CN106407588B - Spacecraft thermal agitation responds Simulation Platform - Google Patents
Spacecraft thermal agitation responds Simulation Platform Download PDFInfo
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- CN106407588B CN106407588B CN201610865924.7A CN201610865924A CN106407588B CN 106407588 B CN106407588 B CN 106407588B CN 201610865924 A CN201610865924 A CN 201610865924A CN 106407588 B CN106407588 B CN 106407588B
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- G06F30/20—Design optimisation, verification or simulation
- G06F30/23—Design optimisation, verification or simulation using finite element methods [FEM] or finite difference methods [FDM]
Abstract
The invention belongs to high-precision Spacecraft guidance and control and dynamics simulation and control technology fields, specially spacecraft thermal agitation responds Simulation Platform, successively leads including data input modeling module, the in-orbit thermal analysis module of flexible accessory, equivalent heat load and calculates module, model analysis module, Coupled Dynamics modeling module, Coupling Dynamic Model solution module and post-processing module；The Simulation Platform establishes spaceborne large-scale flexible attachment and celestial body Coupling Dynamic Model, proposes a kind of spaceborne large-scale flexible attachment and the dedicated emulated analysis platform of celestial body coupling response.The Simulation Platform can be used to solve the attitudes vibration that the spacecraft with complicated flexible accessory is induced by the thermotropic dynamic deformation of its flexible accessory, can for flexible accessory spacecraft heat-driven oscillation response under practical situation prediction and assessment a kind of simple and effective model, Method and kit for are provided, can be used as the reference in real satellite design process.
Description
Technical field
The invention belongs to high-precision Spacecraft guidance and control and dynamics simulation and control technology field, specially spacecraft heat is disturbed
Dynamic response Simulation Platform.
Background technique
When spacecraft passes in and out earth's shadow area, spatial heat environment changes, and the acute variation of temperature can not only make soft
Property attachment biggish thermal deformation occurs, induce thermal vibration, and perturbed force effect can be also transmitted in spacecraft main body.Due to angle
The conservation of momentum, the vibration of flexible accessory will lead to spacecraft subjective posture and shake, and then influence spacecraft payload
Pointing accuracy and attitude stability cause spacecraft to can not work normally or disabler.
NASA has had observed that influence of the heat-driven oscillation to spacecraft orbit posture, also delivered in the world it is many because
Flexible accessory occurs heat-driven oscillation and causes the example of spacecraft operational failure.With later batch scholar to Thermal Load to space flight
The influence of device attitude dynamics is studied.Corresponding Correlation Analysis Technique has carried out a large amount of research work from component-level angle
Make.However most of documents considerations is all influence of the quasi-static thermal deformation of attachment to the attitude of satellite, that is, has ignored thermal response
Transient term, can not embody the hot induced vibration problem caused by day alternates with night because of track in a model.
Influence of the flexible structure thermally-induced motion to attitude motion of spacecraft is analyzed from total system level, it is both domestic and external
Research work is not very abundant.The thermotropic dynamic deformation of flexible accessory induces the analysis and emulation of spacecraft attitude variation, is related to
Various calculating such as the in-orbit Orbital heat flux/ascent of flexible accessory, transient state temperature field, mode, dynamic deformation, coupling response belong to
In typical multi-crossed disciplines problem, single simulation analysis system, which is completed to calculate, analyzes extremely difficult, dedicated system, simulation software
System also lacks very much.In spacecraft rigid body-attachment coupled system heat-dynamics research field relatively new development as number Johnston
With the research of Thornton.Johnston and Thornton discusses flexible accessory heat-driven oscillation with the non-coupled method of heat-structure
Influence to spacecraft dynamic response, for the satellite system simplified model that one kind is made of Rigid Base and flexible accessory, hair
A kind of two-dimensional surface motion theory analysis method that heat-Structural Dynamics is non-coupled has been opened up, the heat of satellite flexible accessory is had studied
Influence of the induced movement to satellite plane attitude dynamics.Shortcoming is that flexible accessory is regarded as to simple beam to handle, for
The labyrinth form of practical attachment does not discuss.
The spacecraft more demanding for pointing accuracy and attitude stability, if only with simplified model come approximate complexity
Flexible accessory structure, it is clear that be not suitable for the Interaction Mechanics characteristic for inducing the thermotropic dynamic deformation of flexible accessory spacecraft attitude variation
Simulation analysis is carried out, it can not the influence of Accurate Prediction and the spaceborne flexible member heat-driven oscillation of assessment to spacecraft attitude.Therefore, compel
The theory and numerical analysis model that can solve the spacecraft thermal agitation response with complicated flexible accessory structure, development will be established by being essential
Corresponding solution technique and software systems.
Summary of the invention
In view of the above technical problems, the purpose of the present invention is induce space flight to solve the thermotropic dynamic deformation of large-scale flexible attachment
The analysis and emulation of device attitudes vibration propose that a kind of spacecraft thermal agitation that spaceborne large-scale flexible attachment is coupled with celestial body response is special
Use Simulation Platform.
Design principle is: Rigid Base-flexible accessory class spacecraft is directed to, using this kind of underexcitation of equivalent heat load as disturbing
Dynamic source carries out modelling to system using hybrid coordinate method and FInite Element, establishes the full star kinetic energy and gesture of meter and hot load action
Energy model, establishes the thermotropic micro-vibration Coupling Dynamic Model of spacecraft using Lagrange method.After numerical discretization, coupling
Kinetic model is eventually exhibited as one group of nonlinear equation, and Newmark method combination Newton iteration method is recycled to be solved.This
Inventing targeted research object is the spacecraft with flexible accessory.Compared to flexible accessory, the rigidity of spacecraft center cabin
It is much bigger, therefore can be approximately the rigid body with lumped mass and rotary inertia center cabin, thus entire spacecraft
Centered on rigid body-flexible accessory coupled system, and ignore the translation displacements of Rigid Base.
Specific technical solution are as follows:
Spacecraft thermal agitation responds Simulation Platform, successively includes that data input modeling module, the in-orbit heat of flexible accessory
Analysis module, equivalent heat load are led calculation module, model analysis module, Coupled Dynamics modeling module, Coupling Dynamic Model and are asked
Solve module and post-processing module；
(1) data input modeling module: it is soft to establish spacecraft Rigid Base-in the way of interactive mode combination automatic conversion
The finite element model and in-orbit thermal model of property attachment coupled system；
(2) the in-orbit thermal analysis module of flexible accessory: existed using the spacecraft Rigid Base-flexible accessory coupled system established
Rail thermal model carries out the in-orbit heat analysis of Flexible appendages of spacecraft, obtains the transient state temperature field on flexible accessory；
The heat transfer type that the in-orbit heat analysis of flexible accessory structure is related to is mainly heat transfer and heat radiation.The present invention series of fortified passes
Note be spacecraft disengaging earth's shadow when due in the short time high temperature variation cause flexible accessory vibration problem, because
This, Orbital heat flux suffered by flexible accessory mainly considers solar radiation hot-fluid.The heat transfer fundamental equation of in-orbit heat analysis and logical
The equation of heat conduction is identical under normal radiation heat transfer, but increases orbit computation, ascent calculating, the calculating of Orbital heat flux.
(3) equivalent heat load leads calculation module: the equivalent heat of transient state temperature field on flexible accessory is carried out using initial strain method
Load leads calculation, obtains the equivalent nodal force changed over time on each node and torque；
Using spacecraft Rigid Base-flexible accessory coupled system finite element model of foundation, by wink on flexible accessory
The hot load in state temperature field is equivalent to the joint load on flexible accessory.Boom, beam, plate shell are generally comprised in flexible accessory structure
Equal components need temperature change suffered on these member units being equivalent to the nodal force load changed over time, so as to
Carry out the coupled system micro-vibration time-histories data analysis of next step.Using initial strain (temperature strain) method in finite element method
To calculate the equivalent nodal force of temperature load.According to the deformation pattern of unit, the temperature load of boom unit can be equivalent to axis
Xiang Li, the temperature load of beam element can be equivalent to axial force and torque, and the temperature load of Shell Finite Element can be equivalent in face
Power and face moment of face can complete Equivalent Calculation by the integral in unit.
(4) it model analysis module: is carried out using structural finite element model using the direct Superposition Method of iteration WYD-Ritz vector
The model analysis of Rigid Base-flexible accessory coupled system, obtains period and the vibration shape of coupled system；
Spacecraft Rigid Base-flexible accessory coupled system mould is carried out using the direct Superposition Method of iteration WYD-Ritz vector
State analysis.Improved efficiency using multinomial technologies such as grouping shift frequency, mode error convergence criterion, the sparse Fast Direct Methods of cell,
Solving precision and reliability.The scale of solving a problem of eigenvalue problem, can up to 30 to 50 ten thousand freedom degrees on common computer at present
Accurately solve up to several hundred a low side mode.Mode error convergence criterion makes the process of model analysis become steady.Test knot
Fruit shows the precision that mode error bit value indicative error can more reflect that eigenvalue problem calculates.When calculating more multi-modal, mode
Error should be as preferred convergence criterion.
(5) Coupled Dynamics modeling module: utilizing modal expanding and Lagrange equation, according to finite element data and equivalent
Hot payload data establishes the thermotropic micro-vibration Coupling Dynamic Model of spacecraft coupled system；
The kinetic energy expression of flexible accessory and celestial body coupled system are as follows:
M is Space Vehicle System quality in formula；Mode exhibition is carried out to the malformation of flexible accessory in above formula with modal coordinate
After opening, the kinetic energy equation of Flexible appendages of spacecraft and celestial body coupled system are as follows:
Φ, η are respectively the modal matrix and modal coordinate battle array of flexible accessory structure in formula；
For Space Vehicle System it is quiet away from；
Rotary inertia for Space Vehicle System with respect to mass center
Battle array；
The flexibility being translatable for the heat-driven oscillation of flexible accessory structure to spacecraft
Coefficient of coup matrix；
Spacecraft is rotated for the heat-driven oscillation of flexible accessory structure
Flexible couplings coefficient matrix；
For the rigid body mode battle array of flexible accessory structure；
It organizes after collecting all unit strain energies of flexible accessory structure, then the spacecraft potential energy equation of meter and hot load effect are as follows:
K, r in formulaTThe respectively Stiffness Matrix of flexible accessory structure and hot load battle array, Λ are that attachment is rigid in Modal Space
Spend battle array；
The Lagrange function representation of Space Vehicle System are as follows:
According to the Lagrange equation of nonholonomic nonconservative, the translation of spacecraft is not considered, and ignores some higher order terms, then band
Spacecraft Rigid Base-flexible accessory coupled system the coupled dynamical equation of single flexible attachment finally simplifies are as follows:
Coefficient matrix in above formula can use the geological information of coupled system finite element model, mass matrix and intrinsic
Frequency and the vibration shape etc. obtain, and hot load equivalent nodal force can also utilize initial strain method by equivalent after obtaining temperature field data
Hot load is led calculation module and is obtained.
(6) Coupling Dynamic Model solves module: utilizing Newmark method combination Newton iteration method, carries out spacecraft coupling
The solution of the thermotropic micro-vibration Coupling Dynamic Model of collaboration system, obtains the time-histories data result and spacecraft appearance of flexible accessory
The time-histories data result at state angle；
The coupled dynamical equation is rewritten are as follows:
The state variable of definition system are as follows:
Then the coupled dynamical equation may be expressed as:
Using Newmark method in time-domain discrete, numeric format are as follows:
Wherein,
As δ >=0.5, γ >=0.25 (0.5+ δ)2When, algorithm is unconditional stability, allow in this way using it is biggish when
Between step-length, such as be selected as several points of the structure minimum period one.Therefore Newmark method can be used to solve time-histories longer
Time-histories data, and biggish time step can also omit the influence that the inaccurate characteristic solution of high-order responds system.In the present invention
In, select δ=0.5, γ=0.25.
The coupled dynamical equation is Nonlinear System of Equations, needs to change using Newton iteration method within each time step
In generation, solves, Iteration are as follows:
Above formula is substituted into load item, and it is decomposed into two parts:
Wherein,Unknown quantity q is required with current time stepn+1It is related,It is related with upper time step known quantity.Again
After arrangement, it is denoted as:
Ψ(qn+1)=F
Using Newton iteration method, the iterative formula solved are as follows:
In actually calculating, only need to restrain for iteration 3-5 times in each time step.
When solving Dynamics Coupling equation, due to using mode superposition method, after the modal information for obtaining model
Greatly reduce the order of equation group, while dynamic response solved using the Newmark method of unconditional stability, allow using compared with
Big time step helps to simulate the longer time or solves more massive problem, restrains in the iteration of each step
Also quickly, the efficient and reliability of derivation algorithm is illustrated.
(7) post-processing module: extracting and shows, the temperature variation curve that exports each node of flexible accessory, equivalent load become
Change the calculated results such as curve, micro-vibration time-histories data curve and spacecraft attitude angle change curve.
Post-processing module mainly requires to extract and show, exports relevant calculated result according to user, such as flexible attached
The temperature variation curve of each node on part, equivalent load change curve, flexible accessory micro-vibration time-histories data curve, spacecraft
Attitude angle change curve.
Spacecraft thermal agitation provided by the invention responds Simulation Platform, carries on the back by engineering of high pointing accuracy spacecraft
Scape establishes spaceborne large-scale flexible attachment and celestial body Coupling Dynamic Model, proposes a kind of spaceborne large-scale flexible attachment and star
The dedicated emulated analysis platform of body coupling response.The Simulation Platform can be used to solve band complexity flexible accessory spacecraft by
The attitudes vibration that the thermotropic dynamic deformation of its flexible accessory is induced can be the band thermotropic vibration of flexible accessory spacecraft under practical situation
The prediction and assessment of dynamic response provide a kind of simple and effective model, Method and kit for, can be used as real satellite design process
In reference.
Detailed description of the invention
Fig. 1 is schematic structural view of the invention；
Fig. 2 is the LEO spacecraft simplified structure diagram of embodiment；
Fig. 3 is end angle point side to light, shady face temperature variation curve and its difference variation curve of embodiment；
Fig. 4 is the equivalent moment curve of the endpoint node of embodiment；
Fig. 5 is the end angle point of embodiment along axial displacement of the lines；
Fig. 6 is the end angle point of embodiment around axial angular displacement；
Fig. 7 is the Rigid Base attitude angle change curve of embodiment.
Specific embodiment
Objects and advantages in order to further illustrate the present invention with reference to the accompanying drawing make into one the present invention with specific example
The explanation of step.
The spacecraft thermal agitation response Simulation Platform structure of the present embodiment is as shown in Figure 1, successively include that data input
The in-orbit thermal analysis module of modeling module, flexible accessory, equivalent heat load are led calculation module, model analysis module, Coupled Dynamics and are built
Mould module, Coupling Dynamic Model solve module and post-processing module.
Use Low Earth Orbit (LEO) spacecraft with single-blade sun battle array structure as shown in Figure 2 for research object.Needle
Enter the deformation of flexibility sun battle array structure dynamics caused by the acute variation thermal environment of shadow region and the vibration of celestial body posture to from sunshine area
Response has carried out Numerical Simulation Analysis.
Using platform proposed by the present invention, implement numerical simulation according to the following steps:
(1) finite element model and in-orbit thermal model of coupled system are established using data input modeling module；
(2) the in-orbit heat analysis that solar battery array is carried out using the in-orbit thermal analysis module of flexible accessory, obtains solar cell
Battle array flexible structure in-orbit period is especially into and out transient temperature field data when earth's shadow area；
(3) calculation module is led using equivalent heat load, the transient temperature field data that previous step obtains is equivalent to be applied to
Hot load on solar battery array flexible structure node；
(4) model analysis module is utilized, model analysis is carried out to coupling system model, obtains preceding 20 rank mode；Entire boat
Rigid body-flexible accessory coupled system centered on its device only constrains three translational degree of freedom of central point；
(5) it is built using Coupled Dynamics modeling module according to the finite element data of coupled system and equivalent heat payload data
The thermotropic micro-vibration Coupling Dynamic Model of vertical spacecraft coupled system；
(6) module is solved using Coupling Dynamic Model, carries out the thermotropic micro-vibration Coupling Dynamic Model of coupled system
Solution, obtain the time-histories data result of flexible accessory and the time-histories data result at spacecraft attitude angle.
Fig. 3 is that solar battery array end angle point is entering earth's shadow moment corresponding temperature variation curve and difference variation
Curve.From in Fig. 3 it can be found that solar battery array structure difference variation curve after initially entering earth's shadow by about
The time of 20s is intended to stablize.Therefore, the heat structure response time that can assert the solar battery array is approximately 20s, while it is tied
The structure response time (first step mode corresponding period) is 8.63s, and the two is not much different.The flexibility sun battle array structure is entering ground
Thermotropic dynamic deformation is likely to occur during ball shade.
Fig. 4 is the equivalent moment curve of the solar battery array end angle point.To illustrate the 0s moment as original state, and assume
Displacement at this time, equivalent heat load are zero.Available sun battle array end angle point is solved by the coupled dynamical equation and center is rigid
The corresponding thermotropic response curve of body posture, is shown in Fig. 5, Fig. 6 and Fig. 7 respectively.
It was found from above-mentioned simulation analysis result: the vibration that the acute variation of spatial heat environment causes the satellite system is rung
It answers.Wherein, sun battle array structure angle point axial deformation maximum amplitude is 0.95mm, bending deformation maximum amplitude in 0.54arcsec,
Celestial body attitude angle generates the maximum variation close to 0.042arcsec, and the thermotropic micro-vibration response level of solar battery array is smaller, but
Induce the flutter response of the attitude of satellite.
Claims (3)
1. spacecraft thermal agitation responds Simulation Platform, which is characterized in that successively include data input modeling module, flexible attached
The in-orbit thermal analysis module of part, equivalent heat load, which are led, calculates module, model analysis module, Coupled Dynamics modeling module, coupling power
Learn model solution module and post-processing module；
(1) data input modeling module: it is attached that spacecraft Rigid Base-flexibility is established in the way of interactive mode combination automatic conversion
The finite element model of part coupled system and in-orbit thermal model；
(2) the in-orbit thermal analysis module of flexible accessory: the spacecraft Rigid Base-in-orbit heat of flexible accessory coupled system established is utilized
Analysis model carries out the in-orbit heat analysis of Flexible appendages of spacecraft, obtains the transient state temperature field on flexible accessory；
(3) equivalent heat load leads calculation module: the equivalent heat load of transient state temperature field on flexible accessory is carried out using initial strain method
Calculation is led, the equivalent nodal force changed over time on each node and torque are obtained；
(4) center model analysis module: is carried out using the direct Superposition Method of iteration WYD-Ritz vector using structural finite element model
The model analysis of rigid body-flexible accessory coupled system, obtains period and the vibration shape of coupled system；
(5) Coupled Dynamics modeling module: modal expanding and Lagrange equation are utilized, according to finite element data and equivalent heat lotus
Data are carried, the thermotropic micro-vibration Coupling Dynamic Model of spacecraft coupled system is established；
(6) Coupling Dynamic Model solves module: utilizing Newmark method combination Newton iteration method, carries out spacecraft coupled systemes
The solution of the thermotropic micro-vibration Coupling Dynamic Model of system, obtains time-histories data result and the spacecraft attitude angle of flexible accessory
Time-histories data result；
(7) temperature variation curve, the equivalent load variation song that post-processing module: extracting and shows, exports each node of flexible accessory
Line, micro-vibration time-histories data curve and spacecraft attitude angle change curve calculated result.
2. spacecraft thermal agitation according to claim 1 responds Simulation Platform, which is characterized in that the coupling is dynamic
Mechanical modeling module establishes the thermotropic micro-vibration Coupling Dynamic Model of spacecraft coupled system specifically:
The kinetic energy expression of flexible accessory and celestial body coupled system are as follows:
M is Space Vehicle System quality in formula；After carrying out modal expanding to the malformation of flexible accessory in above formula with modal coordinate,
The kinetic energy equation of Flexible appendages of spacecraft and celestial body coupled system are as follows:
Φ, η are respectively the modal matrix and modal coordinate battle array of flexible accessory structure in formula；
For Space Vehicle System it is quiet away from；
Rotary inertia battle array for Space Vehicle System with respect to mass center；
The flexible couplings system being translatable for the heat-driven oscillation of flexible accessory structure to spacecraft
Matrix number；
The flexibility that spacecraft is rotated for the heat-driven oscillation of flexible accessory structure
Coefficient of coup matrix；
For the rigid body mode battle array of flexible accessory structure；
It organizes after collecting all unit strain energies of flexible accessory structure, then the spacecraft potential energy equation of meter and hot load effect are as follows:
K, r in formulaTThe respectively Stiffness Matrix of flexible accessory structure and hot load battle array, Λ are Stiffness Matrix of the attachment in Modal Space；
The Lagrange function representation of Space Vehicle System are as follows:
According to the Lagrange equation of nonholonomic nonconservative, the translation of spacecraft is not considered, and ignores some higher order terms, then band is single
Spacecraft Rigid Base-flexible accessory coupled system the coupled dynamical equation of flexible accessory finally simplifies are as follows:
Coefficient matrix in above formula utilizes geological information, mass matrix and the intrinsic frequency and vibration of coupled system finite element model
Type obtains, and hot load equivalent nodal force is led calculation module by equivalent heat load using initial strain method after obtaining temperature field data and obtained
It arrives.
3. spacecraft thermal agitation according to claim 1 responds Simulation Platform, which is characterized in that the coupling is dynamic
Mechanical model solves module, solution procedure are as follows:
The coupled dynamical equation is rewritten are as follows:
The state variable of definition system are as follows:
Then the coupled dynamical equation may be expressed as:
Using Newmark method in time-domain discrete, numeric format are as follows:
Wherein,
Select δ=0.5, γ=0.25；
The coupled dynamical equation is Nonlinear System of Equations, needs to be iterated using Newton iteration method within each time step and ask
Solution, Iteration are as follows:
Above formula is substituted into load item, and it is decomposed into two parts:
Wherein,Unknown quantity q is required with current time stepn+1It is related,It is related with upper time step known quantity；It rearranges
Afterwards, it is denoted as:
Using Newton iteration method, the iterative formula solved are as follows:
In actually calculating, only need to restrain for iteration 3-5 times in each time step.
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| CN109828477B (en) * | 2018-12-13 | 2021-12-21 | 上海航天控制技术研究所 | Vibration suppression method for large flexible spacecraft of Stewart platform |
| CN111814378B (en) * | 2020-07-14 | 2024-02-13 | 北京卫星环境工程研究所 | Environmental effect simulation method and device integrating temperature cycle and triaxial six degrees of freedom |
| CN112364571B (en) * | 2020-10-09 | 2023-01-13 | 天津大学 | Large complex coupling spacecraft dynamics model modeling method |
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| CN115017682B (en) * | 2022-05-10 | 2023-04-25 | 西北工业大学 | Mechanical behavior analysis method of stretching integral module in space force thermal environment |
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