Picosat System Design Course -

1

• Picosat System Design Course -
• Satellite Thermal Control Design Introduction
• ???(J.D. Huang)
• ?????????????
• October 16, 2008

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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

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Space Environments - Satellite Thermal Radiation
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Space Environments Distinguished environmental
conditions

• Thermal Cycling Conditions
• Extremely hot on satellite surface (gt150oC) in
  the daytime because of facing the environmental
  heat sink and sources in an orbit
• Extremely cold on satellite surface (lt-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

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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)

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

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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

Resistance element
Kapton insulation
Lead wire
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Design Example - Thermal Analysis Concepts
GdsPAS 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
• s 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
Qscs?sc,spAscTsc4
GeraFerAscHsu
GetFetAscHet
Earth

• Energy balance
• Qabsorbed Qpower generation Qemitted
• Qds Qer Qet Qinternal s?sc,spAscTsc4
• ? Gds ? Ger ? Get Qinternal
  seFsc,spAscTsc4

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Design Example - Thermal Analysis Concepts(Cont)

• External energy
• Direct solar
• GdsPAS 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
• GetFetAscHet , 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
• GeraFerAscHsu , 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
  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 2.100 Sun
Figure 1. Minimum Sun
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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/ReZ) ? ? - 90
90
?
0
?
Z 1000 Km Re 6371 Km ? 30.2 A 2m2
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Design Example- Thermal Parameters(Cont)

• Earth thermal energy
• Qet ? Het A Fet , for a vertical plate at Z/Re
  1000/63710.157, 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 ? 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 0.192 x 0.318
  229.6 ? (W)
• Maximum sun, ? 90, cos? 0
• Qer 0 (W)

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Design Example- Thermal Parameters(Cont)

• Summary
• Minimum sun
• Qenv Qds Qet Qer 1250 ? 321 ? 229.6 ?
  1479.6 ? 321 ?
• Maximum sun
• Qenv 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
  seFsc,spAscTsc4 , Fsc,sp1.0
• 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

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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

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Design Example- Temperature Change for Power
Change

• Temperature change for a change in power
• ?T/ ?Q-23-(-25)/(90-45)0.044 ?/W in the hot
  case
• ?T/ ?Q0.076 ?/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 72-2547 ?.
• The change due to degradation alone by
  calculating the maximum sun case with new
  (undegraded a0.21) coatings.
• Tsc
• The result is Tsc -52 ? and by difference the
  change due to surface degradation is 27
  ?(-2552). So the environmental changes alone,
  are 20 ?.
• To find the a needed in the minimum sun case, at
  minimum power, the heat balance is solved for a
  with the same temperature as maximum sun, minimum
  power, undegraded(-52 ?)
• (-52273)4x5.67x10-8x2x4x0.851479.6
  a321x0.8545
• a 0.407

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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 RTe-Tsc

• The required thermal resistance in the cold
  case(Te at least 0 ?) is
• R0-(-52)/45
  1.16 ?/W
• The maximum temperature for the hot degraded case
  would be
• Te,max90x1.16(-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 a further so that the minimum-sun
  minimum-power temperature is the same as the
  maximum-sun maximum-power degraded coatings case
  (-25?)
• (-25273)4x5.67x10-8x2x4x0.851479.6
  a321x0.8545
• a 0.771
• The effective thermal resistance required in this
  case for a minimum temperature of 0 ? is
• R0-(-25)/45 0.56 ?/W
• and the maximum temperature is
• Te,max90x0.56(-25) 25.4 ?
• This is a considerable improvement over either of
  the other cases.

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Example of FORMOSAT-2 Thermal Design
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
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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
Temperature
Thermal Balance and Performance Cycle
Pump Down Cold Wall Fill
Thermal Cycling
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 gt 24 hrs Dwell
,

Transient Cool-Down
Ambient
Cold Performance Test gt 24 hrs Dwell
Cold Acceptance
Cold Balance
Heater Check
Cold Proto-flight
2 hrs Soak
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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
  articles 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
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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?