CN102981081A - Evaluation method of thermal vacuum environmental adaptability of elements and components for spacecraft - Google Patents
Evaluation method of thermal vacuum environmental adaptability of elements and components for spacecraft Download PDFInfo
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- CN102981081A CN102981081A CN2012105112131A CN201210511213A CN102981081A CN 102981081 A CN102981081 A CN 102981081A CN 2012105112131 A CN2012105112131 A CN 2012105112131A CN 201210511213 A CN201210511213 A CN 201210511213A CN 102981081 A CN102981081 A CN 102981081A
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Abstract
The invention relates to the technical field of the evaluation of thermal vacuum environmental adaptability, and discloses an evaluation method of the thermal vacuum environmental adaptability of elements and components for a spacecraft. The method comprises the following steps: S1, performing stress analysis to the thermal vacuum environment of the elements and components for the spacecraft, so as to fix the vacuum stress, thermal stress and electric stress of thermal vacuum environmental ground tests, S2, fixing the cycle index of the thermal vacuum environmental ground tests, S3, performing thermal vacuum ground testing to the elements and components for the spacecraft according to test conditions fixed through S1 and S2, so as to judge whether samples of the elements and components for the spacecraft are qualified, and S4, evaluating that the thermal vacuum environmental adaptability of the elements and components meets task requirements if the quantity of unqualified samples is smaller than or equal to the acceptance number regulated in a sampling scheme, or evaluating that the thermal vacuum environmental adaptability of the elements and components does not meet the task requirements. The method can rapidly and accurately evaluate performances of the elements and components for the spacecraft under the thermal vacuum environment, and provides a basis for reasonably selecting the elements and components for the spacecraft.
Description
Technical field
The present invention relates to the adaptive assessment technique of hot vacuum environment field, particularly relate to a kind of spacecraft adaptive evaluation method of component thermal vacuum environment.
Background technology
Components and parts lost efficacy under hot vacuum environment and cause the case of satellite mission failure that report is arranged repeatly, for example, the International Ultraviolet Explorer (IUE) of emission in 1978 is owing to reasons such as thermal design mistakes, so that the 4K of computing machine and 8K storer are cracked on this ultraviolet detector satellite; Japan's broadcasting SDI (BSE) of emission in 1978 because the TWT Power high pressure arc discharge causes holding circuit to lose efficacy, can't be powered to travelling-wave tube, causes satellite end-of-life in 1980; Nineteen eighty-three the U.S. Landsat (Landsat-4), have 2 feed cables to damage in 4 solar battery arrays, failure cause is that thermal cycle makes conductor stress occur, then causes cable bad; The U.S. Geostationary Operational Environmental Satellite GOES-7 of emission in 1987, the data collection platform of satellite is inquired the appearance of (DCPI) system unusually in April, 1993, at stellar eclipse every day after date about 1 hour, No. 1 S-wave band receiver can not receive from instruction and data and catch the interrogating signal that (CDA) stands, the frequency stability of rear discovery receiver exceeds the restriction of desired value ± 5kHz, and reason is that stellar eclipse after date satellite temperature is lower; The European Meteorological Satellite (Meteosat) of European Space Agency in 1993 emission, radiometer recurs fault, is the optical surface that the ice that forms on the instrument has destroyed instrument by analysis.
Thermal vacuum test is a very important test in the space environment test, and its defective to hiding in the exposing product guarantees and the q﹠r of raising product plays a part very important.China in the evaluation check with components and parts, according to space flight user's requirement, has carried out the thermal vacuum test evaluation to portioned product at the product such as military electronic devices and components new product, type spectrum and aerospace engineering.Because the lacuna of components and parts level thermal vacuum test method standard, when formulating testing program, main reference be the thermal vacuum test method of satellite component, subsystem and whole star.The thermal vacuum test stress condition of satellite component, subsystem and whole star arranges the characteristics of Main Basis satellite task, adds accordingly tight examination according to the highest and minimum pre-apparent temperature, as among the GJB1027 to the regulation of test period: " at least 3 temperature cycles.Each circulation generally respectively keeps 4 ~ 8h at the highest (low) temperature value." and, for long-range mission, the existing test standard such as GJB1027 does not provide the corresponding relation between proof stress and the task environment stress.
Spacecraft mainly is subjected to spacecraft periodically to enter the earth's shadow district with the temperature cycles stress of components and parts between must in office to modulate, when the model duty cycle is longer, if by 1:1 components and parts are carried out thermal vacuum test, the cycle index of test is many, not only length consuming time, and cost is high, the result is inaccurate.
Summary of the invention
The technical matters that (one) will solve
The technical matters that the present invention at first will solve is: how a kind of method that can estimate quickly and accurately components and parts performance under hot vacuum environment is provided.
(2) technical scheme
In order to solve the problems of the technologies described above, the invention provides a kind of spacecraft adaptive evaluation method of component thermal vacuum environment, may further comprise the steps:
S1, spacecraft is carried out stress analysis with the component thermal vacuum environment, determine vacuum stress, thermal stress and the electric stress of hot vacuum environment ground experiment;
S2, determine the cycle index of thermovacuum ground experiment;
S3, the test condition of determining according to step S1 and S2 carry out the thermovacuum ground experiment to spacecraft with components and parts, and whether qualified with the components and parts sample to judge spacecraft, described test condition comprises described vacuum stress, thermal stress, electric stress and cycle index;
If S4 failed test sample number is less than or equal to the acceptance number of sampling plan regulation, the hot vacuum environment adaptability of then estimating tested this batch components and parts satisfies mission requirements; Do not satisfy mission requirements otherwise be evaluated as.
Preferably, step S2 specifically comprises:
S21: determine the temperature range Δ T in the most bad situation in the spacecraft module
2
S22: according to described thermal stress and described temperature range Δ T
2Calculate the speedup factor AF of hot vacuum environment test;
S23: according to the spacecraft cycle of operation and task execution time, and described speedup factor AF determines the cycle index N of thermovacuum ground experiment.
Preferably,
Wherein, Δ T
1The limit temperature difference for the components and parts test obtains according to described thermal stress, and q represents the constant relevant with the components and parts material technology.
Preferably,
N=（y/x）/AF
Wherein, x represents the spacecraft cycle of operation, and y represents the spacecraft task execution time.
Preferably, when (y/x)/AF calculated the result and is decimal, the principle of taking only to enter not give up rounded.
Preferably, among the step S3, each test is carried out with components and parts for a collection of spacecraft, and during test, components and parts adopt the lot tolerance percent defective LTPD methods of sampling, select according to task components and parts in latter stage survival probability.
Preferably, among the step S3, if one of following situation occurs in the process of the test or after the test, judge that then components and parts are as defective: 1) occur discharge, arcing or harmful corona in the process of the test; 2) it is undesired or parameter is overproof to monitor the function of components and parts in the process of the test; 3) find after the test that components and parts shell, lead-in wire or sealing occur and damage, or blurring; 4) the components and parts terminal point is measured or disqualified upon inspection after the test; 5) components and parts to sealing carry out leak test after the test, and the result is defective.
(3) beneficial effect
Technique scheme has following advantage: the present invention is on the basis of analyzing hot vacuum environment stress, determine the cycle index of ground experiment by the physical failure acceleration model, thereby test, can estimate fast and accurately spacecraft with the performance of components and parts under hot vacuum environment, for the choose reasonable spacecraft provides foundation with the components and parts product.
Description of drawings
Fig. 1 is method flow diagram of the present invention;
Fig. 2 is component thermal vacuum environment test profile synoptic diagram;
Fig. 3 is the most bad situation temperature range synoptic diagram in the spacecraft module;
Fig. 4 is star geothermal vacuum environmental stress factor synoptic diagram;
Fig. 5 is DC/DC hot vacuum environment test profile synoptic diagram.
Embodiment
Below in conjunction with drawings and Examples, the specific embodiment of the present invention is described in further detail.Following examples are used for explanation the present invention, but are not used for limiting the scope of the invention.
As shown in Figure 1, the invention provides a kind of spacecraft adaptive evaluation method of component thermal vacuum environment, may further comprise the steps:
S1, spacecraft is carried out stress analysis with the component thermal vacuum environment, determine vacuum stress, thermal stress and the electric stress of hot vacuum environment ground experiment, i.e. proof stress section, as shown in Figure 2;
The device on spacecraft surface (group) part is directly exposed to the cosmic space.Therefore, very large for the actual residing vacuum tightness variation range of the outer components and parts of not co-orbital spacecraft.For being installed in the device in each non-pressurized interior in the spacecraft, because the giving vent to anger of various materials in the restriction of vent port conductance and the cabin compared the cabin internal pressure and exceeded several orders of magnitude out of my cabin.Material outgassing in cabin internal pressure and orbit altitude, vent port admittance, the cabin, orbital motion time are relevant, after entering the orbit in the hundreds of hour, and general about 10
-1～10
-6Pa.For the components and parts in the pressurized capsule, although there is certain gas in the cabin, because be in state of weightlessness in orbit, convection heat transfer' heat-transfer by convection is inoperative.
Spacecraft mainly contains continuous working and discontinuous operation two states with components and parts, carries out the open and close machine operation for assembly and subsystem during the task, and components and parts will suffer the open and close electric stress.
The thermal environment of spacecraft in the space mainly refers to cold black and solar irradiation environment: do not consider the radiation of the sun and spacecraft, the energy density in cosmic space is about 1 * 10
-5W/m
2, be equivalent to temperature and be the energy that the black matrix of 3K sends.Heat radiation at spacecraft is absorbed by space entirely, does not have secondary reflection, and this environment is cold darkness environment, also cries heat sink.Main outer thermal source is solar electromagnetic radiation (blackbody radiation that is equivalent to a 6000K), can produce 100 ℃ high temperature during solar radiation, drop to the ultralow temperature below-200 ℃ during without solar radiation, when the spacecraft negative and positive are changed, the temperature fluctuation that the cabin is inside and outside, components and parts will suffer temperature cycles stress.In the cabin of spacecraft temperature as required general control in-10 ℃～55 ℃ scope, be exposed to out of my cabin electronic equipment temperature range generally at-120 ℃～150 ℃.Some material can produce aging and embrittlement in cold darkness environment, affect the performance of components and parts.And space flight with integrated circuit according to the orbit of spacecraft, position and mode of operation in spacecraft, its thermal environment is different again.
Thermal vacuum test is a kind of environmental simulation test, also is combined environment test, as a rule, and the main following points of test key element: 1) test vacuum tightness (being vacuum stress); 2) ultimate temperature (being thermal stress); 3) the ultimate temperature retention time; 4) warm variable Rate; 5) electric stress and working time; 6) test cycle number of times; 7) monitoring parameter; 8) failure criteria (being defective criterion).
According to the space environment CALCULATION OF THERMAL, test vacuum tightness adopts and is better than 1.3 * 10
-3Pa vacuum tightness, heat-conduction coefficient are ten thousand/condition of normal pressure, can estimate spacecraft component thermal physical behavior effects.
On definite in the ultimate temperature retention time, select according to the weight of tested components and parts, so that tested device can reach thermal equilibrium under ultimate temperature.The ultimate temperature retention time should be greater than the value in the table 1.
Test period under table 1 ultimate temperature
Electric stress should adopt the exemplary operation state of components and parts between must in office, and carries out respectively the switching on and shutting down test under ultimate temperature.
Other parameter (except cycle index and failure criteria) is selected according to conventional method.
In the present embodiment, DC/DC hot vacuum environment ground experiment stress profile as shown in Figure 5.Test condition is air pressure≤1.3 * 10
-3Pa; Test maximum temperature+125 ℃, minimum temperature-55 ℃, warm variable Rate 〉=1 ℃/min is by being installed in sample surfaces nonthermal source place temperature probe record test temperature; The residence time: stop 6.5h, minimum temperature stop 2.5h in maximum temperature; Electric stress: in maximum temperature, power up behind the minimum temperature outage 0.5h, sample is applied 28V voltage, load is fully loaded (50W).
S2, determine the cycle index of thermovacuum ground experiment:
S21: determine the temperature range Δ T in the most bad situation in the spacecraft module
2
The temperature variation of spacecraft internal electronic component is controlled at less scope, and representative value is Earth's orbit 5-10 ℃, 15-20 ℃ on Mars track.Yet, require device to have wider temperature range, it was not lost efficacy in unexpected energy loss or overheated situation.Fig. 3 has provided the most bad high low temperature limit of some tasks.-10-55 ℃ has comprised all tasks, can be used as the typical environment of orbiter and spacecraft inside.
In the present embodiment, the DC/DC present position temperature difference is 10 ℃.
S22: according to described thermal stress and described temperature range Δ T
2Calculate the speedup factor AF of hot vacuum environment test;
Because the vacuum stress in the proof stress and electric stress are equal to spacecraft substantially with the stress condition of components and parts between must in office, therefore only consider temperature cycles stress at this.
Wherein, Δ T
1The limit temperature difference for the components and parts test obtains according to described thermal stress, and q represents the constant relevant with the components and parts material technology, and generally between 3-9, conservative value is 3 to span.
In the present embodiment,
S23: according to the spacecraft cycle of operation and task execution time, and described speedup factor AF determines the cycle index N of thermovacuum ground experiment.
N=（y/x）/AF （2）
Wherein, x represents the spacecraft cycle of operation, and y represents the spacecraft task execution time.
Preferably, when (y/x)/AF calculated the result and is decimal, the principle of taking only to enter not give up rounded, and for example, calculating the result is 3.4, and then N is taken as 4, and calculating the result is 3.6, and then N also gets 4.
In the present embodiment, suppose that the spacecraft cycle of operation is 2h, task execution time is 12 years, and the cycle index of calculating DC/DC thermovacuum ground experiment according to formula (2) is:
The cycle index of determining DC/DC thermovacuum ground experiment is 10 times.
S3, the test condition of determining according to step S1 and S2 carry out the thermovacuum ground experiment to spacecraft with components and parts, and whether qualified with the components and parts sample to judge spacecraft, described test condition comprises described vacuum stress, thermal stress, electric stress and cycle index;
Among the step S3, each test is carried out with components and parts for a collection of spacecraft, and during test, components and parts adopt the lot tolerance percent defective LTPD methods of sampling, select according to task components and parts in latter stage survival probability.For the spacecraft components and parts, reliability requirement is high, can adopt under 90% degree of confidence scheme of acceptance number c=0, random sampling in the qualified parent of electric performance test.Sampling number and task survival probability in latter stage are as shown in table 2.
Table 2LTPD sampling plan table (degree of confidence: 90%, N＞200)
Hookup interface unit according to the rules, whether the monitoring device function is normal in the whole process of test.Among the step S3, if one of following situation occurs in the process of the test or after the test, judge that then components and parts are as defective: 1) occur discharge, arcing or harmful corona in the process of the test; 2) it is undesired or parameter is overproof to monitor the function of components and parts in the process of the test; 3) find after the test that components and parts shell, lead-in wire or sealing occur and damage, or blurring; 4) the components and parts terminal point is measured or disqualified upon inspection after the test; 5) components and parts to sealing carry out leak test after the test, find defective.
If S4 failed test sample number is less than or equal to the acceptance number of sampling plan regulation, the hot vacuum environment adaptability of then estimating tested this batch components and parts satisfies mission requirements; Do not satisfy mission requirements otherwise be evaluated as.
In the present embodiment, the test condition of determining according to above-mentioned steps carries out thermal vacuum test to DC/DC.Mission requirements this device task thermovacuum in latter stage effect fiduciary level under 90% degree of confidence is selected 22(0 more than or equal to 90% according to table 2) sampling plan.Input current and output voltage in the process of the test behind the monitoring device steady operation, in the test vacuum by the visual sample of thermal vacuum test case and continue to monitor input current to find possible vacuum discharge phenomenon.Test findings shows 22 and was not all lost efficacy by test agent, so evaluation result satisfies mission requirements for this batch DC/DC product.
As can be seen from the above embodiments, the present invention is on the basis of analyzing hot vacuum environment stress, determine the cycle index of ground experiment by the physical failure acceleration model, thereby test, can estimate fast and accurately the spacecraft performance of components and parts under hot vacuum environment, saved the time, reduced cost, for the choose reasonable spacecraft provides foundation with the components and parts product.
The above only is preferred implementation of the present invention; should be pointed out that for those skilled in the art, under the prerequisite that does not break away from the technology of the present invention principle; can also make some improvement and replacement, these improvement and replacement also should be considered as protection scope of the present invention.
Claims (7)
1. a spacecraft is characterized in that with the adaptive evaluation method of component thermal vacuum environment, may further comprise the steps:
S1, spacecraft is carried out stress analysis with the component thermal vacuum environment, determine vacuum stress, thermal stress and the electric stress of hot vacuum environment ground experiment;
S2, determine the cycle index of thermovacuum ground experiment;
S3, the test condition of determining according to step S1 and S2 carry out the thermovacuum ground experiment to spacecraft with components and parts, and whether qualified with the components and parts sample to judge spacecraft, described test condition comprises described vacuum stress, thermal stress, electric stress and cycle index;
If S4 failed test sample number is less than or equal to the acceptance number of sampling plan regulation, the hot vacuum environment adaptability of then estimating tested this batch components and parts satisfies mission requirements; Do not satisfy mission requirements otherwise be evaluated as.
2. the method for claim 1 is characterized in that, step S2 specifically comprises:
S21: determine the temperature range Δ T in the most bad situation in the spacecraft module
2
S22: according to described thermal stress and described temperature range Δ T
2Calculate the speedup factor AF of hot vacuum environment test;
S23: according to the spacecraft cycle of operation and task execution time, and described speedup factor AF determines the cycle index N of thermovacuum ground experiment.
3. method as claimed in claim 2 is characterized in that,
Wherein, Δ T
1The limit temperature difference for the components and parts test obtains according to described thermal stress, and q represents the constant relevant with the components and parts material technology.
4. method as claimed in claim 3 is characterized in that,
N=（y/x）/AF
Wherein, x represents the spacecraft cycle of operation, and y represents the spacecraft task execution time.
5. method as claimed in claim 4 is characterized in that, when (y/x)/AF calculated the result and is decimal, the principle of taking only to enter not give up rounded.
6. method as claimed in claim 4 is characterized in that, among the step S3, each test is carried out with components and parts for a collection of spacecraft, during test, components and parts adopt the lot tolerance percent defective LTPD methods of sampling, select according to task components and parts in latter stage survival probability.
7. such as each described method in the claim 1 ~ 6, it is characterized in that, among the step S3, if one of following situation occurs in the process of the test or after the test, judge that then components and parts are as defective: 1) occur discharge, arcing or harmful corona in the process of the test; 2) it is undesired or parameter is overproof to monitor the function of components and parts in the process of the test; 3) find after the test that components and parts shell, lead-in wire or sealing occur and damage, or blurring; 4) the components and parts terminal point is measured or disqualified upon inspection after the test; 5) components and parts to sealing carry out leak test after the test, and the result is defective.
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| CN103868750A (en) * | 2014-03-20 | 2014-06-18 | 航天东方红卫星有限公司 | Asymmetry hot test method suitable for repaired onboard product |
| CN104155200A (en) * | 2014-07-23 | 2014-11-19 | 西安空间无线电技术研究所 | Method for resisting frosting and condensation in rapid temperature change experiment |
| CN105158417A (en) * | 2015-08-18 | 2015-12-16 | 中国空间技术研究院 | Structure analysis method for SiP (system in package) device |
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| CN107907761A (en) * | 2017-11-03 | 2018-04-13 | 北京空间技术研制试验中心 | Test method for the component of spacecraft |
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| CN106569055A (en) * | 2016-10-19 | 2017-04-19 | 哈尔滨工业大学 | Electronic material and device heat cycle and charged particle irradiation combined environment test method |
| CN107907761A (en) * | 2017-11-03 | 2018-04-13 | 北京空间技术研制试验中心 | Test method for the component of spacecraft |
| CN108151994A (en) * | 2017-12-07 | 2018-06-12 | 北京空间技术研制试验中心 | The detection method of equipment limit capacity in spacecraft |
| CN109655683A (en) * | 2018-12-13 | 2019-04-19 | 广州广电计量检测股份有限公司 | A kind of automobile electronics thermal fatigue life accelerated test method |
| CN111006774A (en) * | 2019-12-06 | 2020-04-14 | 北京振兴计量测试研究所 | System and method for testing calibration blackbody radiation source manufactured by MEMS (micro-electromechanical systems) process |
| CN111006774B (en) * | 2019-12-06 | 2021-05-07 | 北京振兴计量测试研究所 | System and method for testing calibration blackbody radiation source manufactured by MEMS (micro-electromechanical systems) process |
| CN112781900A (en) * | 2020-12-15 | 2021-05-11 | 兰州空间技术物理研究所 | Inflation unfolding test method for flexible spacecraft in thermal vacuum environment |
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