0. 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

1. Contents 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

2. Space Environments - Satellite Thermal Radiation

3. Space Environments – Distinguished environmental conditions
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

4. Satellite Thermal Control Requirements
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)

5. Satellite Thermal Control Requirements (Continued)
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
25% excess heater control authority (or duty cycle < 80%)
(3) Unit design margin
temperature difference between acceptance and qualification test levels, usually 10oC

6. Requirements of Satellite Thermal Control Predictions and Tests

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

8. Satellite Thermal Control Design Strategy
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

9. Satellite Thermal Design Parameters
Description
Input
Source
Thermal Output
Orbit characteristic
Altitude, inclination, beta angle
SYS
External heater flux, radiator allocations
Environmental heat sources on satellite
Orientation, attitude, operation scenario
Design life
Max. operation time after launch
Thermal-optical characteristics(ELO & BOL)
Thermal margin
Uncertainty margins, thermal design margins, heater margins
TCS
Allowable predicted temperature limits
Thermal range
Temperature limits (Operating/non-operating),
Power dissipations(Max/Min)
Optics,
EE,
SMS
Selection of thermal control materials
Outgassing and degradation criteria
TCS, Optics
Material characteristics
Minimize the temperature gradients
Temperature stability requirements
Optics, SMS
Temperature control set-points

10. Satellite Thermal Design Parameters(cont’)
Description
Input
Source
Thermal Output
Thermal-physical property,
coating
Conductivity (k), emissivity (), absorptivity ()
Optics,
EE,
SMS
Conductance,
emittance,
absorptance
Geometry layout
Locations, dimensions, mass
View factors, thermal capacity, allocations for radiators, heaters, and thermistors
Power budget
Available heater power
EPS
SYS
Required heater power for thermal controls
Allocated numbers of heater line
Available numbers for applying heater lines EE
Allocations of heaters

11. Typical Thermal Control Hardware
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)

12. Typical Satellite Thermal Control Hardware (Continued)
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
Resistance element
Kapton insulation
Lead wire

13. 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
Energy balance:
Qabsorbed + Qpower generation = Qemitted
Qds + Qer + Qet + Qinternal = σsc,spAscTsc4
 Gds +  Ger +  Get + Qinternal = σεFsc,spAscTsc4

14. Design Example - Thermal Analysis Concepts(Cont’)
External energy:
Direct solar
Gds=PAS Hsu , PAS is the projected area in the direction of the sun vector, Hsu is the solar constant (1300 ~1400 W/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, Fer is the configuration factor to sunlit part of the Earth
Internal Energy:
Internal heat input
Qinternal is the energy generated internally as heat and conducted and radiated to the external surface

15. Design Example - Thermal Parameters
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
Z, Up
MLI(top& bottom) Y
1 m
X, Velocity E
Figure 1. Minimum Sun
Figure 2.100% Sun

16. Design Example - Thermal Parameters(Cont’)
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) Z Aa Ab  s
 =cos-1 (Re/Re+Z)
=  - 90
+90  
Z= 1000 Km
Re= 6371 Km
= 30.2
A= 2m2

17. 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 = 321  (W)
Maximum sun, (4 surfaces +X, -X, +Y, -Y)
Qet =  x 224 x (4 x2) x =344.1  (W)
Earth reflected solar energy
Qer =  Hsu a A Fer , the approximation Fer  Fet cos 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 x 0.318=  (W)
Maximum sun, = 90, cos = 0
Qer = 0 (W)

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

19. 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= K or –72.0 °C, at minimum sun and minimum power, 45W
For comparison at maximum power and minimum sun, the temperature is
Tsc= K or –68.6 °C, at minimum sun and maximum power, 90 W

20. 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= 248 K or –25 °C, at maximum sun, degraded coatings, and minimum power, 45 W

21. Design Example - Temperature Change for Power Change
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

22. Design Example - Improving the Temperature Control
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 ℃)
( )4x5.67x10-8x2x4x0.85= α+321x
α = 0.407

23. Design Example - Internal Mass to External Radiator Resistance
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
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.

24. Design Example - Internal Mass to External Radiator Resistance
We increase the α further so that the minimum-sun minimum-power temperature is the same as the maximum-sun maximum-power degraded coatings case (-25℃)
( )4x5.67x10-8x2x4x0.85= α+321x
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.

25. 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

26. Satellite Thermal Control System Verification
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

27. Thermal Balance / Thermal Vacuum Test Temperature Profile
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
Hot
Proto-flight
2 hrs Soak
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

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

29. Satellite Thermal Balance Test
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 for hot and cold balance phases, respectively

30. Satellite Thermal Balance Test (Continued)
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.

31. Satellite Thermal Balance Test (Continued)
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

32. Satellite Thermal Balance Test (Continued)
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.

33. Example of Transient Cooling – FORMOSAT-1

34. Thermal Vacuum Test Requirements

35. Satellite Thermal Vacuum Test
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

36. Satellite Thermal Vacuum Test (Continued)
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

37. Comments and Conclusions
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.

38. TCS Homework
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?