General

JERG-2-310

DESIGN STANDARDSPACECRAFTTHERMAL CONTROL SYSTEM

July 8 ,2009

Japan Aerospace Exploration Agency

JERG-2-310

This is an English translation of JREG-2-310.

Whenever there is anything ambiguous in thisdocument, the original document (the Japanese version) shall be used to clarify the intent of
the requirement.

Disclaimer
The information contained herein is for general informational purposes only.

JAXA makes

no warranty, express or implied, including as to the accuracy, usefulness or timeliness of any
information herein.

JAXA will not be liable for any losses relating to the use of theinformation.

Published by
Japan Aerospace Exploration Agency
Safety and Mission Assurance Department
2-1-1 Sengen Tsukuba-shi,Ibaraki 305-8505, Japan

JERG-2-310

CONTENTS

1 GENERAL RULES .......................................................................................................... 1
1.1 Purpose .................................................................................................................................... 1
1.2 Scope ........................................................................................................................................ 1

2 RELATED DOCUMENTS ............................................................................................... 1
2.1 Applicable documents ............................................................................................................. 1
2.2 Foreign standard etc. .............................................................................................................. 1

3 DEFINITION OF TERMS AND ABBREVIATION ........................................................ 2
3.1 Definition of terms ................................................................................................................... 2
3.1.1 Temperature........................................................................................................................... 2
3.1.2 Construction equipment for spacecraft system ..................................................................... 2
3.1.3 Tests ...................................................................................................................................... 3
3.1.4 Thermal control technology ................................................................................................... 4
3.2 Definition of abbreviation ........................................................................................................ 44 GENERAL REQUIREMENTS........................................................................................ 6
5 DESIGN REQUIREMENTS ........................................................................................... 6
5.1 Environmental conditions ....................................................................................................... 6
5.1.1 Fundamental rule................................................................................................................... 6
5.1.2 Ground environment .............................................................................................................. 6
5.1.3 Launch and ascent environment ........................................................................................... 7
5.1.4 Orbital environment ............................................................................................................... 9
5.1.4.1 External thermal input conditions ................................................................................... 9
5.1.4.2 Orbital conditions .......................................................................................................... 10
5.1.4.3 Attitude conditions......................................................................................................... 10
5.1.5 Thermal environment for spacecraft .....................................................................................115.1.5.1 Special quality of thermo-optics .....................................................................................11i

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5.1.5.2 Internal generation of heat by a spacecraft .................................................................. 14
5.2 Thermal design temperature conditions .............................................................................. 18
5.2.1 Classification of temperature range ..................................................................................... 18
5.2.2 Design of the estimated temperature range ........................................................................ 18
5.2.3 Maximum estimated temperature range.............................................................................. 19
5.2.4 Allowable temperature ranges ............................................................................................. 19
5.2.5 Estimated tolerance ∆T1 ...................................................................................................... 19
5.2.5.1 Meaning of estimated tolerance ∆T1 ............................................................................. 19
5.2.5.2 Procedure to set up estimated tolerance∆T1 ................................................................ 20
5.2.6 Design margin ∆T2 ............................................................................................................... 20
5.2.6.1 Meaning of design margin ∆T2 setting .......................................................................... 20
5.2.6.2 Procedure to set up design margin ∆T2 ........................................................................ 20
5.2.7 ∆T1 and ∆T2 of component thermally controlled in active type ............................................ 21
5.2.7.1 Approach to thermal control .......................................................................................... 21
5.2.7.2 Requirements for ∆T1 and ∆T2 ...................................................................................... 22
5.2.8 Thermal control of thermally critical component .................................................................. 28
5.3 Thermal interface ................................................................................................................... 29
5.3.1 Related document with thermal interface ............................................................................ 29
5.3.2 Thermal interface demarcation ............................................................................................ 29
5.3.2.1 Definition of thermal interface demarcation .................................................................. 29
5.3.2.2 Items to be defined as thermal interface conditions ..................................................... 31
5.3.3 Thermal interface design prepared by the system side ...................................................... 33
5.3.4 Thermal interface design prepared by component side ...................................................... 33
5.3.4.1 Requirements for internal onboard equipment ............................................................. 33
5.3.4.2 Requirement for exposed onboard equipment ............................................................. 35
5.3.5 Interface with propulsion system ......................................................................................... 37
5.3.5.1 Attitude control system.................................................................................................. 37
5.3.5.2 Large-sized liquid thruster ............................................................................................ 38
5.3.5.3 Ion engine system......................................................................................................... 40
5.3.6 Thermal interface between system and launch vehicle ...................................................... 41
5.4 Thermal control device/material technology ....................................................................... 43
5.4.1 Thermal control technology ................................................................................................. 43
5.4.1.1 The requirements for the thermal control system ......................................................... 43
5.4.1.2 Selection of thermal control method ............................................................................. 44
5.4.2 Thermal control device/material .......................................................................................... 45
5.4.2.1 Criteria for selection of thermal control devices and/or material .................................. 45
5.4.2.2 Criterion for selection of process .................................................................................. 46
5.4.2.3 Points to note for applying the thermal control device/material process ...................... 46
5.4.2.4 Production of thermal control parts/material ................................................................. 48ii

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5.4.2.5 Handling of MLI............................................................................................................. 48
5.5 Thermal design verification .................................................................................................. 49
5.5.1 Verification by thermal analysis ........................................................................................... 49
5.5.2 Verification by test ............................................................................................................... 49
5.5.2.1 Classification of tests .................................................................................................... 49
5.5.2.2 Approach regarding the selection of thermal balance test conditions .......................... 50
5.5.2.3 Procedure and selection of thermal balance test ......................................................... 50
5.5.2.4 Evaluation items ........................................................................................................... 51iii

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1 General rules
1.1 Purpose
The purpose of this document is to apply a design standard to the thermal control
design of the spacecraft (satellite/space probe) developed by Japan Aerospace
Exploration Agency (hereafter called “JAXA”). Moreover, this standard is defined in
anticipation of being applied not only to JAXA, but also to other organizations andcompanies.
1.2 ScopeThis standard defines requirements for the thermal design, thermal design analysis
and thermal tests of the spacecraft system and its components (hereafter called“spacecraft etc.”) and requirements for the thermal interface between spacecraft andlaunch vehicles etc.
.
Additionally, the subject of this standard is the general thermal range, between
approximately 100K and 500K.
For temperatures outside of this range, i.e.
temperatures higher than 500K or lower than 100K, this standard is not required, but canbe used just for reference.
2 Related documents
2.1 Applicable documents
The following normative documents contain provisions which, through reference inthis text, constitute provisions of this standard.
(1) JERG-2-311 Design Standard to prevent peeling of MLI(2) JERG-2-211 Design Standard for electrical charge/discharge(3) JERG-2-002 Standard of General Tests for Spacecraft(4) JERG-2-141 Standard of Space Environment(5) JERG-2-212 Design Standard of Wire Derating(6) JMR-010 Standard for Contamination Control
2.2 Foreign standard etc.
(1) ECSS-E-30 Mechanical Part 1A : Thermal Control(2) NASA GSFC GETS (ELV)-1 General Environmental Test Specification for Spacecraft
and Component(3) ISO 14644-1 Clean rooms and associated controlled environments
Part1 : Classification of airborne particulates(4) ISO 14952-2 Space systems – surface cleanliness of fluid systems Part2 :cleanliness levels(5) NASA TM X-64627 “Space and planetary environment criteria guidelines for use inspace vehicle development 1971 Revision”

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3 Definition of terms and abbreviation
3.1 Definition of terms
The terms used in this design standard are defined below.
3.1.1 Temperature(1) Allowable temperature range
The allowable temperature range shows the resistance to the environmental
temperature of a component, and classified under the operating phase (maintaining
performance and function) of equipment, the non-operating phase (preservation) andturn-on. See section 5.2 for more detail.
(2) Temperature reference point
The temperature reference point refers to the allowable temperature range for acomponent on a piece of equipment. Generally the position of the temperaturereference point is important in the context of interface. The overall temperature level may
be determined by the temperature reference point, or in some cases, a selected
thermally critical point . See section 5.3 for more detail.
(3) Design margin
This refers to the margin of safety in the area of thermal design required to keep the
spacecraft in orbit, certainly within the allowable temperature range as defined in (1)above, and with consideration of any unpredictable in-orbit phenomenon. See section
5.2 for more detail.
(4) Prediction error
Prediction error refers to an error of estimated temperature caused by an error inmathematical modeling and/or thermal analytical tool, and included in the measuringsystem and thermal input system for the in-orbit temperature prediction. See section
5.2 for more detail.
(5) Range of test temperatureThis refers to the temperature range loaded onto the model at the qualification test,
acceptance test or PFT (Proto Flight Test).
(6) Design prediction temperature range
This refers to the predicted temperature range as the result of thermal analysis.
section 5.2 for more detail.

See

3.1.2 Construction equipment for spacecraft system
The components of a spacecraft system comprise part of a larger system, to include
the combination of parts, devices and structures, and perform a function independently
within the whole operation of a larger system. Further, the components of spacecraft
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system are classified as being either internal onboard equipment or exposed onboard
equipment, as follows:(a) Internal onboard equipment
Refers to equipment mounted within the spacecraft and not exposed to outer space.
(b) Exposed onboard equipment
Refers to equipment either mounted outside the spacecraft, or when a portion of theequipment is exposed to the outer space and/or installed in an opening.
3.1.3 Tests
The terms concerned with test, used in this design standard, is based on the definition
mentioned on “JERG-2-002 Standard of General Tests for Spacecraft” of 2.1 (3). The
outline is shown below.
(1) Development Phase Tests<1> qualification test
The test to prove that the design and production procedure is suitable to obtain the
results thatthe hardware (and software) satisfy the requirements of the specification.
<2> acceptance test
The test to determine acceptance of goods for supply as flight hardware. The
purpose of the acceptance test is to detect any faults in the materials or manufacturing ofproducts which were produced according to the qualified design.
<3> PFT (Proto Flight Test)
Following the qualification and acceptance tests, the Proto Flight Test determines thequality of design and production procedure, as well as the possibility of acceptance as
the flight model.
(2) Verification of thermal design and the proof of function and performance tests<1> thermal balance test
The thermal balance test includes placing the space system or components into aspace chamber and simulating the thermal vacuum environment found in orbit. This
test confirms and verifies the thermal mathematical model of the component, the validityof the thermal control design and the performance of the thermal control hardware,
based on the acquired thermal data.
In this test, the confirmation and verification of the external thermal input conditions is
included, providing its own thermal vacuum test used for the development of the thermal
control system of the spacecraft.
<2> thermal vacuum test
In a more narrow sense, the thermal vacuum test, where the spacecraft or component
is placed into a space chamber and exposed to a thermal environment as if it were in
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orbit or even under more harsh conditions, is designed to prove the spacecraft orcomponent works normally and satisfies the requirement of function and performance.
In a wider sense, the thermal vacuum test is included as part of thermal balance test as
described above. The general, wider sense of the term, which includes both the thermalbalance test and the thermal vacuum test, is the standard term unless the thermal
vacuum test is specified.
3.1.4 Thermal control technology(1) Thermal control device
The constituent components of thermal control hardware (element, parts etc.) aredefined as such until its function is lost when it is further subdivided.
(2) Passive thermal control method
The passive thermal control method is the procedure to control the temperature of thecomponent within the specified range by adjusting the paths of conduction and radiation,and by the selection of geometric form of each surface and thermo-physical property ofthe spacecraft. The typical thermal control elements used for the passive thermal
control method include multi-layer insulation, thermal coating, heat sink and thermal
doubler (thermal diffusion) etc.
Furthermore, the fixed conductance heat pipe
corresponds to this type of thermal control method.
(3) Active thermal control method
The active thermal control method is the procedure to control the temperature using
mechanical mobile components or fluid, using electric energy from a heater, changing thecomponent’s thermo-physical property, or utilizing another technology to change/control
the temperature. The heater and thermal louver is the typical thermal control element
used in the active thermal control method. In addition, the variable conductance heatpipe corresponds to this type of thermal control method.
(4) Thermal mathematical model
The thermal mathematical model refers to the mathematical representation of thethermal characteristics of the spacecraft or components, which is used for temperature
prediction within the thermal environmental conditions. It is also sometimes called
thermal analysis model.
3.2 Definition of abbreviation
The abbreviation used in this design standard is shown as follows.
AT=Acceptance test
CVCM=Collected Volatile Condensable Materials
EM=Engineering Model
FM=Flight Model
ICD=Interface Control Drawing
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MLI=Multi Layer Insulation
PFM=Proto-flight Model
OSR=Optical Solar Reflector
PM=Prototype Model
QT=Qualification Test
SLI=Single Layer Insulation
TML=Total Mass Loss

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4 General requirements
Every spacecraft is required to have a thermal control design to keep every componentof spacecraft within a specified temperature range. To achieve this requirement, it isnecessary to clarify thermal interface conditions through the manufacturing, test, launchand operation phases, and to implement and to verify the thermal control design throughtest or analysis, considering the thermal interface between spacecraft system and its
components, as well as the launch vehicle.

5 Design requirements
Each requirement below must be considered for the thermal control design of aspacecraft system, to include:(1) environmental conditions(2) temperature conditions of thermal design(3) thermal interface(4) thermal control device, material technology(5) thermal design verification
5.1 Environmental conditions
5.1.1 Fundamental rule
Every spacecraft must meet various thermal environmental requirements at eachphase, from production to in-orbit operation. Every spacecraft must also consider thethermal environment during all phases of operation (on the ground, in the earth andmoon/planet orbit and interplanetary) in their designed service lives.
On the ground, spacecraft must be maintained within the allowable temperature range,
including consideration of temperature and humidity from production to launch. Eachspacecraft has its own uniquely defined thermal environment.
During the launch and ascent phases, spacecraft are heated through the fairing(asdescribed in section5.1.3). These heating conditions are specific to each spacecraft
and launch vehicle.
Spacecraft fly in orbit of outside the earth’s atmosphere, are heated externally by solar
radiation, by albedo (the reflection ingredients of solar radiation from the earth andmoon/planet) and by infrared radiation from the earth and moon/planet. These heat
inputs are taken under consideration, including potential error and fluctuation, througheach spacecraft’s entire designed service life.
5.1.2 Ground environment
Before launch, every thermal environmental condition must be considered for every
spacecraft at every phase of production, test, handling, storage, transportation andpreparation before launch. If any suitable thermal environmental condition (thermal
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conditions of batteries etc.) is necessary for spacecraft, a study using ground support
equipment shall be performed. Thus, the thermal environmental conditions must be
studied in an early stage of a spacecraft’s development as possible.
5.1.3 Launch and ascent environment
Monitoring of the thermal environmental conditions and rapid decompression
environmental conditions during launch and ascent are important. Changes in theenvironmental conditions caused by spacecraft itself, however, are not taken into
account.
(1) thermal environmental conditions
During the launch phase, the conditions are considered separately (see below)
because the thermal environment of a spacecraft is different before and after fairing
separation.
<1> There is heating of the spacecraft during the fairing<2> There is heating of the spacecraft directly after fairing separation
The above thermal environmental conditions must be clearly understood at an earlystage of its development as possible.
<1> heating to spacecraft through fairing
Spacecraft are heated by radiation from inside the fairing and as a result, thetemperature of the fairing increases. Aerodynamic heating, solar radiant heat, albedo
and earth infrared radiant heat increase the temperature of the exterior surface offairing before fairing separation.
It is necessary to consider the heat exchange (through conduction and radiation)
resulting from the interface between spacecraft and launch vehicle. Generally, thetemperature record of the fairing interior is shown by the launch vehicle side.
<2> heating to spacecraft directly after fairing separation
After fairing separation, the exterior surface of the spacecraft is heated by the outsideenvironment directly. The spacecraft is heated directly by the free molecular flow ofheat by rarefied gas, solar radiant heat, albedo and earth infrared radiant heat. In
some cases, it is necessary to consider the heat exchange (through conduction andradiation) resulting from the interface between spacecraft and launch vehicle.
Additionally, with respect to the launch vehicles which use solid fuel motors for theupper-stage propulsion, it is necessary to take into account of the heat conducted fromthe motor case and the heat from the firing and plume from the upper-stage vehicle.
(2) rapid decompression environmental conditions
It is necessary to consider the blistering of MLI and the reduction of internal gas
temperature in the fairing caused by the rapid decompression in the fairing during thevehicle’s ascent. In order to prevent the blistering and the separation of MLI, which
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considerably affects the thermal control system, see section 2.1 (1) below “JERG-2-311
Design Standard to prevent peeling of MLI”.

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5.1.4 Orbital environment
5.1.4.1 External thermal input conditions(1) solar constant
The solar constant is defined as the emission of energy from the unit’s surface area
vertical to the sunlight per hour, to the outside air at a distance of 1 AU (Astronomical
Unit: distance between earth and sun), or 1,366 W/m2. Though the intensity of solar
radiation is in inverse proportion to the square of the distance from the sun, this formula
must be adjusted seasonally by about ±3.5 % on average because of the earth’s ellipticalorbit.
Furthermore, please refer to section 2.1 (5) “JERG-2-141 Standard of Space
Environment” for further details of the solar constant.
(2) albedo and infrared radiation<1> albedo
The albedo is the ratio of sunlight which reflects to the surface of the earth andmoon/planet. The earth albedo changes approximately 0.05~0.6 according to thelatitude, season and weather condition etc. at the point directly below of spacecraftaround the earth. The annual average of albedo over the global surface is 0.3.
<2> infrared radiation
The infrared radiation is the thermal radiation that the earth and moon/planet emit.
The
Earth’s infrared radiation changes approximately 150~300 W/m2 according to thelatitude,season and weather condition etc. at the point directly below of spacecraft around theearth.
The annual average of infrared radiation over the global surface is 237 W/m2.
<3> application of spacecraft to thermal design
Regarding the application of a thermal design to spacecraft, usually the Earth’s
albedo and the Earth’s infrared radiation are treated as constants in time and in space,
and the following value is applied according to the time constant of each component.
Table 5.1.4 Value of the earth albedo and the earth infrared radiation
applied to the thermal design
time constant of thecomponent

earth albedo

earth infrared radiation(W/m2)∆t < 0.3 hr+0.30 +30/-0.15

237 +28/-97

0.3 <∆t < 3 hr

0.30±0.10

237 +24/-48

3 hr <∆t0.30±0.05

237±21

The time constant of Table 5.1.4 is the value determined based upon the relation
between the heat capacity of the spacecraft and the thermal resistance for the (internaland external) thermal input. So, if the time constant of spacecraft is known, the9

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estimation of the thermal response of the system shall be possible. If the values from
Table 5.1.4 are used for the thermal design of spacecraft, in cases when the tolerance of
the Earth’s infrared radiation is positive, the tolerance of the Earth’s albedo is negative
and when the tolerance of the Earth’s infrared radiation is negative, the tolerance of the
Earth’s albedo is positive.
In consideration of components which are sensitive to the change of the external
thermal input (such as an external component with a small heat capacity, a small amountof self-heating, or with a large surface area), occasionally it may be necessary tosimulate the in time and in space change of the Earth’s albedo and Earth’s infrared
radiation. If so, please refer to section 2.1(5) “JERG-2-141 Standard of Space
Environment” to investigate this compensation of external thermal input in detail ifnecessary.
(3) moon/planetary albedo and infrared radiation
A moon/planet’s albedo and infrared radiation with the exception of the Earth andMoon shall be based on current research results.
(4) shade
Shade is defined as the condition of solar light being obstructed by the earth or planet
etc. Shade can be classified into two categories: umbra (when solar light is completely
obstructed); and penumbra (when the obstruction of the solar light is not complete).
When the shade is penumbra for a significant duration, a separate thermal design must
be determined separate from the thermal design during the umbra.
5.1.4.2 Orbital conditions
To decide the strength and the direction of the orbital thermal input, the conditions
below are must be considered within the parameter of the orbit.
(1) orbital altitude : altitude of apogee and perigee (or semi-major axis and eccentricity)(2) orbit inclination(3) annual variation of orbital plane sun angle (so to called β angle)(4) annual variation of shadow cover rate (or shade time)
5.1.4.3 Attitude conditions
The attitude conditions determine the incident direction of each surface of spacecraft ofthe orbital heat input. The attitude conditions are defined by the orbital heat input of thesolar light, as well as the albedo and infrared radiation of the Earth or lunar/planet.
These are expressed by the following:(1) The solar orientation based on the spacecraft coordinate system; and(2) The direction for the earth or moon/planet based on the spacecraft coordinate
system.
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With regard to these items, the time variation profile for each attitude control phase or
the duration of the control phase and the amplitude range for the duration is defined.
The attitude control phase is mostly divided as follows.
(a) each phase of the position capture(b) each position at each intermediate orbit (transfer orbit of stationary spacecraft)(c) steady position(d) recapture at the abnormal position (in case of defined) (safe hold)5.1.5 Thermal environment for spacecraft
5.1.5.1 Special quality of thermo-optics
The special quality of thermo-optics for spacecraft is defined here and is an importantfactor for forming the thermal environment of spacecraft.
The special quality of thermo-optics for the skin of spacecraft is the primary factor todetermine the heat exchange between spacecraft and space. Specifically, the surface
temperature of spacecraft is determined by the balance between infrared radiation fromthe spacecraft to space and conversely, the absorbed amount of incident heat flux oforbital heat input from space to the spacecraft. So, in the design of thermal control ofspacecraft etc., the values of the following optical characteristics of each surface must be
selected properly.
(1) solar light absorptivity (αS)(2) infrared emissivity (ε)
Further, for the complex synthetic aspect, for example an aperture plane with thermal
blanket (MLI) and cavity, it is necessary to define the equal surface characteristics
converting to a simple surface below. In determining these values, it is necessary toclearly define and understand the conditions for which they will be used..
<1> effective solar light absorptivity (αSeff)<2> effective infrared emissivity (εeff)
The following are the considerations in the design of thermal control.
(1) solar light absorptivity (αS)
Solar light absorptivity αS (θ) is defined as the ratio of energy absorbed by the bodyrelative to the energy of the total solar light radiation incident to an arbitrary body with anangleθ. The incident angleθ is defined as the angle from the normal direction of thesurface of the body. The αS (θ) is defined as the following formula, where the spectral
radiant intensity of solar light is Js(λ) , the spectral absorption factor of the body for the
direction ofθ is α(λ,θ). Here, λ is the wavelength of the light.

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   ,    J   d    J   d


S

0

（1）



S

0

S

Generally, the solar light absorptivity is measured on the condition that the angle ofincidence is within 5°and the range of wavelength is 0.25~2.50 μm corresponding to 96%(the range of wavelength of 0.115~100.00μm is 100 %).
Furthermore, it is necessary to occasionally consider a range over 25μm ofwavelength.
(2) infrared emissivity (ε)
Generally, the infrared emissivity has the following 3 kinds of characteristic values:.
<1> whole hemisphere emissivity εH(T)<2> spectral hemisphere emissivity εH(λ, T)<3> vertical spectral emissivity εN(λ, θ, T)
The whole hemisphere emissivity εH(T) is defined as the ratio of the energy emitted
from the surface of a body with T representing surface temperature to hemisphericalspace with unit time and unit area relative to the energy emitted from the surface of a
black body with the same surface temperature as hemispherical space with unit time andunit area, and is expressed with following formula.

  , T E , T d T   E , T d  , T E , T d



0

H

b



H

0

b



0

H

b

T

（2）

4

Here,

 



 



spectralhemisphereemissive
W m2  menergy of black body
 H  , T : spectral hemisphere emissivity
Stefan-Boltzmann constant 5.67 X
:W m2  K4
10-8
Generally, the infrared emissivity is measured at a temperature of 300K and with arange of wavelength within 2.5-25μm. Furthermore, it is necessary to occasionallyconsider a wider range of wavelength.
Additionally, this value is temperature dependent. It is generally understood that thevalue usually obtained is around the normal temperature. However, it is necessary toconsider this temperature dependence when the surface temperature is not regarded as
normal.
E b  , T :

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Generally, attention must be paid to the value of hemisphere emissivity of the dielectric(insulator) surface when it is smaller than the value of vertical emissivity. With respect
to the general radiating surface controlled by the emissivity, the characteristics of heatradiation is the dielectric surface., However, as the value of emissivity obtained by thesimple procedure is mostly vertical emissivity, it is necessary to understand properly theorigin of the value to be used for the sizing design of the radiating surface.
(3) transmissivity (τ)
The transmissivity (τ) is the ratio of penetration energy to whole incident energy on thesurface with permeability for incidence and radiation. There are two kinds ofcharacteristic values as follow:<1> sunlight transmissivity (τS: transmissivity of sunlight spectrum)<2> infrared transmissivity (τIR: transmissivity of black radiation spectrum of generally
300K)
The reflectivity (ρ) is also used, except in terms of the characteristics about theradiation described above. This characteristic is defined for each solar light and infrared
radiation and is equal to absorptivity and emissivity. The following relation comes toexistence between reflectivity, absorptivity and transmissivity in terms of radiation andincidence at the same wavelength.
(reflectivity) + (absorptivity) + (transmissivity) = 1
In terms of reflectivity, as the ratio of ingredient of diffuse reflection and of regularreflection occurs, this value is defined as the specularity. Regarding this value, thefollowing two kinds are also defined according to incidence and radiation spectrum.
<1> solar light specularity (ρS)
<2> infrared specularity (ρIR)
Here, the definition of specularity is shown as the following equation.

（ingredien to fregular reflection）（specularit y）
（reflectivity）（ingredient of diffuse reflection） (ingredient of regular reflection）
As mentioned above, the thermo-optical properties on each surface change due todegradation and pollution from the thruster exhaust. It is therefore necessary to pay
attention to the range of fluctuation (mainly concerning solar light) caused by thedegradation and the pollution during its life time, in addition to its fabrication tolerance.
Moreover, it is necessary to consider the dispersion between the various measured
samples of αS and εH etc.

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5.1.5.2 Internal generation of heat by a spacecraft
The internal generation of heat by a spacecraft is caused by the operation of the eachcomponent. The component generating the internal heat is roughly classified into
electrical and electronic components, propulsion components, solar cell panel and othercomponents, according to its heat source. The following points regarding spacecraftcomponents must be considered for the thermal control design.
(1) electrical and electronic component
Included in this group are the components of the power supply system including
battery, communication system and attitude control system which require actuators.
Heat is generated by converting electrical energy to thermal energy during its operation.
Additionally, in case of the battery, the endothermic generation of heat caused bychemical reaction is added.
Generally, the caloric value of equipment changes according to the operating mode ofcomponent. The operating mode of each component is determined by the operating
mode of the spacecraft. Therefore, the following must be considered with respect toelectrical and electronic components on the spacecraft: .
<1> operating mode of spacecraft (what component, how it operates)<2> caloric value of each component by operating mode
The specific points of concern for each item are shown below.
<1> operating mode of spacecraftThough the component to be operated is determined at each phase during mission
operation, there are multiple alternatives about what component is operated,because of its redundant systems and multiple options for the mission operation.
Among them, the operating mode is considered as the worst case, selecting all
combined mode including the worst conditions of high/low temperature.
The following constraint conditions determine the operation alternatives of eachcomponent.
(a) whether component is to be operated at the same time or not to be operated at
the same time(b) whether there is a condition requiring a component to be used with theredundant system(c) whether there are constraints of mission operations (such as parallel operation,
operating duty and constraints on operation sequence etc.)<b> caloric value of each component in operating mode
As the caloric value of each component changes according to various conditions,the following items shall be defined.
14

JERG-2-310(a) classification of operating mode and caloric value of each mode(b) correlation equation in case the generation of heat depends upon temperatureand timing etc.
(2) propulsion component
The secondary propulsion system used for attitude control and trajectory control isanother component for consideration of heat generation. The secondary propulsion
system is classified to the tank storing the propellant and the plumbing (including
devices) transporting propellant to the thrusters.
With respect to the thermal design for the tank and plumbing (including devices), it isnecessary to consider maintaining the temperature of the propellant within the allowable
range, to avoid a rise or drop in the partial temperature distribution.
The thrusters are classified into two categories: chemical propulsion and non-chemical
propulsion (including electrical propulsion). Chemical propulsion creates propulsion by
generating the gas by chemical reaction or combustion, such as a liquid propellant
thruster. An electrical propulsion system, such as an ion engine, creates propulsion by
converting the electrical energy into kinetic energy.
These components are primarily mounted on the outside of the spacecraft, andfunction as the thermal interface design of the thermal control system design following its
confirmation of the feasibility of the exhaust heat design component.
The following items must be considered in terms of propulsion components.
<1> pollution [The plume of thruster exhaust] affects thermal control material<2> excessive heat inflow to spacecraft at soak back(3) solar cell panel
The solar cells attached to the solar cell panel supplies the electric power converted
from theincidental solar energy for the spacecraft. During this process, all of the energy notconverted to electric power becomes thermal energy.
In the thermal design of solar cell panel on orbit, careful consideration must be given
to the significant temperature changes in the external thermal environment caused by the
trajectory conditions of solar rays during periods of sunlight and shade and the efficiencyof power conversion of the solar cell during those periods.
The temperature level of the solar cell panel in-orbit changes according to the extent
of the power conversion of the incidental solar energy, and depends on the heat balanceof the incidental energy to the solar cell, the radiant energy in space and the generationof electric power.
Please refer to section2.1(6) “JERG-2-212 Design Standard of Wire Derating”
regarding the thermal design of the wire harness.

15

JERG-2-310(4) other components<1> A rotating component (e.g. an antenna with a driving motorized mechanism)
For a rotating component, which has a driving motorized mechanism for rotating on
the spacecraft, it is necessary to consider to the following items.
(a) The temperature stability pertaining to the driving motorized mechanism while itis rotating.
(b) The existence of temperature dependence pertaining to the torque performanceof the driving motorized mechanism(c) The secure exhaust of heat from the driving motorized mechanism (exothermicbody)(d) others<2> refrigerator
The refrigerator used in order to cool down the sensing port of the optical sensor etc.,
is classified into several categories according to its type of cooling mechanism. The
issues which must be address when mounting it to a spacecraft include themechanical environment during the launch and ascending phases andmeasurements to control vibration and the heat exhaust of the driving mechanism(e.g., the compressor etc.) while in-orbit. The refrigerator’s cooling mechanism must
reduce the temperature to the extent possible.
In terms of thermal control design, the following must be considered:(a) The exhaust heat must be designed to be compatible with the mechanicalenvironment(i.e. a secure exhaust heat pass etc.)(b) There must be built-in protection against thermal disturbance to the coolingportion (cold head)(including measures to mitigate the presence of heat)(c) There must be measures against thermal deformation to maintain alignment(d) There must be temperature stability of the refrigerator body for protection
against the thermal environment.
(e) others<3> observation equipment
In case of mounting observation equipment on spacecraft, the following items must
be considered with respect to the thermal design because the performance ofobservation equipment depends to a degree on the amount of thermal noise from thesensor itself. This includes the following considerations:(a) The temperature compensating for heat fluctuation while the observationequipment operating(b) The stability of the internal thermal environmental temperature of theobservation device
16

JERG-2-310(c) The measures designed to mitigate thermal deformation and to maintainalignment(d) others

17

JERG-2-310

5.2 Thermal design temperature conditions
5.2.1 Classification of temperature range
The temperature range of component is classified below.
<1> design estimated temperature range<2> maximum estimated temperature range<3> allowable temperature range
Fig. 5.2 shows mutual relationship of temperature range above mentioned. The
details are described hereafter.
allowable temperature range of component
maximum estimated temperature range

over

over
ΔT2

over

design estimatedtemperature range

ΔT2

over
ΔT1

ΔT1

nominalestimated tolerance

estimated tolerance

design margin

design margin

Estimated tolerance : tolerance caused by uncertainty of analysis technique and input parameter
Design margin : margin to correspond to the unforeseen phenomenon in advanceFig 5.2 Mutual relationship of allowable temperature range and estimatedFig. 5.2 Mutual relationship of allowable temperature range and estimated temperature range
temperature range
5.2.2 Design of the estimated temperature range
The design of the estimated temperature range is the nominal temperature range
considering the heat input fluctuation etc. caused by seasonal variation, operation anddegradation of thermal control devices. Specifically, it means the temperature range
obtained by thermal analysis considering the following factors.
(a) The seasonal variation and short-range variation of external heat source strength.
(b) The variation of heat input by the change of the spacecraft’s posture(c) The changes in the operating condition of component which brings about changes
in the output of heat (including degradation of a component’s heating element)(d) The deterioration of thermal control devices through age(e) The temperature changes to the thermal properties of materials uncertainties as aresult of contamination etc. which must be addressed using results from flighttesting(f) The change of the radiation boundary temperature (i.e. the temperature of a bodywhile releasing radiation which is exchanged with a component) and the
conductive interface temperature (i.e. the temperature of a body while releasing
18

JERG-2-310

conducted heat which is exchanged with a component)
5.2.3 Maximum estimated temperature range
The maximum estimated temperature range means the in-orbit maximum estimatedtemperature range which includes a predicted error of ∆T1 to become the “design
estimated temperature range”.
5.2.4 Allowable temperature ranges
The allowable temperature range is classified below according to the operating
condition and service condition of the component.
(a) The operational temperature range (i.e. temperature range for maintaining acomponent’s efficiency)
This is also the temperature range to satisfy its specifications and performance.
(b) The temperature range for maintaining a component’s function
This temperature range refers to a-drop in-orbit performance drops below its
specification, but the components still functions and satisfies its performancerequirements without any malfunction nor degradation during operation.
(c) inactive temperature range (preserved temperature range)
The temperature range not create a malfunction or permanent degradation
during inactivity while in orbit.
(d) Turn-on temperature range
The temperature range able to turn on the components and does not create a
non-recoverable degradation while in orbit.
5.2.5 Estimated tolerance ∆T1
5.2.5.1 Meaning of estimated tolerance ∆T1
Estimated tolerance ∆T1 means the temperature tolerance considering theuncertainty caused by analysis technique and input parameter for analysis.
Furthermore, estimated tolerance ∆T1 is able to be reviewed after evaluating
quantitatively the tolerance using thermal balancing testing etc. and assessing its validity.
(1) uncertainty caused by analysis technique<1>uncertainties of mathematical modeling: definition of node, approximation ofalbedo/infrared radiation of earth or moon/planet, approximation of form ofradiation plane/flow channel of heat conduction, modeling of MLI<2> uncertainties of analysis tool: uncertainty of Monte Carlo method/numericalintegration
method(2) uncertainty caused by input parameter for analysis<1> measurement error: uncertainty of thermal environment simulation, uncertainty ofmethod of measurement and measuring instrument
19

JERG-2-310<2> uncertainty regarding the calorific value of a component<3> fabrication tolerance: uncertainty of the MLI’s effective emissivity and conduct type<4> uncertainty regarding the amount of degradation due to age: uncertainty in theamount of material degradation
5.2.5.2 Procedure to set up estimated tolerance∆T1
The empirical orientation method is mainly used to establish estimated tolerance ∆T1
because of the uncertainty of analysis method and its parameters.
(1) empirical orientation method
Considers orbit, position, thermal control method and similarities based on past
resultsof spacecraft etc., ∆T1 establishes a value that is even based on the past experience.
(2) probability method
Performing error analysis for each parameter and searching for temperature error
regardingprobability, as defined below.
∆T1 = (absolute sum of errors) ･････worst evaluationor∆T1 = (RSS of errors) ･････to be set up so that errors can be addressed completely
independently
5.2.6 Design margin ∆T2
5.2.6.1 Meaning of design margin ∆T2 setting
Design margin∆T2 is set up to consider speculation using a demonstration test model
while including phenomena difficult to quantify.
5.2.6.2 Procedure to set up design margin ∆T2
For the design margin ∆T2 based on experience to be set up, the following items must
be considered.
(a) levels of past results (factoring in new developments, quality of past results)(b) criticality in case of deviating temperature range
Criticality caused by deviation(c) risk tolerance limit for project20

JERG-2-310

5.2.7 ∆T1 and ∆T2 of component thermally controlled in active type
5.2.7.1 Approach to thermal control
For defining ∆T1 and ∆T2 of a component (thermally controlled in active type) and ∆T1
and ∆T2 of a component (thermally controlled in passive type), the standards forconsideration are described below.
Concerning to the heat input Q (which includes self-heating) to a spacecraft in-orbit,
the temperature of the component must be maintained within the allowable temperature
range by setting up the thermal coupling using conductivity between the component andits surroundings.
Therefore, for the following formula, T is set up within the specified temperature range
within a variable range of Q by selecting the fixed value Q properly. Here, T0 is thetemperature limit (This is also a variable value).
Q=G∆T=G(T-T0)(5.1)

Fig. 5.2.1 shows the conceptual diagram of passive type thermal control.
Q
T

G

T0
Q

Target component
for thermal control

Conceptual diagram of passive type thermal control
OSR, heat insulating spacer etc.
Explanatory note of figure･ Though the relation between Q and T is notstraight line generally, treated as straight line
because of simplification to show its concept.

Heat input range

on orbit

Similarly, the active type thermal control also
treated as straight mentioned later.
・This shows that the variation of heat input on orbit
turns just like the variation of temperature.

So, theheat input range on orbit directly corresponds tothe design estimated temperature range.

T
Over ΔT1+ΔT2

Design estimated

Over ΔT1+ΔT2

Allowable temperature range

Fig 5.2.1 Conceptual diagram and design margin of passive type thermal control

21

JERG-2-310

5.2.7.2 Requirements for ∆T1 and ∆T2With respect to the aforementioned formula at section (5.1), T shall be set up within
designated temperature range by regarding Q or G, or both as variable.
The sample of G as variable is thermal louver and a variable conducting type of heatpipe (VCHP) etc., with Q as variable is thermal control by the heater.
The following diagram shows the active type thermal control concept and design
requirement.
(1) conceptual diagram of active type thermal control (in case of Q as variable [part 1])
when the spare control power is available only to the lower limit of the estimateddesign temperature range.
The sample diagram below is a thermal control heater to maintain the temperature oflower limit.

qConceptual diagram of active type thermal control
in case of Q variable thermostat heater etc.
T

G
Q

Target component
for thermal control

Q

TON：Temperature at thermostat turns ON
TOFF：Temperature at thermostat turns OFF

Maximum heat input
on orbitexample：NiH2 battery ON/OFF = 5/8℃

qtemperature TOFF becomesto be able to maintain

Sparecontrol
power

Active impression

Calorific value

q=0

By adding this value,

TON

TOFF

T

Temperature range of active type thermal control
Over ΔT2

Design estimated

Over ΔT1+ΔT2

temperature range

Allowable temperature range

Fig. 5.2.2 Conceptual diagram and design margin of active type thermal control(Q is variable)

22

JERG-2-310

Explanatory note of figure(1) This is conceptual diagram in case of thermal controlling the lower limit of allowabletemperature range by active type thermal control.
Specifically, this casecorresponds to the thermal control of a battery mounted to a high capacity
geostationary satellite and propulsion system (tank and plumbing) with a highallowable lower limit as compared to other components.
(2) For such a component, usually the size of the radiating surface is measured basedon the time of maximum generation of heat (maximum heat input) for the upper limit
temperature. Conversely, the lower limit temperature is necessary to maintain aminimum generation of heat with a heater.
Design requirement
Though the upper temperature limit is designed as a passive type thermal control, forthe following describes the lower temperature limit.
(1) Concerning the requirement for setting up the heater capacity, it should be larger thanboth the amount of calories able to rise from heater OFF temperature to over ∆T1℃
with the value multiplied to account for the safety factor and heater capacity
necessary to maintain the heater OFF position temperature.
(2) At the lower temperature limit with active type control, the designed estimatedtemperature range should be secured to the margin over ∆T2℃ of the allowabletemperature range. (Further, in some cases, ∆T2℃ secures its margin through theheater capacity.)
Example for reference
DRTS NiH2 battery･ allowable temperature (-10~25[35(*1)]℃) thermostat : ON/OFF = 5/8 (±1℃)(*1) at electric discharge and 3 hours after charge start･ in-orbit data
･ At trickle charging, it shall be maintained within the thermostat control temperature
range [5~8℃]･ At discharge, the temperature increase above the temperature of thermostat OFF(all heaters OFF)[approximately 18℃ at DC arc jet operation, approximately 10℃ at maximum
shade in equinox]

23

JERG-2-310(2) conceptual diagram of active type thermal control (in case of Q as variable [part 2])
when the spare control power is available for both the lower and the upper limit of theestimated design temperature range..
q

T

Conceptual diagram of active type thermal control
in case of Q variable thermostat heater etc.

G
Q

Target component
for thermal control

q

Q

Spare

control

By adding this value,
temperature TOFF becomesto be able to maintain

Active impression

caloric value

TON：Temperature at thermostat turns ON
TOFF：temperature at thermostat turns
example：propellant tank ON/OFF=21/24℃

q=0

TON

T

TOFF
Controlled

Over ΔT2

Design estimated

Over ΔT2

temperature range

Allowable temperature range

Fig. 5.2.3 Conceptual diagram and design margin of active type thermal control (Q isvariable)24

JERG-2-310

Explanatory note of figure(3) This is a conceptual diagram when there is thermal control to the upper and lower
limit of allowable temperature ranges by an active type of thermal control.
Specifically, for example, this is the case for the thermal control component for abattery and optical sensor with a narrow allowable temperature range. The
component that the allowable temperature range is narrower than the design margin(upper and lower limit) of the passive type of thermal design.
Design requirement
The upper temperature limit, principally the value over ∆T1+∆T2℃ from upper limit ofdesign temperature, should be secured (secure of radiating surface), when the active
type thermal control of heater etc. is turned to OFF.
The design margin of lower limit temperature is set up as follows.
<1> Concerning to the requirement for setting up heater capacity, it is the principal that itshould be larger than both the amount of calories able to rise from heater OFF
temperature to over∆T1 ℃ and the value multiplied to account for the safety factor andheater capacity necessary to maintain the heater OFF position temperature.
<2> At the lower temperature limit with active type thermal control, the design estimatedtemperature range should be the secured margin over ∆T2℃ to the allowabletemperature range. (Further, in some case, ∆T2℃ secures its margin through theheater capacity.)25

JERG-2-310(3) conceptual diagram of active type thermal control (in case of G as variable [part 1])
when the spare control power is available for both the lower and the upper limit of theestimated design temperature range
Q

Conceptual diagram of activity type thermal control
In case of G variable
Thermal louver, VCHP etc.

G

T

Q

T0
Target component
for thermal control

D

G=G2

QC

C

Controllable calorie range

Heat input range
on orbit

Spare control power

G=G1

QB

B
TB

A

TC

T

Controlled

Design estimated temperature range
Over ΔT2

Over ΔT2
Allowable

Fig. 5.2.4 Conceptual diagram and design margin of active type thermal control (G isvariable)
Explanatory note of figure･ The following is an explanation of the figure as an example of the thermal louver. (in principal)･ In the range between line segment AB, the blade is closed and the radiating surface is constant with minimum
emissivity (for example ε=0.1).
･ In the range between line segment CD, conversely the blade is hull open and the radiating surface is constantwithmaximum emissivity (for example ε=0.8).
･ In the range between line segment BC, ε changes from 0.1 to 0.8 according to the temperature of radiating
surface.
･ When the calorie variation is from QB to QC, the radiating surface is maintained within the controlled temperature
range from TB to TC.
･ At above figure, as the heat input range on orbit is between QB and QC, it indicates being maintained within thecontrolled temperature range.
･ Because of the factors that the temperature specified point and controlling sensing point is different and/or
setting tolerance of temperature of radiating surface and divergence angle of the blade, the range of estimateddesign temperature of the component does not necessarily correspond to the heat input range on orbit.

26

JERG-2-310

Design requirement<1> In principle, the spare control power (calorie) shall be the calorie to be able tochange the temperature over ∆T1℃ with the calorie from the upper and lower limit ofdesign estimated range.
<2> The estimated design temperature range shall be a secured margin over ∆T2℃ fromthe allowable temperature range.
(4) conceptual diagram of active type thermal control (in case of G as variable [part 2])
when the spare control power is not available for both the lower limit and the upper limitof the estimated design temperature range
Q
G=G2

Q

Controllable calorie range

Heat input range on orbit

T

G

T0
Target component
for thermal control

G=G1

T

ControlledtemperatureOver ΔT1+ΔT2Over ΔT1+ΔT2
Design estimated range
Allowable temperature range

Fig. 5.2.5 Conceptual diagram and design margin of active type thermal control (G isvariable)

Explanatory note of figure･ Same as Fig. 5.2.4, explanation of the figure using the example of the thermal louver. (in principal)･ Although in the previous figure, the heat input range on orbit was within the controllable calorie range ofthe louver, as compared to this figure which was over the range.
･ So, upper temperature limit is passively thermal controlled under the fixed value of ε=0.8 blade full open,
lower temperature limit is also passively thermal controlled under the fixed value of ε=0.1 blade close.
･ Therefore, at the upper and lower limits, there is similarity for the procedure to secure the margin of
temperature using passive type thermal control.

27

JERG-2-310

Design requirement<1> Because the heat input range in-orbit is over the controllable heat input range, theprocedure using the passive type thermal control is applied. That is, the estimateddesign temperature range shall be secured over the temperature margin from allowabletemperature range of prediction error of ∆T1℃ and temperature margin of ∆T2℃.
<2> Additionally, when there exists spare of control power at the upper or lower limits ofthe estimated temperature shown in Fig. 5.2.4, the active type thermal control shall beapplied.
5.2.8 Thermal control of thermally critical component
Because thermal control components are sensitive to temperature, it is necessary to
pay special attention to their thermal control design (e.g. battery, propellant plumbing of
attitude control system etc.) which are defined as a thermally critical components.
Normally, the following items are considered thermally critical components･ components with narrow allowable temperature ranges･ components with wider allowable temperature ranges which diverge significantly
from average temperature level of a usual spacecraft
As the thermally critical component is difficult to secure within the temperature margin(∆T1+∆T2) as a general component, it is required that they satisfy one of the following.
(1) to apply active thermal control(to keep margin not to temperature, but control ability, refer to 5.2.7)(2) to design the small prediction error (to adopt a validated/proven control system, to let
it thermally separate from others, and to isolate the influence of uncertainty fromother portions of the spacecraft,)28

JERG-2-310

5.3 Thermal interface
For the purpose of implementing the design of both a spacecraft system and its
components smoothly, the thermal interface requirements are defined below.
5.3.1 Related document with thermal interface(1) thermal interface condition document
This document defines the thermal interface condition of the system, necessary to
implement the thermal design of a spacecraft system and its components (equipment).
This document is prepared by the system side.
(2) interface control document
This document is called ICD, and defines the thermal interface condition of thecomponent necessary for the system side to implement the thermal design of aspacecraft system. This document is prepared by the component side based on the
instruction defined at section 5.3.1 (4).
(3) interface thermal mathematical modelThis is the information which properly expresses the component necessary to
implement the thermal analysis by the system side, in the format of a thermalmathematical model. The provision of this information is prepared using electronic data
and the document, based on the instructions defined at section (4) below. This
information shall be attached to the manual describing the contents of the thermalmathematical model. This thermal mathematical model and manual is prepared by thecomponent side.
(4) instruction for interface thermal mathematical model and ICD
This document describes the instructions to prepare the above interface thermalmathematical model and ICD by the component side, and is prepared by the systemside.
5.3.2 Thermal interface demarcation
5.3.2.1 Definition of thermal interface demarcation
The purpose is to define the most suitable border region for the thermal design
conditions of the spacecraft system and components, as well as to define the conditionsin this region.
This border region is named “Thermal Interface Demarcation”, and is principally thefollowing region. Concretely, it is described as the thermal interface condition document
and ICD of each component defined in section 5.3.1.
(1) internal onboard equipment(a) thermal interface demarcation concerning conductive heat coupling
will be placed at the clamp face on the system side and component side of eachcomponent.
Furthermore, the internal onboard equipment (which includes
insulated components) will be treated the same as exposed onboard equipment.
(b) thermal interface demarcation concerning radiation heat coupling29

JERG-2-310

In cases when the thermal design of a component is not implemented by the
information of conductive heat coupling, it shall be implemented as the need arises.
In terms of setting, it shall be the outer surface of each component.
(2) exposed onboard equipment(a) thermal interface demarcation concerning conductive heat coupling
will be placed on the clamp face on the system side and the component side ofeach component.
(b) thermal interface demarcation concerning radiation heat coupling
will be placed on the clamp face on the component side, on the surface ofcomponent and on the outer surface, opposite the spacecraft, along with othercomponents.
Fig.5.3 and Fig. 5.4 show the conceptual diagram of thermal interface demarcation ofinternal onboard equipment and exposed onboard equipment

To be implemented by component side
■ Thermal design of component

To be implemented by system side(including outer surface)■ Thermal control at specified point

Internal onboard
equipment

Clamp face

Thermal interface demarcation of component
Thermal interface demarcation of system side
Heat conduction
Fig. 5.3 Thermal interface demarcation of internally onboard equipment and spacecraft

30

JERG-2-310
To be implemented by systemside
To be implemented by component side

■ Set up of onboard thermal environmental■ Thermal design of component (including
outer surface)conditions■ Thermal control at clamp face of component■ Calculation and set up of on board heat

■ Set

input

surfaceandtemperatureofspacecraft■ Design of conduction and radiation■ Set up heat radiation conditions fromconductance to design heat flow underallowable

up

heat

exchange

spacecraftvolume(including another equipment) to exposed
onboard equipment■ Presentation of amount of allowable heat

Exposedonboardequipment

Clamp face

Thermal interface demarcation of component
Thermal interface demarcation of component
Heat radiation

Fig. 5.4 Thermal interface demarcation of exposed onboard equipment and spacecraft
5.3.2.2 Items to be defined as thermal interface conditions(1) thermal interface between system and internal onboard component
The following items shall be defined as the thermal interface between the spacecraft
and the internal on board component. The concrete contents of these items shall bedefined in ICD and prepared by the component side.
(a) geometries and contact area
Regarding the clamp face of a component, the diameter of a mounting hole, theposition of mounting hole, the shape of the part of clamp screw including thickness,dimensions and contact area shall all be defined. Regarding the outer surface except
the attaching portion, the shape and dimensions shall be specified in ICD.
(b) finishing treatment for thermal control
The clamp face of the component, employed material, surface treatment, planeroughness anddeviation from flatness shall be defined according to official standard as JIS and MIL
31

JERG-2-310

standard. Additionally, with respect to the outer surface except the clamp face, theemployed material, external surface treatment and emissivity shall be defined separately.
(c) calorific value
Calorific value of the component shall be defined with consideration to its operating
condition and include its tolerance..
(d) heat capacity
The value of heat capacity (mass X specific heat) shall be calculated for eachcomponent, and defined to include its tolerance.
(e) average heat generating density
The average heat generating density (W/cm2) of the clamp face of the component
shall be calculated according to the definitional equation indicated at 5.3.4.1 (5), anddefined.
(f) temperature specified point
As a general rule, the point of base plate (mounting flange etc.) shall be selected as atemperature specified point of the component, and the details of the position shall bedefined in the ICD.
(g) conditions for mounting
Meeting the demand for mounting, thermal filler or heat insulator shall be used as aninterface at the attaching portion of the component. The conditions regarding its
mounting shall be defined.
(h) active thermal control
Regarding the component’s applied active thermal control, control system, calorificvalue, mounting position of heater etc., the controlled temperature range and operating
method shall be defined.
(i) interface thermal mathematical model
When using the interface thermal mathematical model defined at 5.3.1 (3) as thethermal interface condition, the node splitting, the shape and dimensions of the outersurface node, the conduction /radiation combined network between nodes and heat
capacity of each node etc. shall be defined.
In creating this mathematical model, instruction for the interface of the thermalmathematical model shall be based on the model defined at section 5.3.1 (4).
(2) thermal interface between the system and the exposed onboard component
The thermal interface of an exposed onboard component is different from the interface
of an internal onboard component because of the thermal interface direct exposure withouter space, as well as because the thermal coupling of outer space is generally stronger
than the coupling of a spacecraft system. Moreover, large portions of the outer surfaceof the spacecraft cannot control freely, and exist under an exposed visual field etc.
Therefore, the thermal control design is mainly the thermal interface control for outerspace, which has the high independence. Moreover, most of the surface temperature of
the component is not uniform nor is it regarded as one node. In principle, the interface isdefined with the thermal mathematical model which includes the inside of the component.
32

JERG-2-310

5.3.3 Thermal interface design prepared by the system side
In implementing the thermal interface design of the component by system side, internal
onboard equipment is designed to be able to let calorific value emit from clamp face ofcomponent by thermal conduction.
Regarding the exposed component, the thermal balance of exposed in-orbit equipment
must be considered at the most suitable temperature, under the trajectory conditions of
the spacecraft, the installation condition of exposed onboard equipment with thespacecraft system and the thermal condition of internal heat generation etc. by exposed
onboard equipment.
The system side implements thermal analysis for acquiring the information andimplementing the interface thermal mathematical model into the system thermalmathematical model described in ICD and indicated by the component side, which
conducts the thermal design for the system.
5.3.4 Thermal interface design prepared by component side
The component side implements the thermal design analysis based on the thermalinterface condition document defined at 5.3.1 (1) as indicated by the system side, andconfirms the suitability of the thermal design of the component. Furthermore, theinterface thermal mathematical model (defined at 5.3.1 (3)) prepared by the component
side and provided to the system side, shall be validated as to its suitability through tests
and analysis implemented by the component side.
5.3.4.1 Requirements for internal onboard equipment
Internal onboard equipment shall be designed to satisfy the following requirements.
In case it does not satisfy these requirements, the component side shall coordinate withthe system side and specify the discrepancies to ICD.
(1) approach to radiation
In the thermal design of the internal onboard equipment, it shall be the premise that
self-heating is released by conductive heat transfer only from the clamp face to thespacecraft.
In cases when another heat abstraction is necessary without the conductive heattransfer from clamp face, the component side shall specify ICD after coordinating withthe system side, providing the interface thermal mathematical model defined at 5.3.1 (3)to the system side.
(2) homogenization of surface temperature
In the design of a component, the temperature, not only at the clamp face but also ateach section of the component casing including the clamp face, shall be homogenized.
(3) means of attachment
The component is joined to the system’s body structure by means of screw fastening inprinciple. Flanges or tabs are principally used along the component. The pitch of themounting hole for the component with the generation of heat shall be determined not only33

JERG-2-310

considering mechanical environmental conditions, but also self-heating properties,
because thermal contact resistance depends on the pitch of mounting hole.
The hole or clearance gap on the clamp face shall be prohibited from infiltrating thethermal filler, because sometimes thermal filler is applied during component installation.
The system side will determine if the thermal filler is applied or not.
(4) surface roughness and flatness
The surface roughness and flatness factors shall be specified as necessary to heattransfer at the plane of union of the subject component.
The recommendation value is provided as follows.
<1> surface roughness : better than 1.6 Ra “JIS B 0601” (Ra : arithmetical meanroughness)<2> flatness factor : under 0.001 mm/1mm
Flatness shall be calculated according to the following formula, and indicated by markto ICD. But flatness factor shall be defined by rising with 0.05mm pitch.
(flatness) = (flatness factor) X (length of the longest part of contact surface)(5) average heat generation density
The average heat generation density shall be under 0.06 W/cm2 for contact surface ofclamp face. But the definition of average heat generation density is as follows.
(average heat generation density) =(maximum heat generation: note 1)/(contact area of clamp face)
Note 1 : except maximum heat generation within 15 minutes
When the component does not satisfy this rule, there shall be coordination with thesystem side and specified to ICD.
(6) distributed heat source
The deviation of heat generation density of component shall be under ±50% of
average heat generation density. When the component does not satisfy this rule, there
shall be coordination with the system side and specify to ICD. The definition of
deviation of heat generation density is as follows.
(deviation of heat generation density) = (partial heat generation density – average heatgeneration density)/(average heat generation density)
The definition of partial heat generation density is as follows. (refer to Fig. 5.5)
(partial heat generation density) = (heat generation density corresponding to eachnode dividing clamp face of component in case of generation of thermal mathematicalmodel of component)

34

JERG-2-310

b1[cm]

Q1[W]

Q2[W]

Q3[W]

a1[cm]Q4[W]

Q = Q1+Q2+Q3+Q4(Q is calorific value getting away from each node to satellite panel at eachclamp face)qi[W/cm2]=Qi/(aiXbi)
Vl

d
i lh
i d i
Fig. 5.5 Definition of partial heat generation density(7) shape of clamp face
The clamp face of component shall not be prepared for counterbore. However, if thecomponent is under 0.06 W/cm2 of average heat generation density and under ±50% ofdeviation of heat generation density, this rule shall not be applied.
(8) surface treatment of clamp face
The clamp face of each component shall not be part of the surface treatment bringing
heat insulation such as painting. The surface treatment of the clamp face shall be
determined by evaluating level of generated corrosion (electrolytic corrosion) etc. causedby contact with dissimilar metals.
(9) surface treatment of outer surface except clamp face
Emissivity of outer surface shall be over 0.8 of hemispherical emissivity in principle.
In terms of area being restricted for surface treatment, the surface treatment of the range
and and the hemispherical emissivity shall be described to ICD beforehand.
(10) other cases
In case there are any concerns or doubts regarding the requirements for internal
onboard equipment, the system side and the component side shall coordinate witheach other.
5.3.4.2 Requirement for exposed onboard equipment
The special feature of exposed onboard equipment is its requirement to function andperform outside of the spacecraft exposed to outer space,, usually for observation rather
than to perform a thermal control function. The degree of freedom for its mounting
position is relatively limited. Occasionally a suitable arrangement for thermal control
purposes becomes an unsuitable position for observation purposes. Most importantly,
35

JERG-2-310

however, it is directly exposed to the hard space thermal environment.
The thermal control of exposed onboard equipment must perform taking into account abalance of the space thermal environment and self heating. Of primary importance isthe design of the external form and the suitableness of the position of primary thermal
control face. For this reason, the design of the internal structure for all exposed
onboard equipment shall be started at first.
The thermal control for the exposed onboard equipment often becomes untenablebecause of the rules affecting the clamp face and the easy adjustment of thermo-opticalproperties such as the internal onboard equipment. Sometimes there is a complicatedouter surface shape of exposed onboard equipment, or because of the mounting positionof exposed onboard equipment and system, the multiple reflection of external heat input
exists between the external surfaces of the component or between the component andthe system.
Therefore, it is necessary that the exposed onboard equipment side shall be designed
to fit in and around the thermal environment condition in the thermal design of the
exposed onboard equipment.
In principal, the thermal design shall be implemented according to the following
standards based on thermal interface condition document defined at 5.3.1 (1) andprepared by the system side.
(1) volume of heat exchange
The exposed onboard equipment shall be insulated against radiation and andconduction within system’s structure. However, the volume of heat exchanged shall bedefined using a suitable value by the system side.
(2) surface reflection properties
The surface shall diffuse solar light, albedo and earth and moon/planet infrared
radiation shine to the utmost. Unless there is a process for radiation treatment with
OSR etc., this shall be excluded.
(3) heater for temperature control
In principle, when there is thermal controlling of exposed onboard equipment with aheater, s the exposed onboard equipment shall have its own heater with a control circuit.
This will especially be the case when the temperature directly effects the efficiency of thecomponent (for example, when the temperature dependence of efficiency is large, andthere is interfere with the unit’s efficiency without a control circuit), the control circuit shallnevertheless be included.
However, when the thermal control or operation of the component is not considered
an obstacle, it may be possible to request the heater control function from the systemside. The system side may sometimes require mounting a survivable heater (including
thermostat) or replacement heater (i.e. a heater that generates heat equally to thecomponent) for the component side if necessary.
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JERG-2-310

5.3.5 Interface with propulsion system
It is necessary to address the propulsion system for the thermal interface, with theinterface requirements defined below.
5.3.5.1 Attitude control system
Mainly the attitude control system is composed by following component.
Attitude control system

Thruster module
Propellant feed system
Propellant tank
Plumbing
FilterFill and drain valve
Isolation valve
Measurement equipment

The following shall be considered as thermal interface items between spacecraft
system and attitude control system･ thermal characteristic value of component･ allowable temperature range of component･ caloric value of component･ electric power for thermal control of component･ thermal environmental conditions of component(1) thruster module
Because the thruster module generates so much heat and is affects the heat of aspacecraft both outside and inside it, the interface information shall be provided tospacecraft system side with the thermal mathematical model if necessary. The
following shall be considered as thermal interface items for the thruster module.
<1> thermal characteristic value : defined using the thermal mathematical model<2> allowable temperature range : to be defined for component classified without
injection (at inert and at stand by), at injection and at soak back heating<3> caloric value : to be defined using heat input (at injection, at no injection and at soakback heating) to spacecraft system<4> electric power value for thermal control<5> thermal environmental conditions : heat input from external environment(2) propellant feed system
Generally the components of the propellant feed system do not generate heat and is37

JERG-2-310

thermally controlled by a heater to keep it within allowable temperature range ofpropellants. The following shall be considered as thermal interface items of the propellant
feed system.
<1> thermal characteristic value : to be defined by the area of the base plate, materials,
flatness, surface roughness, geometries, infrared emissivity of outer surface and heat
capacity to each component<2> allowable temperature range : to be defined by the propellant and each component<3> electric power value for thermal control<4> thermal environmental conditions : to be defined by the inner temperature of thespacecraft system surrounding the component with the propellant feed system
Every definition shall be considered under the baseline of worst high temperature caseand worst low temperature case for the propellant system.
5.3.5.2 Large-sized liquid thruster
The following items shall be considered as the thermal interface of the large-sizedliquid thruster (to be called liquid engine hereafter) such as a thruster used for orbitaltransformation and the apogee engine.
<1> radiation heat input from liquid engine to spacecraft<2> conduction heat input from liquid engine to spacecraft<3> convection heat input from liquid engine to spacecraft<4> allowable temperature range<5> electric power for thermal control<6> contamination
The outline of each item is explained as follows.
(a) radiation heat input from liquid engine to spacecraft
The heat flux flowing into the spacecraft by the combustion of the liquid engine shall bedefined by thermal mathematical model, if necessary.
(b) conduction heat input from liquid engine to spacecraft
The heat flux flowing into the spacecraft through an attached portion of the spacecraft
during the liquid engine firing and after engine firing stops, shall be defined by thethermal conductivity and temperature change at the attached portion of the spacecraft.
using the thermal mathematical model, if necessary.
(c) convection heat input from liquid engine to spacecraft
The convection heating from the plume (backflow) from the combustion of the liquid
engine, shall be defined, if necessary. Plume shield shall be adjusted, if necessary.
(d) allowable temperature range
The allowable temperature range of the component containing the liquid engine shallbe required to have a thermal control design.
(e) electric power for thermal control
38

JERG-2-310

The electric power shall be set up based on the worst low temperature prior to the firingof the liquid engine. The elements are components for thermal control by heater.
･ propellant valve･ injector･ others(f) contamination
Contamination from the exhaust plume of a liquid engine shall be defined, if necessary.

39

JERG-2-310

5.3.5.3 Ion engine system
Ion engine system is one of the electrical propulsion systems, and the composition is
described below.
Ion engine system

Ion engine thruster
Power supply device
Propellant storage and feed system
Valve driving electronics
Thruster controller

It is necessary to consider the following thermal interface items between spacecraft
system and ion engine system.
<1> thermal characteristic value or thermal mathematical model of the component<2> allowable temperature range of the component<3> external heat input to the component<4> calorific value and heat soak-back of the component<5> electric power for the thermal control of the component
The thermal interface of the component for the ion engine system is described below.
(1) ion engine thruster
The ion engine thruster (called “thruster” hereafter) is a component which generates
high amount of heat that is exposed to outer space. Because it is basically isolated fromthe spacecraft, it has its own independent temperature control. The radiation on theexternal surface of spacecraft, the presence of heat from incidental solar light, and heatsoak-back etc. to the spacecraft through the clamp face are important parameters ofthermal control design for the spacecraft and the thruster. The interface information isprovided to the spacecraft system side with a thermal mathematical model, if necessary.
The following shall be necessary to consider as the thermal interface items for thethruster.
<1> thermal characteristic value<2> node tearing diagram<3> node heat capacity and calorific value<4> conduction resistance and radiation resistance between nodes<5> external heat input : amount of external heat input for each node
<6> thermal control heater : lower/upper limit temperature of heater calorific value<7> allowable temperature : allowable temperature range at inactive/turn-on/active<8> amount of heat soak back : amount of conduction/radiation soak-back(2) power supply device, thruster controller and valve driving electronics
The main feature of this power supply component is its generation of high amounts ofheat and heat density, and the large change of calorific value when it is active versus
40

JERG-2-310

when it is inactive. In spite of above features, because the power supply device islocated inside of spacecraft, it is necessary to consider an equal thermal interface of
these items to the other equipment on board the spacecraft. The thruster controller
and valve driving electronics have relatively small calorific value and it is possible tohandle it in the same manner as the other equipment on board the spacecraft.
Therefore, the thermal interface items of these types of components can be treated thesame way as the items described in section 5.3.2.2.
(3) propellant storage and feed system
The propellant storage and feed system stores the propellant and supplies the thruster.
The structure of the system, therefore, is different according to the kind of propellant
used. In the interface to the spacecraft, usually heat insulation to spacecraft isprocessed and thermal control by the heater is performed, especially for the thermal
control at low temperature. The following shall be considered as the thermal interface.
<1> thermal characteristic value : heat capacity of parts, area of attached portion,materials, flatness, surface roughness, shape of parts, level, dimensions, infraredemissivity of the surface and contact heat transfer coefficient.
<2> heat generation distribution : heat generation distribution of parts, heat generation
profile<3> active thermal control : controllable lower/upper limit and heat generation
distribution of thermal control heater<4> allowable temperature : allowable temperature range when it is active and when it
is inactive (definition of temperature specified point is necessary)
5.3.6 Thermal interface between system and launch vehicle
The following items shall be considered regarding the thermal interface between thespacecraft and the launch vehicle.
<1> air conditioned temperature inside of the fairing at launch site
<2> radiation heat from fairing<3> free molecular flow heating (after fairing separation)<4> heat conduction and radiation heat at separation section of the spacecraft<5> decompression profile in fairing(1) air conditioned temperature inside of fairing at launch site
Thermal environment in the fairing shall be controlled by an air conditioner while thespacecraft is encapsulated into the fairing.
(2) free molecular flow heating
It shall be set up using the free molecular flow heating rate after the fairing
separation.
(3) heat conduction and radiation heat at separation section of the spacecraft
It shall be set up with temperature history, heat transfer coefficient and emissivity ofeach section of the spacecraft separation part from launch to spacecraft separation.
41

JERG-2-310(4) decompression profile in the fairing
It is necessary to define the decompression profile in the fairing, considering toprevent the expansion and peeling of MLI.

42

JERG-2-310

5.4 Thermal control device/material technology
5.4.1 Thermal control technology
It is necessary to select the most suitable method of thermal control technology for the
spacecraft. The technology to maintain the spacecraft at a suitable temperature iscalled thermal control technology.
The requirements for the thermal control system is mentioned in section 5.4.1.1below,
and of the various types of thermal control methods are located in 5.4.1.2.
5.4.1.1 The requirements for the thermal control system
The function of thermal control system of the spacecraft is to maintain the componentof spacecraft within a specified temperature range and temperature rate change, fromlaunch and throughout its entire mission. The general requirements from the spacecraft
system side regarding mass, electrical power and reliability for the thermal control
system are described below.
(1) to be light weight and have a simple composition(2) to be small utilizing the necessary electric power to the utmost(3) to have stable performance and high reliability under the harsh environment of space,accounting such things as radial rays, ultra violet rays, atomic oxygen, high vacuum,
and high and low temperatures.
Fig. 5.4 shows the general classification of the thermal control device and material tosatisfy above requirement. The thermal control device/material is classified roughly into
passive type thermal control material, such as coating and multi layer insulation (MLI) etc.
and active type thermal control device, such as a heater, a heat pipe or a thermal louver
etc.

Passive typeCoating (paint, OSR, SLI, sheet)thermal control material

Separator
Thermal filler

ThermalMLI (multi layer insulation)

controlHeat sink thermal doubler
Fixed conductance type heat pipe
Active typeHeaterthermal control device

Variable conductance type heat pipe
Thermal louver
Fluid loop

Fig. 5.4 Classification of thermal control device/material

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JERG-2-310

5.4.1.2 Selection of thermal control method
The thermal control methods for spacecraft are classified into two types - the passive
type and the active type based on the thermal control device or material, which comprisethe thermal control system. The passive type thermal control method is the systemdesigned to control the temperature of the component by coordinating the route ofradiation and conduction with the physical characteristics (such as the shape) that are
inherent in the material.
Specifically, despite the change to its physical property, it
has high reliability. Additionally, the heat input on orbit and the change of calorific value
directly affects the temperature change of the component. The physical property hardly
changes the temperature as compared to heat input on orbit or the change of calorificvalue. The active type thermal control method is a system to control temperature of thecomponent using a technology which has mechanical and/or moving parts or by using
fluid, and which utilizes the change of heat with electric energy. The typical thermal
control device used for the active type thermal control method is a heater, thermallouver, variable conductance type heat pipe, or a fluid loop.
In spite of the attitude control method (spin or tri-axial control) of a spacecraft, thepassive type thermal control material such as coating or multi layer insulation isconsidered basic, while the active type thermal control device is complex and used for acomponent with extreme requirements of to maintain the temperature at a suitable
thermal control range. The following describe the various kinds of passive type thermal
control materials and active type thermal control devices.
[passive type thermal control material](1) coating (paint, OSR, SLI and sheet)･ Used on the outside of spacecraft structure, with white paint for heat insulation
control･ Used on the inside of spacecraft structure, with black paint for homogenization oftemperature in the interior of the structure･ Used on the radiating surface, with glass system rigid OSR, macro-molecularsystem flexible OSR and Teflon with silver deposits etc.
･ Used as thermal control tape with adhesives on film or sheet etc.
(2) separator･ Used as heat insulation control between spacecraft structure and component etc.
･ Used with materials with low thermal conductivity(3) thermal filler･ Used for heat transfer control between spacecraft structure and component etc.
･ Used for materials with high thermal conductivity(4) MLI･ Used for heat insulation control of spacecraft structure･ polyimide and Beta-cloth® etc. used as an external layer(5) heat-sink thermal doubler･ Used for thermal diffusion of high heat element
44

JERG-2-310(6) fixed conductance type heat pipe
･Used for transportation of heat exhaust from onboard equipment with high calorificvalue or heat generation density, such as relay equipment for telecommunication etc.
･ Used for homogenization of temperature of the spacecraft structure or panel for
onboard equipment･ Used for development of heat radiation capability by heat pipe and radiator[active type thermal control device](1) heater･ Used for thermal control of thermally critical component such as attitude control
system and battery･ heater generating heat similar to component when the component is OFF(2) variable conductance type heat pipe･ thermal control of high heat onboard equipment used to maintain within a fixed andnarrow temperature range(3) thermal louver･ Used for thermal control of spacecraft, external heat input changes for a relatively
short period such as a satellite in low earth orbit･ Used to control heat discharge from spacecraft, calorific value of internal onboard
equipment changes for a relatively short period･ thermal control of component used to maintain temperature in narrow temperature
range(4) fluid loop
Thermal control when long-distance thermal transportation is required, when the
ON/OFF of thermal transfer function is required, when gravity must be controlledduring the ground test or when flexibility of exhaust heat route is required.
5.4.2 Thermal control device/material
The criteria for selection and points to note regarding thermal control devices and/or
materials and its processes are described below.
5.4.2.1 Criteria for selection of thermal control devices and/or material
The thermal control device and/or material must satisfy the requirements for
performance, physical properties, reliability and environmental resistance for both thecomponent and the spacecraft and are generally selected according to following
sequence.
(1) recommended parts/entry item in list of material by public institution (such as
JAXA,NASA and ESA etc.).
(2) material produced domestically with an established evaluation for space use for JAXA
or with results, an/or with the possibility of getting technical documents confirming theircapabilities.
45

JERG-2-310(3) parts/material guaranteed or recommended by foreign spacecraft manufacturers orcompanies for which JAXA has confirmed its quality.
(4) parts/material with a newly prepared specification (or technical documents) andguaranteed quality and reliability have been conducted through evaluation and tests
etc.
5.4.2.2 Criterion for selection of process
The process for the selection of guaranteed thermal control parts and materials shallbe repeated as many times as is necessary to satisfy the performance, physicalproperty, reliability and environmental resistance requirements, and will be generally
selected according to sequence below.
(1) Selection according to standardized specification (or procedure manual) from a public
institution(2) Selection according to the already established specifications for space use(3) Selection according to a newly prepared specification and guaranteed repeatability by
evaluation and tests etc.
5.4.2.3 Points to note for applying the thermal control device/material process
The special points to note for the thermal control device/material process for acomponent and spacecraft are described below.
(1) traceability
The parts/material and process applied to spacecraft (PM, PFM and FM) shall include
a traceability of parts, material and its process by lot number etc. defined byparts/material specification and process specification or procedure manual.
(2) outgas
When outgas is expelled from the thermal control device/material, it clings to thesurface of a sensor and radiating face, causes characteristic degradation by pollution(contamination).
Therefore, the expulsion of outgas from the thermal control
device/material must occur as little as possible in space. Generally, the following value
is applied as the outgas requirement of parts/material. (refer to JMR-010 Standard forContamination Control)(a) total mass loss (TML)･･････････under 1%(b) collected volatile condensable materials (CVCM)･･････････under 0.1%
(3)cleanliness
The thermal control device/material for spacecraft use shall be controlled forcleanliness in all aspects of manufacturing/integration, testing, transportation, storage
and pre-launch operations. The target of cleanliness control includes the ambientenvironment of where the spacecraft is placed as well as the surface of thermal control
parts/material.
(a) cleanliness in clean room
The requirement for cleanliness during manufacturing/integration shall be generally
over ISO class 8 (equivalent to FED-STD-209E class 100,000).
46JERG-2-310(b) surface cleanliness of thermal control device/material
The surface cleanliness shall be defined by the size and number of particulate
residual material and mass (mg) of non-volatile residue, removed from area of thethermal control device/material per 1 square feet by solvent. (ISO14952-2)
Requirement for above (a) and (b) shall be defined for each spacecraft according tothermal control parts/material and work-in-progress.
(4) cleaning
When the surface of the thermal control device/material becomes contaminated,cleaning is performed as a part of cleanliness control. At this time, consideration forcleanliness of the cleaning tool and solvent, as well as the compatibility of the solvent forsurface of thermal control parts/material etc., shall be considered.
(5) refurbishment
Thermal control device/material of the spacecraft, damaged or contaminated byhandling after testing or before launch, shall be refurbished.
However, therefurbishment shall not damage the effectiveness of the design and tests that have been
performed until that time.
(6) radiation-resistant characteristic
The surface characteristics (thermo-optical properties and light transmittance) andmechanical strength of the thermal control device/material deteriorates because of age
and because of radial rays (mainly electrons and protons), ultraviolet rays and the heat
cycle. The surface characteristics of the thermal control parts/material are required tohave minimal deterioration due to age and be of stable quality.
For thermal control
devices/materials of the spacecraft projected to orbit at a low altitude (200-700 km)
above earth, extra care shall be taken because of expected degradation from atomic
oxygen.
(7) environmental characteristics for heat resistance
The surface temperature of antenna and solar panel exposed in outer space,fluctuates between about -160 ~ 100℃. Therefore, the mechanical properties of thethermal control materials used for the antenna and panel, such as paint and adhesive
agents are not desirable because they cannot significantly affect this wide temperature
range. Though paint and adhesive agents are classified roughly into categories such as
epoxy system, acrylic, urethane system, silicon system and inorganic system, they shallbe used according to a usable environment (operating temperature range) and according
to the strength of their adhesive bonding.
(8) grounding
When the integration of thermal control device/material to the spacecraft occurs, thethermal control parts/material shall be grounded to the spacecraft structure to prevent
electrification or an electric discharge, as well as to avoid electromagnetic interference.
The steps to be taken for grounding is shown below.
(a) conductive coating or conductive paint(b) conductive adhesive(c) ground strap
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JERG-2-310

There will be a requirement for grounding resistance which shall be defined individuallyfor each spacecraft.
(9) service life
The service life of paint, adhesive and tape etc. is limited, and therefore must meet or
exceed the service life of the spacecraft. A longer service life is desirable. .
(10) corrosion protection
Beryllium, silver, aluminum and magnesium are necessary and suitable treatments forcorrosion protection.
(11) metals with significant outgas
As cadmium and zinc hold relatively high levels of outgas, these metals shall not be
used for space parts/materials.
(12) germanium
As germanium degrades easily from moisture in the air, the temperature in the clean
room shall be carefully controlled and special measures for storage shall be taken(13) ITO
When ITO (indium-tin oxide) coating is applied, it shall be carefully handled so as not todamage the ITO coating, which is used for securing conductivity (or resistance to atomic
oxygen) on the surface of the thermal control material. Because ITO is transparent andany damage is difficult to detect, careful attention is required, to prevent abrasion or
bending. Also, because it has a low adhesion quality with materials containing fluorinesuch as Teflon with silver deposits it can easily be peeled off.
5.4.2.4 Production of thermal control parts/material
The system side and on board component side shall conduct coordination andmanufacturing of hardware for the thermal control parts/materials as necessary forthermal control.
5.4.2.5 Handling of MLI(1) designed to prevent MLI from peeling off
For the purpose of preventing MLI from peeling off, the design of MLI shall be
conducted according to section 2.1 (1) “JERG-2-311 Design Standard to prevent peelingof MLI”.
(2) precautions regarding the design of MLI
The system and onboard equipment shall be designed with consideration for bothexperiencing a loss of interface with MLI. When the MLI must be installed over aninterface demarcation point, intervention by the system side and by the on board
component side shall be started as quickly as possible, because it may be a complicatedprocedure and may take significant amount of time. The result of intervention shall be
described to ICD. In principle, ground handling of MLI shall be performed based on
section 2.1 (2) “JERG-2-211 Design Standard for electrical charge/discharge”.

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5.5 Thermal design verification
The thermal design shall be verified as conforming to the requirement of specified
performance, as described below.
(1) Thermal analysis or test shall verify that the results of the design adapts to therequired function and performance.
(2) Though verification by test is desirable, there must be an explanation if anothermethod is used for verification.
(3) Verification shall be performed to the component and the spacecraft system.
5.5.1 Verification by thermal analysis
In the verification by thermal analysis, conformity to the design requirement shall beverified using suitable analysis modeling and performance prediction.
In the verification by thermal analysis, every condition of test environment and test
configuration restricted by constraint of test facility etc. shall be included.
5.5.2 Verification by test
Verification by test means verification of conformity to the design requirement, placing
the spacecraft system and components under the conditions which re-create the actual
environment that is expected. The demonstration of requirement for thermal control
design is called “thermal balance test”, and the demonstration of requirement for
behavior of spacecraft is called “thermal vacuum test”. Classification of these tests isshown in section 5.5.2.1.
During testing, the test condition shall be similar to the acceptance level or qualificationlevel, depending upon its required verification level.
5.5.2.1 Classification of tests
Thermal vacuum tests of spacecraft are classified into the thermal balance test whichis performed simulating the thermal environment on orbit for the purpose of confirming
the appropriateness of thermal control design of the spacecraft system and components,
and the thermal vacuum test, which is performed for the narrower purpose of proving thatthe spacecraft system and components can function and perform under the thermal
vacuum condition.
■ thermal balance test
■ thermal vacuum test
■ thermal vacuum test
(thermal vacuum test in narrow sense)
The standard regarding the thermal vacuum test of spacecraft shall follow 2.1 (3)“JERG-2-002 Standard of General Tests for Spacecrafts”, to establish the thermal
control design standard, while the thermal balance test is performed for the purpose ofvalidating the thermal control system.

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5.5.2.2 Approach regarding the selection of thermal balance test conditions
Although the thermal balance test is conducted to confirm the validity of the thermal
control design and the performance of thermal control hardware, also it is also used toconfirm the validity of thermal mathematical model applied during the design phase. The
following shall be generally be considered for the selection of test conditions.
(a) Using a test condition so that the thermal control design that can be easily
evaluated･ including spacecraft configuration and thermal environmental conditions during its
operation phase･ including worst conditions concerning both the external and internal thermal
conditions･ including thermally critical conditions which are peculiar to the specific spacecraft･ including suitability during steady state/non-steady state test conditions(b) Using a test condition that thermal control hardware that can be easily evaluated･ including being able to evaluate the performance of insulation, coating, heat sink,thermal louver, heat pipe and heater etc.
(c) Using a test condition that thermal mathematical model that can be easily evaluated･ including being able to evaluate uncertain parameters of the thermal mathematicalmodel(d) Using a test condition to confirm external heat input during a thermal vacuum test･ including being able to validate external heat input conditions by setting up atthermal vacuum test in a narrow sense
5.5.2.3 Procedure and selection of thermal balance test
The thermal vacuum test is conducted by placing the spacecraft system andcomponents into a space chamber and simulating the high vacuum low temperatureenvironment of outer space using the gas exhaust in the space chamber and cooling of
the shroud. The environmental condition in space chamber during the thermal balance
test, shall be set according to the value shown in Table 5.5.2 in principal.
Table 5.5.2 Environmental condition in space chamber at thermal vacuum test
Item
Pressure
Shroud Temperature

Value
-3

Under 1.3 X 10 Pa *1 (1X10-5 Torr)
Under 100 K (-173 ℃)

Note : *1 for the purpose to simulate heat transfer appearance in outer space
A thermal sensor shall be installed at each section of the spacecraft system andcomponent to monitor the temperature and provide thermal data. The onboard heat
input simulation during the thermal balance test shall be performed according to theheating method described below.

50

JERG-2-310(1) solar method(2) IR method (IR lamp method, IR panel method and IR shroud method)(3) skin heater method(4) A combination of the above three methods
The irradiation intensity or heat input by these heating methods shall be monitored
using a radiation intensity meter or calorie meter or wattmeter etc. The heating method
to test the spacecraft during the thermal balance test shall be set up as the most suitable
method according to purpose of test, and considering the following.
(1) simulation accuracy is necessary (especially for thermal balance test)(2) the influence of gravity (for testing with a heat pipe etc., the properties of whichchange according to the direction of gravity)(3) Consideration for the influence on the spacecraft etc. according to the establishmentof test configuration (e.g.in cases when a skin heater is installed on a spacecraft
directly etc.)(4) facility capability(5) cost for test
5.5.2.4 Evaluation items
For the purpose of an evaluation to be performed after thermal balance test of thespacecraft system etc., following items and contents shall at least be included.
Regarding the sequence of an evaluation for the thermal balance test, there is first an
evaluation performed for the setting up of test conditions, while there is a continuous
performance evaluation of the thermal control hardware and an evaluation of thermalmathematical model as follows.
(a) evaluation of setting up of test condition･ evaluation of environment of the space chamber (degree of vacuum andtemperature of shroud)･ evaluation of set up of external heat input (including heat input from jig)･ evaluation of set up of internal heat input･ evaluation of set up of boundary temperature･ evaluation of measurement system (temperature, degree of vacuum and external
heat input)(b) performance evaluation of the thermal control hardware･ performance evaluation of the thermal control hardware such as insulation, coating,heat sink, thermal louver, heat pipe and heater etc.
(c) evaluation of the thermal mathematical model･ evaluation of the thermal mathematical model by comparing the estimated to theactually measured temperature･ evaluation of uncertain parameters of thermal mathematical model(d) evaluation of the thermal design･ direct thermal design evaluation from the measured temperature when test condition
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JERG-2-310

is equal･ thermal design evaluation from flight prediction temperature by evaluated thermalmathematical model, when the test condition is different from flight condition(e) evaluation of thermal vacuum test (narrowly defined)･ validation of external heat input condition(f) evaluation of contamination･ evaluation of the existence of a contamination occurrence･ evaluation of the contamination source･ evaluation of the influence of contamination on the spacecraft

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