ISO 24412:2023

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

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ISO/FDIS 24412:2022(E)
© ISO 2022
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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
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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
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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
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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
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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.
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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.
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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
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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.
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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.
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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.
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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 r
...