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Picosat System Design Course - Satellite Thermal Control Design Introduction 黃正德 (J.D. Huang)

Picosat System Design Course - Satellite Thermal Control Design Introduction 黃正德 (J.D. Huang) 國家太空中心機械組熱控小組 October 16, 2008. Space Environments Satellite Thermal Control Requirements Satellite Thermal Design Philosophy Satellite Thermal Control Design Strategy

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Picosat System Design Course - Satellite Thermal Control Design Introduction 黃正德 (J.D. Huang)

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  1. Picosat System Design Course- Satellite Thermal Control Design Introduction 黃正德(J.D. Huang) 國家太空中心機械組熱控小組 October 16, 2008

  2. Space Environments Satellite Thermal Control Requirements Satellite Thermal Design Philosophy Satellite Thermal Control Design Strategy Satellite Thermal Design Parameters Typical Satellite Thermal Control Hardware Design Example Satellite Thermal Control System Verification Satellite Thermal Balance Test Satellite Thermal Vacuum Test Comments and Conclusions Contents

  3. Space Environments - Satellite Thermal Radiation

  4. Thermal Cycling Conditions: Extremely hot on satellite surface (>150oC) in the daytime because of facing the environmental heat sink and sources in an orbit Extremely cold on satellite surface (<-150oC) in the eclipse because of facing the environmental heat sink and sources in an orbit (Approximate) Vacuum Condition: Almost no medium and the convection heat transfer can be neglected Outgassing effect must be avoided or may cause contamination on some thermal control and optical areas Micro-gravity Condition: Any unit design with flow inside being different from ground use Space Environments – Distinguished environmental conditions

  5. The purpose of thermal control system is to maintain all the elements of a satellite system within their temperature limits (operating and non-operating) for all mission phases. Two top level thermal requirements, i.e., unit temperature limits and design margins should be defined before starting to develop a satellite thermal control for the sake of predictions and tests. Unit Temperature Limits (1) Operating limits unit operating ranges (ex. electronics: -10oC to +40oC; battery:-5oC to +25oC; hydrazine propellant elements: +10oC to +50oC; solar array panels: -100oC to +110oC; etc.) (2)Non-operating limits unit non-operating ranges (ex. electronics: -20oC to +50oC; most others same as operating limits) Satellite Thermal Control Requirements

  6. Design Margins (1) Uncertainty thermal design margin applied on the region where there is no thermal control or only passive thermal control 11oC for military and 5oC for other commercial and scientific satellites (2) Heater margin applied on the region where there is a heater 11oC for military and 5oC for other commercial and scientific satellites 25% excess heater control authority (or duty cycle < 80%) (3) Unit design margin temperature difference between acceptance and qualification test levels, usually 10oC Satellite Thermal Control Requirements(Continued)

  7. Requirements of Satellite Thermal Control Predictions and Tests

  8. Satellite Thermal Design Philosophy Radiation Property Radiation Execution Factors Radiation Computer Program – TRASYS, TSS Orientation & Attitude External Heater Flux View Factors Configuration Electrical Power Dissipation Thermal Analyzer Program – SINDA Selection of Thermal Control Materials and Hardware Elements Thermo-physical Property Geometry Predicted Thermal Performance Requirements System-level Test Comparison

  9. The satellite or spacecraft thermal control is quite unique and its design strategy is listed in the following: Predictions for Worst Hot and Cold Temperatures Temperatures predicted from the thermal mathematical model by considering extreme (worst hot and cold) thermal environmental effects including equipment operation, internal power dissipation, satellite attitudes, environmental heating (direct solar, earth infrared, and albedo radiation), etc. Cold-Bias Design Method Passive thermal control (ex. SSM / white paint and MLI) used first to lower all unit temperatures under their allowable upper limits Heaters used to raise some unit temperatures if they are lower than their allowable lower limits Active thermal control (ex. heat pipe and louver) used if cold-bias does not work Satellite Thermal Control Design Strategy

  10. Satellite Thermal Design Parameters

  11. Satellite Thermal Design Parameters(cont’)

  12. Multi-layered Insulation (MLI): to keep satellite warm by reducing conduction and radiation leaks (i.e., clothes of spacecraft) Second-surface Mirror (SSM): to reflect incident solar radiation (with low as) and radiate satellite excessive internal heat (with high e) into the space (i.e., radiator) Typical Thermal Control Hardware

  13. Heater: to keep satellite units warm and make up heat loss from the radiator during eclipse Heat Pipe: to transfer heat efficiently by using phase change between gas and liquid flow in a pipe container Typical Satellite Thermal Control Hardware (Continued) Resistance element Kapton insulation Lead wire

  14. Energy balance: Qabsorbed + Qpower generation = Qemitted Qds + Qer + Qet+Qinternal=σsc,spAscTsc4  Gds+ Ger +  Get+Qinternal=σεFsc,spAscTsc4 Design Example - Thermal Analysis Concepts Gds=PAS Hsu : gray body interchange factor er: earth reflected et: earth thermal sp: space sc: spacecraft su: sun ds: direct solar a : albedo  :solar absorptivity  :IR emissivity σ:Stefan-Boltzmann =5.67x10-8 W/m2K4 Gds:direct solar Ger:earth-reflected solar energy Get:earth-emitted thermal energy Qinternal s(sun vector) Asc Qsc=σsc,spAscTsc4 Ger=aFerAscHsu Get=FetAscHet Earth

  15. External energy: Direct solar Gds=PAS Hsu , PAS is the projected area in the direction of the sun vector, Hsuis the solar constant (1300 ~1400W/m2/oC) Earth-Emitted Thermal Energy Get=FetAscHet , Fet is the configuration factor to the Earth, Asc is the satellite area, Het is the Earth constant (198 ~274 W/m2/oC) Earth-Reflected Solar Energy Ger=aFerAscHsu , albedo a( 0.2 ~ 0.4) is the average fraction of the solar energy that is reflected by the earth, Feris the configuration factor to sunlit part of the Earth Internal Energy: Internal heat input Qinternalis the energy generated internally as heat and conducted and radiated to the external surface Design Example - Thermal Analysis Concepts(Cont’)

  16. Descriptions: Box shaped satellite with the - Z side always facing nadir (down) Dimension : 2 x 2 x 1 (L x W x H) m3 Top and bottom are covered with insulations (MLI) and sides may be considered isothermal Maximum power = 90 W Minimum power = 45 W Hsu = 1306 ~ 1400 W/m2 Het = 209 ~ 224 W/m2 Albedo a= 0.36 Design Example- Thermal Parameters Z, Up MLI(top& bottom) Y 1 m X, Velocity E Figure 2.100% Sun Figure 1. Minimum Sun

  17. External Heat Inputs Direct solar energy Qds =PAS Hsu, where PAS = Minimum sun, sun vector parallel to orbit plane(Fig. 1) PASAa= = = 0.478 By symmetry, PASAa= PASAb , Hsu= 1306 (W/m2) Qds =(PASAa+ PASAb)PAS Hsu= 1250 (W) Maximum sun, sun perpendicular to the orbit plane(Fig. 2), the sun is perpendicular to the +Y side, Hsu= 1400 (W/m2) Qds = PAS Hsu = 2 x 1400  = 2800  (W) Design Example- Thermal Parameters(Cont’) Z Aa Ab  s  =cos-1 (Re/Re+Z) = - 90 +90  0  Z= 1000 Km Re= 6371 Km = 30.2 A= 2m2

  18. Design Example- Thermal Parameters(Cont’) • Earth thermal energy Qet = Het A Fet , for a vertical plate at Z/Re =1000/6371=0.157, Fet = 0.192 • Minimum sun, (4 surfaces +X, -X, +Y, -Y) Qet =x 209 x (4 x2) x 0.192= 321 (W) • Maximum sun, (4 surfaces +X, -X, +Y, -Y) Qet =x 224 x (4 x2) x 0.192=344.1 (W) • Earth reflected solar energy Qer = Hsu a A Fer, the approximation Fer Fetcos will be used cos = =0.318 • Minimum sun, (4 sides, top and bottom surfaces are insulated)) Qer =x 1306 x 0.36 x (4 x2) x 0.192x 0.318= 229.6 (W) • Maximum sun, = 90, cos = 0 Qer =0(W)

  19. Summary Minimum sun Qenv =Qds + Qet + Qer = 1250  + 321  + 229.6  = 1479.6 + 321  Maximum sun Qenv =Qds + Qet + Qer=2800  + 344.1  Design Example- Thermal Parameters(Cont’)

  20. Design Example- Worst Case Temperature Predictions • Worst case cold • Consider the satellite to be an isothermal body with minimum power dissipation, minimum sun, undegraded thermal control surface (white paint, = 0.21, = 0.85) Qds + Qer + Qet + Qinternal = Qenv + Qinternal=σεFsc,spAscTsc4,Fsc,sp=1.0 Tsc= Tsc= 201.0 K or –72.0 °C, at minimum sun and minimum power, 45W • For comparison at maximum power and minimum sun, the temperature is Tsc= Tsc= 204.0 K or –68.6 °C, at minimum sun and maximum power, 90 W

  21. Design Example- Worst Case Temperature Predictions(Cont’) • Worst case hot • The worst case hot consists of maximum power, maximum solar input, and degraded thermal control coatings. The degraded solar absorptivity, , is 0.4 and the emissivity, , is unchanged. Tsc= Tsc= 250 K or –23.0 °C, at maximum sun, degraded coatings, and maximum power, 90 W • For comparison at maximum sun and minimum power, the temperature is Tsc= Tsc= 248 K or –25 °C, at maximum sun, degraded coatings, and minimum power, 45 W

  22. Temperature change for a change in power: ΔT/ ΔQ=[-23-(-25)]/(90-45)=0.044 ℃/W in the hot case ΔT/ ΔQ=0.076 ℃/W in the cold case In this case the design is not very sensitive to change in power, because the environmental inputs are much larger than the internal power Design Example- Temperature Change for Power Change

  23. For minimum power the change in temperature due to the environment and thermal control surface degradation is 72-25=47 ℃. The change due to degradation alone by calculating the maximum sun case with new (undegraded α=0.21) coatings. Tsc= The result is Tsc= -52 ℃ and by difference the change due to surface degradation is 27 ℃(-25+52). So the environmental changes alone, are 20 ℃. To find the α needed in the minimum sun case, at minimum power, the heat balance is solved for α with the same temperature as maximum sun, minimum power, undegraded(-52 ℃) (-52+273)4x5.67x10-8x2x4x0.85=1479.6 α+321x0.85+45 α = 0.407 Design Example- Improving the Temperature Control

  24. Based on a two-node model consisting of an outer shell and an inner electronics mass, we can calculate the required effective thermal resistance to raise the inner mass to the desired temperature. The effective thermal resistance is defined as Qint R=Te-Tsc Design Example - Internal Mass to External Radiator Resistance • The required thermal resistance in the cold case(Te at least 0 ℃) is • R=[0-(-52)]/45= 1.16 ℃/W • The maximum temperature for the hot degraded case would be • Te,max=90x1.16+(-23)= 81.4 ℃ • The maximum temperature is much higher than is desirable.

  25. We increase the α further so that the minimum-sunminimum-power temperature is the same as the maximum-sunmaximum-powerdegraded coatings case (-25℃) (-25+273)4x5.67x10-8x2x4x0.85=1479.6 α+321x0.85+45 α = 0.771 The effective thermal resistance required in this case for a minimum temperature of 0 ℃ is R=[0-(-25)]/45= 0.56 ℃/W and the maximum temperature is Te,max=90x0.56+(-25)= 25.4 ℃ This is a considerable improvement over either of the other cases. Design Example - Internal Mass to External Radiator Resistance

  26. Example of FORMOSAT-2 Thermal Design RSI Housing / FPA - MLI and radiator (outside) - heater (inside) - black paint (inside) IRU - radiator and MLI - heater ISUAL-S/P,A/P,CCD ISUAL-AEP - radiator and MLI - heater Star-Tracker - radiator and MLI Payload Platform - MLI - thermal isolation Bus Panel with Components - radiator and MLI (outside) - heater (inside) - black Kapton (inside) Solar Array - backside with Carbon Adapter Cone - MLI

  27. The satellite thermal system verification usually consists of thermal balance test and thermal vacuum test: Thermal Balance Test: To verify satellite thermal control system design adequacy by a simulated hot/cold space thermal environments To obtain thermal data for the correlation and correction of the thermal analytical models Thermal Vacuum Test: To demonstrate the ability to meet system design requirements under the specified hot/cold temperature extremes in a vacuum condition To demonstrate the system-level workmanship Satellite Thermal Control System Verification

  28. Thermal Balance / Thermal Vacuum Test Temperature Profile Temperature Pump Down & Cold Wall Fill Thermal Cycling Thermal Balance and Performance Cycle Return To Ambient Chamber Environments: Cold Wall Temp.  -173oC, Pressure  1.0 x 10-5 Torr 2 hrs Soak Hot Proto-flight Hot Acceptance Hot Balance Hot Performance Test > 24 hrs Dwell , Transient Cool-Down Ambient Cold Performance Test > 24 hrs Dwell Cold Acceptance Cold Balance Heater Check Cold Proto-flight 2 hrs Soak

  29. Example of FORMOSAT-2 Thermal Vacuum / Balance Test at NSPO

  30. Hot and cold balance phases: Objective: To achieve thermal equilibrium states in test article under simulated space hot and cold conditions to verify G (conductance) and Gr (radiation conductance) values assumed in TMM Conditions: Maximum and minimum orbit-averaged power dissipation of each unit applied for hot and cold balance phases, respectively Heating sources (for test) set to simulate hot and cold orbit-averaged heating loads on test article’s surface forhot and cold balance phases, respectively Satellite Thermal Balance Test

  31. Model correlation: TMM for test predictions in hot and cold steady-state conditions should be correlated to test results in hot and cold balance phases, respectively. The errors should be identified and corrected either from TMM or test itself if pre-test predictions are significantly deviated from the test results. The correlated predictions should agree within ±3oC of test data in general before the correlated TMM is used to make final temperature predictions for various satellite mission phases during the flight. Satellite Thermal Balance Test (Continued)

  32. Transient heating (warm-up) and cooling (cool-down) phases: Objective: To achieve transient heating and cooling in test article under simulated space warm-up and cool-down conditions to verify C (thermal capacitance) values assumed in TMM Conditions: Turning on and off all units in the test article for warm-up and cool-down phases, respectively, to speed heating and cooling rates Maximum and minimum heating powers applied on the external surfaces of the test article for warm-up and cool-down phases, respectively Satellite Thermal Balance Test (Continued)

  33. Model correlation: TMM for test predictions in transient-state heating and cooling conditions should be correlated to test results in warm-up and cool-down phases, respectively. In addition to the accuracy requirement same as hot and cold balance phases, the unit temperature curve from pre-test model should not cross or intercept with that from test result. Satellite Thermal Balance Test (Continued)

  34. Example of Transient Cooling –FORMOSAT-1

  35. Thermal Vacuum Test Requirements

  36. The satellite thermal vacuum test usually consists of ordinary and long thermal cycling phases in a vacuum condition: Ordinary Thermal Cycling Phase: Objective: To achieve unit hot and cold temperature extremes with hot and cold dwells, respectively, based on a specified test level for test article Control Requirements: Heating and cooling of test article controlled by heating sources and cold wall of the T/V chamber, respectively At least one component in each equipment zone reaching its specified hot and cold temperature limits; then dwelling Completion Criteria: Test article dwelling at hot and cold temperature limits (i.e.,unit temperature change is less than 2oC/hr) for at least 2 hours Satellite Thermal Vacuum Test

  37. Long Thermal Cycling Phase: Objective: To achieve unit hot and cold temperature extremes with hot and cold performance tests, respectively, during dwells based on a specified test level for test article Control Requirements: Heating and cooling of test article controlled by heating sources and cold wall of the T/V chamber, respectively At least one component in each equipment zone reaching its specified hot and cold temperature limits; then dwelling Completion Criteria: Test article dwelling at hot and cold temperature limits (i.e.,unit temperature change is less than 2oC/hr), and hot and cold performance tests conducted for at least 24 hours Satellite Thermal Vacuum Test(Continued)

  38. The satellite thermal control is an important task that can protect a satellite from a hostile thermal environments and keep it working well and surviving in all mission phases. The goal of developing a satellite thermal control should be achieved by considering cost, schedule, and technical aspects simultaneously although the thermal control technology is only mentioned here. In other words, we need a cheap, fast developed, and capable thermal control system in a satellite program. The thermal analysis work is usually going through the entire thermal control development from the beginning of design to the end of verification (by testing) phases. It is the most powerful supporting while developing a satellite thermal control system. The verification (by thermal balance test and thermal vacuum test) is the most complex and formidable task during the entire satellite development process. The performance in the thermal verification is a good indication if a satellite has a good thermal control in the space. Comments and Conclusions

  39. What kinds of thermal environments and thermal specifications should be considered during satellite design phase? Why? Is there any thermal design difference between LEO satellites and GEO satellites? Why? TCS Homework

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