to combine with local stream piston theory.For the Arbitrary 3 D object, suppose that air-flow is along the either direction of object plane tangent plane, set up the lower expression formula of washing that causes perpendicular to the object plane vibration at tangent plane, utilize the geometric relationship between practical structures direction of vibration and the object plane outer normal direction, set up the expression formula that local stream piston theory calculates the non-Unsteady Flow of random appearance, and the locality of complex appearance stream obtains by the CFD method of finding the solution the Euler equation, and the large angle of attack, assembly flow field are disturbed, washed down etc. and all reflected in the stream of locality.Namely in formula (2), non-steady pressure distributes and uses following local stream piston theory to calculate:

Δp(x，θ，t)＝Δp _b(x，θ，t)+Δp _w(x，θ，t)+Δp _t(x，θ，t) (6)

The non-steady pressure distribution of bomb body computing formula:

$\mathrm{\Δ}{p}_b(x,\mathrm{\θ},t)=\frac{2q_L}{{M_L}^2{c}_L}{\mathrm{\μ}}_b\left(x\right)\sin \mathrm{\θ}\·(\frac{\∂}{\∂t}+M_{\mathrm{xL}}{c}_L\frac{\∂}{\∂x})Z(x,t)---\left(7\right)$

The non-steady pressure distribution of missile wing computing formula:

$\mathrm{\Δ}{p}_w(x,y,t)=\±\frac{{2q}_L{\mathrm{\μ}}_w}{M_L^2{c}_L}(\frac{\∂}{\∂t}+M_{\mathrm{xL}}{c}_L\frac{\∂}{\∂x}+M_{\mathrm{yL}}{c}_L\frac{\∂}{\∂y})Z(x,y,t)---\left(8\right)$

The non-steady pressure of empennage distributes and comprises two parts, induces non-permanent lower gas washing power two parts of generation to be comprised of aerodynamic force and missile wing trailing vortex that the structural vibration of empennage own produces at empennage:

$\mathrm{\Δ}{P}_t(x,y,t)=\mathrm{\Δ}{P}_t^{\left(t_0\right)}(x,y,t)+\mathrm{\Δ}{P}_t^{\left(d\right)}(x,y,t)---\left(9\right)$

Wherein

$\mathrm{\Δ}{p}_t^{\left(t_0\right)}(x,y,t)=\±\frac{2q_{\∞}}{M_{\∞}}{\mathrm{\μ}}_t(\frac{1}{M_{\∞}{c}_{\∞}}{B}_t(x,y)\frac{\∂}{\∂t}+C_{\mathrm{tX}}(x,y)\frac{\∂}{\∂x})Z(x,y,t)---\left(10\right)$

$\mathrm{\Δ}{p}_t^{\left(d\right)}(x,y,t)=\stackrel{\‾}{+}\frac{q_{\∞}}{2M_{\∞}^2}{k}_f^{\left(d\right)}\cos \mathrm{\α}\·{\mathrm{\μ}}_t{B}_t(x,y).$

$\left.\begin{array}{c}B(x,\mathrm{\θ})={(p_L/p_{\∞}\·{\mathrm{\ρ}}_L/{\mathrm{\ρ}}_{\∞})}^{1/2}\\ C(x,\mathrm{\θ})=M_{\mathrm{xL}}/M_{\∞}\·p_L/p_{\∞}\end{array}\right\}---\left(12\right)$

In above-mentioned formula, μ is that body surface outer normal is to the direction cosine of z axle; Q is dynamic pressure; C is the velocity of sound; P is pressure distribution; ρ is Density Distribution; M is Mach number; B and C are the dimensionless Aerodynamic Coefficient; Φ is each branch vibration shape; τ washes time delay under being.Subscript L represents local stream, and subscript ∞ represents the infinite distance incoming flow.So far, structure motion equation (1) can be found the solution fully.

4) set up integrated simulation analysis techniqueflow

As shown in Figure 2, for integrated hypersonic aeroelasticity mathematical model and each single technology, set up the technical Analysis flow process of many integrated simulation analysis systems of Coupled Dynamics.It has provided realization approach and the mutual relationship of each individual event functional module and whole integrated analysis platform from the angle of integrated many Coupled Dynamics simulation analysis.In Fig. 2, integrated Coupled Dynamics comprises that specifically FLUTTER CALCULATION (analysis of structure aeroelastic stability), heat flutter, aeroelastic divergence analysis and hot gas move servo flexibility analysis four partial contents.When carrying out concrete Dynamics Simulation Analysis, can select corresponding step and flow process according to this figure.Detailed process is as follows:

When carrying out flutter analysis, need the formal parameter of guided missile and trajectory parameter to calculate local stream parameter, recycling step 3) the Computation of Unsteady aerodynamic force of improving one's methods; On the other hand, obtain the modal parameters such as the vibration shape of structure and frequency by FEM (finite element) calculation or test; Then, utilize step 1) mathematical model carry out the aeroelastic stability analysis.

When carrying out heat flutter calculating, the process that aerodynamic force calculates link and FLUTTER CALCULATION is similar, Aerodynamic Heating calculates yet modal parameters then needs to be coupled, namely utilize step 2) Innovative method calculating Aerodynamic Heating environment, find the solution the guided missile three-dimensional temperature field, thermal stress and hot-die attitude by the FEM (finite element) calculation body structure recycle step 1 at last) mathematical model consider aerothermal aeroelastic stability analysis.Thermal effect and trajectory time correlation are calculated and need to be become along the flutter examining computation of trajectory time from the flutter examining computation of single state of flight so carry out heat flutter.

When carrying out the aeroelastic divergence analysis, finish the process of front FLUTTER CALCULATION needs except needs, also need by step 1) set up control system model, calculate Generalized Control power, finish at last the aeroelastic divergence analysis of considering that structure, three pneumatic and control system are coupled.

When carrying out the moving servo elasticity calculating of hot gas, then need all structures-pneumatic-Re and four kinds of applied fields of control system are carried out the Coupled Dynamics simulation analysis.Aerodynamic Heating is by step 2) method calculate, non-Unsteady Flow is by step 3) method calculate, modal parameters coupling Aerodynamic Heating structure obtains by FEM (finite element) calculation, Generalized Control power is by step 1) control system model calculate and obtain, the hot gas that is coupled at last moves servo flexibility analysis.

5) design integration simulation analysis job model and data relationship

As shown in Figure 3, for four functions of specific implementation Fig. 2, designed job model and the data relationship of integrated simulation analysis system.The major function of each operation and data relationship and implementation method are described as follows among Fig. 3:

Operation 01: modal calculation.According to aircraft profile and structural parameters, set up respectively the finite element model of each structure branch, calculate natural frequency, the vibration shape and the modal mass of each branch.Modal parameter is saved in the database for operation 02 and operation 08.This operation is realized by finite element business software NASTRAN.

Operation 02: non-Unsteady Flow calculates.According to the locality non-steady pressure of stream calculation of parameter of each branch vibration shape of operation 01 output and operation 03 output distribute and integral and calculating corresponding to the broad sense aerodynamic coefficient matrix of each branch.Adopt the CFD numerical simulation to calculate the non-Unsteady Flow that is caused by elastic deformation with the technology that local stream piston theory combines.The present invention is directed to the Arbitrary 3 D object, suppose that air-flow is along the either direction of object plane tangent plane, set up the lower expression formula of washing that causes perpendicular to the object plane vibration at tangent plane, utilize the geometric relationship between practical structures direction of vibration and the object plane outer normal direction, set up the expression formula that local stream piston theory calculates the non-Unsteady Flow of random appearance, and the locality of complex appearance stream obtains by the CFD method of finding the solution the Euler equation, and the large angle of attack, assembly flow field are disturbed, washed down etc. and all reflected in the stream of locality.Last result of calculation exports operation 08 to.This operation is realized by self-compiling program.

Operation 03: local airflow calculates.Read the complex flowfield aerodynamic parameter of being found the solution by CFD software from database, flow in coefficient and preservation or the database for operation 02,04 and 07 locality that calculates fuselage, wing and empennage.This operation is realized by fluid mechanics business software FLUENT.

Operation 04: calculation of aerodynamic heating.The Aerodynamic Heating fast algorithm that adopts CFD to combine with engineering method, finding the solution the Euler equation without the viscosity flow Flow Field Numerical and determine boundary layer outer rim parameter the outside, then adopt reference enthalpy method gauging surface hot-fluid in the active inside boundary of viscosity, both overcome the difficulty that Engineering Algorithm is difficult to find the solution complex appearance boundary layer outer rim parameter, and avoided again numerical solution Navier-Stokes equation to need the problem of huge calculated amount and computational resource.By the method for this combination, calculate the wing-body-tail assembly under the different angles of attack of different Mach number heat flux and export to operation 05 and 06.This operation is realized by self-compiling program.

Operation 05: temperature field, thermal stress and hot-die attitude are calculated.Heat flux result of calculation or operation 06 heat shield back side heat flow density with operation 04 are calculated the result as input, utilize finite element software to carry out Analysis of Three-Dimensional Temperature along trajectory, on this basis, analyze the aerofoil thermal stress, form hot stiffness matrix and finally find the solution the hot-die attitude.Export each branch hot-die attitude vibration shape to operation 02, the fuselage hot-die attitude vibration shape is to operation 07, and each branch hot-die attitude natural frequency and modal mass are to operation 08.This operation is realized by finite element business software NASTRAN.

Operation 06: heat shield ablation Calculation.According to the heat flow density of operation 04 output, set up the layering ablating model and calculate heat transfer and loss in the heat shield, calculate the heat flow density of passing to load-carrying construction by the heat shield back side.Output reaches the heat flow density on load-carrying construction surface and is saved in database for operation 5 through heat shield.This operation is realized by self-compiling program.

Operation 07: control is calculated.Set up the servo-control system state-space model, according to the fuselage vibration shape value of the sensitive element position of operation 01 output and the mode participation value of operation 08 output, calculate the angle of rudder reflection of servo control mechanism output, then calculate the Generalized Control power of being given birth to by the rudder oblique presentation.Export operation 08 to.This operation is realized by self-compiling program.

Operation 08: Coupled Dynamics emulation and critical parameters identification.Modal parameter according to operation 01 or 05 output, the broad sense aerodynamic coefficient of operation 02 output, set up the state space aeroelasticity coupling model of full machine structural system, can realize the Coupled Dynamics emulation of aeroelasticity or aerothermoelasticity, find the solution flutter or heat flutter critical parameters; Form the closed-loop path with operation 07, form the aeroelastic divergence coupling model, can realize that aeroelastic divergence or hot gas move servo flexible Coupled Dynamics emulation, find the solution the critical parameters of servo flutter or hot servo flutter.Exporting each time step modal coordinate participation is worth to operation 07.Net result output: dynamic response time history diagram, critical dynamic pressure and threshold frequency.This operation is realized by self-compiling program.

6) realize integrated simulation analysis system

Can find from the job model of Fig. 3, have the data relationship of more complicated in the system between each operation, an operation often needs to carry out exchanges data with a plurality of operations.As shown in Figure 4, in order to realize the exchanges data of this Method In The Whole-process Analysis, set up unified database, each operation is carried out data storage and the exchange of (comprising the forms such as figure, text) by data base read-write, avoids the immediate data exchange between each operation as far as possible.Except non-Unsteady Flow calculates, control is calculated and three modules of Coupled Dynamics emulation because Relationship Comparison closely (three operations can be regarded as an integrated operation module with larger function), exist between operation outside the exchanges data, between each operation and the original input, exchanges data is the implementing reading and writing by database all.As shown in Figure 5, set up the integrated emulated data library structure of many Coupled Dynamics of full bullet.

At last, whole integral method is integrated business software technology and self-compiling program technology two large divisions.The self-compiling program technology adopts the database visualization programming technique based on SQLServer2005 and VC++.The main menu of master routine man-machine interface comprises the data typing, sets up model, program solution, data base querying and result of calculation output etc., can carry out separate part or the moving servo elastokinetics emulation of flutter, heat flutter analysis, aeroelastic divergence dynamics simulation and the hot gas of full machine assembly under different Mach number and different angle of attack conditions.

The above has illustrated the whole step of the integrated simulation analysis system of realization Coupled Dynamics and process, is the correctness of checking this method, and it is carried out wind tunnel test and two link checks of pure CFD contrast.The comparative result of the rudder model wind tunnel flutter test that table 1 was finished for this integrated simulation analysis system result of calculation and nineteen ninety-five, its critical velocity deviation is no more than 15.6%, and the two result quite coincide.

Table 1 rudder the model calculation and results of wind tunnel are relatively

Such as Fig. 6 (a) with (b), except the present invention and wind tunnel experiment relatively, also compare with pure CFD numerical method.Fig. 6 (a) and (b) be CFD/CSD coupling comparison of computational results figure by the result of calculation of this method and the non-Unsteady Flow of pure CFD numerical solution.Can see, the deviation of the two is very little in the low mach district, and then deviation is bigger in the High Mach number district, but the maximum deviation of critical dynamic pressure also only has 15.53% (conversion only has 7.48% for the critical velocity deviation) during Ma=7.Use the time of a Mach number point of this method calculating about 1 hour, and surpass 100 hours the computing time of pure CFD numerical value emulation method, in the situation that satisfy accuracy requirement, counting yield improves more than 100 times.

Contrast with wind tunnel test and two links of pure CFD numerical method by top, verified correctness and the validity of this integrated simulation analysis invention.

Such as Fig. 7 (a), (b) with (c), as example, provided the full fuselage-wing of certain type aircraft finished by the present invention-rudder assembly each branched coordinate displacement response under a certain operating mode.Under this dynamic pressure condition, flutter at first occurs in rudder face as seen from the figure, and 1,2 order frequencies of rudder face are tending towards synchronously, and the vibration on fuselage, wing and rudder face 3 rank still is tending towards convergence at this moment.

Be to examine Aerodynamic Heating, temperature field and thermal stress on the impact of structural modal, and verify validity of the present invention, calculated and analyzed the hot gas dynamic elasticity stability of full machine assembly on the high speed change angle of attack trajectory of imagination.Fig. 8 (a) and be two kinds of full machine assemblys of different structure material of certain model (b) along the critical dynamic pressure calculation result of the heat flutter of trajectory relatively.Can find, Aerodynamic Heating is very serious on the impact of flutter under the High Mach number condition, the result of calculation of aluminium alloy structure particularly, although taked the solar heat protection measure, solar heat protection and hot strength technical requirement have been satisfied, but still there is significantly decline on the heat flutter border, can not satisfy the flutter designing requirement.Another phenomenon is that the most dangerous heat flutter border does not appear at the highest Trajectory-terminal of structure temperature, the result of calculation of titanium alloy structure particularly, the Flutter Boundaries of trajectory back segment is greatly improved on the contrary, even is higher than the normal temperature state of not considering Aerodynamic Heating.This is because the impact of Aerodynamic Heating not only makes each rank model frequency descend, and affect the degree of closeness of frequency between the vibration shape and the mode, when these two factors so that during the degree of coupling reduction, just above-mentioned situation may occur.So for hypersonic aircraft, the heat flutter analysis is absolutely necessary, and must carry out transient state temperature field and corresponding hot-die attitude and heat flutter analysis along trajectory.

Claims (3)

1. the integrated simulating analysis of many Coupled Dynamics of hypersonic missile is characterized in that comprising the steps:

1) sets up the structure motion equation that is coupled;

2) improve the calculation of aerodynamic heating method;

3) improve non-Unsteady Flow computing method;

4) set up integrated simulation analysis techniqueflow;

5) design integration simulation analysis job model and data relationship;

6) realize integrated simulation analysis system.

2. the integrated simulating analysis of many Coupled Dynamics of hypersonic missile according to claim 1, it is characterized in that: described step 1-4), set up integrated simulation analysis techniqueflow, comprise that flutter analysis techniqueflow, heat flutter analytical technology flow process, aeroelastic divergence analytical technology flow process and hot gas move servo flexibility analysis techniqueflow.

3. the integrated simulating analysis of many Coupled Dynamics of hypersonic missile according to claim 1, it is characterized in that: described step 1-5), set up integrated simulation analysis job model and data relationship, comprise modal calculation operation module, non-Unsteady Flow computational tasks module, local airflow computational tasks module, calculation of aerodynamic heating operation module, temperature field, thermal stress and hot-die attitude computational tasks module, heat shield ablation Calculation operation module, control computational tasks module and Coupled Dynamics emulation and critical parameters identification operation module.

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