CN104071360B - A kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation and system - Google Patents
A kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation and system Download PDFInfo
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Abstract
A kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation of disclosure and system, described method includes: (1) calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；(2) solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is calculated, to obtain the simulation heating performance number of one second Orbital heat flux；(3) the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container is calculated, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；(4) the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtains an Orbital heat flux simulation heating total power value；(5) at least one surface configuration heater of satellite, and export a current load in described heater by program-controlled d. c voltage-stabilized, be heated according to described Orbital heat flux simulation heating total power value simultaneously, to realize radiation Coupled Heat Transfer equivalent simulation.
Description
Technical field
The present invention relates to technical field of aerospace control, specifically, be a kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation and system.
Background technology
Current heat balance test can be divided into according to trystate: steady-state heat balance test and transition heat balance test.The test of described steady-state heat balance is that in virtual space vacuum, low temperature environment, subjects temperature is by the test of initial temperature changes in distribution to the equilibrium temperature profile tended towards stability under the simulation of track cycle integrated average Orbital heat flux and constant endogenous pyrogen mode of operation.The simulation of described transient heat flux is the checking test under the real space orbit external thermal flux environment and transient working pattern of satellite flight track, it is obtained in that the maximum temperature and minimum temperature, and temperature transient situation of change that satellite changes about the transient temperature in the actual operation on orbit process of assembly.In heat balance test process, it is necessary to carry out Orbital heat flux simulation, and can be divided into according to Orbital heat flux analog form: cold quartz mercury vapor lamp simulation heating simulation test, infrared radiation heating simulation test and absorption heat flux simulation test.Described cold quartz mercury vapor lamp simulation test refers in Vacuum Heat device, utilizes solar simulator and motion simulator, shines upon operating mode under analog satellite different flight attitudes in-orbit, the method have directly perceived, true, data reliable and precision high.The test of described absorption heat flux simulation can simulate transient condition effectively.Described infrared radiation heating simulation test, due to radiation heating time-lag effect reason, should not simulate transient condition.
When the heat balance test of satellite, the Application comparison of described steady state thermal flow field simulation is extensive, but described steady state thermal flow field simulation is not satellite real work state in orbit, but situation that is a kind of approximate and that simplify.Although this approximate and simplification can obtain satellite temperature scope substantially, but cannot reflect that each instrument on satellite is because of transient temperature situation of change produced by the impact by Orbital heat flux or endogenous pyrogen, therefore the high temperature residing for each instrument is in-orbit, two temperature extremals of low temperature cannot be determined, it is impossible to can the temperature of interpretative instrument meet index request.
Although it addition, the cold quartz mercury vapor lamp comparison used in above-mentioned cold quartz mercury vapor lamp simulation test is true, but it exists the situations such as equipment volume is huge, technology is more complicated, operation and maintenance cost is high, therefore realize transitional heat balance simulation relatively difficult.And the infrared cage used in above-mentioned infrared radiation heating simulation test or infrared lamp arrays etc., heat time heating time, hysteresis quality was bigger, satellite surface is easily produced bigger radiation and blocks impact by its auxiliary equipment, and simulation heating deviation and inhomogeneities bigger, realize the simulation of transient state Orbital heat flux relatively difficult, and simulation error wants big.
Simultaneously, along with the development of little micro-satellite technology, inside satellite heat power consumption is little, and satellite temperature is affected greatly by external environment condition change, heat balance test needs the thermal transient environmental change in-orbit of true analog satellite, and therefore the demand realizing simulation transitional heat balance being seemed becomes increasingly conspicuous.In view of this, a kind of transitional heat balance test method of offer and system are needed badly to solve the problems referred to above.
Summary of the invention
It is an object of the invention to, a kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation is provided, its can in the limited vacuum test container of size analog satellite Orbital heat flux transient changing situation in-orbit, thus obtaining satellite true temperature data in-orbit, verify correctness and the accuracy of thermal design.
For achieving the above object, the present invention provides a kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation, comprise the following steps: (1) calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；(2) solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is calculated, to obtain the simulation heating performance number of one second Orbital heat flux；(3) the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container is calculated, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；(4) the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtains an Orbital heat flux simulation heating total power value；(5) at least one surface configuration heater of satellite, and the current load exported by program-controlled d. c voltage-stabilized is in described heater, is heated according to described Orbital heat flux simulation heating total power value simultaneously, to realize radiating Coupled Heat Transfer equivalent simulation.
As optional technical scheme, step (2) farther includes step: calculate the transient temperature of solar energy sailboard according to default solar energy sailboard exterior-heat flow valuve and default operating mode transient state satellite total heat power consumption number；
According to the solar energy sailboard radiation observability position relationship at least one surface of described satellite, calculate the solar energy sailboard radiation exchange coefficient at least one surface of described satellite；
The solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is obtained according to described transient temperature and described radiation exchange coefficient.
As optional technical scheme, in step (5), described heater is arranged on radiating surface region or the insulating assembly outer layer at least one surface of described satellite.
As optional technical scheme, described heater arranges a thermal control coating.
Another object of the present invention is to, a kind of transitional heat balance pilot system based on radiation Coupled Heat Transfer equivalent simulation is provided, including: one first Orbital heat flux simulation heating power acquisition module, for calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；
One second Orbital heat flux simulation heating power acquisition module, for calculating the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite, to obtain the simulation heating performance number of one second Orbital heat flux；
One the 3rd Orbital heat flux simulation heating power acquisition module, for calculating the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；
One Orbital heat flux simulation heating total power value acquisition module, it is connected with described first Orbital heat flux simulation heating power acquisition module, the second Orbital heat flux simulation heating power acquisition module and the 3rd Orbital heat flux simulation heating power acquisition module respectively, for the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtain an Orbital heat flux simulation heating total power value；
One radiation Coupled Heat Transfer module, it is connected with described Orbital heat flux simulation heating total power value acquisition module, for being heated according to the heater on the described Orbital heat flux simulation heating total power value at least one surface to being arranged on satellite, to realize radiation Coupled Heat Transfer equivalent simulation, the current load wherein exported by program-controlled d. c voltage-stabilized is in described heater.
As optional technical scheme, described second Orbital heat flux simulation heating power acquisition module farther includes: a transient state temperature calculation unit, for according to presetting solar energy sailboard exterior-heat flow valuve and the transient temperature of default operating mode transient state satellite total heat power consumption number calculating solar energy sailboard；
One radiation exchange coefficient calculation unit, for calculating the solar energy sailboard radiation exchange coefficient at least one surface of described satellite according to solar energy sailboard to the radiation observability position relationship at least one surface of described satellite；
One radiation heating hot-fluid acquiring unit, it is connected with described transient temperature computing unit and described radiation exchange coefficient calculation unit respectively, for obtaining the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite according to described transient temperature and described radiation exchange coefficient.
As optional technical scheme, described heater is arranged on radiating surface region or the insulating assembly outer layer at least one surface of described satellite.
As optional technical scheme, described heater arranges a thermal control coating.
It is an advantage of the current invention that (1) can pass through analog satellite space orbit Orbital heat flux transient changing in-orbit and the transient working pattern of the instrument and equipment arranged on satellite, it is thus achieved that satellite true temperature data in-orbit, to verify correctness and the accuracy of thermal design.(2) present invention need not be arranged around support at satellite, it is not required that installs heat-flow meter at satellite surface, it is to avoid satellite surface is produced radiation and blocks.(3) present invention is applicable to low heat flow density heated mould plan, such as shadow region Orbital heat flux simulation etc..(4) heater being arranged on satellite surface can be carried out high-precision power output loading by the present invention, to promote the accuracy of Orbital heat flux simulation, coordinates program-controlled d. c voltage-stabilized to optimize the realization of transient heat flow simulation simultaneously.(5) present invention is capable of substituting under the solar energy sailboard whole star transitional heat balance test situation of participation, realize satellite Orbital heat flux transient changing in-orbit and radiation Coupled Heat Transfer hot-fluid equivalent simulation, and the transient working pattern examination of instrument and equipment on satellite, thus saving test space and cost.(6) present invention has the features such as strong operability, expense be relatively low, will become a kind of main radiation Coupled Heat Transfer equivalent simulation method of Orbital heat flux simulation from now on and a kind of development trend.
Accompanying drawing explanation
Fig. 1 is the flow chart of steps of the transitional heat balance test method described in one embodiment of the invention based on radiation Coupled Heat Transfer equivalent simulation.
Fig. 2 is the Organization Chart of the transitional heat balance pilot system described in one embodiment of the invention based on radiation Coupled Heat Transfer equivalent simulation.
Detailed description of the invention
Below in conjunction with accompanying drawing, the transitional heat balance test based on radiation Coupled Heat Transfer equivalent simulation provided by the invention and system detailed description of the invention thereof are elaborated.
Shown in Figure 1, a kind of transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation, comprise the following steps:
Step S110, calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；Step S120, the calculating solar energy sailboard radiation heating hot-fluid at least one surface of described satellite, to obtain the simulation heating performance number of one second Orbital heat flux；Step S130, calculate the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；Step S140, simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtain an Orbital heat flux simulation heating total power value；Step S150, at least one surface configuration heater of satellite, and the current load exported by program-controlled d. c voltage-stabilized is in described heater, it is heated according to described Orbital heat flux simulation heating total power value, to realize radiation Coupled Heat Transfer equivalent simulation simultaneously.
Hereinafter above-mentioned steps will be elaborated.
Step S110: calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux.
Satellite is a kind of spacecraft, when satellite orbits, is being continually changing relative to the sun, the position of the earth and direction.Therefore, solar irradiation, earth infra-red radiation and the earth light hot-fluid that each position, celestial body surface receives is also in change, and namely the space orbit Orbital heat flux on celestial body surface is as satellite and flies in-orbit and be continually changing.Therefore, adopting transient heat flux analogy method to carry out thermal balancing test of satellite, ratio adopts steady state thermal flow field simulation method to carry out thermal balancing test of satellite, closer to the true thermal environment that satellite flies in-orbit；Meanwhile, satellite equipment, according to transient working mode operation in-orbit, thus can obtain the maximum temperature of the associated component of satellite and the transient change situation of minimum temperature and temperature.
In the invention process, the quantity of described satellite surface is multiple, and situation hereafter is identical, and in other embodiments, the quantity of satellite surface is not limited to multiple, it is also possible to be one.
Structural model information according to satellite, the orbit information of satellite and flight attitude information calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface so that obtaining the simulation heating performance number of one first Orbital heat flux.
Step S120: calculate the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite, to obtain the simulation heating performance number of one second Orbital heat flux.
In the step s 120, optional step is farther included:
(a): calculate the transient temperature of solar energy sailboard according to default solar energy sailboard exterior-heat flow valuve and default operating mode transient state satellite total heat power consumption number.
The simulation heating performance number presetting solar energy sailboard Orbital heat flux is the structural model according to described satellite and is determined, and can pass through special-purpose software and calculate and obtain.
The function of solar energy sailboard is to provide the energy (here for electric energy) to the instrument of inside satellite, and it works required electric energy by the solar energy of absorption is partially converted into the instrument of inside satellite, and all the other are changed into heat at self；Simultaneously, it is considered to the earth infra-red radiation heating to windsurfing, it is possible to obtain solar energy sailboard is space orbit Orbital heat flux in-orbit.Under satellite space vacuum low temperature environment in-orbit, space orbit Orbital heat flux heating suffered by the heating of space orbit Orbital heat flux, celestial body outer surface suffered by solar energy sailboard, and the release of endogenous pyrogen under the default mode of operation of satellite, heat transfer model (including radiation Coupled Heat Transfer between windsurfing and satellite body) according to whole star (including solar energy sailboard), analyzes satellite and presets the transient temperature of solar energy sailboard under operating mode.
Passing in and out shadow region due to satellite and the attitude of satellite points to change, cause celestial body space surface orbit external thermal flux to change, so that making whole star transient temperature change, therefore transient temperature can be a changing value.
(b): the radiation observability position relationship at least one surface of described satellite is calculated the solar energy sailboard radiation exchange coefficient at least one surface of described satellite according to solar energy sailboard.
In embodiments of the present invention, owing to the quantity of satellite surface is multiple, the quantity of solar energy sailboard is multiple, and each solar energy sailboard is different to the position relationship on multiple surfaces of satellite, it is therefore desirable to calculate all solar energy sailboards radiation exchange coefficient to multiple surfaces of described satellite.
(c): obtain the solar energy sailboard radiant heat transfer hot-fluid at least one surface of described satellite according to the radiation exchange coefficient between the transient temperature of described solar energy sailboard and described solar energy sailboard surface and satellite surface.
Owing to satellite has the surface heat radiation characteristic difference of the various location on even same of multiple surface, it is therefore desirable to process respectively, that is calculate all solar energy sailboards radiation heating hot-fluid to each surface of satellite.
Calculating by the radiation exchange coefficient between transient temperature and solar energy sailboard surface and each surface of celestial body to solar energy sailboard, to obtain the solar energy sailboard radiant heat transfer hot-fluid to each surface of satellite, wherein said radiant heat transfer hot-fluid can be represented by below equation, G=σ ε A F (T1 4-T2 4), wherein σ is Boltzmann constant, and ε is the infrared hemispherical radiation rate on solar energy sailboard surface, F is the solar energy sailboard radiation exchange coefficient to satellite surface, and T1 is the transient temperature of solar energy sailboard, and T2 is the transient temperature of satellite surface, subscript 4 represents that biquadratic, A are windsurfing areas.
Pass through above-mentioned steps, it is capable of substituting solar energy sailboard and participates in whole star heat balance test, the radiant heat transfer hot-fluid of satellite surface is superimposed to by solar energy sailboard in the simulation of satellite surface space orbit Orbital heat flux and simulates in the lump, whole-satellite experiment is participated in, thus saving test space and cost without band windsurfing.
Step S130: calculate the radiant heat transfer hot-fluid on heat sink at least one surface to described satellite of vacuum test container, to obtain the simulation heating performance number of one the 3rd Orbital heat flux.
Satellite space low temperature background ambient temperature in-orbit is close to 4K, and under the true deep cooling spatial environments of 4K, it is only small that the background radiation of satellite surface is added heat by spatial environments, it is possible to is similar to and ignores.And in ground space environmental simulation equipment, ambient temperature that its low temperature is heat sink is close to 100K, now the heat sink radiation heating amount to celestial body surface be can not ignore, need the heat sink background deducting vacuum test container from the Orbital heat flux simulation loading step radiation heating hot-fluid (i.e. the simulation heating performance number of the 3rd Orbital heat flux) at least one surface of described satellite, wherein Orbital heat flux load step refer to by accurately control program control DC voltage-stabilizing export an electric current with control be actually loaded on heater power, be described in detail in step S150.
Step S140: the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtains an Orbital heat flux simulation heating total power value.
That is, the simulation heating performance number of simulation heating performance number-three Orbital heat flux of simulation heating performance number+the second Orbital heat flux of simulation heating general power=the first Orbital heat flux.
Satellite is process in-orbit, and the space orbit Orbital heat flux of satellite surface is continually changing.Under the heat balance test operating mode set, the Orbital heat flux of satellite surface is generally with an orbit time as the cycle, cyclically-varying, or the large period change of multiple time orbital period.The radiation heating hot-fluid on multiple surfaces of described satellite is changed by solar energy sailboard with the variations in temperature of solar energy sailboard, namely determined by the radiation Coupled Heat Transfer Process between space orbit Orbital heat flux and solar energy sailboard and the satellite body of solar energy sailboard, be also mechanical periodicity.And satellite different surfaces is determine value by the heat sink background radiation heating hot-fluid of vacuum test container.Therefore, Orbital heat flux simulation heating total power value is also for cyclically-varying value, it is possible to achieve periodically transient simulation.
Step S150: at least one surface configuration heater of satellite, and the current load exported by program-controlled d. c voltage-stabilized is in described heater, is heated according to described Orbital heat flux simulation heating total power value, to realize radiating Coupled Heat Transfer equivalent simulation.
Wherein, described heater can be arranged on the radiating surface region at least one surface of described satellite, with the equivalent simulation Orbital heat flux in relevant position.In embodiments of the present invention, described heater is a heating plate.At the thermal control coating designed by described heating plate surface spraying satellite corresponding region.Certainly, described heating plate can also adopt the alternative coating of equivalent infrared hemispherical radiation rate.Described alternative coating is the succedaneum of ground level heat blance test, and it is less costly, but its infrared signature is basically identical with actual thermal control coating.
It addition, described heater (such as heating plate) can also be arranged on the multilayer insulation assembly outer layer at least one surface of described satellite, with the Orbital heat flux of equivalent simulation relevant position multilayer insulation assembly.Inner surface at the outermost layer thin film (being equivalent to thermal control coating) of multilayer insulation assembly arranges (such as passing through bonding method) heating plate, with the equivalent simulation Orbital heat flux at relevant position multilayer insulation assembly outer layer.
In the invention process, each surface of satellite is equipped with heating plate.Described heating plate can not only be provided at the radiating surface region of satellite surface, it is also possible to is arranged on the multilayer insulation assembly outer layer of satellite surface.
The heater (i.e. the heating plate of radiating surface region or multilayer insulation assembly outer layer) of satellite surface is carried out Orbital heat flux heating and simulates by the electric current exported by accurately controlling program-controlled d. c voltage-stabilized, and it is heated controlling according to described Orbital heat flux simulation heating total power value, to reach radiation Coupled Heat Transfer equivalent simulation.Wherein, described radiation Coupled Heat Transfer, except including the Orbital heat flux suffered by each surface of satellite, also includes two parts coupling heating hot-fluid.A part is the heat transfer model (including radiation Coupled Heat Transfer between solar energy sailboard and satellite body) according to whole star (including solar energy sailboard), obtains the transient temperature of solar energy sailboard under the default operating mode of satellite by analyzing；And according to the solar energy sailboard radiation observability position relationship at least one surface of described satellite, calculate the solar energy sailboard radiation exchange coefficient at least one surface of described satellite；The solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is obtained by the radiation exchange coefficient on the transient temperature of described solar energy sailboard and described solar energy sailboard and at least one surface of satellite.Another part is the radiation exchange coefficient to satellite surface of the vacuum test container low temperature background according to calculating, in conjunction with the heat sink ambient temperature of low temperature in Dewar vessel (100K), obtains the low temperature heat sink background radiation heating hot-fluid to satellite surface in vacuum test container.
Control to have the advantage that one by adopting program-controlled d. c voltage-stabilized to carry out long-range electric current, can ensure that output current precision controls error less than ± 1%；Its two, adopt electric current to control to ignore the control error that remote wire transmission causes, it is to avoid to cause the impact that remote pressure drop brings because adopting Control of Voltage；Its three, adopt program-controlled d. c voltage-stabilized can according to the transient state Orbital heat flux of practical situation real-time loading celestial body outer surface.
Prior art is to realize Orbital heat flux heating simulation by the irradiation of infrared cage, infrared lamp arrays or cold quartz mercury vapor lamp, it has time delay, the deviation that controls and to surface heated inhomogeneities relatively big (being typically in about 5% ~ 10%), for low heat flow density (< 50W/m2) deviation of simulation then can be bigger.Compared with prior art, the present invention carries out long-range electric current control by program-controlled d. c voltage-stabilized, to realize Orbital heat flux simulation, namely by the size of control output current value in real time to obtain corresponding required heating power value, owing to it is real-time, therefore the present invention can pass through to control output electric current over time, it is achieved Orbital heat flux transient changing simulation in time, to embody the transient response of the present invention.The present invention can not only carry out high accuracy control to being arranged on performance number produced by the heater of satellite surface, to promote the accuracy of Orbital heat flux simulation, and coordinates program-controlled d. c voltage-stabilized can optimize the realization of transient heat flow simulation.
It addition, in the infrared radiation heating of prior art is simulated, it is necessary to infrared cage or infrared lamp arrays being arranged on around satellite by support, and at satellite surface, heat-flow meter is installed, being used for measuring incident hot-fluid, thus controlling the actual loaded hot-fluids such as infrared cage.But satellite surface actually needs region be directly facing heat sink (the vacuum test container inner wall) of vacuum test container, if infrared cage or infrared lamp arrays are arranged on around satellite, then can block the satellite surface heat sink area towards vacuum test container to a certain extent, so that creating a difference with truth.And the present invention is it can be avoided that produce to block to satellite surface, and current control accuracy is high, higher for low heat flux simulation occasion control accuracy, the especially low heat flux simulation in shadow region.Wherein above-mentioned low heat flux simulation refers to the situation that Orbital heat flux is less, if Orbital heat flux is less, corresponding loading current value is also less, therefore the present invention can carry out high accuracy control to being arranged on performance number produced by the heater of satellite surface, to promote the accuracy of Orbital heat flux simulation.Certainly, the invention is not restricted to the test of low heat flux simulation.
Another object of the present invention is to, a kind of transitional heat balance pilot system based on radiation Coupled Heat Transfer equivalent simulation is provided, including: one first Orbital heat flux simulation heating power acquisition module 210,1 second Orbital heat flux simulation heating power acquisition module 220, the 3rd Orbital heat flux simulation heating power acquisition module 230, Orbital heat flux simulation heating total power value acquisition module 240 and a radiation Coupled Heat Transfer module 250, as shown in Figure 2.
Wherein, described first Orbital heat flux simulation heating power acquisition module 210, for calculate satellite in orbit time satellite the Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux.
Described second Orbital heat flux simulation heating power acquisition module 220, for calculating the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite, to obtain the simulation heating performance number of one second Orbital heat flux.
Described 3rd Orbital heat flux simulation heating power acquisition module 230, for calculating the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container, to obtain the simulation heating performance number of one the 3rd Orbital heat flux.
Described Orbital heat flux simulation heating total power value acquisition module 240, it is connected with described first Orbital heat flux simulation heating power acquisition module 210, described second Orbital heat flux simulation heating power acquisition module 220 and described 3rd Orbital heat flux simulation heating power acquisition module 230 respectively, for the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtain an Orbital heat flux simulation heating total power value；
Described radiation Coupled Heat Transfer module 250, it is connected with described Orbital heat flux simulation heating total power value acquisition module 240, for being heated according to the heater on the described Orbital heat flux simulation heating total power value at least one surface to being arranged on satellite, to realize radiation Coupled Heat Transfer equivalent simulation, the current load wherein exported by program-controlled d. c voltage-stabilized is in described heater.
In the present embodiment, described second Orbital heat flux simulation heating power acquisition module 220 can include following optional assembly: transient state temperature calculation unit 221, radiation exchange coefficient calculation unit 222 and a radiation heating hot-fluid acquiring unit 223.Wherein, described transient temperature computing unit 221, for according to presetting solar energy sailboard exterior-heat flow valuve and the transient temperature of default operating mode transient state satellite total heat power consumption number calculating solar energy sailboard.Described radiation exchange coefficient calculation unit 222, for calculating the solar energy sailboard radiation exchange coefficient at least one surface of described satellite according to solar energy sailboard to the radiation observability position relationship at least one surface of described satellite.Described radiation heating hot-fluid acquiring unit 223, it is connected with described transient temperature computing unit 221 and described radiation exchange coefficient calculation unit 222 respectively, for obtaining the solar energy sailboard radiant heat transfer hot-fluid at least one surface of described satellite according to described transient temperature and described radiation exchange coefficient.
It addition, in the present embodiment, described heater (such as heating plate) can not only be provided at the radiating surface region at least one surface of described satellite, it is also possible to is arranged on the multilayer insulation assembly outer layer at least one surface of described satellite.By accurately controlling the electric current of program-controlled d. c voltage-stabilized output, the heating plate being arranged on satellite surface carried out remote power loading, and be heated according to described Orbital heat flux simulation heating total power value, to reach radiation Coupled Heat Transfer equivalent simulation.
Compared with prior art, the present invention carries out long-range electric current control by program-controlled d. c voltage-stabilized, to realize Orbital heat flux simulation, namely by changing current value size in real time to obtain corresponding required performance number, owing to its real-time is high, therefore the present invention can pass through to control output electric current over time, it is achieved Orbital heat flux transient changing simulation in time, to embody the transient response of the present invention.The present invention can not only carry out high accuracy control to being arranged on performance number produced by the heater of satellite surface, to promote the accuracy of Orbital heat flux simulation, and coordinates program-controlled d. c voltage-stabilized can optimize the realization of transient heat flow simulation.
The above is only the preferred embodiment of the present invention; it should be pointed out that, for those skilled in the art, under the premise without departing from the principles of the invention; can also making some improvements and modifications, these improvements and modifications also should be regarded as protection scope of the present invention.
Claims (6)
1. the transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation, it is characterised in that comprise the following steps:
(1) calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；
(2) solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is calculated, to obtain the simulation heating performance number of one second Orbital heat flux；
(3) the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container is calculated, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；
(4) the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtains an Orbital heat flux simulation heating total power value；
(5) at least one surface configuration heater of satellite, and the current load exported by program-controlled d. c voltage-stabilized is in described heater, with control output electric current realize over time Orbital heat flux in time transient changing simulation, it is heated according to described Orbital heat flux simulation heating total power value, to realize radiation Coupled Heat Transfer equivalent simulation simultaneously；
Step is farther included in step (2):
The transient temperature of solar energy sailboard is calculated according to default solar energy sailboard exterior-heat flow valuve and default operating mode transient state satellite total heat power consumption number；According to the solar energy sailboard radiation observability position relationship at least one surface of described satellite, calculate the solar energy sailboard radiation exchange coefficient at least one surface of described satellite；
The solar energy sailboard radiation heating hot-fluid at least one surface of described satellite is obtained according to described transient temperature and described radiation exchange coefficient.
2. the transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation according to claim 1, it is characterized in that, in step (5), described heater is arranged on radiating surface region or the insulating assembly outer layer at least one surface of described satellite.
3. the transitional heat balance test method based on radiation Coupled Heat Transfer equivalent simulation according to claim 2 a, it is characterised in that thermal control coating is set on described heater.
4. the transitional heat balance pilot system based on radiation Coupled Heat Transfer equivalent simulation, it is characterised in that including:
One first Orbital heat flux simulation heating power acquisition module, for calculate satellite in orbit time satellite the space orbit Orbital heat flux at least one surface, to obtain the simulation heating performance number of one first Orbital heat flux；
One second Orbital heat flux simulation heating power acquisition module, for calculating the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite, to obtain the simulation heating performance number of one second Orbital heat flux；
One the 3rd Orbital heat flux simulation heating power acquisition module, for calculating the radiation heating hot-fluid on heat sink at least one surface to described satellite of vacuum test container, to obtain the simulation heating performance number of one the 3rd Orbital heat flux；
One Orbital heat flux simulation heating total power value acquisition module, it is connected with described first Orbital heat flux simulation heating power acquisition module, the second Orbital heat flux simulation heating power acquisition module and the 3rd Orbital heat flux simulation heating power acquisition module respectively, for the simulation heating performance number according to the simulation heating performance number of described first Orbital heat flux, the simulation heating performance number of the second Orbital heat flux and the 3rd Orbital heat flux, obtain an Orbital heat flux simulation heating total power value；
One radiation Coupled Heat Transfer module, it is connected with described Orbital heat flux simulation heating total power value acquisition module, for being heated according to the heater on the described Orbital heat flux simulation heating total power value at least one surface to being arranged on satellite, to realize radiation Coupled Heat Transfer equivalent simulation, the current load wherein exported by program-controlled d. c voltage-stabilized, in described heater, is realized Orbital heat flux transient changing in time over time and simulates controlling output electric current；
Described second Orbital heat flux simulation heating power acquisition module farther includes:
One transient state temperature calculation unit, for according to presetting solar energy sailboard exterior-heat flow valuve and the transient temperature of default operating mode transient state satellite total heat power consumption number calculating solar energy sailboard；
One radiation exchange coefficient calculation unit, for calculating the solar energy sailboard radiation exchange coefficient at least one surface of described satellite according to solar energy sailboard to the radiation observability position relationship at least one surface of described satellite；
One radiation heating hot-fluid acquiring unit, it is connected with described transient temperature computing unit and described radiation exchange coefficient calculation unit respectively, for obtaining the solar energy sailboard radiation heating hot-fluid at least one surface of described satellite according to described transient temperature and described radiation exchange coefficient.
5. the transitional heat balance pilot system based on radiation Coupled Heat Transfer equivalent simulation according to claim 4, it is characterised in that described heater is arranged on radiating surface region or the insulating assembly outer layer at least one surface of described satellite.
6. the transitional heat balance pilot system based on radiation Coupled Heat Transfer equivalent simulation according to claim 5 a, it is characterised in that thermal control coating is set on described heater.
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