ISO 24412:2023

INTERNATIONAL ISO
STANDARD 24412
First edition
2023-01
Space systems — Thermal vacuum
environmental testing
Systèmes spatiaux — Essais environnementaux sous vide thermique
Reference number
© ISO 2023
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Published in Switzerland
ii
Contents Page
Foreword .v
Introduction . vi
1 Scope . 1
2 Normative references . 1
3 Terms and definitions . 1
4 Symbols and abbreviated terms.2
5 Test purpose . 2
5.1 Thermal balance test . 2
5.2 Thermal vacuum test . 3
5.2.1 General purpose . 3
5.2.2 Qualification test . 3
5.2.3 Proto-flight test . 3
5.2.4 Acceptance test . 3
6 Test methods . 3
6.1 Thermal balance test . 3
6.1.1 Test description . 3
6.1.2 Test conditions . 6
6.1.3 Basic requirements of test facilities . 7
6.1.4 Monitoring during TBT . . 7
6.2 Thermal vacuum test . 7
6.2.1 Test description . 7
6.2.2 Test conditions . 10
6.2.3 Basic requirements for test facilities . 13
6.2.4 Monitoring during TVT . 13
7 Test facility . .13
7.1 Laboratory environment . 13
7.2 Laboratory infrastructure . 14
7.3 Test system . 14
7.3.1 Overview . 14
7.3.2 Chamber system . 14
7.3.3 Vacuum system . 15
7.3.4 Thermal system .15
7.3.5 Data acquisition system . 18
7.3.6 MGSE . 18
7.3.7 Contamination measurement and control system . 18
8 Test requirements .19
8.1 Test tolerance and accuracy . 19
8.2 Test configuration. 19
8.3 Temperature and heat flux measurement . 20
8.3.1 General .20
8.3.2 Location of temperature monitoring point for test article .20
8.3.3 Location of temperature monitoring point for test equipment .20
8.4 Heating device selection .20
8.5 Safety requirements and recommendations . 21
9 Test procedure .21
9.1 Test flow . 21
9.2 Test procedure . 21
9.2.1 General . 21
9.2.2 Before test . 22
9.2.3 Test implementation . 23
9.2.4 After test .23
iii
10 Test interruption and handling .24
10.1 Interruption . 24
10.1.1 Test facility malfunction . 24
10.1.2 Test article malfunction . 24
10.2 Interruption handling . 24
11 Test documentation .24
Annex A (informative) Main characteristic of a solar simulator .25
Annex B (informative) An example of IR heater design flow for absorbed flux simulation
method in TBT .27
Bibliography .30
iv
Foreword
ISO (the International Organization for Standardization) is a worldwide federation of national standards
bodies (ISO member bodies). The work of preparing International Standards is normally carried out
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ISO collaborates closely with the International Electrotechnical Commission (IEC) on all matters of
electrotechnical standardization.
The procedures used to develop this document and those intended for its further maintenance are
described in the ISO/IEC Directives, Part 1. In particular, the different approval criteria needed for the
different types of ISO documents should be noted. This document was drafted in accordance with the
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This document was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles,
Subcommittee SC 14, Space systems and operations.
Any feedback or questions on this document should be directed to the user’s national standards body. A
complete listing of these bodies can be found at www.iso.org/members.html.
v
Introduction
The on-orbit environments of spacecraft, with their vacuum state, cryogenic and black background, and
complex heat transfer, are harsher and more complex than the ground environment. They have a strong
impact on the success of spacecraft mission. Thermal balance tests (TBT) and thermal vacuum tests
(TVT) at spacecraft level are conducted to ensure the units in spacecraft operate normally in specified
pressure and thermal range.
This document provides methods and specifies general requirements for spacecraft level thermal
balance tests and thermal vacuum tests. However, the technical requirements in this document can
be tailored by the parties for some special spacecraft, such as manned vehicle, deep space explorer,
extra-terrestrial body lander or the satellites with emphasis on low-cost and fast delivery, which are
characterized by extensive use of non-space-qualified commercial-off-the-shelf (COTS) units.
This document acts as a supplement to ISO 15864 and ISO 19683. It is applicable to test project
designers and test organizations. It also serves as a reference for spacecraft designers and test facility
manufacturers.
vi
INTERNATIONAL STANDARD ISO 24412:2023(E)
Space systems — Thermal vacuum environmental testing
1 Scope
This document provides methods and specifies general requirements for spacecraft level thermal
balance tests (TBT) and thermal vacuum tests (TVT). It also provides basic requirements for test
facilities, test procedures, test malfunction interruption emergency handling and test documentation.
The methods and requirements can be used as a reference for subsystem-level and unit-level test article.
2 Normative references
The following documents are referred to in the text in such a way that some or all of their content
constitutes requirements of this document. For dated references, only the edition cited applies. For
undated references, the latest edition of the referenced document (including any amendments) applies.
ISO 15864:2021, Space systems — General test methods for spacecraft, subsystems and units
ISO 17566:2011, Space systems — General test documentation
3 Terms and definitions
For the purposes of this document, the following terms and definitions apply.
ISO and IEC maintain terminology databases for use in standardization at the following addresses:
— ISO Online browsing platform: available at https:// www .iso .org/ obp
— IEC Electropedia: available at https:// www .electropedia .org/
3.1
maximum predicted temperature
highest temperature that can be expected to occur during the entire life cycle of the subsystem (3.4)/
unit (3.8) in all operational modes plus an uncertainty factor
3.2
minimum predicted temperature
lowest temperature that can be expected to occur during the entire life cycle of the subsystem (3.4)/unit
(3.8) in all operational modes plus an uncertainty factor
3.3
spacecraft
integrated set of subsystems (3.4) and units (3.8) designed to perform specific tasks or functions in
space
3.4
subsystem
assembly of functionally related units (3.8), which is dedicated to specific functions of a system
3.5
thermal balance test
test conducted to verify the adequacy of the thermal model and the adequacy of the thermal design
3.6
thermal uncertainty margin
temperature margin included in the thermal analysis of units (3.8), subsystems (3.4) and spacecraft
(3.3) to account for uncertainties in modelling parameters such as complex view factors, surface
properties, contamination, radiation environments, joint conduction and interface conduction and
ground simulation
3.7
thermal vacuum test
test conducted to demonstrate the capability of the test item to operate according to requirements in
vacuum at predefined temperature condition
Note 1 to entry: Temperature conditions can be expressed in terms of temperature level, gradient, variation and
number of high-low temperature cycles.
3.8
unit
lowest level of hardware assembly that works with specified complex electrical, thermal and/or
mechanical functions
4 Symbols and abbreviated terms
AT acceptance test
EGSE electrical ground support equipment
FM flight model
IR infrared
MGSE mechanical ground support equipment
OSR optical solar reflector
PFT proto-flight test
QT qualification test
TBT thermal balance test
TQCM temperature-controlled quartz crystal microbalances
TVT thermal vacuum test
UPS uninterruptible power supply
UV ultraviolet
5 Test purpose
5.1 Thermal balance test
The purpose of the thermal balance test is to provide the data necessary to verify the analytical
thermal model and demonstrate the ability of the spacecraft thermal control subsystem to maintain
the specified operational temperature limits of the units throughout the entire spacecraft.
5.2 Thermal vacuum test
5.2.1 General purpose
The purpose of the thermal vacuum test is to demonstrate the ability of the test item and its units to
meet the design requirements under vacuum conditions and temperature extremes that simulate those
predicted for flight. TVT detects material, process and workmanship defects that would respond to
vacuum and thermal stress conditions.
The test level and test duration are described in 6.2.2.1 and 6.2.2.2 respectively.
5.2.2 Qualification test
During the qualification test (QT), the thermal vacuum test serves to validate the performance of the
qualification model (QM) in the intended environments with the specified qualification margins.
5.2.3 Proto-flight test
During the proto-flight test (PFT), the thermal vacuum test serves to validate the performance of the
proto-flight model (PFM) on the first flight in the intended environments with the specified proto-flight
margins.
5.2.4 Acceptance test
During the acceptance test (AT), the thermal vacuum test serves to validate the performance of the
flight model (FM), except the one used as pro-flight, in the intended environments with the specified
acceptance margins.
6 Test methods
6.1 Thermal balance test
6.1.1 Test description
The on-orbit external thermal flux simulation can be conducted by one of the following methods:
a) Incident flux method
The intensity, spectral content and angular distribution of the incident solar, albedo and planetary
irradiation encountered by on-orbit spacecraft are simulated by using solar simulator system, shown in
Figure 1 or using the other method (e.g. with axial location of solar simulator).
The solar simulator is composed of the xenon lamp, the filter and the collimator. Generally, the test
article is installed on a motion simulator (rotating platform) to simulate the different attitudes on orbit.
For the requirements of a solar simulation system, see 7.3.4.5. For the main characteristic of a solar
simulator, see Annex A.
Key
1 shroud 2 motion simulator 3 test article
4 solar simulator 5 vacuum chamber 6 collimator
Figure 1 — Solar simulation method
This method is suitable for spacecraft with complex shapes and large differences in surface thermal
characteristics. It can provide incident illumination with matching spectral, uniformity and stability
of irradiance, divergence angle for the thermal test of the spacecraft. However, it is difficult to simulate
the effects for performance degradation of thermal control coatings at end of lifetime. This method may
be restricted for the effect of reflection light or heat from surfaces of shroud and MGSE, large operating
cost and heat pipes on-board normally working horizontally.
b) Absorbed flux method
The absorbed solar, albedo and planetary irradiation for on-orbit spacecraft, are simulated by using
infrared (IR) heaters (cage, lamps, calrods and thermal plate) with their spectrum adjusted to the
external thermal coating properties, or by using film heaters attached to spacecraft surfaces with the
absorbed heat flux controlled by electrical power, shown in Figure 2. For the requirements for IR heater
and film heater, see 7.3.4.3 and 7.3.4.4. Annex B describes the design flow of an IR heater in the absorbed
flux method in TBT.
Key
1 vacuum chamber 2 shroud 3 IR cage or IR thermal plate
4 test article 5 IR lamp/calrod array
Figure 2 — Absorbed flux method
This method is suitable for spacecraft with simple shapes and similar in surface thermal characteristics.
It has the advantage of high reliability, low manufacturing and operation cost. It may be restricted for
the containment released from MGSE, limited temperature ramp and the numbers of heating loops or
electrical power.
c) The combination of methods a) and b)
The combination of the methods a) and b) can be used for heat flux simulation of different surfaces of
the test article in TBT.
Generally, the following shall be considered during test article design:
— The profile, structures, materials, instrument and device layout, cable network, various thermal
control measures, envelop dimension, surface state, installation and connection mode, internal heat
sources, thermal capacity shall meet the requirements of thermal design and simulation.
— The thermal simulation model of spacecraft or its units may be designed specially, whose thermal
capacity and heat consumption are in accord with that on orbit.
— The large antenna, solar array and other external components may not participate in the test, but
their radiation heat effects shall be evaluated. Conduction heat shall be simulated on installation
interfaces by proper heat insulation, heat leakage compensation, or constant temperature.
— Additional radiation flux created by thermal vacuum chamber, MGES and heating devices frames
shall be taken into account.
— If the natural convection effects cannot be ignored under the ground gravitation condition,
pressurized cabin convection boundary shall be simulated by adjusting the gas temperature,
pressure and velocity on the units’ surface to ensure the heat transfer is equivalent.
— The propellant tank is filled with protective gas.
6.1.2 Test conditions
6.1.2.1 Test cases design
TBT cases depend on the mission, spacecraft design, spacecraft operational modes, and times required
to reach stabilization. According to the internal heat source heating mode, orbital heating mode and
other thermal boundary conditions, there are four types of operating cases.
a) Case 1
Internal heat source, simulative orbital heating and other thermal boundary conditions are
constant;
b) Case 2
Internal heat source works in a set periodic change mode, while the simulative orbital heating and
other thermal boundary conditions are constant;
c) Case 3
Internal heat source works in a set periodic change mode; the simulative orbital heating and other
thermal boundary conditions are in the periodic orbit change mode;
d) Case 4
Internal heat source, simulative orbital heating mode or other thermal boundary conditions are in
the aperiodic change during the specified phase.
For b) and c), the cyclic test for several periods can be repeated either with the heat source operating
mode and simulative orbital heating mode in one orbit period until the temperature of test model
is steady periodically, or with several orbit periods as one test period until the temperature of test
model is steady periodically.
The design principles of the test cases are as follows.
— Test phases shall simulate cold and hot conditions to verify all aspects of the thermal hardware and
software, including heater operation, radiator sizing, and critical heat transfer paths.
— Test cases shall obtain sufficient critical parameters required for thermal analytical model
verification and flight mission indication.
— To validate the adequacy of the thermal control design, the cases shall contain hot case and cold
case at least. Consideration should be given for testing an “off-nominal” case such as a safehold or a
survival mode.
— Generally, the test for the only purpose of verifying thermal analytical model shall contain transient
case.
— Transient case shall be set when the influence of on-orbit heat flux or other thermal boundary
conditions on spacecraft temperature increases with time.
6.1.2.2 Temperature stabilization
The exposure shall be long enough for the test article to reach temperature stabilization so that
temperature distributions are ensured in the steady-state conditions. The test temperature shall be
considered as stabilized, in case that
a) temperature monitored at the test article is within the allowed tolerance around the specified test
temperature;
b) temperature change rate is lower than the value allowed for stable conditions.
Steady-state conditions shall be defined in test specification. The temperature fluctuation should be
within ±0,5 °C over 4 h; or monotonous change should be less than 0,1 °C/ h over 4 h. Meanwhile the
fluctuation of other temperature points can be used as a reference.
6.1.3 Basic requirements of test facilities
-2
a) The test pressure should be no higher than 1,33 × 10 Pa.
b) The shroud surface temperature should be no higher than 100 K.
c) The distance between testing equipment and a test item shall ensure:
— convenience while performing preparation and completion operations with a test item;
— availability of required uniformity of heat fluxes, incident on a test item surface when performing
tests.
d) The shroud surface shall be painted with high-emissivity black coating whose solar absorption
ratio shall be higher than 0,95 and hemispheric emissivity shall be higher than 0,9.
e) The recommendations in a) and b) should be reassessed according to the specified elements such as
external and internal thermal and pressure environment, operational modes of spacecraft and its
units, and flight mission.
6.1.4 Monitoring during TBT
The test article shall be operated and monitored throughout the test. Functional tests shall be
conducted before, during, and after the test for flight model. Sufficient and timely measurements shall
be made on the major internal and external units to verify the major units’ thermal design, hardware,
and analyses. The heat flux, temperature, unit’s operation mode and other performance parameters
shall be controlled to meet the requirements of the specified case.
The modification of the thermal analytical model is applicable to all test cases. The modification
parameters shall be within the acceptable range. After modification of the thermal analytical model,
the modification parameters shall be configured to the thermal analytical model to indicate the
temperature of spacecraft flying on orbit.
After the test, a comprehensive analysis shall be made on energy balance in test cases for test error
sources. The absorbed and irradiated heat by the test model shall be compared, whose difference is
generally controlled within ±10 %. Test errors generally are derived from limitations of the heat flux
simulation mode, deviation between the test model and actual spacecraft, measurement accuracy of
heat flux and temperature.
6.2 Thermal vacuum test
6.2.1 Test description
Spacecraft shall be placed in a thermally controlled vacuum chamber having the capability to expose
the test article at or beyond the minimum and maximum test temperatures.
The following should be considered.
a) Units of spacecraft should be flight products (except qualification test).
b) Some units may be replaced by qualification parts, process parts or simulation parts with their
thermal performance and electrical performance parameters conforming to the test requirements.
c) Large units such as large antenna and solar array may not participate in the test, or may be set
apart from spacecraft in test with cable connection.
d) The propellant tank is generally filled with protective gas.
The temperature can be controlled by the following two measures.
— Controlled by test facility
1) Temperatures of shroud and heating devices shall be controlled by adjusting the flow of gas/
liquids or the electrical power to assure that the test article reaches the required temperature.
2) The temperature on main functional areas (e.g. light entrance, OSR radiation surface) should
be consistent with the extreme external temperature on orbit.
3) The surface protrusion, surface coating, cables temperature change shall be taken into
consideration during adjusting to prevent the test article from being damaged.
— Controlled by inner heat source
Internal heat sources include heat consumption of active units or active control heaters.
1) In cooling down the test article, its units may be powered off or work in a minimum power
consumption state.
2) In heating up the test article, its units may be powered on and work in a maximum power
consumption state.
3) Temperature control thresholds of the active temperature controller should be extended at
both ends when the unit temperature is required to be reached faster.
4) Temperature should be monitored for temperature-sensitive units during the thermal cycling
process.
The temperature profile for TVT is shown in Figure 3. The durations of thermal soak, thermal
stabilization and thermal dwell for different units of spacecraft depend on their operation modes,
heat inertia characteristics, lifetime, etc. Functional/performance testing should be performed after
adequate thermal stabilization during thermal soak.
Key
X time Y temperature 1 hot case
2 ambient temperature 3 cold case 4 thermal soak
5 thermal dwell 6 thermal stabilization 7 functional/performance testing
8 hot test temperature 9 within test tolerance
Figure 3 — Temperature profile for testing
Consideration should be given to conducting the thermal balance test in conjunction with the thermal
vacuum test program. A combined test is often technically and economically advantageous. It shall,
however, satisfy the requirements of both tests. Examples for combined test temperature profiles are
shown in Figure 4, where TBT and TVT are performed in sequence, and in Figure 5, where TBT hot and
cold cases are integrated into the first TVT cycle. TBT may be performed after TVT in order to test the
thermal-cycling effect on thermal interfaces.
Key
X time Y temperature 1 ambient temperature
2 TB hot case 3 TB cold case 4 TV hot case1
5 TV cold case1 6 TV hot case2 7 TV cold case2
8 TV hot case3 9 TV cold case3 10 TV hot case4
11 TV cold case4
Figure 4 — TBT and TVT are carried out successively
Key
X time Y temperature 1 ambient temperature
2 TB hot case 3 TV hot case1 4 TB cold case
5 TV cold case1 6 TV hot case2 7 TV cold case2
8 TV hot case3 9 TV cold case3 10 TV hot case4
11 TV cold case4
Figure 5 — TBT is integrated in first TVT cycle
6.2.2 Test conditions
6.2.2.1 Test levels
In TVT, the temperature of all units to be verified can be confirmed on the basis of the temperature
result from TBT, or the temperature result predicted by the thermal model analysis, as shown in
Figures 6 and 7. TVT temperature margin of unit is determined in Table 1 according to the test level.
Key
1 TBT result 2 AT margin 3 maximum TBT temperature
4 minimum TBT temperature 5 AT margin 6 AT temperature range
7 PFT margin 8 PFT margin 9 PFT temperature range
10 QT margin 11 QT margin 12 QT temperature range
Figure 6 — Temperature range and margin of TVT based on TBT result
Key
1 AT margin 2 maximum predicted temperature 3 minimum predicted temperature
4 AT margin 5 PFT margin 6 AT temperature range
7 PFT margin 8 QT margin 9 PFT temperature range
10 QT margin 11 QT temperature range
Figure 7 — Temperature range and margin of TVT based on predicted temperature analysis
Table 1 — Determination of TVT temperature margin
Test Margin Margin
level
based on TBT result based on predicted temperature analysis
QT 5 °C to 10 °C beyond acceptance tem- 5 °C to 10 °C beyond acceptance tempera-
perature ture
PFT 0 °C to 5 °C beyond acceptance temper- 0 °C to 5 °C beyond acceptance tempera-
ature ture
AT 0 °C to 5 °C beyond the maximum and 3 °C to 11 °C as thermal uncertainty mar-
minimum temperatures of TBT gin beyond thermal model analysis
The thermal uncertainty margin shall be evaluated by importance, complexity, type, operation
temperature of units, and whether the TBT has been carried out or not.
For passive thermal control, the thermal uncertainty margin is a temperature added to worst-case
temperature predictions. For active thermal control, the thermal uncertainty margin is a power margin
to ensure thermal control stability. When the margin is added to worst-case temperature predictions,
the resulting temperature forms the basis for the acceptance temperature range.
Generally, for units that adopt active thermal control, the minimum thermal uncertainty margin
can be reduced to ±3 °C to ±5 °C. For units that meet more uncertain operational or environmental
conditions, the thermal uncertainty margin may be larger than ±7 °C to ±11 °C. Lower thermal margins
may be adopted for units with an allowable operating temperature below −170 °C or above 120 °C.
Specific estimated uncertainty values are set in a spacecraft thermal system specification. For active
thermal control units at the cold extreme, 25 % excess heater control authority is used in lieu of 11 °C
temperature margins. For unit controlled by heat pipes, the heat transfer margin should be considered
if one of the heat pipes fails.
Test temperatures are restricted as follows.
a) The temperature of the working units should not exceed the temperature range for the
corresponding test levels; and the temperature range of the non-working units should not exceed
the storage temperature range.
b) The upper and lower temperature limits in the same thermal control zone shall be within the
temperature envelope of the test units.
c) The temperature ramp rate should be equal to or higher than the maximum predicted ramp rate. In
case that the test system is not capable to achieve the ramp rates, rates may be lower at no expense
of the test objective and test validity.
6.2.2.2 Test duration
The number of thermal cycles and the duration of thermal soak shall be specified with consideration of
mission requirements.
Generally, the numbers of cycles for spacecraft are:
a) QT: 8 cycles;
b) PFT: 4 cycles;
c) AT: 4 cycles.
The thermal soak is suggested to be over 8 h at each temperature extreme during the first and last
cycle, and over 4 h for the intermediate cycles.
Test temperature shall be considered as stabilized, in case:
— the test article temperature is within the allowed test tolerance at the specified test temperature;
— the temperature change rate is less than 3 °C per hour.
It is allowed to shorten the measure time or reduce the cycles for some special units such as short-time
or disposable products. Their temperatures are not required to reach the stability value. They shall be
sufficiently tested before the subsystem TVT or unit TVT.
6.2.3 Basic requirements for test facilities
-2
a) The test pressure shall be no higher than 1,33 × 10 Pa.
b) The shroud temperature and heating devices shall be controlled to meet the requirements for the
maximum and minimum temperature and ramp rate of the test article.
c) The size of the thermal vacuum chamber shall be sufficient to safely perform installation activities
of the test article and MGSE.
-4
d) For spacecraft with high-voltage and high-power units, the test pressure should be at 10 Pa level
due to the internal and external pressure difference.
e) For some special spacecraft such as Mars lander, the test pressure shall be redetermined according
to specified mission environment.
6.2.4 Monitoring during TVT
a) The performance and function of the test article shall be tested in hot and cold temperature
stabilization.
b) Before, during and after the TVT, functional tests shall be conducted and the results meet its
specification.
c) Within thermal soak at cold and hot test temperature of each cycle, the performance of the test
article shall be tested in each operational mode that is foreseen during its service life.
d) If parts of units are sensitive to vacuum environment, vacuum gauges shall be mounted inside the
spacecraft. These units shall not be powered unless pressure is within acceptable range.
e) Low-pressure discharge testing (to simulate the pressure decrease during launch) should be
conducted during the pumping or pressurization process.
f) Temperatures of high-power units during test deserve special attention to avoid damage from
excessive temperature.
g) Thermal sensors for testing shall be installed as indicated by the thermal analysis for monitoring
and acquiring thermal data.
h) Backup units and redundant circuits shall be tested at hot and cold temperature. Their test duration
shall be the same as that of the primary units and shall be equally distributed in each cycle.
i) Collect and analyse the test environment data, including vacuum degree, shroud temperature,
composition and content of released gas, contamination fluctuation.
j) If required, the motion property or thermal deformation should be measured or monitored when
checking mechanism or thermal structure.
7 Test facility
7.1 Laboratory environment
The laboratory environment shall meet the requirements of the specified test article, which include
temperature, humidity, cleanliness, etc.
7.2 Laboratory infrastructure
The laboratory infrastructure supports the test system operation and test article handling by providing
e.g. water, pressurized air, power, lifting and transport means and communication interfaces.
The laboratory infrastructure shall provide all means to safely:
a) operate the test system;
b) handle the test article;
c) conduct tests.
7.3 Test system
7.3.1 Overview
TBT and TVT shall be performed in a thermal vacuum chamber.
An example for a typical composition of a thermal vacuum chamber is shown in Figure 8. A thermal
vacuum chamber usually includes:
a) chamber system;
b) vacuum system;
c) thermal system (cryogenic/heating);
d) data acquisition system;
e) mechanical spacecraft ground support equipment (MGSE);
f) contamination measurement and control equipment.
Figure 8 — Typical composition of a test system
7.3.2 Chamber system
a) The chamber shall provide sufficient number and performance of feedthrough connectors for
measurement channels, power supply, EGSE communication, to comply with test requirements.
b) Cables and connectors permanently used inside the chamber shall be suitable for use in extreme
high and low temperature, and vacuum environment.
c) Leakage of flanges and connector/pipe feedthroughs shall be minimized.
d) Temperature sensors on shrouds shall be sufficient and distributed uniformly to ensure
temperature measurement accuracy. The measurement acquisition period for chamber parameters
should not exceed 1 min.
2 2
e) Heat flux from sink background shall be less than 10 W/m to 30 W/m based on the relative sizes
of the tested articles and a vacuum chamber.
f) The chamber shall provide a test article grounding interface. The grounding resistance should be
less than 1,0 Ohm.
7.3.3 Vacuum system
a) The vacuum system shall be capable of achieving and maintaining the vacuum pressure required
for testing under consideration of leakage, heat devices, MGSE, test article. Pressures as low as
-3 -4
10 Pa to 10 Pa are readily obtainable. The vacuum system should be entirely dry and should
have “dry” vacuum pumps both in its fore-vacuum part (e.g. dry screw and mechanical booster
roots-type vacuum pumps) and high-vacuum part (e.g. turbomolecular and cryogenic vacuum
pumps).
b) The pressure change rates should simulate the pressure environment change during the spacecraft
ascent and descend.
c) the vacuum system should provide turbomolecular pumps to evacuate some released gas during
testing like hydrogen and helium, provide mass spectrometer to measure the partial pressures
of each composition.
d) Pressurization of the chamber is accomplished with dry nitrogen. This allows the chamber to be
returned to ambient pressure at any time that the contaminant cleaning equipment and all major
equipment in the chamber are above the minimum allowable temperature of the satellite. Moisture
condensation is preven
...

ISO/FDIS 24412:2022(E)
ISO TC 20/SC 14/WG 2
Date: 2022-09-1910-25
Secretariat: ANSI/AIAA
Space systems — Thermal vacuum environmental testing

ISO/FDIS 24412:2022(E)
All rights reserved. Unless otherwise specified, or required in the context of its implementation, no
part of this publication may be reproduced or utilized otherwise in any form or by any means,
electronic or mechanical, including photocopying, or posting on the internet or an intranet, without
prior written permission. Permission can be requested from either ISO at the address below or ISO’s
member body in the country of the requester.
ISO copyright office
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CH-1214 Vernier, Geneva
Phone: +41 22 749 01 11
Email: copyright@iso.org
Website: www.iso.org
Published in Switzerland
ii © ISO 2022 – All rights reserved

ISO/FDIS 24412:2022(E)
Contents
Foreword . v
Introduction . vi
1 Scope . 1
2 Normative references . 1
3 Terms and definitions . 1
4 Symbols and abbreviations . 2
5 Test purpose . 2
5.1 Thermal balance test . 2
5.2 Thermal vacuum test . 2
5.2.1 General purpose . 2
5.2.2 Qualification test . 3
5.2.3 Proto-flight test . 3
5.2.4 Acceptance test . 3
6 Test methods . 3
6.1 Thermal balance test . 3
6.1.1 Test description . 3
6.1.2 Test conditions . 5
6.1.3 Basic requirements of test facilities . 6
6.1.4 Monitoring during TBT . 7
6.2 Thermal vacuum test . 7
6.2.1 Test description . 7
6.2.2 Test conditions . 9
ISO/FDIS 24412:2022(E)
6.2.3 Basic requirements of test facilities . 11
6.2.4 Monitoring during TVT . 11
7 Test facility . 12
7.1 Laboratory environment . 12
7.2 Laboratory infrastructure . 12
7.3 Test system . 12
7.3.1 Overview . 12
7.3.2 Chamber system . 13
7.3.3 Vacuum system . 13
7.3.4 Thermal system . 14
7.3.5 Data acquisition system . 16
7.3.6 MGSE . 17
7.3.7 Contamination measurement and control system . 17
8 Test requirements . 17
8.1 Test tolerance and accuracy . 17
8.2 Test configuration . 18
8.3 Temperature and heat flux measurement . 18
8.3.1 General . 18
8.3.2 Location of temperature monitoring point for test article . 19
8.3.3 Location of temperature monitoring point for test equipment . 19
8.4 Heating device selection . 19
8.5 Safety requirements . 19
9 Test procedure . 20
9.1 Test flow . 20
9.2 Test procedure . 20
9.2.1 Before test . 20
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ISO/FDIS 24412:2022(E)
9.2.2 Test implementation . 21
9.2.3 After test . 22
10 Test interruption and handling . 22
10.1 Interruption . 22
10.1.1 Test facility malfunction . 22
10.1.2 Test article malfunction . 23
10.2 Interruption handling . 23
11 Test documentation . 23
Annex A (informative) Main characteristic of a solar simulator . 24
Annex B (informative) An example of IR heater design flow for absorbed flux simulation
method in TBT . 26
Bibliography . 30
ISO/FDIS 24412:2022(E)
Foreword
ISO (the International Organization for Standardization) is a worldwide federation of national
standards bodies (ISO member bodies). The work of preparing International Standards is normally
carried out through ISO technical committees. Each member body interested in a subject for which a
technical committee has been established has the right to be represented on that committee.
International organizations, governmental and non-governmental, in liaison with ISO, also take part in
the work. ISO collaborates closely with the International Electrotechnical Commission (IEC) on all
matters of electrotechnical standardization.
The procedures used to develop this document and those intended for its further maintenance are
described in the ISO/IEC Directives, Part 1. In particular, the different approval criteria needed for the
different types of ISO documents should be noted. This document was drafted in accordance with the
editorial rules of the ISO/IEC Directives, Part 2 (see www.iso.org/directives).
Attention is drawn to the possibility that some of the elements of this document may be the subject of
patent rights. ISO shall not be held responsible for identifying any or all such patent rights. Details of
any patent rights identified during the development of the document will be in the Introduction and/or
on the ISO list of patent declarations received (see www.iso.org/patents).
Any trade name used in this document is information given for the convenience of users and does not
constitute an endorsement.
For an explanation of the voluntary nature of standards, the meaning of ISO specific terms and
expressions related to conformity assessment, as well as information about ISO's adherence to the
World Trade Organization (WTO) principles in the Technical Barriers to Trade (TBT), see
www.iso.org/iso/foreword.html.
This document was prepared by Technical Committee ISO/TC 20, Aircraft and space vehicles,
Subcommittee SC 14, Space systems and operations.
Any feedback or questions on this document should be directed to the user’s national standards body. A
complete listing of these bodies can be found at www.iso.org/members.html.
vi © ISO 2022 – All rights reserved

ISO/FDIS 24412:2022(E)
Introduction
The on-orbit environments of spacecraft, with their vacuum state, cryogenic and black background, and
complex heat transfer, are harsher and more complex than the ground environment. They have a strong
impact on the success of spacecraft mission. Thermal balance tests (TBT) and thermal vacuum tests
(TVT) at spacecraft level are conducted to ensure the units in spacecraft operate normally in specified
pressure and thermal range.
This document provides methods and specifies general requirements for spacecraft level thermal
balance tests and thermal vacuum tests. However, the technical requirements in this document can be
tailored by the parties for some special spacecraft, such as manned vehicle, deep space explorer, extra-
terrestrial body lander or the satellites with emphasis on low-cost and fast delivery, which are
characterized by extensive use of non-space-qualified commercial-off-the-shelf (COTS) units.
This document acts as a supplement to ISO 15864 and ISO 19683. It is applicable to test project
designers and test organizations. It also serves as a reference for spacecraft designers and test facility
manufacturers.
FINAL DRAFT INTERNATIONAL STANDARD ISO/FDIS 24412:2022(E)

Space systems — Thermal vacuum environmental testing
1 Scope
This document provides methods and specifies general requirements for spacecraft level thermal
balance tests (TBT) and thermal vacuum tests (TVT). It also provides basic requirements for test
facilities, test procedures, test malfunction interruption emergency handling and test documentation.
The methods and requirements can be used as a reference for subsystem-level and unit-level test
article.
2 Normative references
The following documents are referred to in the text in such a way that some or all of their content
constitutes requirements of this document. For dated references, only the edition cited applies. For
undated references, the latest edition of the referenced document (including any amendments) applies.
ISO 15864:2021, Space systems — General test methods for spacecraft, subsystems and units
ISO 17566:2011, Space systems — General test documentation
3 Terms and definitions
For the purposes of this document, the following terms and definitions apply.
ISO and IEC maintain terminology databases for use in standardization at the following addresses:
— ISO Online browsing platform: available at https://www.iso.org/obp
— IEC Electropedia: available at https://www.electropedia.org/
3.1
maximum predicted temperature
highest temperature that can be expected to occur during the entire life cycle of the subsystem
(3.4)/unit (3.8) in all operational modes plus an uncertainty factor
3.2
minimum predicted temperature
lowest temperature that can be expected to occur during the entire life cycle of the subsystem (3.4)/unit
(3.8) in all operational modes plus an uncertainty factor
3.3spacecraft3
spacecraft
integrated set of subsystems (3.4) and units (3.8) designed to perform specific tasks or functions in
space
3.4subsystem4
subsystem
assembly of functionally related units (3.8), which is dedicated to specific functions of a system
ISO/FDIS 24412:2022(E)
3.5
thermal balance test
test conducted to verify the adequacy of the thermal model and the adequacy of the thermal design
3.6
thermal uncertainty margin
temperature margin included in the thermal analysis of units (3.8), subsystems (3.4) and spacecraft (3.3)
to account for uncertainties in modelling parameters such as complex view factors, surface properties,
contamination, radiation environments, joint conduction and interface conduction and ground
simulation
3.7
thermal vacuum test
test conducted to demonstrate the capability of the test item to operate according to requirements in
vacuum at predefined temperature condition
NOTENote 1 to entry: Temperature conditions can be expressed asin terms of temperature level, gradient,
variation and number of high-low temperature cycles.
3.8
unit
lowest level of hardware assembly that works with specified complex electrical, thermal and/or
mechanical functions
4 Symbols and abbreviated terms
AT acceptance test
AT acceptance testelectrical ground support equipment
EGSE
FM flight model
IR infrared
MGSE mechanical ground support equipment
OSR optical solar reflector
PFT proto-flight test
QT qualification test
TBT thermal balance test
TQCM temperature-controlled quartz crystal microbalances
TVT thermal vacuum test
UPS uninterruptible power supplyultraviolet
UV
UV ultraviolet
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ISO/FDIS 24412:2022(E)
5 Test purpose
5.1 Thermal balance test
The purpose of the thermal balance test is to provide the data necessary to verify the analytical thermal
model and demonstrate the ability of the spacecraft thermal control subsystem to maintain the
specified operational temperature limits of the units throughout the entire spacecraft.
5.2 Thermal vacuum test
5.2.1 General purpose
The purpose of the thermal vacuum test is to demonstrate the ability of the test item and its units to
meet the design requirements under vacuum conditions and temperature extremes that simulate those
predicted for flight. TVT detects material, process and workmanship defects that would respond to
vacuum and thermal stress conditions.
The test level and test duration isare described in subclause 6.2.2.1 and 6.2.2.2 respectively.
5.2.2 Qualification test
During the qualification test (QT), the thermal vacuum test serves to validate the performance of the
qualification model (QM) in the intended environments with the specified qualification margins.
5.2.3 Proto-flight test
During the proto-flight test (PFT), the thermal vacuum test serves to validate the performance of the
proto-flight model (PFM) on the first flight in the intended environments with the specified proto-flight
margins.
5.2.4 Acceptance test
During the acceptance test (AT), the thermal vacuum test serves to validate the performance of the
flight model (FM)，), except the one used as pro-flight, in the intended environments with the specified
acceptance margins.
6 Test methods
6.1 Thermal balance test
6.1.1 Test description
The on-orbit external thermal flux simulation can be conducted by one of the following methods:
a) Incident flux method
The intensity, spectral content and angular distribution of the incident solar, albedo and planetary
irradiation encountered by on-orbit spacecraft are simulated by using solar simulator system, shown in
Figure 1 or using the other method (e.g.,. with axial location of solar simulator).
The solar simulator is composed of the xenon lamp, the filter and the collimator. Generally, the test
article is installed on a motion simulator (rotating platform) to simulate the different attitudes on orbit.
For the requirements of a solar simulation system, see 7.3.4.5. For the main characteristic of a solar
simulator, see Annex A.
ISO/FDIS 24412:2022(E)
Key
1 shroud 2 motion simulator 3 test article
4 solar simulator 5 vacuum chamber 6 collimator
Figure 1 — Solar simulation method
This method is suitable for spacecraft with complex shapes and large differences in surface thermal
characteristics. It can provide incident illumination with matching spectral, uniformity and stability of
irradiance, divergence angle for the thermal test of the spacecraft. However, it is difficult to simulate the
effects for performance degradation of thermal control coatings at end of lifetime. This method may be
restricted for the effect of reflection light or heat from surfaces of shroud and MGSE, large operating
cost and heat pipes on-board normally working horizontally.
b) Absorbed flux method
The absorbed solar, albedo and planetary irradiation for on-orbit spacecraft, are simulated by using
infrared (IR) heaters (cage, lamps, calrods and thermal plate) with their spectrum adjusted to the
external thermal coating properties, or by using film heaters attached to spacecraft surfaces with the
absorbed heat flux controlled by electrical power, shown in Figure 2. For the requirements for IR heater
and film heater, see 7.3.4.3 and 7.3.4.4. TheAnnex B describes the design flow of an IR heater in the
absorbed flux method in TBT can be referred to Annex B.
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ISO/FDIS 24412:2022(E)
Key
1 vacuum chamber 2 shroud 3 IR cage or IR thermal plate
4 test article 5 IR lamp/calrod array
Figure 2 — Absorbed flux method
This method is suitable for spacecraft with simple shapes and similar in surface thermal characteristics.
It has the advantage of high reliability, low manufacturing and operation cost. It may be restricted for
the containment released from MGSE, limited temperature ramp and the numbers of heating loops or
electrical power.
c) The combination of methods a) and b)
The combination of the methods a) and b) can be used for heat flux simulation of different surfaces of
the test article in TBT.
Generally, the following shall be considered during test article design:
— The profile, structures, materials, instrument and device layout, cable network, various thermal
control measures, envelop dimension, surface state, installation and connection mode, internal heat
sources, thermal capacity shall meet the requirements of thermal design and simulation.
— The thermal simulation model of spacecraft or its units may be designed specially, whose thermal
capacity and heat consumption are in accord with that on orbit.
— The large antenna, solar array and other external components may not participate in the test, but
their radiation heat effects shall be evaluated. Conduction heat shall be simulated on installation
interfaces by proper heat insulation, heat leakage compensation, or constant temperature.
— Additional radiation flux created by thermal vacuum chamber, MGES and heating devices frames
shall be taken into account.
— If the natural convection effects cannot be ignored under the ground gravitation condition,
pressurized cabin convection boundary shall be simulated by adjusting the gas temperature,
pressure and velocity on the units’ surface to ensure the heat transfer is equivalent.
— The propellant tank is filled with protective gas.
ISO/FDIS 24412:2022(E)
6.1.2 Test conditions
6.1.2.1 Test cases design
TBT cases depend on the mission, spacecraft design, spacecraft operational modes, and times required
to reach stabilization. According to the internal heat source heating mode, orbital heating mode and
other thermal boundary conditions, there are four types of operating cases.
a) Case 1
Internal heat source, simulative orbital heating and other thermal boundary conditions are
constant;
b) Case 2
Internal heat source works in a set periodic change mode, while the simulative orbital heating and
other thermal boundary conditions are constant;
c) Case 3
Internal heat source works in a set periodic change mode; the simulative orbital heating and other
thermal boundary conditions are in the periodic orbit change mode;
d) Case 4
Internal heat source, simulative orbital heating mode or other thermal boundary conditions are in
the aperiodic change during the specified phase.
For b) and c), the cyclic test for several periods can be repeated either with the heat source
operating mode and simulative orbital heating mode in one orbit period until the temperature of
test model is steady periodically, or with several orbit periods as one test period until the
temperature of test model is steady periodically.
The design principles of the test cases are as follows.
— Test phases shall simulate cold and hot conditions to verify all aspects of the thermal hardware and
software, including heater operation, radiator sizing, and critical heat transfer paths.
— Test cases shall obtain sufficient critical parameters required for thermal analytical model
verification and flight mission indication.
— To validate the adequacy of the thermal control design, the cases shall contain hot case and cold
case at least. Consideration should be given for testing an “off-nominal” case such as a safehold or a
survival mode.
— Generally, the test for the only purpose of verifying thermal analytical model shall contain transient
case.
— Transient case shall be set when the influence of on-orbit heat flux or other thermal boundary
conditions on spacecraft temperature increases with time.
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ISO/FDIS 24412:2022(E)
6.1.2.2 Temperature stabilization
The exposure shall be long enough for the test article to reach temperature stabilization so that
temperature distributions are ensured in the steady-state conditions. The test temperature shall be
considered as stabilized, in case that
a) temperature monitored at the test article is within the allowed tolerance around the specified test
temperature;
b) temperature change rate is lower than the value allowed for stable conditions.
Steady-state conditions shall be defined in test specification. The temperature fluctuation should be
within ±0,5 °C over 4 h; or monotonous change should be less than 0,1 °C/ h over 4 h. Meanwhile the
fluctuation of other temperature points can be used as a reference.
6.1.3 Basic requirements of test facilities
-2
a) The test pressure should be no higher than 1,33 × 10 Pa.
b) The shroud surface temperature should be no higher than 100 K.
c) The distance between testing equipment and a test item shall ensure:

- — convenience while performing preparation and completion operations with a test item;
- — availability of required uniformity of heat fluxes, incident on a test item surface when performing
tests.
d) ShroudThe shroud surface shall be painted with high-emissivity black coating whose solar
absorption ratio shall be higher than 0,95 and hemispheric emissivity shall be higher than 0,9.
e) The requirementsrecommendations in a) and b) should be reassessed according to the specified
elements such as external and internal thermal and pressure environment, operational modes of
spacecraft and its units, and flight mission.
6.1.4 Monitoring during TBT
The test article shall be operated and monitored throughout the test. Functional tests shall be
conducted before, during, and after the test for flight model. Sufficient and timely measurements shall
be made on the major internal and external units to verify the major units’ thermal design, hardware,
and analyses. The heat flux, temperature, unit’s operation mode and other performance parameters
shall be controlled to meet the requirements of the specified case.
The modification of the thermal analytical model is applicable to all test cases. The modification
parameters shall be within the acceptable range. After modification of the thermal analytical model, the
modification parameters shall be configured to the thermal analytical model to indicate the
temperature of spacecraft flying on orbit.
After the test, a comprehensive analysis shall be made on energy balance in test cases for test error
sources. The absorbed and irradiated heat by the test model shall be compared, whose difference is
generally controlled within ±10 %. Test errors generally are derived from limitations of the heat flux
simulation mode, deviation between the test model and actual spacecraft, measurement accuracy of
heat flux and temperature.
ISO/FDIS 24412:2022(E)
6.2 Thermal vacuum test
6.2.1 Test description
Spacecraft shall be placed in a thermally controlled vacuum chamber having the capability to expose the
test article at or beyond the minimum and maximum test temperatures.
The following should be considered.
a) Units of spacecraft should be flight products (except qualification test).
b) Some units may be replaced by qualification parts, process parts or simulation parts with their
thermal performance and electrical performance parameters conforming to the test requirements.
c) Large units such as large antenna and solar array may not participate in the test, or may be set
apart from spacecraft in test with cable connection.
d) The propellant tank is generally filled with protective gas.
The temperature can be controlled by the following two measures.
— Controlled by test facility
1) Temperatures of shroud and heating devices shall be controlled by adjusting the flow of gas/liquids or
the electrical power to assure that the test article reaches the required temperature.
2) The temperature on main functional areas (such ase.g. light entrance, OSR radiation surface) should be
consistent with the extreme external temperature on orbit.
3) The surface protrusion, surface coating, cables temperature change shall be taken into consideration
during adjusting to prevent the test article from being damaged.
— Controlled by inner heat source
Internal heat sources include heat consumption of active units or active control heaters.
1) In cooling down the test article, its units may be powered off or work in a minimum power consumption
state.
2) In heating up the test article, its units may be powered on and work in a maximum power consumption
state.
3) Temperature control thresholds of the active temperature controller should be extended at both ends
when the unit temperature is required to be reached faster.
4) Temperature should be monitored for temperature-sensitive units during the thermal cycling process.
The temperature profile for TVT is shown in Figure 3. The durations of thermal soak, thermal
stabilization and thermal dwell for different units of spacecraft depend on their operation modes, heat
inertia characteristics, lifetime, etc. Functional/performance testing should be performed after
adequate thermal stabilization during thermal soak.
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ISO/FDIS 24412:2022(E)
Key
X time Y temperature 1 hot case
12 ambient temperaturehot case 23 ambient cold casetemperature 34 cold casethermal soak
45 thermal soak dwell 56 thermal dwellstabilization 67 functional/performance testingthermal
stabilization
78 hot test8 9 hotwithin test temperaturetolerance 9 Within test tolerance
temperaturefunctional/pe
rformance testing
Figure 3 — Temperature profile for testing
Consideration should be given to conducting the thermal balance test in conjunction with the thermal
vacuum test program. A combined test is often technically and economically advantageous. It shall,
however, satisfy the requirements of both tests. Examples for combined test temperature profiles are
, where TBT hot and
shown in Figure 4, where TBT and TVT are performed in sequence, and in Figure 5
cold cases are integrated into the first TVT cycle. TBT may be performed after TVT in order to test the
thermal-cycling effect on thermal interfaces.

Key
X time Y temperature 1 ambient temperature
ISO/FDIS 24412:2022(E)
12 TB hot caseambient2 3 TB hotcold case 34 TB cold caseTV hot case1
temperature
45 TV hotcold case1 56 TV cold case1hot case2 67 TV hotcold case2
78 TV cold case2hot case3 89 TV hotcold case3 910 TV cold case3hot case4
101 TV hotcold case4 11 TV cold case4
Figure 4 — TBT and TVT are carried out successively

Key
X time Y temperature 1 ambient temperature
12 TB hot caseambient2 3 TBTV hot casecase1 34 TV hot case1TB cold case
temperature
45 TBTV cold casecase1 56 TV cold case1hot case2 67 TV hotcold case2
78 TV cold case2hot case3 89 TV hotcold case3 910 TV cold case3hot case4
101 TV hotcold case4 11 TV cold case4
Figure 5 — TBT is integrated in first TVT cycle
6.2.2 Test conditions
6.2.2.1 Test levels
In TVT, the temperature of all units to be verified can be confirmed on the basis of the temperature
result from TBT, or the temperature result predicted by the thermal model analysis, as shown in
FigureFigures 6 and 7. TVT temperature margin of unit is determined in Table 1 according to the test
level.
a) Based on TBT result
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ISO/FDIS 24412:2022(E)
Key
1 TBT result 2 AT margin 3 maximum TBT temperature
4 minimum TBT temperature 5 AT margin 6 AT temperature range
7 PFT margin 8 PFT margin 9 PFT temperature range
10 QT margin 11 QT margin 12 QT temperature range
Figure 6 — Temperature range and margin of TVT based on TBT result

Key
1 AT margin 2 maximum predicted temperature 3 minimum predicted temperature
4 AT margin 5 PFT margin 6 AT temperature range
7 PFT margin 8 QT margin 9 PFT temperature range
ISO/FDIS 24412:2022(E)
10 QT margin 11 QT temperature range
b) Based on predicted temperature analysis
Figure 67 — Temperature range and margin of TVT based on predicted temperature analysis
Table 1 — Determination of TVT temperature margin
Test Margin Margin
level
based on TBT result based on predicted temperature analysis
QT 5 °C to 10 °C beyond acceptance 5 °C to 10 °C beyond acceptance
temperature temperature
PFT 0 °C to 5 °C beyond acceptance 0 °C to 5 °C beyond acceptance
temperature temperature
AT 0 °C to 5 °C beyond the maximum and 3 °C to 11 °C as thermal uncertainty
minimum temperatures of TBT marginbeyondmargin beyond thermal
model analysis
The thermal uncertainty margin shall be evaluated by importance, complexity, type, operation
temperature of units, and whether the TBT has been carried out or not.
For passive thermal control, the thermal uncertainty margin is a temperature added to worst-case
temperature predictions. For active thermal control, the thermal uncertainty margin is a power margin
to ensure thermal control stability. When the margin is added to worst-case temperature predictions,
the resulting temperature forms the basis for the acceptance temperature range.
Generally, for units that adopt active thermal control, the minimum thermal uncertainty margin can be
reduced to ±(±3 - °C to ±5 ) °C. For units that meet more uncertain operational or environmental
conditions, the thermal uncertainty margin may be larger than ±(±7 - °C to ±11)° °C. Lower thermal
marginmargins may be adopted for units with an allowable operating temperature below -−170 °C or
above 120 °C. Specific estimated uncertainty values are set in a spacecraft thermal system specification.
For active thermal control units at the cold extreme, 25 % excess heater control authority is used in lieu
of 11 °C temperature margins. For unit controlled by heat pipes, the heat transfer margin should be
considered if one of the heat pipes fails.
Test temperatures are restricted as follows.
a) The temperature of the working units should not exceed the temperature range for the
corresponding test levels; and the temperature range of the non-working units should not exceed
the storage temperature range.
b) The upper and lower temperature limits in the same thermal control zone shall be within the
temperature envelope of the test units.
c) The temperature ramp rate should be equal to or higher than the maximum predicted ramp rate. In
case that the test system is not capable to achieve the ramp rates, rates may be lower at no expense
of the test objective and test validity.
6.2.2.2 Test duration
The number of thermal cycles and the duration of thermal soak shall be specified with consideration of
mission requirements.
Generally, the numbers of cycles for spacecraft are:
a) QT: 8 cycles;
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ISO/FDIS 24412:2022(E)
b) PFT: 4 cycles;
c) AT: 4 cycles.
The thermal soak is suggested to be over 8 h at each temperature extreme during the first and last cycle,
and over 4 h for the intermediate cycles.
Test temperature shall be considered as stabilized, in case:
— the test article temperature is within the allowed test tolerance at the specified test temperature;
— the temperature change rate is less than 3 °C per hour.
It is allowed to shorten the measure time or reduce the cycles for some special units such as short-time
or disposable products. And theirTheir temperatures are not required to reach the stability value. They
shall be sufficiently tested before the subsystem TVT or unit TVT.
6.2.3 Basic requirements for test facilities
-2
a) The test pressure shall be no higher than 1,33 × 10 Pa.
b) The shroud temperature and heating devices shall be controlled to meet the requirements for the
maximum and minimum temperature and ramp rate of the test article.
c) The size of the thermal vacuum chamber shall be sufficient to safely perform installation activities
of the test article and MGSE.
-4
d) For spacecraft with high-voltage and high-power units, the test pressure should be at 10 Pa level
due to the internal and external pressure difference.
e) For some special spacecraft such as Mars lander, the test pressure shall be redetermined according
to specified mission environment.
6.2.4 Monitoring during TVT
a) The performance and function of the test article shall be tested in hot and cold temperature
stabilization.
b) Before, during and after the TVT, functional tests shall be conducted and the results meet its
specification.
c) Within thermal soak at cold and hot test temperature of each cycle, the performance of the test
article shall be tested in each operational mode that is foreseen during its service life.
d) If parts of units are sensitive to vacuum environment, vacuum gauges shall be mounted inside the
spacecraft. These units shall not be powered unless pressure is within acceptable range.
e) Low-pressure discharge testing (to simulate the pressure decrease during launch) should be
conducted during the pumping or pressurization process.
f) Temperatures of high-power units during test deserve special attention to avoid damage from
excessive temperature.
g) Thermal sensors for testing shall be installed as indicated by the thermal analysis for monitoring
and acquiring thermal data.
ISO/FDIS 24412:2022(E)
h) Backup units and redundant circuits shall be tested at hot and cold temperature. Their test duration
shall be the same as that of the primary units and shall be equally distributed in each cycle.
i) Collect and analyse the test environment data, including vacuum degree, shroud temperature,
composition and content of released gas, contamination fluctuation.
j) If required, the motion property or thermal deformation should be measured or monitored when
checking mechanism or thermal structure.
7 Test facility
7.1 Laboratory environment
The laboratory environment shall meet the requirements of the specified test article, which include
temperature, humidity, cleanliness, etc.
7.2 Laboratory infrastructure
The laboratory infrastructure supports the test system operation and test article handling by providing
e.g. water, pressurized air, power, lifting and transport means and communication interfaces.
The laboratory infrastructure shall provide all means to safely:
a) operate the test system;
b) handle the test article;
c) conduct tests.
7.3 Test system
7.3.1 Overview
TBT and TVT shall be performed in a thermal vacuum chamber.
An example for a typical composition of a thermal vacuum chamber is shown in Figure 78. A thermal
vacuum chamber usually includes:
a) chamber system;
b) vacuum system;
c) thermal system (cryogenic/heating);
d) data acquisition system;
e) mechanical spacecraft ground support equipment (MGSE);
f) contamination measurement and control equipment.
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ISO/FDIS 24412:2022(E)
Figure 78 — Typical composition of a test system
7.3.2 Chamber system
a) The chamber shall provide sufficient number and performance of feedthrough connectors for
measurement channels, power supply, EGSE communication, to comply with test requirements.
b) Cables and connectors permanently used inside the chamber shall be suitable for use in extreme
high and low temperature, and vacuum environment.
c) Leakage of flanges and connector/pipe feedthroughs shall be minimized.
d) Temperature sensors on shrouds shall be sufficient and distributed uniformly to ensure
temperature measurement accuracy. The measurement acquisition period for chamber parameters
should not exceed 1 min.
2 2
e) Heat flux from sink background shall be less than 10 W/m to 30 W/m based on the relative sizes
of the tested articles and a vacuum chamber.
f) The chamber shall provide a test article grounding interface. The grounding resistance should be
less than 1,0 Ohm.
7.3.3 Vacuum system
a) VacuumThe vacuum system shall be capable of achieving and maintaining the vacuum pressure
required for testing under consideration of leakage, heat devices, MGSE, test article. Pressures as
-3 -4
low as 10 Pa to 10 Pa are readily obtainable. VacuumThe vacuum system should be entirely dry,
and should have “dry” vacuum pumps both in its fore-vacuum part (e.g. dry screw and mechanical
booster Rootsroots-type vacuum pumps) and high-vacuum part (e.g.,. turbomolecular and
cryogenic vacuum pumps).
b) The pressure change rates should simulate the pressure environment change during the spacecraft
ascent and descend.
c) Vacuumthe vacuum system should provide turbomolecular pumps to evacuate some released gas
during testing like hydrogen and helium, provide mass spectrometer to measure the
partial pressures of each composition.
ISO/FDIS 24412:2022(E)
d) Pressurization of the chamber is accomplished with dry nitrogen. This allows the chamber to be
returned to ambient pressure at any time that the contaminant cleaning equipment and all major
equipment in the chamber are above the minimum allowable temperature of the satellite. Moisture
condensation is prevented with this method.
e) Equipment shall be above the dew point if ambient air is pumped into the chamber. The rate should
be controllable.
f) Vacuum gauges shall be selected in accordance with the test pressure range. Pressure acquisition
period shall not exceed 5 s.
7.3.4 Thermal system
7.3.4.1 General requirements
The thermal system shall be able to provide the required thermal conditions both in the volume and on
the surface of test article. Thermal conditions typically are generated by:
a) shroud system;
b) film heaters;
c) IR heaters;
d)
...