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

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

3

Space Environments - Satellite Thermal Radiation

4

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

5

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)

6

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 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

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Requirements of Satellite Thermal Control Predictions and Tests

Acceptance Hot

AnalyticalRange

Prediction Test

Thermal UncertaintyMargin (11oC or 5oC)

Heater Margin(11oC or 5oC; 25%excess control)

Acceptance Cold

Protoflight Hot

5oC

5oC

10oC

10oC Protoflight Cold

Qualification Hot

Qualification Cold

Operating/Non-operatingHot

Operating/Non-operatingCold

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Satellite Thermal Design Philosophy

Radiation PropertyRadiation Property

Orientation & AttitudeOrientation & Attitude

Configuration Configuration

Radiation Execution FactorsRadiation Execution FactorsRadiation Computer Program – TRASYS,

TSS

Radiation Computer Program – TRASYS,

TSS

Electrical Power Dissipation Electrical Power Dissipation

Thermo-physical Property Thermo-physical Property

Requirements Requirements

Geometry Geometry

Thermal Analyzer Program – SINDAThermal Analyzer Program – SINDA

External Heater FluxExternal Heater Flux

View FactorsView Factors

Selection of Thermal Control Materials and Hardware

Elements

Selection of Thermal Control Materials and Hardware

Elements

Predicted Thermal Performance Predicted Thermal Performance

Comparison Comparison System-level Test System-level Test

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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

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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

SYS External heater flux, radiator allocations

Design life Max. operation time after launch

SYS 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

Allowable predicted temperature limits

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

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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 SMS 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

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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 s) and radiate satellite excessive internal heat (with high ) into the space (i.e., radiator)

10 x VDA/0.25 milMylar/VDA,perforated

11 SpacersDacron B4ANET

2 mil Mylar / VDA, perforated, To structure

2 mil Kapton / VDA, perforated, To space

5 mil Telfon, facing to space

966 acrylic adhensive

Aluminum or Silver

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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

Kapton insulationLead wire

Resistance element

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Design Example - Thermal Analysis Concepts

Energy balance:

Qabsorbed + Qpower generation = Qemitted

Qds + Qer + Qet + Qinternal = σsc,spAscTsc4

Gds + Ger + Get + Qinternal = σεFsc,spAscTsc4

Get=FetAscHet

Qinternal

Ger=aFerAscHsu

s(sun vector)Asc

Gds=PAS Hsu

Qsc=σsc,spAscTsc4

Earth

: gray body interchange factorer: earth reflectedet: earth thermalsp: spacesc: spacecraftsu: sunds: direct solara : albedo : solar absorptivity : IR emissivity σ : Stefan-Boltzmann =5.67x10-8 W/m2K4 Gds: direct solarGer: earth-reflected solar energyGet: earth-emitted thermal energy

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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

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Design Example- Thermal Parameters

Descriptions:

Box shaped satellite with the - Z side always facin

g 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

X, Velocity

Y

Z, UpMLI(top& bottom)

1 m

E

Figure 1. Minimum Sun Figure 2.100% Sun

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Design Example- Thermal Parameters(Cont’)

External Heat Inputs Direct solar energy

Qds = PAS Hsu , where PAS =

a) 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)

b) Maximum sun, sun perpendicular to the orbit plane(Fig. 2), t

he sun is perpendicular to the +Y side, Hsu = 1400 (W/m2)

Qds = PAS Hsu = 2 x 1400 = 2800 (W)

2

02

1sdA

Z

Ab

Aa

s

=cos-1 (Re/Re+Z)= - 90

0

Z= 1000 KmRe= 6371 Km= 30.2A= 2m2

90

2.30

0

90

2.30

00coscos

2

1

dadA

0sinsin2

2.300

090

A

+90

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Design Example- Thermal Parameters(Cont’)

Earth thermal energyQet = Het A Fet , for a vertical plate at Z/Re =1000/6371=0.157, Fet = 0.192

a) Minimum sun, (4 surfaces +X, -X, +Y, -Y)

Qet = x 209 x (4 x2) x 0.192 = 321 (W)

b) Maximum sun, (4 surfaces +X, -X, +Y, -Y)

Qet = x 224 x (4 x2) x 0.192 =344.1 (W)

Earth reflected solar energyQer = Hsu a A Fer , the approximation Fer Fet cos will be used

cos =

= 0.318

a) Minimum sun, (4 sides, top and bottom surfaces are insulated))

Qer = x 1306 x 0.36 x (4 x2) x 0.192 x 0.318= 229.6 (W)

b) Maximum sun, = 90, cos = 0

Qer = 0 (W)

2/

0

0

2/3cos0cos

2

1

dd

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Design Example- Thermal Parameters(Cont’)

• Summary Minimum sun

Qenv =Qds + Qet + Qer = 1250 + 321 + 229.6 = 1479.6 + 321

Maximum sunQenv =Qds + Qet + Qer =2800 + 344.1

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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 temperat

ure is

Tsc=

Tsc= 204.0 K or –68.6 °C, at minimum sun and maximum power, 90 W

4/1

8 4285.01067.5

4585.032121.06.1479

4/1

8 4285.01067.5

9085.032121.06.1479

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Design Example- Worst Case Temperature Predictions(Cont’)

• Worst case hot The worst case hot consists of maximum power, maximum solar inp

ut, 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 po

wer, 90 W

For comparison at maximum sun and minimum power, the temperat

ure is

Tsc=

Tsc= 248 K or –25 °C, at maximum sun, degraded coatings, and minimum powe

r, 45 W

4/1

8 4285.01067.5

9085.01.3444.02800

4/1

8 4285.01067.5

4585.01.3444.02800

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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

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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 )℃

(-52+273)4x5.67x10-8x2x4x0.85=1479.6 α+321x0.85+45 α = 0.407

4/1

8 4285.01067.5

4585.01.34421.02800

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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.

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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 )℃

(-25+273)4x5.67x10-8x2x4x0.85=1479.6 α+321x0.85+45α = 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.

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Example of FORMOSAT-2 Thermal Design

Payload Platform- MLI- thermal isolation

ISUAL-S/P,A/P,CCD ISUAL-AEP- radiator and MLI- heater

Bus Panel with Components- radiator and MLI (outside)- heater (inside)- black Kapton (inside)

IRU- radiator and MLI- heater

RSI Housing / FPA- MLI and radiator (outside)- heater (inside)- black paint (inside)

Star-Tracker- radiator and MLI

Solar Array- backside with Carbon

Adapter Cone- MLI

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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

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Thermal Balance / Thermal Vacuum Test Temperature Profile

2 hrs Soak

Thermal Balance and Performance Cycle

Pump Down & Cold Wall Fill

Chamber Environments: Cold Wall Temp. -173oC, Pressure 1.0 x 10-5 Torr

Return To Ambient

Ambient

ColdProto-flight

2 hrs Soak

ColdAcceptance

HotProto-flight

HotAcceptance

Hot Performance Test > 24 hrs Dwell

Cold Performance Test > 24 hrs Dwell

Hot Balance

Transient Cool-Down

Heater Check

Cold Balance

Temperature Thermal Cycling

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

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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

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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.

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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

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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.

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Example of Transient Cooling –FORMOSAT-1

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Thermal Vacuum Test Requirements

Thermal Vacuum Test Parameter

Acceptance Test Level

Protoflight Test Level

Qualification Test Level

Temperature Range and Extremes

Operating/non-operating temperature range, for at least one component in each spacecraft equipment area

Acceptance test temperature range 5oC, for at least one component in each spacecraft equipment area

Acceptance test temperature range 10oC, for at least one component in each spacecraft equipment area

Number of Cycles

Minimum of 4 cycles

Same as acceptance level or half of

qualification level

Minimum of 13 cycles

Dwell

Minimum of 2 hours soak at each

temperature extreme of each

cycle

Minimum of 2 hours soak at each

temperature extreme of each

cycle

Minimum of 2 hours soak at each

temperature extreme of each

cycle Chamber Pressure

10-5 Torr or less 10-5 Torr or less 10-5 Torr or less

Cold Wall Temperature

-173oC or less -173oC or less -173oC or less

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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

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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

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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.

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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?