General

JERG-2-241

DESIGN STANDARD
EMC

May 10, 2012

Japan Aerospace Exploration Agency

JERG-2-241
This is an English translation of JERG-2-241. Whenever there is anything ambiguous in this document, theoriginal 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.
will not be liable for any losses relating to the use of the information.

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

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JAXA

JERG-2-241
Contents
EMC

1

1. GENERAL PROVISIONS1

1. 1 Objective

1

1. 2 Scope

1

2 RELATED DOCUMENTS2

2. 1 Applicable documents2

2. 2 Reference document

3

3 DEFINITION OF TERMS AND ABBREVIATED TERMS

5

3. 1 Definition of terms

5

3. 2 Abbreviated terms

8

4 BASIC REQUIREMENTS

9

4. 1 Total EMC control

94. 1. 1 Summary

9

4. 1. 2 EMC control organization

9

4. 1. 3 EMC program10

4. 1. 4 Identification of critical portions

12

4. 1. 5 Securing safety margins

12

5. INDIVIDUAL DESIGN REQUIREMENTS

13

5. 1 System requirements

13

5. 1. 1 External electromagnetic environment

13

5. 1. 2 Intrasystem EMC

13

5. 1. 3 EMI control

13

5. 1. 4 Grounding and wiring design13

5. 1. 4. 1 Grounding

13

5. 1. 4. 2 Wiring

14

5. 1. 5 Electrical bonding

15

5. 1. 5. 1 General

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5. 1. 5. 2 Power current feeder and return paths

15

5. 1. 5. 3 Shock and safety hazards

15

5. 1. 5. 4 Antenna counterpoise15

5. 1. 5. 5 RF potentials

15

5. 1. 5. 6 Electrostatic discharge

15

5. 1. 5. 7 Explosive atmosphere protection

15

5. 1. 5. 8 Structure shielding

15

5. 1. 5. 9 Filtering

16

5. 1. 6 Antenna-to-antenna (RF) compatibility

16

5. 1. 7 Lighting discharge

16

5. 1. 8 Spacecraft and static charging

16

5. 1. 8. 1 General

16

5. 1. 8. 2 Differential charge/discharge

17

5. 1. 8. 3 Internal charging

17

5. 1. 8. 4 Charging of fluid lines (liquid fuel lines)

17

5. 1. 8. 5 Electric discharge prevention under low vacuum environment

17

5. 1. 9 Hazards of electromagnetic radiation

17

5. 1. 10 Consideration of life cycle

17

5. 1. 11 External grounds

18

5. 1. 12 Test equipment/facility interface

18

5. 1. 13 Spacecraft magnetic emissions

18

5. 1. 13. 1 Magnetic material

18

5. 1. 13. 2 Current loop

19

5. 1. 13. 3 Magnetic moment

1920

5. 2 Subsystem/component requirement
5. 2. 1 General requirement20

5. 2. 2 Grounding and insulation of on-board equipment

20

5. 2. 3 Primary power line conduction emissions and power source20

5. 2. 4 Primary power line conduction emissions and load

20

5. 2. 5 Primary power line switching transient22

5. 2. 5. 1 General requirement22

5. 2. 5. 2 Inrush current

22

5. 2. 5. 3 Current changing rate22

5. 2. 6 Primary power-line ripple

22

5. 2. 7 Signal line conduction emissions

23

5. 2. 8 Antenna terminal spurious emissions

23

5. 2. 9 Magnetic field radiated emissions

23
23

5. 2. 9. 1 Magnetic materialiii

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65. 2. 9. 2 Current loop

23

5. 2. 9. 3 Magnetic moment24

5. 2. 10 Radiated electric field emissions

24

5. 2. 11 Immunity to power-line ripple

24

5. 2. 12 Immunity to power line switching transient24

5. 2. 13 Immunity to the conducted effects of radiated electromagnetic fields

24

5. 2. 14 Immunity to magnetic field emissions

25

5. 2. 15 Immunity to electric field emission

25

5. 2. 16 Immunity to signal line magnetic field conduction

25

5. 2. 17 Antenna terminal out-of-band immunity25

5. 2. 18 Immunity to electrostatic discharge

25

VERIFICATION

26

6. 1 General requirement26

6. 1. 1 General26

6. 1. 2 Verification scenario

26

6. 1. 2. 1 Subsystem/component level26

6. 1. 2. 2 Space system-level26
27

6. 1. 3 System-level EMC verification plan
6. 1. 3. 1 System-level verification method27

6. 1. 3. 2 Test condition27

6. 1. 4 EMC verification report27

6. 2 System-level verification

27

6. 2. 1 General27

6. 2. 2 Safety margins of critical/EED circuit

28

6. 2. 3 External electromagnetic field environment28

6. 2. 4 Intrasystem EMC

28

6. 2. 5 EMI control

28

6. 2. 6 Grounding and wiring design28

6. 2. 6. 1 Grounding28

6. 2. 6. 2 Wiring design28

6. 2. 7 Electrical bonding29

6. 2. 7. 1 General29

6. 2. 7. 2 Power current feeder and return paths

29

6. 2. 7. 3 Shock and safety hazards

29

6. 2. 7. 4 Antenna counterpoise29

6. 2. 7. 5 RF potential29

6. 2. 7. 6 Electrostatic discharge

29
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29

6. 2. 7. 7 Explosive atmosphere protection
6. 2. 8 Antenna-to-antenna compatibility

29

6. 2. 9 Lighting discharge

30

6. 2. 10 Spacecraft and static charging

30

6. 2. 10. 1 General30

6. 2. 10. 2 Differential charge/discharge

30

6. 2. 10. 3 Internal charging

30

6. 2. 10. 4 Charging of fluid lines (liquid fuel lines)

30

6. 2. 11 Hazards of electromagnetic radiation30

6. 2. 12 Consideration of life cycle31

6. 2. 13 External grounds

31

6. 2. 14 Test equipment/facility interface

31

6. 2. 15 Spacecraft magnetic emissions

31
31

6. 3 Subsystem-level verification

31

6. 3. 1 General
6. 3. 2 Power-induced power line conductivity emission, time and frequency domain characteristics(Primary power line conduction emissions and power source)32

6. 3. 3 Load-induced power line conductivity emission, frequency domain characteristics (Primary powerline conduction emissions and load)

32

6. 3. 4 Load-induced power line switching transient (primary power line switching transient)32

6. 3. 4. 1 Application

32

6. 3. 4. 2 Control of long-duration load-induced switching transients (inrush current)

32

6. 3. 4. 3 Control of fast load-induced switching transients (current changing rate)33

6. 3. 5 Load-induced power-line ripple (primary power-line ripple)

33

6. 3. 6 Signal line conduction emissions

33

6. 3. 7 Antenna connection port spurious emissions

33

6. 3. 8 Magnetic field radiated emissions

33

6. 3. 9 Radiated electric field emissions

33

6. 3. 10 Immunity to power-line ripple

34

6. 3. 11 Immunity to power line switching transient34

6. 3. 12 Immunity to the conducted effects of radiated electromagnetic fields

34

6. 3. 13 Immunity to magnetic emissions

34

6. 3. 14 Immunity to field emission

34

6. 3. 15 Immunity to signal line magnetic field conduction

34

6. 3. 16 Antenna terminal out-of-band immunity34

6. 3. 17 Immunity to electrostatic discharge

34
35

6. 4 Measurement instrument35

6. 4. 1 Immunity test equipmentv

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6. 4. 1. 1 Immunity to conduction

35

6. 4. 1. 2 Immunity to field emission

35

6. 4. 1. 3 Immunity to magnetic emissions

35

6. 4. 1. 4 Immunity to electrostatic discharge

35
35

6. 4. 2 EMC sensor
6. 4. 2. 1 Conduction emissions

35

6. 4. 2. 2 Radiated electric field emissions

36

6. 4. 2. 3 Magnetic field radiated emissions

36
36

6. 4. 3 Shield room/anechoic chamber
6. 4. 3. 1 Shield room36

6. 4. 3. 2 Anechoic chamber

36

6. 4. 4 Background noise reduction

36

APPENDIX

38

A1 system requirements

38

A2 Measures against magnetic field of spacecraft

39

A3 System verification/test related

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1. General provisions
The general requirements and design policy related to the design of electromagnetic
compatibility (hereinafter referred to as EMC) of the spacecrafts developed by the Japan
Aerospace Exploration Agency (hereinafter referred to as JAXA) are specified to the electric
design standard (JERG-2-200). This EMC Design Standard (hereinafter referred to as "thisdesign standard") is a summary of unified descriptions of electric design standard as a design
standard for achieving EMC and specifies the details of the EMC requirements, general test
conditions, test methods and verification requirements in the system-level andsubsystem/component level for achieving the spacecraft EMC.
This design standard is reestablished based on the JAXA technologicalrequirements/guideline document "EMC Design Standard (JERG-0-028)" in light of the latest
technological information and the international trend of the Europe standards (such as ECSS)
and international standards (such as ISO). It is expected that this standard will be utilized as adevelopment guideline for future development.

1. 1 Objective
This design standard aims at defining the common process related to the EMC design,
control, analysis, test and verification of the system, subsystem and component for ensuring theelectromagnetic compatibility (EMC: Electro Magnetic Compatibility) of the spacecrafts
developed by JAXA.
The requirement values and standard values during test, which are not specified in thisdesign standard, shall be specified properly in the EMC control plan in individual projects.

1. 2 Scope
This design standard establishes the basic requirements related to the EMC design andverification of the spacecrafts developed by JAXA as a common standard related to EMC and isapplied to the spacecraft development, launch vehicle, ground support facility and test
equipment interface. However, if the JAXA specifications or other documents are applied in thescope to which this design standard is called and the details contradict between the two, theformer shall deserve higher priority unless otherwise specified.

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2 Related documents
2. 1 Applicable documents
The following documents shall be applied to the range to which this design standard is called.
If a contradiction arises, this design standard shall deserve higher priority unless otherwise
specified.
(1) JMR-001
System safety standard(2) JMR-002A
Rocket payload safety standard(3) JERG-2-200
Electric design standard(4) JERG-2-211
Charge/discharge design standard(5) JERG-2-213
Insulation design standard(6) JERG-2-410
RF communication system design standard(7) ISO14302
Electromagnetic compatibility requirements(8) IEC 61002-4-2Electromagnetic compatibility (EMC) 4-2 Electrostatic discharge immunity test method

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2. 2 Reference document
The reference documents related to this design standard shall be as follows.
(1) ANSI/TIA/EIA-422-B-1994
Electrical Characteristics of Balanced Voltage Digital Interface Circuit(2) MIL-HDBK-83575
GENERAL HANDBOOK FOR SPACE VEHICLE WIRING HARNESS DESIGN ANDTESTING(3) JERG-2-400
Communication design standard(4) JERG-2-143
Radiation-proof design standard(5) DOD-E-8983C
Electronic Equipment, Aerospace, Extended Space Environment, General Specificationfor(6) JAXA-QTS-2060D
Reliability assurance connector common specification for space development(7) SAE-AS50881
WIRING, AEROSPACE VEHICLE(8) SAE ARP 1172:1972 (R1991)
Filters, conventional, electromagnetic interference reduction, general specification(9) MIL-W-16878/4
WIRE, ELECTRICAL, POLYTETRAFLUOROETHYLENE(PTFE) INSULATED, 200 DEG.
C, 600 VOLTS, EXTRUDED INSULATION(10) MIL-HDBK-4001
ELECTRICAL GROUNDING ARCHITECTURE FOR UNMANNED SPACECRAFT(11) MIL-STD-464A
ELECTROMAGNETIC ENVIROMENTAL EFFECTS REQUIREMENTS FOR SYSTEMS(12) NASA-HDBK-4002
AVOIDING PROBLEMS CAUSED BY SPACECRAFT ON-ORBIT INTERNALCHARGING EFFECTS(13) NASA Technical Paper 2361
DESIGN GUIDELINES FOR ASSESSING AND CONTROLLING SPACECRAFTCHARGING EFFECTS(14) NASA H-29919D [12]SCIENTIFICA AND TECHNICAL AEROSPACE REPORTS VOLUME 38(15) MIL-STD-1686C
ELECTROSTATIC DISCHARGE CONTROL PROGRAM FOR PROTECTION OF
ELECTRICAL AND ELECTRONIC PARTS, ASSEMBLIES AND EQUIPMENT(EXCLUDING ELECTRICALLY INITIATED EXPLOSIVE DEVICES)(16) JIS C 0617
Graphical symbols for electrical apparatus(17) ISO/TC20/SC14 international space standard(a) ISO/DIS 24637 Electromagnetic interference (EMI) test reporting requirements

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JERG-2-241(b) ISO/DIS 26871 Explosive systems and devices(c) ISO 14621-1 EEE parts - Parts Management(d) ISO 14621-2 EEE parts - Control program requirement(e) ISO 15389 Flight-to-ground umbilical(18) ISO 7137-3.6 (Section 20.5)
Aircraft - environmental conditions and test procedures for airborne equipment

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3 Definition of terms and abbreviated terms
3. 1 Definition of terms
The definition of terms related to this design standard shall be as follows.
(1) Electromagnetic compatibility (EMC)
The ability of a space equipment or system to function satisfactorily in its
electromagnetic environment without introducing intolerable electromagnetic
disturbances to anything in that environment.
(2) Electromagnetic environment (EME)Electromagnetic environment formed by natural phenomenon and other equipment.
(3) Electromagnetic interference (EMI)
Degradation of the performance of a space equipment, transmission, channel, or systemcaused by an electromagnetic disturbance(4) Immunity
The ability of a device, equipment, or system to perform without degradation in thepresence of an electromagnetic disturbance (insusceptibility to interference).
Susceptibility used in MIL-STD-461 is the degree of sensibility to interference. Immunity
is used in this design standard.
(5) Emission
Electromagnetic energy transmitted from equipment by radiated emission or conduction.
(6) Radio frequency interference (RFI)
Degradation of the reception of a fundamental signal caused by a radio frequency
disturbance.
(7) Safety margin
Ratio of circuit threshold of immunity to induced circuit noise under worst-case expectedenvironmental conditions (intra and intersystem).
(8) Spacecraft
The space vehicles including artificial satellites and probes.
(9) Space system
Normally the spacecraft or launch vehicle itself but used as a generic term forspacecrafts and related facilities and test equipment in this design standard.

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JERG-2-241(10) Bus equipment
Generic term for primary subsystem configuring the space system.
(11) Mission equipment
Generic term for the primary subsystem that governs observation and communication,
which are objectives of spacecrafts.
(12) Subsystem
Functionally-partitioned system elements that configure the spacecraft. From the point of
EMC, no significant differences exist between the bus-subsystem andmission-subsystem in this design standard. Both are referred to as subsystems.
(13) Component
Generic term for units configuring the subsystems in this design standard.
(14) Electric power system
Subsystems that govern such functions as power generation, power control, power
storage, and power distribution (except for harnesses supplying power to loading units).
(The electric power system does not necessarily mean the combination of the solar cell
paddle system and the power source system.)(15) Primary power source
Power source that is generated/controlled in the electric power system of the spacecraft
and distributed commonly to loading units in the spacecraft. (The power that is stepped
down or boosted from the primary power source bus in the electric power system anddistributed commonly to the loading units in the spacecraft shall be the secondary powersource.)(16) Secondary power source
The power that is stepped down or boosted from the primary power source bus in andoutside the electric power system of the spacecraft.
(17) Multibus
The multibus is the system having more than one power source bus that is functional
independently. Other methods are available to connect or disconnect more than onebus.
(18) Hazard
Human damage, physical damage to public or third-party private property, system, range
facility, and environmental implications.
(19) Malfunction/damage
Malfunction means the degradation of the performance and damage means the loss of6

JERG-2-241
the function.
(20) Breakout box
A piece of non-flight test equipment that is connected in series with the cables and allows
the measuring external connections (normally, connection terminals) or serial-parallel
test networks to be connected to the cables.
(21) Internal charging
Static charging phenomenon generated in metallics or dielectrics having no electricalinterconnection with the surrounding. A phenomenon when secondary radioactive rays,
which are caused by penetration of cosmic radiation or cosmic radiation throughspacecraft structures and/or component walls, are remained in the object or electrons in
the object are ejected outside.
(22) Power source quality requirement
Requirements for power subsystems in the spacecraft defining the supply voltage, noise(caused by load stabilization, spike or hanging), and impedance required by powersource users.
(23) Intermodulation characteristic
A characteristic that the amplitude of fundamental wave is influenced by the change in
the amplitude of the distorted wave component (third-order harmonics component inparticular) of the interfering wave generated by the nonlinearity of the amplifier in thereceiver when both desired wave and interfering wave are input at the same time to thereceiver input.
(24) Spurious modulation characteristic
A characteristic that a new unremovable frequency component is formed in the reception
bandwidth by the distorted wave component of the interfering wave generated by thenonlinearity of the amplifier in the receiver when both desired wave and interfering wave
are input at the same time to the receiver input.

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3. 2 Abbreviated terms
Abbreviated terms related to this design standard shall be as follows.
AC
BCI
CE
CS
DC
DSO
EED
EMC
EME
EMI
ESD
EUT
FFT
GSE
I/F FMEA
LISN
RE
RF
RFI
RS
SDR
SRR

Alternative Current
Bulk Current Injection
Conducted Emissions
Conducted Susceptibility
Direct Current
Digital Storage OscilloscopeElectro Explosive Device
Electro-Magnetic Compatibility
Electro-Magnetic Environment
Electro-Magnetic InterferenceElectrostatic Discharge
Equipment Under Test
Fast Fourier Transform
Ground Support Equipment
Interface Failure Mode Effects Analysis
Line Impedance Simulation Network
Radiated Emissions
Radio Frequency
Radio Frequency interference
Radiated Susceptibility
System Design Review
System Requirement Review

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4 Basic requirements
The space system shall be electromagnetically compatible with the spacecraft itself andpredefined external facilities/other systems, and electromagnetic environment in nature during
all phases and orbital operations of its mission including design, manufacturing, test andlaunching. The EMC control of the space system shall be achieved by designing of all
subsystems/components within the space system and of the interface design with external
electromagnetic environment.
The objectives of EMC control are to maintain the cost and schedule for manufacturing andtesting of the spacecraft and eliminate electromagnetic interference during operation, ensuring
the spacecraft-level function and performance.
In this chapter, general requirements for EMC control related to the entire space system are
specified. Individual design requirements are specified in chapter 5: space system-level EMC
requirements and subsystem/component-level EMC requirements are specified in Section 5. 1
and 5. 2, respectively. Requirements on verification are specified in Chapter 6.

4. 1 Total EMC control
4. 1. 1 Summary
In the spacecraft project, the EMC program shall be specified in its early stage.
The program shall be a comprehensive plan based on suitable interface control
specifications in consideration of requirements from the bus equipment to the mission
equipment, requirements from the mission equipment to the bus equipment, all other related
external facilities/other systems, and electromagnetic environment in nature.
The EMC between these mission equipment/bus equipment (intersubsystem) andsystems/subsystems shall be proved to be free from functional problems with a suitable safetymargins.
The EMC control shall be implemented based on the EMC plan by the EMC control
organization. The EMC control organization shall be specified in the following section.

4. 1. 2 EMC control organization
Comprising an organization mainly to investigate the system-level EMC plan, totalrequirements and verification plans, and solve the issues over the entire period of the project
ranging from the initial phase of the space system development to the integration phase,
implementing the EMC control program smoothly. The organization shall consist of spacecraft
project and spacecraft developers (including the manufacturer), mission equipment developers
and independent intellectuals under the supervision of the project manager and be committed
to solving the issues in implementing the entire EMC control program. The EMC control
organization shall select and organize well-equipped personnel for performing EMC control asappropriate according to the spacecraft development phase and the size and complexity of thesystem.

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4. 1. 3 EMC program
The EMC program in the spacecraft project shall be developed in the following steps andcontrolled by the EMC control organization.
a) EMC program control requirements

The EMC program shall be documented as an EMC control plan. The EMC control plan shall
include the following control requirements.
1) Each constituent component in the space system shall be controlled by relating the
responsibility of EMC requirements to the manufacturing and verification plans andrequired design change process.
2) The EMC control plan shall include the following items;(a) The structure of facilities and personnel required for succeeding the EMC program(b) The method and procedure for implementing the design review and adjustment related
to EMC(c) Specify EMC requirements and standards (including the setting of specific domain
protection related to radio communication equipment)(d) Weber, deviation process
3) Program and schedule:
The EMC program, schedule and milestone in the program development masterschedule shall be specified.
b) System-level function and design requirements
To achieve EMC as a space system, the following requirements shall be considered.

1) Identification of related environment
As spacecraft-level EMC requirements, consideration shall be given to all electronic
devices in the system, test facilities used in testing and electronic devices in the launchsite, launch vehicle and risk of electromagnetic radiation to explosive devices including
launch site environment.
2) Allocation of EMC performance
EMC performance shall be allocated to each component from the standpoint of system
feasibility in the light of space system-level EMC requirements.
c) Subsystem/component requirements

In deciding subsystem/component EMC requirements, consideration shall be given to thefollowing points.
1) Space system-level requirements shall be allocated properly tosystems/subsystems/components and ground support facilities defined in this design
standard. In light of these performance requirements, the EMC test level for EMC test
facilities shall be set.
2) In conducting EMC verification, the conditions including the procedure and method for

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adjusting the verification test range and test method with the EMC test facility side shallbe specified.
3) Subsystem/component EMC test result summary(a)Subsystem/component-level EMC test results shall be summarized.
(b)All problems judged to be acceptable shall be recorded in detail. The analysis results inlight of conditions of general EMC performance shall be presented as a criterion for
decision. The test results shall be investigated in consideration of cost, schedule,
reliability, system operability and other factors.
d) EMC analysis requirements
Subsystem/component EMC analysis shall be conducted as follows:

1) For possible interference in the space system, EMC analysis shall be conducted at thesystem, subsystem and component levels to check for interrupted operation.
2) In performing design to cope with possible or actual interferences, environment at thetest facility level shall be considered as conditions for input impedance coupling(conductivity radiated emissions) and electromagnetic field radiation of cables and wiring
harness to be wired.
3) In all modes in which the on-orbit space system is expected to be operated, EMC
analysis based on the actual presented values shall be conducted on the interface ofEMI characteristics of the facilities, subsystems and components. The EMI safety
coefficient of the space system EMC shall be determined.
4) If judged to be necessary to achieve the EMI safety coefficient according to the analysis
results, it shall be appropriate to request for addition of filtering and shielding tosubsystems/components and design change with respect to EMI.
e) Verification plan requirements

EMC verification shall be conducted in a comprehensive plan on items necessary forensuring the interfaces, functions and performances in phases including designing,
manufacturing and testing of systems, subsystems and components without any omission.
Notes for EMC analysis shall be included in the verification plan.
This shall include identification and selection evidences of critical portions which are subjectof safety margins validation, and measurement technology of sensitivity to critical portions andEED equipment/circuits.
f) Weber related

Even if the results of confirmation and verification after manufacturingsubsystems/components (including supplied parts) cannot satisfy the requirements for the EMC
standard established in the space system project, the equipment concerned can be applied tothe system as long as the system side judges that it is allowable. However, judgment validity on
the system side shall be discussed by the EMC control organization.

g) Charge/discharge analysis

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JERG-2-241
1)Electrostatic discharge
It shall be verified in analysis or testing that nonenergized electronic devices are notdamaged by electrostatic discharge to the pin or external connector.
2) Charge/discharge
The subsystems/components shall be designed not to be damaged by plasma in any
operational environment of the space system. It shall be indicated in analysis that corona
will not damage the subsystems/components. See Charge/discharge design standard(JERG-2-211) for details.
h) Requirements related to explosive articles
For equipment related to explosive articles, particular attention shall be given to the following
requirements. See Section 5. 5 of Rocket payload safety standard (JMR-002B) for details.

1) Extraction and evaluation of requirements and setting of margin for proper explosive
article-related EMC
2) Design-related technological requirements
3) Verification-related technological requirements

4. 1. 4 Identification of critical portions
The EMC control organization shall identify functional criticality for all subsystems/components.
Criticality categories include the following:A) Category I: Safety Critical - EMI problems could result in loss of life and/or loss of space
platformB) Category II, Mission Critical - EMI problems that could results in injury, damage to space
platform, mission abort or delay, or performance degradation which unacceptably
reduces mission effectivenessC) Category III, Noncritical - EMI problems that could result only in annoyance, minor
discomfort, or loss of performance which does not reduce desired spacecraft
effectivenessTo ensure the success of EMC control, items that should be identified to be of special
importance and verified as a space system shall be extracted. Identification of some critical
portions must be done from the stages before starting designing. Consultation shall be held infull by the EMC control organization from the initial phase of the project.

4. 1. 5 Securing safety margins
Design safety margins (system margins) for important circuits and EED circuits shall be
reviewed and specified by the EMC control organization. Safety margins shall be specified
properly in consideration of the degradation of functions and performance, degradation modedue to secular change and protection method of the systems, subsystems and components inthe spacecraft over the entire planning period. The safety margins related to electromagnetic
interference in critical portions shall be defined based on the comparison of noise level andsensitivity level in these portions.

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JERG-2-241
5. Individual design requirements
5. 1 System requirements
EMC requirements shall be divided into the following two categories according to the
purpose:(1) Equipment malfunction prevention
Requirements to prevent equipment malfunction due to electromagnetic interference
and meet equipment functions. Requirements common to a number of spacecrafts.
(2) Mission requirement achievement
Requirements that are not at the level causing malfunction but required to achieve
mission requirements and meet equipment performance. Requirements unique tospacecrafts.
System requirements and design standard shall be specified by making a close investigation
as to which is stricter in the phases (such as SRR and SDR) before designing.

5. 1. 1 External electromagnetic environment
In system design, the electromagnetic environment due to external interference sources
shall be discussed. The degree of degradation of the system validity caused by external
environment shall be discussed in consideration of mission profile and available
electromagnetic environment data.

5. 1. 2 Intrasystem EMC
The space system shall not interfere with subsystems/components for mission key
requirements. Each subsystem/component shall operate without performance degradation
during concurrent operation of any combination of the remaining subsystems/components,
subject to mission requirements.

5. 1. 3 EMI control
In EMI control, it is important that the system controls the entire spacecraft in a crosssectional manner. The presence of a piece of equipment with a significant noise emission willaffect the entire spacecraft. The project EMC control plan shall be observed.

5. 1. 4 Grounding and wiring design
5. 1. 4. 1 Grounding
As intrasystem grounding points, metal structure portions and grounding bus bars that canbe used as the zero potential reference points provided properly in the spacecraft structure shallbe specified. Details shall be clarified in the EMC control plan. The following items shall be

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JERG-2-241
included:(1) Relationship with the return/standard electrode potential of both power system andsignal system shall be indicated.
(2) The impedance of connections related to the preceding paragraph shall be indicated.
Impedance magnitudes of these connections over the affected signal spectrum shall beconsidered in determining which kind of power and signals share common paths (wire or
structure/enclosure).
Specific resistance and inductance values for each element of electric networks provided as
ground points can be assigned. (See Section A1. 1.) The use of these values allows you to
calculate the common mode voltage at each circuit reference point and compare with
conductivity sensitivity requirements of each equipment.
5. 1. 4. 1. 1 Grounding/insulation system diagram
Because grounding and insulation are related closely to the space system EMC, thegrounding and insulation system diagram for space systems, subsystems and components
shall be created to check for compatibility in the entire system.

5. 1. 4. 1. 2 Grounding and insulation of power source(1) Primary power source
The grounding of power source shall be determined based on the power design policy. The
grounding point to the structure shall be single so that the structure is not used as thecurrent return path.
(2) Secondary power source
The return path of the secondary power source shall be controlled as the space system andgrounded with any of the on-board equipment, subsystem or system.
Grounding points shall be determined in perspective of EMC control including
function/performance requirements for on-board equipment, requirements for otherpieces of on-board equipment to be interfaced with or inter-equipment interfaces as thesubsystem and interfaces in the system. A proper insulation resistance shall be secured
between the primary power source and secondary power source as a rule.
(3) Multibus power source
When setting a multibus power source, grounding policy shall be clarified for eachindependent power source system. When each power system is provided with asingle-point grounding as a policy, a single-point grounding shall be provided for eachsystem. A proper insulation resistance shall be ensured between power systems.

5. 1. 4. 2 Wiring
The design guideline for system wiring and cable separation/signal classification shall be
specified.

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JERG-2-241

5. 1. 5 Electrical bonding
5. 1. 5. 1 General
This paragraph is intended to ensure the equipment performance. The main point is toensure a proper current path and voltage distribution in the electric circuit to be installed.

5. 1. 5. 2 Power current feeder and return paths
Twist pair wires shall be used for the power current feeder and return paths. Attention shallbe given so that direct current magnetic field and electromagnetic wave will not be radiated. The
structure shall not be used as the current return path.

5. 1. 5. 3 Shock and safety hazards
Not applicable.

5. 1. 5. 4 Antenna counterpoise
The structure with the counterpoise attached shall be bonded so that RF currents have a low
impedance path.

5. 1. 5. 5 RF potentials
It is desirable that portions, except for antenna, that are possible subject to electromagnetic
radiation to outside or electromagnetic interference from outside be bonded to the grounding
point with a faying surface bond to present a low impedance. The bonding straps shall be as
short as possible and maintain a low inductance path.
5. 1. 5. 6 Electrostatic discharge
JERG-2-211 shall be applied.

5. 1. 5. 7 Explosive atmosphere protection
Not applicable.

5. 1. 5. 8 Structure shielding
The structure discontinuous portions (such as cover, inspection holes and joints) shall belimited to the minimum. For discontinuous portions at important points with respect to RF
interference noise and interference sensitivity, a low impedance conductive path shall be
provided for conduction of electricity.
Note that the wire harness having an imperfect EMI shield may act as a discontinuous

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JERG-2-241
portion of the structure.

5. 1. 5. 9 Filtering(1) For filtering spacecraft noise, the signal interface cable establishing connection between
on-board components shall be provided with an EMI shield as needed in consideration of
the signal level, interface circuit filter characteristics and radiated electromagnetic field.
(2) The attenuation level of the radiated electromagnetic field with respect to the on-board
equipment shall be analyzed by a standard measurement method or at typical places in
the internal shielding enclosure. The shielding may be done by providing a spacecraft
basic structure designed as a Faraday cage (box) having a minimum number of
openings or through-hole portions, or an enclosure for electronic device, or using the
combination of these shields.
(3) The electronic devices and cables outside the structure shall be shielded so that proper
margins are ensured in consideration of the uncertainty of spacecraft electromagnetic
environment.
To allow the filter to produce an effect, it shall be noted that the impedance at before andbehind the insertion point and the relation between electromagnetic environment
characteristics and impedance are proper. Prepare a mechanism for letting the noise andradiated emissions that should be cut off escape to another path or for heating. (See Section A1.
2.)
5. 1. 6 Antenna-to-antenna (RF) compatibility
Designing and inspection shall be implemented so that no unexpected interference occurs
between any RF devices. This shall be realized not only in the single system but also between
the systems when interface requirements are made. Intermodulation characteristics shall be
included when doing with not inspection but analysis.

5. 1. 7 Lightning discharge
The space system shall be protected from direct and indirect effects of lightning discharge toaccomplish the mission without any performance degradation after exposed to lightning
discharge. A combination of elimination of lighting discharge through operation and electric
overstress design technology shall be used as protective measures.

5. 1. 8 Spacecraft and static charging
5. 1. 8. 1 General
In order to avoid any risk of the human electric shock, fuel ignition, radio wave interference(RFI), dielectric material breakout, and other caused by ground operation or environment prior
to launch and in-orbit high energy plasma environment, the space system shall be controlled its
charging and eliminated.

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5. 1. 8. 2 Differential charge/discharge
Differential static charging phenomenon in components due to plasma, electric discharge
phenomenon and degraded performance of the space system shall be minimized throughpreventive measures for design and integration. However, because complete elimination ofelectric discharge cannot be assured, it shall be required that tolerance of the satellite system
be reinforced and that the spacecraft be verified to be operative without malfunction,
performance degradation and out-of-spec parameter under repeated electrostatic arc
discharge, which is a typical disturbance phenomenon of possible exposure.

5. 1. 8. 3 Internal charging
When the electronic flux event is an orbit parameter enough to cause internal charging, thetolerance reinforcement technology for preventing electrostatic discharge (ESD) from reaching
the limit value shall be adopted to minimize static charging to these surfaces.

5. 1. 8. 4 Charging of fluid lines (liquid fuel lines)
To prevent the pipes, tubes and hoses for liquid from developing arc discharge, static
charging preventive measures shall be implemented on the liquid and its transportation system.

5. 1. 8. 5 Electric discharge prevention under low vacuum environment
For the on-board equipment that may develop glow discharge during vacuum transition(critical pressure), consideration shall be given to electric discharge prevention. The examples
shall be as follows:(a) On-board equipment with the primary power source turned on during launching(b) Any component that may be charged in the capacitor bank, etc(c) Transmission filter (or diplexer) during high-power transmitter output
Consideration shall be given to the interrelation between the emission rate of outgas volume
discharged from the air vent and the vacuum transition rate due to rocket launching elevation.

5. 1. 9 Hazards of electromagnetic radiation
The space system shall be designed so that fuels, personnel, explosive systems (includingEED), and electronically actuated engine/thruster are not exposed to electromagnetic radiation
at a dangerous level. Safety shall be ensured in consideration of possible electromagnetic
interference from the external transmitter.

5. 1. 10 Consideration of life cycle
The electromagnetic environment (EME) protection design shall include the concept ofprotection life cycle (e.g., identification, method, reliability, maintainability, serviceability,

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JERG-2-241
verification and inspection requirements of protection elements).
For maintainability in particular, the EME protection mechanism shall be designed so that it
can be easily accessed, maintained or complete the operating life of the space system without
essential maintenance or inspection. Bonding, shielding and other protection technologies to bedisconnected, pulled out or cancelled during maintenance shall be specified in the maintenancedocument with operation required for reassembly. These protection mechanisms that need tobe repaired during life cycle of the space system shall have the function to test or inspectspecified performance.

5. 1. 11 External grounds
Be sure that the grounding cable can be connected to the space system to equalize theelectrode potential (or diffuse static charging) before applying electric power by otherprocedures or through the interface.
5. 1. 12 Test equipment/facility interface
For the test equipment/facility to be electrically interfaced with the spacecraft, consideration
shall be given to the following points:
For the power lines, there shall be protection function on the test equipment/facility side.
The regulations on disturbance during testing shall be complied with.
Grasp the measures for spacecraft installation facility during lightning in advance and ensure aproper connection.
Grasp the mutual grounding system of spacecraft, test equipment and facility and ensure aproper connection.
Implement I/F-FMEA on the circuit to be interfaced with the spacecraft as needed.
Grasp the electric limitations on the test equipment/facility in advance and clarify theinterface with the spacecraft.

5. 1. 13 Spacecraft magnetic emissions
Control on magnetic emissions shall be required when a small magnetic field of the earth,planet surrounding or inter-planet space is observed, or equipment sensitive to magnetic field isinstalled, or satellite position disturbance must be strictly avoided or controlled under arelatively strong magnetic field environment (low elevation around the earth). Depending on the
purpose of control, proper standards of DC and AC magnetic fields shall be defined for eachcomponent in the spacecraft. It will be necessary to perform control not only on the spacecraft
design and magnetism inherent to parts concerned but also on prevention of subsequent
magnetization. Thus, regulations shall be specified as needed for the magnetic fieldenvironment for spacecraft testing and the tools used for manufacturing and testing. Appendix
A2 shows specific reference information on measures against spacecraft magnetic field.

5. 1. 13. 1 Magnetic material
The spacecraft with magnetic field requirements shall be designed to use magnetic

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JERG-2-241
materials to a minimum. Employed materials shall be specified in the design standard. The
EMC control organization shall grasp and control the use of magnetic materials as well as its
countermeasures if there is no other choice but to use them. (See Section A2. 1.)
5. 1. 13. 2 Current loop
If current loop exists in the spacecraft, magnetic field is developed according to the current
strength and loop area. When designing a spacecraft with magnetic field requirements, properconsideration shall be given. Adequate care shall be given to the harness wiring between
devices and the loop through the ground. (See Section A2. 2. 3.)

5. 1. 13. 3 Magnetic moment
When it is necessary to control the satellite position disturbance due to the environment
magnetic field after launching, the magnetic moment of the spacecraft may be controlled.
Criteria shall be specified in the project EMC control plan to control the magnetic moment.
Measures shall be taken to offset the magnetic moment as needed. (See Section A2. 2. 2.)

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5. 2 Subsystem/component requirement
5. 2. 1 General requirement
System-level EMC requirements shall be distributed to EMI requirements at thesubsystem/component level. Subsystem/component-level EMI requirements shall be specified
basically for each space system development based on the following concepts.
The matrix of the applicability, reference test method and standard forsubsystem/component-level requirements shall be shown in Tables 5. 2-1 and 5. 2-2.
For the immunity requirements for evaluating the effects of radio wave transmission, theradio modulation system used in actual spacecrafts shall be simulated.

5. 2. 2 Grounding and insulation of on-board equipment
Grounding and insulation of on-board equipment shall be designed as follows: (See Section A1. 1.)(1) When the on-board equipment is not connected to the spacecraft power source, proper
insulation resistance shall be provided between the primary power source input line ofeach on-board equipment and the chassis of the on-board equipment.
(2) Proper insulation shall be provided between the primary power line and the low
frequency signal line.
(3) Proper insulation resistance shall be provided between primary power source return andthe signal or secondary power source return of the on-board equipment that generates
high frequency or operates at high frequency.

5. 2. 3 Primary power line conduction emissions for power source
These requirements shall be applied only to the power subsystem. These requirements are
predicated on the power bus characteristics under resistive load.
The power source bus voltage ripple (time domain, frequency domain) shall satisfy thepower source quality requirements under all load conditions. To limit radioactive emission fromthe power source bus, consideration shall be given to the control of conductivity emission.

5. 2. 4 Primary power line conduction emissions for loadThese requirements shall be applied from the power source bus to subsystem/component
operation.
When specifying requirements for conductivity emission, the total voltage noise ofload-induced power source bus shall meet the power source quality requirements. The normal
and common mode ripple noise requirements shall be applied to each subsystem/component
as long as the effect of noise generated by the subsystem will not exceed the power source
quality standard. When specifying the upper limit of conductivity noise, consideration shall begiven to radio emission noise having effect on the receiver.

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Table 5. 2-1 Subsystem/component level test application matrix (emission)
Requirement/verificati
Test categoryonApplicability
Item number
Power-source induced power line
Applied to power
conductivity emission, time andsubsystem
5.2.3/6.3.2
frequency domain characteristics
Load-induced power line conductivity
Application
emission, frequency domain
5.2.4/6.3.3
characteristics
Control of long-duration load-induced
Specified individually
5.2.5.2/6.3.4.2
switching transientsControl of fast load-induced switching
Application
5.2.5.3/6.3.4.3transients
Load-generating power-line ripple
Specified individually
5.2.6/6.3.5
Signal line/conductivity emission
Specified individually
5.2.7/6.3.6
Antenna connection port spurious
Applied to antennaemissionsconnection terminal
5.2.8/6.3.7

Basis for test
methodAppendix

MIL-STD-461
CE101、CE102
Appendix
Appendix

Appendix
Appendix
Guidelineisdescribed
inAppendix.
Magnetic field/radioactive emission
Specified individually MIL-STD-461
5.2.9/6.3.8
RE101
Electric field/radioactive emission
Application
MIL-STD-461
5.2.10/6.3.9
RE102
Note: For items indicated as "specified individually", the necessity of application shall be specified
according to the EMC control plan.
Table 5.2-2 Subsystem/component level test application matrix (immunity)
Requirement/verification
Basis for test
Applicability
Test category
Item number
methodAudio frequency power
Application
MIL-STD-461
5.2.11/6.3.10
line/ripple immunity
CS101
Power line/switching
Application
Appendix
5.2.12/6.3.11
transient/immunity
Radioactive electromagnetic
Specified
MIL-STD-461
5.2.13/6.3.12
field conductivity effect immunity
individually
CS114
Audio frequency radioactive
Specified
MIL-STD-461
5.2.14/6.3.13
electromagnetic field immunity
individually
RS101
Radioactive electromagnetic
Application
MIL-STD-461
5.2.15/6.3.14
field immunity
RE103
Magnetic field induced signal
SpecifiedAppendix
5.2.16/6.3.15
immunity to cable
individually
Antenna connection port
ReceiverDepending on
immunity to out-of-bandthereceiver5.2.17/6.3.16
interference
type
Immunity to electrostatic
Applicationor MIL-STD-1541
discharge
5.2.18/6.3.17
handling
procedureNote: For items indicated as "specified individually", the necessity of application shall be specified
according to the EMC control plan.

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5. 2. 5 Primary power line switching transient
5. 2. 5. 1 General requirement
The effect of two types of switching transients shall be controlled.
Among the two types of switching transients are long-duration switching transient (durationof order of millisecond or more) and rapid switching transient (duration of order of millisecond or
less).
(1) Output impedance
For the power subsystem, specify the power source side output impedance in theloading unit input terminal and define the LISN (Line Impedance Simulation Network)
model.
(2) Maximum input current
The maximum value of on-board equipment side consumption current including theoperational transient during steady operation (excluding during inrush) shall be
specified.

5. 2. 5. 2 Inrush current
The load-induced long-duration switching transient (inrush current) shall be restricted so thatthe voltage level will be maintained within the power source quality allowable range of thepower subsystem.

5. 2. 5. 3 Current changing rate
The envelope range of the rapid switching transient shall be controlled so that an accurate
envelope range of the steady transient current can be specified by the power source quality
specification. The size and duration of both the operational transient and turn-on transient shallbe controlled. These transients can be evaluated separately unless they are repeated transient
disturbance that can be observed in the frequency domain.

5. 2. 6 Primary power-line ripple
The primary power-line ripple induced by load ripple shall be controlled. In consideration of
the power source supply impedance, specify the ripple requirements for components, conduct
time domain ripple measurement using test piece and check that the envelope line satisfies therequirements.

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5. 2. 7 Signal line conduction emissions
When emission to the receiver and high-sensitivity electronic devices must be prevented at
the mission-specific working frequency, the common mode current of the cable bundle shall becontrolled.

5. 2. 8 Antenna terminal spurious emissions
To ensure radio wave frequency domain (RF) compatibility between antennas, control shallbe performed on the spurious emissions at the antenna connection terminal.
When defining the allowable limit and frequency range, consideration shall be given to thefollowing items.
a) Sensitivity including the out-of-band response of the interfered receiver subsystem(receiver, transmission line, antenna)b) Exemption of transmission frequency (of transmitter and transceiver) and modulation
bandwidth for information transmissionc) Highest and lowest frequency used intentionally by the spacecraft intrasystem receiverd) Gain/loss characteristics of antenna connection terminal accessory

5. 2. 9 Magnetic field radiated emissions
When the magnetic field of the spacecraft must be controlled, magnetic emission reduction
measures shall be implemented from the design phase at the subsystem and component level.
When deciding service parts and design, magnetic effect on the spacecraft performance shallbe verified. It is useful, in view of test verification, to simply set a uniform standard (magnetic
field strength generated at a certain distance from each subsystem or component) for allowablemagnetic emission strength used in verification to each subsystem or component. However, theeffects on the spacecraft performance shall be verified eventually in consideration ofcharacteristics (such as spacecraft on-board position and orbital operation method) of eachsubsystem/component. When magnetic field control is implemented, it shall be required toprevent the radiation of magnetic field and perform control for preventing magnetization. Thus,
it may be necessary to set specifications on the magnetic field environment of the test site forsubsystems/components and the tools for manufacturing and testing.
Specific reference information on subsystems/components magnetic field control shall beshown in Appendix A2.

5. 2. 9. 1 Magnetic material
Same as Section 5. 1. 12. 1. (See Section A2. 1.)

5. 2. 9. 2 Current loop
The presence of current loop in the subsystem/component leads to the development of amagnetic field in accordance with the current and loop area. When designing
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subsystems/components installed to the spacecraft with magnetic field requirements, dueconsideration shall be given. Special care must be taken with the equipment requiring large
current or large-area equipment.
(See Section A2. 2. 3.)

5. 2. 9. 3 Magnetic moment
Magnetic components have a large magnetic moment in many cases. In consideration of thelevel to be reduced, presence or absence of time variation in magnetic moment, effect of time
variation in magnetic moment on the spacecraft performance and available weight budget, an
appropriate method shall be selected from magnetism offset between components, magnetic
shield or cancel magnet.
(See Section A2. 2. 2 for specific method for reducing magnetic moment.)

5. 2. 10 Radiated electric field emissions
The radioactive electric field emission of all subsystems/components shall be controlled.
Special care shall be taken with the receiver frequency bandwidth of the spacecraft andlaunch vehicle. For subsystems/components commonly used, radioactive electric fieldemission requirements for the target system side receiver sensitivity bandwidth shall besatisfied. The regulations for allowable range shall reflect the operational sensitivity of theinterfered receiver, the gain, directivity and installation position of the interfered antenna.

5. 2. 11 Immunity to power-line ripple
Subsystems/components shall have immunity to the allowable audio bandwidth ripple inaccordance with the applicable power source quality specifications and conductivity emission
requirements. A proper margin shall be provided between the allowable ripple level on the
power source bus and the sensitivity level of subsystems/components.

5. 2. 12 Immunity to power line switching transient
All subsystems/components shall have immunity to the load-induced switching transient
defined in the power source quality specification.

5. 2. 13 Immunity to the conducted effects of radiated electromagnetic fields
Subsystems/components operated in the spacecraft system exposed to intentional radio
emission shall have immunity to the common mode current induced byequipment-to-equipment connection cables or power lines. These requirements shall only beapplied to the frequency of possible intentional radio wave transmission. These requirements
shall be verified by implementing the bulk current injection (BCI) test up to 400 MHz.

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5. 2. 14 Immunity to magnetic field emissions
The immunity to magnetic field shall be controlled without any hindrance in the presence ofAC magnetic field environment that may cause hindrance to the subsystems/components in thespace system.

5. 2. 15 Immunity to electric field emission
Whether transmission is intentional or not, the immunity of subsystem/component to radio
wave transmission shall be controlled so that space system-level electromagnetic compatibility
is ensured.
Special consideration shall be given to providing immunity to the electromagnetic fieldradiated from other than the spacecraft transmitter operation bandwidth and the space system
when the electromagnetic field must be controlled.

5. 2. 16 Immunity to signal line magnetic field conduction
When high-density wiring cables are used in the space system, immunity to cable-to-cablecoupling shall be verified in inductive noise sensitivity test. In that case, audio bandwidth
magnetic field coupling and transient disturbance coupling shall be applied.

5. 2. 17 Antenna terminal out-of-band immunity
The out-of-band radiation interference characteristics of the radio receiver shall becontrolled. Spurious signal characteristics, cross-modulation characteristics andintermodulation characteristics shall be controlled.

5. 2. 18 Immunity to electrostatic discharge
In consideration of electrostatic discharge, the ESD control method forsubsystem/component immunity and subsystem/component handling shall be introduced. The
actual ESD level at the time of approval shall reflect the environment for the equipment from thefinal assembly to mission completion. Consideration of spacecraft static charging is shown in
Section 5. 1. 8.

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6

Verification

6. 1 General requirement
6. 1. 1 General
The prime contractor or developer shall be responsible to verify the conformance to all
requirements specified in this standard. Specific tasks may be delegated to the sub-contractor
through the EMC control organization as needed. Verification shall be performed properly in thequalification testing, analysis and inspection and equivalent test.

6. 1. 2 Verification scenario
It shall be required to make and implement a verification plan for verifying that eachsubsystem and each on-board component individually meet the electromagnetic compatibility
standard. It shall also be necessary to make and implement a verification plan for the entire
space system.

6. 1. 2. 1 Subsystem/component level(1) All verifications related to electromagnetic compatibility standard shall be conducted oneach subsystem/component under the responsibility of the person in charge ofmanufacturing and development before installation to the space system.
(2) Verification shall be conducted in the development phase of the engineering model
before the flight model so that design can be modified. The verification of the flight model
shall be performed as the last confirmation as far as possible.
(3) If the person in charge of manufacturing and development cannot prepare the special
facility and sensor for radiation noise measurement in conducting verification,verification shall be performed using the facility owned by JAXA or others.
(4) For deviation from the accepted electromagnetic compatibility standard level inverification, the person in charge of manufacturing and development shall perform
analysis of verification results and modify each subsystem/component based on the
results so that the compatibility level is satisfied before the installation of the spacesystem. However, as for the flight model, the necessity of modification shall be discussed
in consultation with the EMC control organization before modification.
(5) Each subsystem/component developer (including manufacturer) shall report the resultsof verification on the electric compatibility standard to the EMC control organization
before installation to the spacecraft.

6. 1. 2. 2 Space system-level(1) Verification of the conformity to the electromagnetic compatibility standard shall be
conducted on the entire space system with all the subsystems/components assembled
under the responsibility of the JAXA project.
(2) In verifying the electromagnetic compatibility in the entire space system, the operation

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quality shall be verified using on-board components themselves in addition tomeasurement using the EMC measurement sensor to check for interference between
components installed in the spacecraft.
(3) The necessity of component modification shall be discussed based on the verification
results by the EMC control organization.

6. 1. 3 System-level EMC verification plan
The prime contractor or developer (including manufacturer) shall develop the EMC
verification planning paper specifying the space system-level EMC verification plan. The detailsof the verification method for each EMC requirement and the acceptance standard for eachsubsystem/component shall be specified. Before starting the EMC verification test, confirmation
for EMC verification plan shall be performed in the spacecraft project.
The EMC verification plan shall at least include the following.

6. 1. 3. 1 System-level verification method
Depending on the fidelity of the engineering model, verification may be conducted
separately in the test for engineering model and the test for flight model/protoflight model.
The method, procedure and measurement equipment for recording the test results shall be
included in system verification target. (See Section A3. 1. 3. 1.)

6. 1. 3. 2 Test condition
The necessary personnel shall be defined. The procurement activity, contractor,
sub-contractor, and quality control representative shall be included. The required test
equipment shall be defined. It shall include the description of specific EMC measurement
equipment such as electric, electronic and mechanical input-output equipment to the test piece,
measurement equipment and LISN used in system-level test. (See Section A3. 1. 3. 2.)

6. 1. 4 EMC verification report
The prime contractor or developer (including the manufacturer) shall create the EMC
verification report. The EMC verification report shall provide evidence of acceptance in eachrequirement defined in this standard. The procedure and objective of each test shall be
described in the report. The summary of results and evaluation indicating the acceptance in thecomponent/subsystem-level verification shall be included. (See Section A3. 1. 4.)

6. 2 System-level verification
6. 2. 1 General
Each subsystem/component shall be assembled in the system to satisfy the acceptancefunction test procedure.

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6. 2. 2 Safety margins of critical/EED circuit
When the safety margin is not determined in advance by the component level test or
analysis, validation shall be performed in combination of components and subsystems bysimulation of actual operation in the system-level integration. The test piece shall be measured
by either directly measuring the actual induced noise or measuring the noise injection that
reduces the S/N ratio to the safety margin (indirect measurement). Select the more appropriate
method in view of technicality and actuality. To validate the safety margin of the circuit (includingEED) having sensitivity to time domain phenomenon, the time domain method shall be used.
(See Section A3. 2. 2.)

6. 2. 3 External electromagnetic field environment
The space system shall be exposed to the external electromagnetic environment defined tocomply with Section 5. 1. 1. When it is not practical to expose the space system to the external
electromagnetic environment, the compatibility to this requirement can be verified by analyzing
the component/subsystem-level test data.

6. 2. 4 Intrasystem EMC
The intrasystem EMC shall be validated by means of a proper combination of the test andanalysis. Before carrying out the system-level EMC test, all component/subsystem shall beassembled in the system and meet the function/performance requirements. (See Section A3. 2.
4.)

6. 2. 5 EMI control
The advance verification results of component/subsystem performance based on therequirement (subsystem design requirement) in Section 5. 2 shall be the basis for system-level
test requirements. The mission-specific or special test procedure for operation for all devices
composing the intrasystem EMC matrix (See Section A3. 2. 4) shall be control target, if any.

6. 2. 6 Grounding and wiring design
6. 2. 6. 1 Grounding
System-level electric grounding and insulation shall be verified by the system-levelgrounding and insulation design and testing during system assembly illustrated in the
grounding diagram (single phase connection diagram is not sufficient).

6. 2. 6. 2 Wiring design
The wiring classification category shall be verified by design review and inspection.

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6. 2. 7 Electrical bonding
6. 2. 7. 1 General
The compatibility to bonding requirements shall be verified by test, analysis, or inspection
appropriate to individual bonding measures. The compatibility to corrosion-proofing control
technology shall be verified by demonstrating that corrosion-proofing treatment is performed asa part of manufacturing process. (See Section A3. 2. 7. 1.)

6. 2. 7. 2 Power current feeder and return paths
The bonding of power current feeder line shall be validated by analytic measurement of thecurrent path, current level and bonding resistance. (See Section A3. 2. 7. 2.)

6. 2. 7. 3 Shock and safety hazards
Not applicable.

6. 2. 7. 4 Antenna counterpoise
The bonding to the antenna counterpoise shall be verified by test, analysis and inspection
appropriate to individual bonding. (See Section A3. 2. 7. 4.)

6. 2. 7. 5 RF potential
Verification on milliohm-level bonding for RF interference control shall be carried out using
the milliohm meter of special AC low voltage output. The meter voltage output shall be
alternating current to eliminate the effect of contact electrode potential. If only the DC output
meter is available, the true value of bonding resistance can be obtained by using the average oftwo measured values with different polarity. The multimeter may be used when the bonding
path is fault/return/pass. However, in that case, measurement at low voltage and low current
shall be finished first. (See Section A3. 2. 7. 5.)

6. 2. 7. 6 Electrostatic discharge
The bonding in the metal portion in which electric discharge element, thermal blanket or
electrostatic potential must be the same shall be verified in testing during assembly to thestructure. (See Section A3. 2. 7. 6.)

6. 2. 7. 7 Explosive atmosphere protection
Not applicable.
6. 2. 8 Antenna-to-antenna compatibility

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The analysis for identifying the frequency with risk element shall be specified as a part of the
EMC control plan. These shall be verified through validation under the operation about thesame as the actual operation. Generally, a combination of interfering elements and interferedelements shall be operated to maximize the potential of interference development. However,
the operation mode shall be limited to the simulated mission operation. In validating thecompatibility between the receiver to be possibly interfered and the interfering elements, theminimum reception capacity of the wanted signal of the receivers shall be included. The
absence of intermodulation interference shall be verified by a combination of analysis andtesting. (See Section A3. 2. 8.)

6. 2. 9 Lightning discharge
Protection against direct and indirect effects of lightning discharge shall be verified by aproper combination of test, analysis and inspection. (See Section A3. 2. 9.)

6. 2. 10 Spacecraft and static charging
6. 2. 10. 1 General
The effects due to electrostatic charge shall be controlled properly by testing, analysis orinspection. (See Section A3. 2. 10. 1.)

6. 2. 10. 2 Differential charge/discharge
The proper control on effects of differential charge/discharge shall be verified by a proper
combination of testing, analysis and inspection. (See Section A3. 2. 10. 2.)

6. 2. 10. 3 Internal charging
The proper control on effects of internal charging shall be verified properly by testing,analysis or inspection. (See Section A3. 2. 10. 3.)

6. 2. 10. 4 Charging of fluid lines (liquid fuel lines)
In the proper control on the static charging of the fluid piping lines (liquid fuel lines), it shall beverified in confirmation that the conductivity of fluid and fluid piping can prevent arc discharge.
(See Section A3. 2. 10. 4.)

6. 2. 11 Hazards of electromagnetic radiation
The safety confirmation of effects due to electromagnetic radiation on the fuel, personnel,
explosive article (including EED), and flight/engine /thruster control shall be validated by aproper combination of testing, analysis and inspection.

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JERG-2-241

6. 2. 12 Consideration of life cycle
The concept of system design introduced for EMC shall be checked for compatibility with thelife cycle requirements of reliability, maintainability and serviceability. The maintainability, easeof test and ease of degradation confirmation shall be validated. The maintenance method andtools shall be defined in the EMC verification plan and proper maintenance document.
6. 2. 13 External grounds
The proper layout and display of external ground position in the space system shall beverified by inspection. The compatibility to bonding requirements shall be verified by testing.

6. 2. 14 Test equipment/facility interface
It shall be verified by a proper combination of testing, analysis and inspection that the test
equipment and facility cause no damage to the space system. It shall be verified that themeasures against thunder stroke and power outage are proper in the EMC verification plan.

6. 2. 15 Spacecraft magnetic emissions
The compatibility to requirements of magnetic emission in the spacecraft shall be validated
by a combination of analysis and testing. (See Section A3. 2. 15.)

6. 3 Subsystem-level verification
6. 3. 1 General
MIL-STD-461 standard test method shall be used in the compatibility verification to Section 5.
2. 1 to 5. 2. 17 (subsystem design requirements) wherever possible. See Table 5. 2-1 and Table
5. 2-2 for MIL-STD-461 application items. The purpose is to minimize the test cost (normally tominimize the procurement of the hard-to-get EMI test equipment).
The standard test method may be inappropriate to EMC requirements in some cases. The
standard test method may not respond to space system-specific EMC problems. These specificissues shall be described in the following items. The test distance between the test piece andthe antenna shall be "1 m" during measurement and peak detection of radiated emissions. Aslong as the test piece is set onto the ground plane as in the case with the actual flight, anotherstandard may be used. Any specification not conforming to this evaluation standard shall be
reviewed by the EMC control organization.
In the electromagnetic field emission and immunity tests, the form in which the harness to beconnected to the test piece is installed to the actual system shall be simulated to the extent
possible.
The measurement bandwidth, sweep rate, or step width shall be controlled in emission
measurement in which the frequency band must be swept. The sweep rate shall be slow
enough so that the band-pass filter of the minimum intermediate frequency can respond

31

JERG-2-241
properly. The step width shall be less than half the measurement bandwidth. For the operationof the test piece, the frequency sweep time shall be taken into consideration. Unwanted
radiation sweep must be repeated several times for all unwanted radiation measurements incomponents/subsystems repeatedly operated.
The test frequency point and application time shall be controlled in the immunity test
requiring frequency band sweep. For the equipment operated repeatedly, the operation modewith the maximum noise shall be identified before conducting the frequency domain test of
conductivity emission and radioactive emission.
The test procedure for component-level test (including setup) shall be evaluated by the EMC
control organization.

6. 3. 2 Power-induced power line conductivity emission, time and frequency domaincharacteristics(Primary power line conduction emissions and power source)
The power-induced power line conductivity interference time domain and frequency domaincharacteristics shall be verified by testing. The voltage ripple time domain characteristics shallbe measured at the both ends of the resistance load using the oscilloscope directly. The voltage
ripple frequency domain characteristics can be measured similarly. (See Section A3. 3. 2.)

6. 3. 3 Load-induced power line conductivity emission, frequency domaincharacteristics(Primary power line conduction emissions and load)
The frequency domain characteristics of load-induced power line conductivity interference
shall be verified by testing. The requirements may be defined as the voltage or current ripple
standard. The frequency domain LISN shall be specified. LISN cannot operate properly at 150
kHz or less. Conducted emissions at 150 kHz or less shall be controlled solely by current
control. At a frequency of 150 kHz or less, the bandwidth shall be 2% or less of the tuning
frequency. However, the bandwidth shall be set so that the noise floor (lowest observable noiselevel) caused by the local oscillator near the DC domain will be at least 6 dB below therequirement specification standard. (See Section A3. 3. 3.)

6. 3. 4 Load-induced power line switching transient(primary power line switching transient)
6. 3. 4. 1 Application
Both the long-duration switching transient caused by current change and resistance change
and the rapid switching transient caused by current change shall be examined. (See Section A3.
3. 4. 1.)

6. 3. 4. 2 Control of long-duration load-induced switching transients (inrush current)

32

JERG-2-241

Verification shall be conducted by the load-induced long-duration switching transient.
(See Section A3. 3. 4. 2.)

6. 3. 4. 3 Control of fast load-induced switching transients (current changing rate)
The load-induced rapid switching transient shall be verified by testing. The transient voltage
within 50 microseconds (μs) can be measured by using the LISN used for measuring thefrequency domain characteristics of conductive emissions. When measuring the transient
voltage of 50 μs or more, the space system-specific power source impedance model shall beestablished. To measure the time domain disturbance from voltage application to steady statewith accuracy, the LISN impedance shall be defined to the DC domain.
The power source impedance in the transient-producing portion of test setup shall be larger
than the estimated impedance of the power line. When the power source switching is notincluded in the test piece, the transient data when the power is turned off shall be regarded as
reference. (See A3. 3. 4. 3)

6. 3. 5 Load-induced power-line ripple (primary power-line ripple)
The load-induced power-line ripple shall be verified by testing. (See Section A3. 3. 5.)

6. 3. 6 Signal line conduction emissions
The signal line conduction emissions and frequency domain characteristics shall be verified
by testing. The compatibility to this requirement shall be verified by mounting the current probearound each cable to be tested. Other verification methods may be adopted as long as the EMC
control organization agrees. (See Section A3. 3. 6.)

6. 3. 7 Antenna connection port spurious emissions
The antenna connection port spurious emissions shall be verified by testing. If possible, thecompatibility to the requirement shall be verified by connecting the coaxial cable/wave guidedirectly between the antenna connection port and the EMI meter. Impedance mismatch shall be
dealt with the impedance conversion transformer or correlation network in consideration of theloss in the relevant portions. If direct connection measurement is not effective, the 2 antenna
test method shall be implemented. In that case, the reception antenna used in the test shallhave characteristics similar to those of the interfered antenna to the extent possible. (See
Section A3. 3. 7.)
6. 3. 8 Magnetic field radiated emissions
Magnetic field radiated emissions shall be verified by testing. (See Section A3. 3. 8.)

6. 3. 9 Radiated electric field emissions

33

JERG-2-241

Radiated electric field emissions shall be verified by testing. (See Section A3. 3. 9.)

6. 3. 10 Immunity to power-line ripple
Immunity to power-line ripple shall be verified by testing. (See Section A3. 3. 9)

6. 3. 11 Immunity to power line switching transient
Immunity to power line switching transient shall be verified by testing.

6. 3. 12 Immunity to the conducted effects of radiated electromagnetic fields
The induction effect immunity in the radiated electromagnetic field shall be verified by testing.
The current/voltage set in the specification shall be injected by the current probe of which loss
frequency characteristics are known.

6. 3. 13 Immunity to magnetic emissions
Immunity to magnetic emissions shall be verified by testing. It is desirable that compatibility
be validated by the Helmholtz coil. The local magnetic field source (portable type) may be used
when the test target is large. These two magnetic field sources shall be calibrated by thephysical dimensions and electric current.

6. 3. 14 Immunity to field emission
Immunity to field emission shall be verified by testing. It is desirable that testing be carriedout in the anechoic chamber.

6. 3. 15 Immunity to signal line magnetic field conduction
Immunity to signal line magnetic field conduction shall be verified by testing.

6. 3. 16 Antenna terminal out-of-band immunity
Antenna terminal out-of-band immunity shall be verified by testing. The test method largely
depends on the types of receivers to be certified. The test method shall be described in a proper
chapter in the EMC control plan before starting the test.

6. 3. 17 Immunity to electrostatic discharge
Immunity to electrostatic discharge shall be verified by testing or controlled by approved
handling procedure. The testing shall be conducted according to IEC 61000-4-2. Because the
ESD test may cause devastating damage to the test target (including progressive and potential

34

JERG-2-241
damage elements), verification shall not be implemented with the flight model but only withengineering model or the prototype model.

6. 4 Measurement instrument
The EMC verification shall be conducted using the proper measurement sensor under theenvironment with the background noise reduced properly.

6. 4. 1 Immunity test equipment
6. 4. 1. 1 Immunity to conduction
To superimpose the conductivity noise with the power line to be measured, the signal
generator of required frequency band, audio isolation transformer (a few tens of kHz or less),high pass filter (a few tens of kHz or more) or current probe for injecting bulk current shall be
used. The power amplifier shall be required depending on the required noise level.

6. 4. 1. 2 Immunity to field emission
The power amplifier shall be necessary depending on the signal source, antenna and levelof the required frequency band. At a low frequency of a few tens of MHz or less, the parallel
element antenna is used in many cases. The biconical antenna is used at a frequency of a few
tens of MHz to a few hundreds of MHz band while the log periodic antenna or horn antenna isused at higher frequency. The electric field probe for monitoring the applied electric field value
shall be required in measurement. To control the electric field strength loaded to the test piece,
the power supplied to the antenna performs feedback control on the electric field strength
obtained from the electric field sensor placed near the test piece.

6. 4. 1. 3 Immunity to magnetic emissions
For the immunity to the magnetic field, the measurement device comprising of the Helmholtz
coil and the signal source for driving the alternating current shall be used. The magnetic fieldsensor for monitoring the magnetic field applied shall be required in measurement.

6. 4. 1. 4 Immunity to electrostatic discharge
The ESD gun shall be used in the immunity test to electrostatic discharge.

6. 4. 2 EMC sensor
6. 4. 2. 1 Conduction emissions
The current probe having a proper sensitivity to the required frequency bandwidth shall be
used for measuring conduction emissions.

35

JERG-2-241

6. 4. 2. 2 Radiated electric field emissions
In measuring the electric field radiation, it is of importance that the sensor in accordance withthe frequency bandwidth is used. Generally, the monopole antenna shall be used in the low
frequency domain of a few tens of MHz or less. The dipole antenna and biconical antenna shallbe used at a few hundreds of MHz or less. In GHz band, the horn antenna, spiral antenna or log
periodic antenna shall be used.

6. 4. 2. 3 Magnetic field radiated emissions
In measuring the magnetic field radiation, it is of importance that the sensor in accordancewith the frequency bandwidth is used. The flux gate magnet meter shall be used in thefrequency ranging from direct current to a super low frequency of a few Hz. In the bandwidth of
100 kHz or less, the search coil magnet meter shall be used. In any higher frequency bandwidth,
the loop antenna shall be used.

6. 4. 3 Shield room/anechoic chamber
EMC measurement shall be carried out under the measurement environment that canrealize the electromagnetic environment satisfying the EMC standard.

6. 4. 3. 1 Shield room
Measurement of radiation noise from the direct current magnetic field to the low frequency
magnetic field shall be conducted in the magnetic shield room that can attenuate the low
frequency magnetic field from the direct current including the magnetic field of earth. When themeasurement of low frequency magnetic field is not necessary, measurement may be
conducted in the radio wave shield room having no magnetic field shield effect.

6. 4. 3. 2 Anechoic chamber
In field emission at a few tens of MHz or more and immunity measurement, measurement
shall be conducted in the anechoic chamber to suppress the wall reflection of radio wave fromthe noise source or radiation source for immunity measurement.

6. 4. 4 Background noise reduction
Background noise shall be reduced enough so that measurement at a level specified in the
EMC standard can be conducted. Depending on the measurement target, the power lines, dataand control lines from the shield room and external anechoic chamber must be drawn and thesignal lines exist for the EMC sensor. After arranging these required cables, measure thebackground noise in advance and check that the noise is enough to carry out the EMC
measurement. If the noise is so large that the measurement is interfered, check the ground36

JERG-2-241
connection and power source grounding method for the shield of the cables drawn from theoutside to reduce the background noise.

37

JERG-2-241
Appendix
A1 system requirements
A1. 1 Grounding and insulation requirement

Specific values are defined in MIL-STD-1541A and MIL-STD-464C.
An insulation resistance of 1 MΩ is used typically.

A1. 2 Filtering criteria

As a shield margin for the structure external, 6 dB or more is typically required.

38

JERG-2-241

A2 Measures against magnetic field of spacecraft
The reference information for spacecraft magnetic field measures shall be as follows:

A2. 1 Magnetic material
A2. 1. 1 General magnetism

Magnet (used for interior of motor, valve and relay) and iron.
Stainless has magnetism depending on the type (such as martensite system) and must betreated as a magnetic material in some cases.

A2. 1. 2 Magnetism requiring care in handling (spacecraft-specific)

Pure nickel (used for coating and plating on the hood of battery electrode and opticalequipment) and magnetic nickel alloy (invar (used for CIC board, etc)). Used for enclosures ofequipment with a high calorific value.)A2. 1. 3 Non-magnetism

Copper, aluminum, titanium, magnesium alloy

A2. 2 Magnetic field reduction measures
A2. 2. 1 Magnetic shield

By surrounding the source of magnetic field with high-magnetic permeability material,
magnetic emissions can be trapped. Steady reduction is possible with the magnetic field sourceof which magnetic field fluctuates. Materials having a high magnetic permeability and relatively
low coercive force (magnetization) such as permalloy are most effective. To suppress the massof shield material, it is effective to cover the minimum portion including the magnetic field source.
Even when using a material having a low coercive force, control on magnetization prevention isnecessary because magnetization is not reduced to zero. When using soft iron having a
relatively high coercive force, pre-operation demagnetization and more careful control onmagnetization prevention are necessary.

A2. 2. 2 Cancel magnet

The force exerted on the spacecraft can be reduced with the external magnetic field by
mounting the magnet in the opposite polarity direction of the spacecraft magnetic moment tocancel the force exerted between the external magnetic field and the magnetic moment of thespacecraft. By mounting a magnet having the same magnetic moment and opposite polarity
with respect to the magnetic field source in the vicinity, the magnetic field placed far enough incomparison with the distance between the source and the added magnet can be limited tominute. However, when developed magnetic field changes in terms of time, the cancel magnet
shall be selected and installed in accordance with the case in which the magnetic field must be
canceled more often.

39

JERG-2-241
A2. 2. 3 Cancel loop

By arranging wiring with the current on the hot side balanced with that on the return side sothat the current follows the same path, the current loop area can be downsized, which leads toreduction in magnetic field developed. If possible, it is desirable that the twist pair wire be
employed. If the current is not balanced, by adding the current path having the same current
strength inversely, the magnetic field developed can be canceled. It is effective as magnetic
field measures by means of current loop with varying current strength.

A2. 2. 4 Scheme of design and installation method

When installing more than one device generating magnetic field, magnetic field can be
reduced by mounting a pair of devices inversely each other. When using more than one magnet
in the component, magnetic field developed can be reduced by arranging the magnets so thatthey can cancel out the magnetic moment each other. For example, it is effective to install therelay and circulator opposite in the equipment. The magnetic field can be canceled out more
effectively by selectively using the magnet having similar strength from plural magnets having
varied strength. Unless the devices are arranged antiparallelly, components will be left
uncanceled despite the same magnetic moment strength. Thus, the mechanism to fine-tune theinstallation direction shall be provided as needed.
For the layout in the spacecraft for measuring weak magnetic field, a proper distance shallbe ensured between the source of the magnetic field and the magnetic field sensor.

A2. 3 Magnetization prevention control(1) On-board equipment may be magnetized depending on the tooling. All tools and diesused shall be made of non-magnetic material or demagnetized. The use of a driver using
magnet on the tip shall be strictly prohibited. To prevent the demagnetized tooling frombecoming magnetized, the tooling shall be stored and kept in a non-magnetic tool box
such as plastic box. The demagnetized tooling shall be distinguished with a label fromothers. Magnetization level shall be measured on the contact surface of the tooling (suchas torque wrench having a ratchet in particular) by using the gauss meter/tesla meter
before use as needed.
(2) Among manufacturing or testing facilities, some are magnetized or generating magnetic
field; care must be exercised. Check the magnetic field strength in advance and ensure
the distance according to the strength.
(3) Special care must be exercised with the carry-on items of workers getting near the
spacecraft. Carry-on items shall be determined depending on the magnetic fieldrequirements.

A2. 4 Equipment apt to radiate magnetic field(1) Power source system equipment (small loop can become a large magnetic field source
due to large current)(2) Battery (large-current loop is formed when assembled in the unit)(3) Solar cell panel (The panel routing becomes the loop.)(4) Isolation valve (Strong magnet is used generally.)(5) Communication system circulator (Strong magnet is used generally.)40

JERG-2-241(6) Propellant valve (Magnetic material is used in some cases.)(7) Communication system switch (Magnet is used in some cases.)(8) Large current relay (Magnet is used in some cases.)(9) Step motor, solar cell paddle or antenna motor (Note that the magnetic field directionchanges according to the drive.)41

JERG-2-241

A3 System verification/test related
A3. 1. 3 System-level EMC verification planA3. 1. 3. 1 System-level verification method
The general system-level verification method shall include the following details:(1) Selection method for the critical circuit of which compatibility to degradation standard
and safety margins must be monitored.
(2) Malfunction judgment criterion and limit determination procedure(Example: In addition to checking that the signal of known strength can be received
during interference verification of the receiver, it is of importance to check the noise
suppression sensitivity without the signal input.)(3) System-level EMC verification method(4) Power line conductivity margin verification method in normal/common mode. (About
distribution output and critical circuit in particular)(5) Verification method for common mode margin of cable bundle. (The cable bundle of thecritical circuit and bundle arranged near the sensor or critical circuit in particular)(6) RF self-compatibility (compatibility between the transmitter, receiver and entire
spacecraft in all operation modes) verification method(7) Static charging measures design verification method and lightning strike measuresprogram(Advance planning for ensuring the safety margins of EED with the safety margins
supersensitized or mounted is necessary for timely implementation of EMC verification.)(8) Simulation/test method for electro-explosive device (EED)(9) Power quality verification and power bus monitoring method(10) EMC verification method in system-to-system interface such as rocket interface.
(11) Effect due to cavity resonanceA3. 1. 3. 2 Test condition
The parameters required for setting test conditions shall include the following:(1) Test conditions for all electric/electronic devices that are assembled in the space systemor required for test procedures. (The intrasystem EMC matrix shown in Section A3. 2. 4
shall be included in a part of system verification.)(2) Implementation and application of test procedure including monitor points for operation
mode and each subsystem/component(3) Utilization of test results that are obtained and judged to be valid in interference test forsubsystem/component.
(4) Data report/record method and analysis method(5) Description on the equivalent test implementation location and installation method when
the operation of the actual unit is unrealistic.
(6) Configuration of subsystem/component operation mode for testing the interferedsubsystems/components at the maximum sensitivity mode and for testing the interfering
subsystems/components at the maximum noise level mode. (Because eachsubsystem/component can be interfered or interfering, it is necessary to conduct testing42

JERG-2-241
in more than one operation mode.)(7) Details on the frequency range, channel and specific combination of test targets.
(Example: Image frequency, intermediate frequency, local exciter frequency, transmitter
fundamental wave and higher harmonics-related frequency, and reception bandwidthand sweep rate of measurement equipment (The high-sensitivity frequency of thesubsystem confirmed in test phase shall be included.))A3. 1. 4 EMC verification report
The items to be included in the EMC verification report shall be as follows:(1) Setting of specific test purpose in consideration of applicable requirements and EMC
verification plan reference.
(2) Test piece information including serial numbers, configurations, figures and pictures.
(3) Description about repair and configuration change of the test piece proved to beincompatible in verification(4) Summary of verification results. (Including the summary of the degree of compatibility torequirement)(5) Differences from the test facility, analysis method/tool and inspection equipment
specified in the EMC verification plan.
(6) Differences from the test procedure specified in the EMC verification plan.
(7) Figures and pictures representing the test setting(8) List of test equipment including adjustment information(9) Record data and log. (Including the measurement equipment measurement value,correction coefficient, and integrated data (The data integration method shall be
described. If the data is compromising due to restriction of test conditions, the reasonsand effect on the data shall be described.))(10) Specification of surrounding and other test conditions.

A3. 2. 2 Safety margins of critical/EED circuit

When the SN ratio can be reduced during operation by applying the safety margin coefficient,
the EUT circuit may not be measured directly. The presence of safety margins shall be validated
when the subsystems/components are operated properly. The typical method is that the EED
circuit is replaced with a high-sensitivity EED (Bridge wire fuse is provided at 1/10 current levelwith the EED unit turned off.) and devices related to thermocouple and other temperature
measurement are connected to the real bridge wire.

A3. 2. 4 Intrasystem EMC

As an effective verification method for intrasystem EMC, the intrasystem EMC matrix iscreated. The intrasystem EMC matrix represents a combination of all subsystems/componentsof the intrasystem EMC to be verified. An example of the basic format of the concerned matrix
shall be shown in Table A3. 2. 4. The test procedure for all matrix equipment shall be included to
conduct testing smoothly. Special equipment required for operation of interfered/interfering
equipment shall be included if any. In any combination of interfered /interfering equipment, it
shall be verified before compatibility validation that the interfering equipment can operate

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properly and that the interfered equipment can operate properly when the interfering equipment
is turned off.

Interfering/interfered
Component A
Component B

Table A3. 2. 4 - Intrasystem EMC matrix
Component
Component
Component
A
B
C
N/A
Test notrequired
Test

Component C-

Component

-

N

Test

Test
N/A

Test
Test notrequired
N/A

Test
-

Test notrequired

-

Test

---

Component
N

-----

Test
Test

---

Test notrequired
-

N／A

---

N/A

a "N/A" means "not applicable".
b "Test not required” means that the two concerned devices cannot be operated at the sametime or there is no operation leading to mutual interference. Thus, validation is not required.

A3. 2. 7 Electrical bonding
A3. 2. 7. 1 General

When verifying a metal or conductive composite structure, verification shall be normally
carried out according to the validated bonding process. For the verification of the dielectricsurface with conductive treatment, the surface electric resistance ratio measurement and theelectric contact of conduction pass shall be required.

A3. 2. 7. 2 Power current feeder and return paths

The final item in power bus power source supply capacity verification is to check all the
on-board equipment for application of proper voltage with all the power subsystems installed.

A3. 2. 7. 3 Shock and safety hazards

Not applicable.
A3. 2. 7. 4 Antenna counterpoise

The antenna counterpoise can be verified by operating the system in the compact range
which is an antenna pattern measurement facility.

A3. 2. 7. 5 RF potential

Low voltage output shall be applied to the milliohm meter so that the bonding causing effecton the EMI induced current is simulated to the bonding measurement value. A high-voltage
source may penetrate the oxide film or contamination layer and supply an excessively highimpedance to EMI signal on milliampere/millivolt scale. Thus, the high voltage short circuitcurrent test shall be conducted after the low voltage RF bonding test. This test procedure must

44

JERG-2-241
be observed. Even if the short circuit current bonding measurement value is below the specified
limit value of the RFI bonding, the bonding proved to be compatible to the short circuit current
requirements may not be compatible to the RFI bonding requirements. The maximum output
voltage of the low voltage output meter is 20 mV, the typical output is 200 μV and the test
current range is 1 μA to 10 mA.

A3. 2. 7. 6 Electrostatic discharge

The bonding shall be verified for compatibility to preventive requirements of electrodepotential formation that serves as the electrostatic discharge source. To ensure that properbonding can be maintained, inspection shall be conducted on the surface treatment applied to
all materials connected to the structure.

A3. 2. 7. 7 Explosive atmosphere protection

Not applicable.

A3. 2. 8 Antenna-to-antenna (RF) compatibility

When there is no signal to be received or simulation signal input, check the noise
suppression sensitivity as needed in addition to checking the reception antenna for noise.
For the reception antenna in which a signal source proper to the interfered reception
antenna test cannot be used or the performance evaluation cannot be completed without thesignal, the reception antenna shall be removed from the feed and replaced with the test
reception system having a similar reception sensitivity. Thus, the test reception system can beused for quantifying the RFI level input to the reception antenna and comparing it to theinterfered reception antenna sensitivity.
However, the noise suppression capacity of the interfered reception antenna cannot be
simulated in this test. Generally, to ensure that the maximum sensitivity to the reception
antenna noise is grasped, each reception antenna shall be operated after setting anotherantenna to the maximum power /antenna direction without implementing intentional signal
input.

A3. 2. 9 Lighting discharge

Refer to SAE ARP 5412[10] for space system test-related recommendation items. The SAE
standard specialized to the test method includes the detailed information about lightning strikeverification.

A3. 2. 10 Spacecraft and static charging
A3. 2. 10. 1 General

Applying the computer model shall be effective to evaluate the static charging degree on thespacecraft surface.

A3. 2. 10. 2 Differential charge/discharge

It is recommended to verify the differential static charging by testing. It is advisable to use theradioactive ESD source. Pulse energy shall be set in consideration of the static

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charging/emission risk specialized to the mission.
If current is possibly injected to the structure depending on the component, the conductive
ESD source shall be applied.
(1) Immunity test of discharging electrostatic caused by static charging
The recommended test method shall be shown in Figure A3. 2. 10. 2-1.

EGSE

EGSE

LISN

Test piece

Spark gap
Grounding surface

Attenuation resistance

Choke resistance

Choke resistanceHigh voltage
sourceFigure A3. 2. 10. 2-1-Immunity test of discharging electrostatic caused by static charging

The explanation of circuit parameter specified in Figure A3. 2. 10. 2 shall be as follows:a) Spark gap
The typical value shall be 6 kV. The pressurized/sealed spark gap with a short
breakdown time shall be desirable. There shall be no cavities.
b) Capacitance
The typical value shall be 100 pF by the low inductance high voltage condenser.
c) Attenuation resistance
Although the typical value is 47 Ω, adjustment is possible in accordance with thecritical attenuation determined by the capacitance and self inductance of the electric
discharge circuit.
d) Choke resistance
It shall be used to prevent the current from flowing into the uncontrolled path due todischarging high-frequency components. If it is noted that the minimum value is 10 kΩ,the electric discharge parameter will not be dependent on the length and position of
the wiring of the high voltage source.
e) High voltage source
Though the high voltage source may be direct current, a choke resistance of 10 MΩ ormore shall be used. In terms of safety, it is desirable to use the ESD generator
specified in the IEC 61000-4-2. This generator corresponds to aerial electric
discharge mode on the precondition that the generator is connected to the electric
discharge return connection of which electric discharge electrodes are connected to
one choke resistance and the other choke resistance.

46

JERG-2-241f) Electric discharge circuit
The circuit shall be isolated and arranged along the EUT harness by 20 cm.
g) Transient current pulse
The goal is that the half width is maintained at 30 A for 30 ns.
(2) Immunity test of field emission caused by arc discharge
Testing shall be conducted by generating pulse electric discharge having a pulse rate of
1 pps (pulse per second) and a voltage of 10 kV from the electrode at a location
separated by 30 cm from the test piece. Testing shall be conducted on each surface ofthe test piece.
If a malfunction occurs at a voltage level of 10 kV, the voltage level shall be lowered until thedevice operates normally. The level shall be recorded. The test equipment thatgenerates the required voltage and arc discharge is shown in Figure A3. 2. 10. 2- 2.
When using other equivalent test equipment, indicate so in the EMC verification plan.
A3. 2. 10. 3 Internal charging

Conducting internal charging test is extremely difficult. It is required to analyze andimplement the design method for shielding with the use of the grounded aluminum having aproper thickness.

A3. 2. 10. 4 Charging of fluid lines (liquid fuel lines)

If possible, the resistivity of the liquid fuel shall be controlled by using additive. All fluid pipes
comprising of metal braiding (external or internal) shall be provided with clarified grounding withrespect to the metal braiding.

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1kΩ

0.1μF

Turns ratio 100:1
50pF（floating capacity）

DC power source
0 to 600 volts
500Ω

Variable gap

SCR

Carbonelectrode

6V

Signalgenerator

Note 1 The signal generator shall have the capacity that can drive the relay at 1 pps.
Note 2 For the gap, the carbon electrode installed to the variable Teflon shaft shall be mounted
onto the phenol plate.

Gap (mm)

Relationship between the electrode gap and breakdown voltage
Breakdown voltage (KV)
Emission energy (W･s)-61
1.5
56.5×10-6
2.5
3.5
305×10-65
6
900×10-6
7.5
92000×10Figure A3. 2. 10. 2-2-Immunity of field emission caused by arc discharge

A3. 2. 15 Spacecraft magnetic emissions(1) Constraint of costa) Because the total moment of the spacecraft is not so high, simplified calculation ispossible. The procedure of the method is as follows: convert the moments of primaryfactors (equipment having high-magnetic materials known) to scalar values and then
combine the values; perform simplified statistical treatment on the vector direction forother factorial equipment (Example: use of root sum square (rss)); and then add thetotal to the former.
b) The equipment shall be verified for compatibility to these requirements by equivalenttest or "rough" test.
c) The moment of primary factorial equipment and the total of statistical distribution (rss)of other equipment shall be added.
(2) Constraint of on-board magnetic field sensora) Main causes related to the sensor involve the sensitivity of the sensor to the magnetic
field. The sensitivity shall be specified as the maximum tolerance requirement withrespect to the magnetic emissions generated from the space system. Measurementof magnetic moments of the three orthogonal axes in the unit allows you to convert
the values to magnetic field strength and determine the sensor position based on the
results. Conducting similar measurement and calculation on all units allows you to
obtain the total magnetic field at the sensor position. It is of importance to separate
48

JERG-2-241
the components properly from the sensor so that the component-level favorable
measurement values and the inverse cube law can be used.
b) Each component shall be verified by testing using more than one Helmholtz
compensation coil.
c) The magnetic flux density and magnetic moment vector at the sensor position that are
obtained from all the test results shall be calculated. This allows you to obtainhigh-precision results.
d) As optional, it is desirable that the space system test is conducted and that thesystem-level calculation results are supported.
A3. 3. 2 Power-induced power line conductivity emission, time and frequency domaincharacteristics(Primary power line conduction interference, power source)

The voltage ripple in the frequency domain can be measured by using the digital
oscilloscope (DSO: Digital Storage Oscilloscope) having the fast Fourier transform (FFT)
capacity or the interface connecting the FFT-implementing computer. Because the noise of
power subsystem under pure resistance load is developed at intervals, there are no problems
as long as the DSO repeated frequency bandwidth is compatible to the power quality
requirements. Capturing the transient response of each single measurement is good enoughfor evaluating the power quality. When using the EMI meter, the high pass filter shall be used toprevent power frequency voltage from damaging the measuring device. When the limit isspecified in that manner, the ripple current in the frequency domain can be measured instead.
The EMI current probe having the high pass filter function inherently may be used. When
applying the conductive emission limit to the frequency domain in order to control radiatedemissions, the EMI current probe shall be used. Because such measurement method is more
likely to be posed as a common mode requirement, the EMI current probe shall be installednear the feeder and return path.
A3. 3. 3 Load-induced power line conductivity emission, frequency domain characteristics(primary power line conduction emissions, load)

LISN in the frequency domain shall be shown in Figure A3. 3. 3 -1. Figure A3. 3. 3 -2 is the
setup of the LISN-based conductivity emission test. The register connected to the LISN EMI
port is 50 Ω. Current control is required at 150 kHz or less. Because, it is impossible tostandardize only the single power source impedance at such low frequency. When conductive
emissions must be controlled at 150 kHz or less, refer to the guideline specified in the section
for audio frequency conductivity emission in SAE ARP 1972 [11].
A3. 3. 4 Load-induced power line switching transient(primary power line switching transient)A3. 3. 4. 1 Application

The direct current bus is adopted in almost space system power buses. Because it allows
the simplification of the transient signal measurement. In the direct current bus, the transient
signal timing is not important and the transient waveform can be identified clearly. If theline-to-line electrostatic capacity is increased and the power input of the test piece is bypassed,
the power supply can be controlled at will for a short period of time.

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Power input

2 電力出力PoweroutputStandard power source

EMI port

Frequency domain LISN
Figure A3. 3. 3 -1 LISN frequency domain

Test piece

50 Ω (EMI meter)
50 Ω (EMI meter)
Figure A3. 3. 3 -2 Example of setup of LISN-based voltage conductive emissions test

A3. 3. 4. 2 Control of long-duration load-induced switching transients (inrush current)
The inrush current from the low impedance source shall be measured. For example, select asource that will not allow the voltage of the total power source to dive when applying power to thetest piece. In other words, when specifying the resistance of the power source for testing, theresistance in the space system power subsystem must not be simulated. The test setting shall beshown in Figure A3. 3. 4. 2. The voltage drop (due to transient rush current) measured in the entire
series resistance when applying power to the test piece shall be smaller than the amount below:∆Vtrans (V nom−Vmin ) ≤

I ss
I bus

Where,
ΔVtrans: maximum allowable voltage drop during rush current
measurement (volt)
Vnom: nominal bus voltage under power quality specification (volt)
Vnom: minimum bus voltage under power quality specification (volt)
Iss: steady-state power pulling of test piece (ampere)
Ibus: space system power bus steady-state maximum current load
(ampere)The internal/series resistance shall conform to the above maximum voltage drop
requirements. The input power condenser of the test piece shall be discharged before transient
signal measurement. By providing the entire direct current power source output with the bypass
capacity, the internal resistance can be reduced effectively.

50

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SwitchSeries resistanceInternal resistance
Power

DSO
Digitalstorageoscilloscope

Test piece

SourceFigure A3. 3. 4. 2 An example of long-duration switching transient test setup

A3. 3. 4. 3 Control of fast load-induced switching transients (current changing rate)

The frequency lower limit is placed on the conductive emission (CE) requirements in thefrequency domain. The LISN impedance cannot be defined under the limit. Figure A3. 3. 4. 3
-1 illustrates the comparison between the frequency domain LISN and time domain LISN.
Because the time domain LISN has no components connecting the line with the ground, thecommon mode impedance of the power source cannot be controlled. Because the time domaintransient signal and the line-to-line ripple of the current return power bus on the ground are
regarded as normal mode phenomenon, measurement is possible. Figure A3. 3. 4. 3 -2
illustrates the measurement test setting using two types of LISN. When applying the frequency
domain test setting to the time domain test, the DSO shall be used instead of the EMI meter. A
load of 50 Ω shall be applied continually to each EMI port. Balanced line-to-line measurement
shall be conducted not to measure the common mode effect of the EMI port. However, note thatthe impedance below 150 kHz cannot be controlled properly with the frequency domain LISN.
(Resonance occurs.)

Power output
Power input
EMI port
Frequency domain LISN

Power
outputPower input
ReturnTime domain LISN

Figure A3. 3. 4. 3 -1 Comparison between the frequency domain LISN and time domain LISN

A3. 3. 5 Load-induced power-line ripple (primary power-line ripple)

When the bandwidth is referenced in the power quality specification and the specific ripple
voltage is defined, the specified bandwidth shall be matched with the bandwidth of the51

JERG-2-241
measurement equipment (digital storage oscilloscope having true rms capacity). The repeatedfrequency bandwidth of DSO shall be the equal to or larger than the required bandwidth of thepower quality specification. The 100-ns waveform capturing capacity is required for the single
bandwidth of DSO. The power source impedance in the entire ripple measurement range shallbe controlled so that it will not exceed the required bandwidth defined in the power quality
specification.
Because both the maximum amplitude and the rms value of the ripple voltage are confirmed,
true rms read values are required when implementing the test. For the power quality
specifications requiring the test pieces receiving the power supply from the direct current bus or
measurement of the time domain ripple at the bandwidth (waveform capture at 100 ns) below
10 MHz, a simplified power source impedance may be established based on the worst-caseinductance and the following formula.

L=d 2N 20.45d + l

Where,
L: inductance (nH)d: coil diameter (mm, however, coil center distance)
l: coil length (mm)
N: number of turns

Figure A3. 3. 5 -1 shows the physical realization of the above inductance formula. The testlayout for direct current equipment time domain ripple measurement shall be shown in Figure
A3. 3. 5 -2. The DSO input impedance shall be 1 MΩ or more. The power source impedanceshall be kept to a minimum so that the ripple voltage measured on the power source side with a
combination of inductance and resistance will be negligibly small in comparison with thespecification limit.

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JERG-2-241

Test piece

Frequency

domainCE

50 Ω (EMI meter)

DSO

Test pieceTime domain CE measurement
Figure A3. 3. 4. 3 -2 An example of frequency/time domain CE test setupdll

Figure A3. 3. 5 -1 Choke coil53

JERG-2-241

Power source impedanceDirect current
power source

Test piece

DSO

Change onrequirement

Figure A3. 3. 5 -2 An example of direct current equipment power-line ripple test setup

A3. 3. 6 Signal line conduction emissions

The bulk current shall be measured in this test. It is desirable measurement be conducted at
a location 5 cm away from the test piece.
The test method shall be shown in Figure A3. 3. 6. This test method is similar to bulk current
injection test method. However, only the current probe for measurement shall be mounted
without mounting the injection probe.
PowersourceShield room

Capacitor
LISN

Current probe for powerline measurement

5cm

Test piece
Current probe for component-to-component
connection line measurement

5cm

Ground planeSpace systemload(shield room floor/side wall)

0.1m

GSE

Spectrum analyzer or EMI receiver
Data recorderFigure A3. 3. 6 An example of signal line/conduction emissions test setup

54

JERG-2-241

A3. 3. 7 Antenna connection port spurious emissions

Prevent the damage caused by transmitter transmitting the current having high-voltage
standing wave ratio. When testing the transmitter in the transmission mode, the notch filter shallbe required. See MIL-STD-462 RE03 for test method.
As long as the antenna is selected and arranged strictly, the limit may be represented not as
effective electric field strength but as the power received by the EMI receiver. This can beobtained by setting the target power to the receiver.

A3. 3. 8 Magnetic field radiated emissions

Low frequency magnetic field emission shall be measured using the electrically shieldedloop antenna or AC magnetometer. The sensor size differs depending on the distance between
the test piece and the loop antenna. The air-core loop having a diameter of 13 cm shall be used
when the interval is 10 cm or less. For the small-sized AC magnetometer, the sensor having a
diameter of 13 cm or less shall be used. The wire used, number of turns, and type of magnetic
material shall be selected so that the measurement system sensitivity will be proper whenconnected to the EMI meter. Because the air-core loop is not effective at low frequency, themost sensitive EMI receiver that can capture the data at 1 kHz or less is required. When thesensor is separated far from the test piece, the size of the sensor shall be matched with that of
the test piece circuit.

A3. 3. 9 Radiated electric field emissions

When the test frequency bandwidth significantly different from the interfered bandwidth (1.5times or more), identification of narrow/broad bandwidth may be necessary. Unless thenarrow/broad bandwidth must be identified individually (special environment), the function ofpeak detector shall be used. The narrow bandwidth data may be evaluated using the averagevalue detector. When interfered antenna characteristics of the space system are adjusted
during testing, the polarization of the test antenna shall be matched with that of the spacesystem antenna. To improve the test quality, both horizontal and vertical polarization shall beapplied.
See MIL-STD-462 RE02 for test method. A typical example of antenna position during
measurement shall be shown in Figure A3. 3. 9. See Section 6. 4. 2. 2 for measurement
antenna selection policy.

A3. 3. 10 Immunity to power-line ripple

In this test, it shall be verified through measurement that the voltage at power input terminalof the test piece is at the specified value. Injection signals shall be applied without modulation.
For the impedance of the ripple generator, select the larger value of the power source
impedances estimated at 0.5 Ω or 1500 Hz or less. The power source impedance increases
when the frequency exceeds 1500 Hz. The requirement compatibility will be validated when any
one of the two conditions is valid.
a) As long as the voltage of the power input terminal of the test piece is at the specified
value, the function/performance of the test piece will not be deteriorated

55

JERG-2-241
unacceptably.
b) Even if the current injected by the ripple generator generates the specified voltage atload impedance regardless of the voltage applied to the power input terminal of thetest piece, the function/performance of the test piece will not be deteriorated
unacceptably.
Note: Unless the power source having enough capacity is used in this test, it may appear that thepower quality is degraded and the EUT sensitivity is high.
Modulated signals may be used. Installing the line-to-line condenser on the power source, voltage
can be applied to the test piece easily.
Groundingplate
Maximum radiatedemissions area

Test pieceequipment

To power source

At least 2 m

1m

Test antenna

10 μF
through-type
capacitor(glued to
groundingplate)

EMI
meter

Figure A3. 3. 9 Antenna position

A3. 3. 11 Immunity to power line switching transient

The power line, switching transient and immunity shall be performed by sending transient
signals from the coupling transformer. A typical test method shall be shown in Figure A3. 3.11.
However, the test method must be changed depending on the power source impedance andtransient signal characteristics of the space system unit specified in the power quality
specification. NASA H-29919D [12] shall be recommended as reference documents on
switching transient immunity.

A3. 3. 12 Immunity to the conducted effects of radiated electromagnetic fields

Care shall be given to the insertion loss requirements of the selected test standard. To apply
the current at the specified upper limit into the system at 50 Ω, the power supplied to the bulk
current injection clamp shall be obtained from the insertion loss.

P = I lim + I L − 73
Where,
P: target power level (decibel × meter (dB･m))
Ilim: specification limit (decibel ×micro ampere (dB･㎂))
IL: BCI clamp insertion loss (decibel (dB))

56

JERG-2-241

Supply the pre-calculated power to the bulk current injection clamp and perform monitoring.
In addition, mount the current probe near the cable to be tested to monitor the injection current.
For this system, set the overcurrent limit of specified upper limit + 10 dB.

Oscilloscope
Pulse generator+ sidePower input

Test piece
10μF

Figure A3. 3.11 Power line/switching transient/immunity test method

A3. 3. 13 Immunity to magnetic emissions

This test shall not be required normally except for science satellite tests using themagnetometer or tests for spacecrafts using the magnetic torque meter.
As long as the coil diameter is electrically shorted and electrode potential generated in theentire coil is negligible (inductive reactance is small), the magnetic field can be calculated withthe current and physical parameters using the following formula. The formulas (1) and (2) shallbe applied to Helmholtz coil and small-sized coil, respectively.

8µNI
125d
µNId 2B=2(d 2 + r 2 )3 / 2

B=

(1)
( 2)

Where,
B: induction field (tesla (T))-7
μ: magnetic permeability in free space (=4×10 H/m)
N: number of turns, I: current (ampere (A))
d: coil diameter (meter)
r: distance along shaft from loop (meter)

The important difference between the two coils is that the internal magnetic field of theHelmholtz coil is even while the magnetic field of the small-sized loop is weakened in proportion
to the distance. The formula of the small-sized loop can only be applied to on the axis because
the magnetic field fluctuation is complex in places other than on the axis. The test using theHelmholtz coil has a far higher repeatability and enables verification under more strict
conditions.

57

JERG-2-241
A3. 3. 14 Immunity to field emission

Only when the test piece is static with the course of time (no repeated cycle, only
steady-state), the mode adjustment method shall be applied. Rather than continuous operation,
step operation is more appropriate for the mode regulator. When the test piece is large andcomplex, the sensitivity can be determined early in a relatively short period of time if mode
adjustment is implemented in advance.
If the sensitivity is unknown in the initial phase, the uncertainty of the aspect angle in the test
stroke ISO 7137-3.6 (Section 20.5) becomes higher and irradiation is repeated often, which
may prolong the test time. The mode adjustment method is useful as a preliminary diagnosis
method toward more strict and higher precision technology such as test stroke ISO 7137-3.6(Section 20.5).
When adjusting the antenna characteristics which will be interfering element during testing,the polarization of the test antenna shall be matched with that of the interfering antenna. In thetest chamber during attenuation, the polarization of the test piece may be different from that of
the test antenna. Thus, to improve the quality of test, both polarizations (horizontal and vertical)
shall be used.

A3. 3. 15 Immunity to signal line magnetic field conduction

Immunity to signal line magnetic field conduction shall be conducted by arranging the line
applied with transient voltage parallel with the cable of the test piece. Typical test method shallbe shown in Figure A3. 3. 15.
A3. 3. 16 Antenna terminal out-of-band immunity

The verification method largely differs depending on the receiver to be tested. For thequalification of the conventional superheterodyne receiver, the test methods specified in CS03,
CS04 and CS05 in MIL-STD-461C shall serve as a guideline. (CS103, CS104, CS105 in
MIL-STD-461E)
A3. 4. 1 Immunity test equipment

It is desirable to verify before testing that the EMI equipment used for measurement in the
immunity test can measure 6 dB below the application standard limit.
Test piece orload simulation
deviceComponent-to-component connection line

<5mm

Length(L)>1m>0.15mTest piece>0.15mI
Figure A3. 3. 15 Signal line magnetic field conduction immunity test method58