General
JERG-2-214

DESIGN STANDARD
POWER SUB SYSTEM

July 8, 2009

Japan Aerospace Exploration Agency

JERG-2-214

This is an English translation of JERG-2-214. Whenever there is anything ambiguousin this document, the original document (the Japanese version) shall be used to clarify
the intent of the requirement.

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

JAXA

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

JAXA 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, Japani

JERG-2-214

Table of contents
1. Scope of application ..................................................................................................................1
2. Applicable documents ...............................................................................................................1
3. Terms, definitions and abbreviations ........................................................................................1
3.1 Definitions of terms ...........................................................................................................1
3.2 Abbreviations and symbols ................................................................................................2
4. Design requirements ...............................................................................................................10
4.1 General .............................................................................................................................10
4.2 Definitions and restraints of items ...................................................................................10
4.2.1

Configuration and components ....................................................................................10

4.2.2

Operation mode and transition conditions ................................................................... 11

4.2.3

Definition of interface..................................................................................................12

4.2.4

Common design standard.............................................................................................20

4.3 Power supply capacity .....................................................................................................27
4.3.1

Required power profile ................................................................................................27

4.3.2

Required size of power source .....................................................................................28

4.4 Power supply characteristics ............................................................................................31
4.4.1

Bus voltage ..................................................................................................................31

4.4.2

Voltage stability ...........................................................................................................31

4.4.3

Ripple and spike ..........................................................................................................32

4.4.4

Output impedance ........................................................................................................33

4.4.5

Transient characteristics...............................................................................................33

4.5 Power subsystem..............................................................................................................36
4.5.1

Interface with solar paddle subsystem .........................................................................36

4.5.2

Battery..........................................................................................................................36

4.5.3

Power control ...............................................................................................................40

4.5.4Power distribution ........................................................................................................43

4.5.4.1

Power distribution pattern.........................................................................................43

4.5.4.2

Operating temperature of power distribution harness...............................................44

4.5.4.3

Inrush current/outrush current ..................................................................................44

4.5.4.4

Specification of power distribution specifications....................................................44

4.5.5

Protection .....................................................................................................................45

4.5.5.1

Overview ..................................................................................................................45

4.5.5.2

Protection in power subsystem .................................................................................46

4.5.5.3

Protection in cooperation with other subsystem .......................................................47

4.5.5.4

Protection against ground equipment abnormalities .................................................50ii

JERG-2-214

5. Design verification ..................................................................................................................51
5.1 General .............................................................................................................................51
5.2 Definitions of and restrictions of goods items .................................................................55
5.2.1

Configuration and components ....................................................................................55

5.2.2

Definition of interface..................................................................................................55

5.2.3

Common design standard.............................................................................................58

5.3 Power supply capacity .....................................................................................................61
5.3.1

Required power profile ................................................................................................62

5.3.2

Required size of power source .....................................................................................62

5.4 Power supply characteristics ............................................................................................63
5.4.1

Bus voltage ..................................................................................................................64

5.4.2

Voltage stability ...........................................................................................................64

5.4.3

Ripple and spike ..........................................................................................................64

5.4.4

Output impedance ........................................................................................................64

5.4.5

Transient characteristics...............................................................................................65

5.5 Power subsystem..............................................................................................................65
5.5.1

Interface with solar paddle subsystem .........................................................................65

5.5.2

Battery..........................................................................................................................66

5.5.3

Power control ...............................................................................................................68

5.5.3.1

Unit test.....................................................................................................................69

5.5.3.2

Combination test .......................................................................................................69

5.5.3.3

System test ................................................................................................................69

5.5.4

Power distribution ........................................................................................................72

5.5.5

Protection .....................................................................................................................72

5.5.5.1

Overview ..................................................................................................................72

5.5.5.2

Protection in power subsystem .................................................................................72

5.5.5.3

Protection in cooperation with other subsystems .....................................................73

5.5.5.4

Protection for abnormality of ground equipment .....................................................73

Appendix-I

Design Overview (as reference materials) ............................................................74

I.1 General .............................................................................................................................74
I.2 System considerations......................................................................................................77
I.2.1

Orbit .............................................................................................................................77

I.2.2

Spacecraft shape ..........................................................................................................77

I.2.3

Required power profile ................................................................................................78

I.2.4

Life requirement ..........................................................................................................79

I.2.5

Design restrictions .......................................................................................................79iii

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I.3 Power supply capacity .....................................................................................................80
I.3.1

Overview......................................................................................................................80

I.3.2

Power budget analysis .................................................................................................81

I.3.3

Related tasks in each development stage .....................................................................84

I.4 Power supply characteristics ............................................................................................86
I.4.1

Bus voltage ..................................................................................................................86

I.4.2

Regulated bus/non-regulated bus .................................................................................87

I.4.3

Relation with electromagnetic compatibility ...............................................................88

I.5 Major components ...........................................................................................................90
I.5.1

Battery..........................................................................................................................90

I.5.3

Power control ...............................................................................................................97

I.5.3

Power distribution ......................................................................................................107

I.5.4

Protection ................................................................................................................... 110

I.5.5

Others......................................................................................................................... 112

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JERG-2-214

1. Scope of application
This Design Standard was established for the purpose of being applied to the powersubsystem design of spacecraft developed by the Japan Aerospace Exploration Agency(JAXA). The contents of this Design Standard specify the major requirements to be applied
to the design and development of a spacecraft power subsystem.

2. Applicable documents
JERG-2-200

Electrical design standard

JERG-2-211

Electrostatic charge and discharge design standard

JERG-2-212

Wire derating design standard

JEGR-2-213

Insulation design standard

JERG-2-215

Solar Paddle Subsystem Design Standard

JMR-001

System safety design

JMR-002

Rocket payload safety standard

3. Terms, definitions and abbreviations
3.1

Definitions of terms
Explanations and definitions of the terms used in this Design Standard are given below.

(1)

Analog TypeIs a type of power control by which the control elements are used in their linear region.

(2)

Albedo (of the earth)
Is the intensity ratio of the light reflected from the earth to the incident light from thesun and varies greatly depending on various factors. Cannot be ignored for a low
orbiting spacecraft as a heat input, etc.

(3)

Regulated Bus
Is a type of bus in which the bus voltage supplied to the load is kept nearly constant by
the power subsystem.

(4)

Umbilical Cable
Is a cable to electrically connect the rocket and AGE (aerospace ground equipment) andhas the function of being disconnected at a prescribed time during launch of the rocket.
In addition to supplying power from a ground power source to the rocket or spacecraft,it transmits various signals necessary for launch operations.

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JERG-2-214(5)

Energy Density (Specific Energy)
One of the performance indices of a power subsystem (power controller, solar array
wing/cell, battery/cell, DC/DC converter, etc.) indicating the stored energy per unit
mass or volume. Its unit is Wh/kg or Wh/liter, sometimes differently used as specificenergy for the former and as energy density for the latter. For a battery/cell, it refers to
the discharge energy in actual operation and is sometimes multiplied by depth ofdischarge to show usable specific energy or usable energy density.

(6)

Ground (Umbilical) PowerIs ground power to supply power to the spacecraft through an umbilical connector until
just before launch of the rocket. It generally supplies bus equipment power and battery
charging power to the spacecraft.

(7)

Pyrotechnics
Refers to an explosive processed in accordance with the intended use; in this Standard,
it refers to such a device that has a built-in bridge wire and is detonated by an electricalsignal {electro-explosive device (EED)}.

(8)

Over-charge
Refers to charging further after fully charged. A sealed battery has an over-charge
resistance in the specified range of current, temperature, etc. An excessive over-charge
exceeding the range lowers the life characteristics.

(9)Over-discharge
Refers to discharging the battery beyond the depth of discharge allowed by the design.
A further discharge can incur the risk of causing a reversal charge.

(10) Conditioning
Refers to the operations to bring the battery to a usable state after it is stored or unused(not charged) for a long time by charging and discharging it in the specified way.
(11) Solar Array Simulator
Is AGE that simulates the voltage-current output characteristics of a solar array as to its
BOL, EOL, etc. In a ground test it supplies power to the spacecraft instead of the panel.
(12) Reversal Charge
If the battery consists of cells of uneven capacity, the cells with small remaining stored
charge reverse their polarity by continuing to discharge even in a state where the battery
sustains the discharge voltage.

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The reversal charge reaction can incur the risk of causing an irreversible change, gas
generation, etc. in such a cell.
(13) Cant
Refers to a wing shape where the wing shaft is not contained in an extension of thepanel plain.
(14) Light Load Mode
A mode to force the on-board equipment load power to a minimum if it is difficult to
supply power from the power subsystem.
(15) Nominal capacity (rated capacity or name plate capacity is also used)
The nominal value used to show the battery capacity. Generally, about 80 to 90% of theinitial capacity under the specified temperature and charge and discharge conditions isused. It is used as a reference for setting the depth of discharge or charge and discharge
rate and is also called rated capacity. Its unit is Ah.
(16) Cold Launch
A launch with each piece of spacecraft equipment not powered up during the flight of
the rocket.
(17) Sequentially Control TypeA type of control in which the control elements, etc. are controlled in a sequence. In thistype of control, for example, the solar array is divided into multiple array circuits, eachof which is equipped with a shunt and controlled in a sequence.
(19) Self discharge
Refers to a decrease in the remaining charge of the battery without current being takenout by an external circuit. The decrease rate is generally larger at higher temperatures.
(20) Shunt (regulator) type
A type of control in which the output current of the solar array is split (and shunted) tothe control elements and the shunt currents are controlled so that the power bus voltage
takes the prescribed value. The solar array output end and a load are directly connectedby a power supply line (Direct Energy Transfer) without passing through a control
element.
This type of control is used to treat the surplus power of the power source, so it is notcommonly used for power regulators of commercial power. A simple constant-voltage(Zener) diode circuit to obtain a reference voltage falls under this type of control.
(21) Shadow diodeIs a diode connected to an array circuit of a potentially shadowed solar array with the3

JERG-2-214

reverse polarity in the series direction of the cells and in parallel with that array circuit.
It lets the current of the array circuit flow while bypassing shadowed solar cells and, in
addition, suppresses the heat generation, etc. of the shadowed array circuit.
(22) Charge/discharge rate
Refers to indicating a charge or discharge current using hour rate and nominal capacity.
The t-hour rate is equal to the current value obtained by dividing the nominal capacity
by t and is expressed as C/t using the nominal capacity C.
(23) Series (regulator) type
A type of control in which a control element is inserted in series in the power supply
line from the solar array to the load side to control the difference between the solar
array end output voltage and the prescribed power bus voltage.
This type of control is widely used for commercial power. Because a control element ofan analog type will adjust the voltage drop, it is sometimes called a dropper.
(24) Switching control typeIs a type of power control in which the control element is used in the ON/OFF region.
In contrast to the analog type, it is sometimes called a digital type.
(25) Skew
A solar array wing shape that does not allow the pitch axis of the spacecraft to coincidewith the rotation axis of the solar array wing.
(26) Sneak circuit analysis
Techniques to ensure the safety and reliability of the system by identifying andremoving a potential state (called a sneak circuit) bringing about a current flow, statuschange, etc. which causes the occurrence of an unintended function, suppression of adesirable function, or the like. A sneak circuit is apt to occur when turning on or off orchanging over the power, return, etc. or changing equipment connections, or connecting
or disconnecting a connector, etc.
(27) Thruster plume
A flow of hot gas particles ejected from the thruster. Becomes a contaminant for a solar
array, thermal bracket, optical sensor and other spacecraft part exposed outside.
(28) Battery cell
Individual cells making up a battery. Generally, it is called simply a cell.
(29) Solar simulator
An artificial light source to simulate and generate the irradiation light intensity andspectral radiant distribution of AM0 (air mass zero) sunlight. Also called an artificial
solar radiation source or simulated sunlight source.

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JERG-2-214(30) Solar constantIs the solar radiation energy received during a unit length of time by a plane of unit area
perpendicular to the solar rays at the top of the earth’s atmosphere (at an altitude of 80
km or more from the earth’s surface) when the earth is at the mean distance from thesun and is called AM0. For details, refer to JERG-2-141 Space Environmental Standard.
The sunlight in the neighborhood of the earth exhibits seasonal variations of ±3%(approximately +: winter solstice / -: summer solstice) due to earth-sun distancevariations. For this reason, the solar array output generally becomes minimum aroundthe summer solstice at EOL. Therefore, the maximum and minimum sunlight intensities
are (1366±10W/m2) × (seasonal variation due to sun-earth distance change) including
the uncertainty in the solar constant. Assuming that on the earth it is 1, the amount ofsolar radiation on other planets is as follows: 6.7 on Mercury, 1.9 on Venus, 0.43 on
Mars, 0.04 on Jupiter, and about 0.01 on Saturn.
(31) solar array wing (paddle)
Is an inclusive term for a solar battery and a structural body supporting it and it has ashape where it juts out of a spacecraft like the wing of a boat, hence the name.
(32) Solar array
Inclusive name for the power generation circuits of a solar array wing.
(33) Majority vote redundancy
One of the redundant configurations. Is employed in the case where normal output
cannot be judged with simple parallel redundancy. In the power subsystem, it is used in
the voltage detection section of voltage control and generally has a configuration where
the output of three detection sections is judged by a majority circuit. In this case, afailure of one detection section is allowed.
Another example is on-board computers, which have a majority configuration of threeunits, five units, etc. depending of the allowed number of failures. Such a redundantconfiguration needs to provide a means to ensure in advance that each individualfunction is normal. The function of making a majority judgment of multiple inputs canbecome critical if it fails.
In a power subsystem, it is used mainly in an error amplifier circuit of bus voltage
control.
(34) Stored energy
Refers to the energy a battery or a battery cell can deliver at a specified or higher
discharge voltage starting from a fully charged state. Its unit is Wh and it is calculated
as the product of mean discharge voltage and capacity.
(35) Charge array
A type of charge control in which the charge current is set to a prescribed value using
the current (limited) region of the voltage-current characteristics. It is often employed ina geostationary orbit spacecraft using a relatively low charge current.

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JERG-2-214(36) Dead band
A non-operating region of control, set between operating regions of charge anddischarge control or solar array output control to prevent operations from interferingwith each other harmfully. A similar term “dead zone” is used in general control theory
and it refers to one of the nonlinear characteristics of a control element. Incidentally, it
is a synonym for dead zone from a comprehensive viewpoint of power control in whichcharge and discharge control and panel output control are considered as control
elements.
(37) DC/DC converter
A circuit or device which makes a control element operate on the input voltage of DC
power on the switching mode to convert it into a different DC output voltage. The DC
power is converted inside by switching into an alternating current, which is converted
again into a direct current and delivered. Multiple output voltages are used in manycases, so the converter is indispensable in a spacecraft for the function of obtaining
output voltages suitable for the operation of IC, operational amplifier, TWT and otherelectronic devices from an input voltage of several ten volts or more. There is theisolated type, in which the output is isolated from the input, and the non-isolated type.
Generally, the isolated type is required for equipment other than a power subsystem
according to the specifications, etc. of the Electrical Design Standard.
(38) Battery charge method(a) Constant current charge
Generally, refers to a charge method in which charging takes place with a constantbattery charge current.
(b) Constant voltage charge/taper charge
Generally, corresponds to constant voltage charging in which charging takes place at aconstant battery voltage. Has this name because the charge current usually decreases
exponentially during a constant voltage charge.
(c) Trickle charge
Refers to charging with a very small current continuously or intermittently to make up
for the self-discharge of the battery and always keep it in a highly charged state.
(39) (Primary) power bus
The line to make a solar array and battery and these power sources’ power control
functions operate in a coordinated way.
The required power of a spacecraft is supplied from this bus through the power
distribution function.
(40) Transient
Is transient characteristics and refers to the voltage/current change occurring at atransition from a steady state to other modes.

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Other modes include also a transient due to an abnormality disturbance and test
disturbance.
(41) Missing number(42) Secondary cell
Refers to a cell that can be used repeatedly by recharging (rechargeable cell).
(43) Partial shunt type
A type of control in which a shunt tap is taken from the middle of the solar cell string to
reduce the number of shunted cell tiers with the aim of reducing shunt power andprocess power.
(44) Battery simulator
AGE that simulates the charge and discharge characteristics of a battery to evaluate thecharge and discharge control by a ground test, regardless of the state of charge of thebattery.
(45) Unregulated bus
A type of bus in which the bus voltage supplied to the load varies in a relatively largeway.
(46) Specific power (or specific output)
Is a performance index of a solar panel (wing) and shows the output power per unit
mass or unit area. Its unit is W/kg or W/m2. For batteries as well, the specific power per
unit mass is sometimes used as a performance index. In this case, the operating
conditions (depth of discharge and discharge time) are involved.
(48) V/T curve
For a charge without imposing excessive stress on the battery, the charge current isreduced after the battery voltage has reached a specified value. The specified voltage
value is shown by this. This is called a voltage-temperature curve because the specified
value is compensated for temperature considering the temperature characteristics of thebattery and its abbreviation is used. Also called a V/T level.
(49) Multiple power bus configuration
A configuration in which there are multiple power lines (power supply buses) and eachline supplies power to its load. Each power line does not necessarily have the same
configuration or characteristics (different bus voltages suitable for the load aresupplied).
(50) Boost converter
Is a kind of DC/DC converter and is a circuit or device to convert the input voltage intoa higher voltage. It is often used for the discharge control function of the regulated bus
type, and the non-isolated type is generally employed for efficiency and other reasons.
(51) Full charge
Refers to charging a battery to the fully charged state. The charge current value at which

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JERG-2-214

the charge proceeds rapidly is called the full charge current. Under actual operation
conditions, if a charge with the full charge current is continued to the fully charged state,
the battery experiences unnecessary stresses due to a temperature rise, etc. Therefore,
except for an evaluation test on a cell proper, the state of charge is monitored and thecharge current is reduced.
(52) Blocking diodeA diode to prevent reverse flow of current, it is also called a reverse current prevention
diode. It is attached to the output end of a solar array or battery connected in parallel.
(53) Depth of discharge
The ratio of the amount of discharge (electricity) to the nominal capacity of the battery,
expressed as a percentage. Also called DOD. SOC (state of charge, remaining stored
charge), which is the ratio of the actual state of charge to the charge complete state(fully charged state) expressed as a percentage, is often used as an index to show thestate of the battery.
(54) Hot launch
Refers to a launch with each piece of spacecraft equipment powered up during the flightof the rocket.
(55) Bonding
Refers to sustaining electrical conductivity between equipment and system structures toprevent accidents associated with a lightning discharge, static electricity, lightning
stroke, or other similar event and suppress electromagnetic interferences.
(56) Full charge state
The state of a battery being fully charged. A charge procedure or other similar
document generally gives a criterion to judge that full charge is reached.
(57) Memory effectA phenomenon in which a battery exhibits a lower discharge voltage than normal
during a discharge due to the operating conditions, hysteresis, etc. Removal of thisphenomenon can be expected by reconditioning.
(58) UVC (under voltage control) function
A function for battery over-charge prevention. It monitors the terminal voltage of abattery, battery cell or battery cell group and reduces the load power when the terminal
voltage falls under a certain value.
(59) Capacity
The amount of electricity a battery (battery cell) can deliver at a specified or higher
voltage (for example, 1.0 V/cell) starting from the charge complete state. Sometimes
called actual capacity. Its unit is Ah.

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JERG-2-214(60) Reconditioning
When the discharge voltage or remaining stored charge of a secondary battery drops by
memory effect or such phenomenon is expected, the battery is temporarily discharged
to a deep depth of discharge with a very small current in order to recover the battery
characteristics.
(61) Ripple and spike noise
Refers to periodical voltage variations superimposed on the bus voltage.
(62) Lock-up
Refers to entering the following situation: the solar array operating point, which isdetermined as the intersection of the characteristics of the solar array, charge anddischarge control, loads, etc., does not transfer to the appropriate position, and while
there is sunshine, discharge control operates regardless of whether the load power andcharge power can be supplied by the panel. The load power is adjusted (reduced) tomake a transition to a normal state or other measures are taken.

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3.2

Abbreviations and symbols[Abbreviations (alphabetical order)]• AGE

Aerospace Ground Equipment

AGE

• AM0

Air Mass Zero

Solar constant in vacuum space(without atmospheric attenuation)• BOL

Beginning of Life

Beginning of life

• DOD

Depth of Discharge

Depth of discharge• EMC

Electro Magnetic Compatibility

Electromagnetic compatibility• EOL

End of LifeEnd of life

• EPS

Electrical Power Subsystem

Electrical power system (it is asubsystem)• ESD

Electrostatic Discharge

Electrostatic discharge• FMEA

Failure Mode and Effects Analysis

Failure mode and effect analysis• GEO

Geosynchronous Earth Orbit

Geostationary orbit• ICD

Interface Control Documentation/Interface Control Drawing• LEO

Low Earth Orbit

Low earth orbit• LLM

Light Load Mode

Light load mode

• SOE

Sequence of Event

A correspondence table of events andtheir times• UVC

Under Voltage Control

Under-voltage control[Symbols]• Ah:

Unit of electrical charge It is the product of current and time (ampere-hours) and is used
as a unit of battery capacity, etc.
It is related to the coulomb, a unit of electrical charge: 1 Ah = 3600 coulombs.

• C:

Symbol to denote the charge and discharge rate using the nominal capacity.

• Wh: Unit of energy. It is the product of electrical power and time (Watt-hours) and is used asa unit of discharge energy, stored energy, etc. of a battery.
The discharge energy of a battery is the product of the amount of discharge and (mean)discharge voltage.
It is related to the joule, a unit of energy: 1 Wh = 3600 joules.

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JERG-2-214

4. Design requirements
4.1

GeneralA spacecraft power subsystem has the role of supplying electrical power necessary for the
spacecraft without interruption within the prescribed voltage characteristics throughout theentire mission period. The definitions, intrinsic properties, design restraints to be consideredwhen designing a power subsystem to implement this requirement and related
considerations are described below.

4.2 Definitions and restraints of items
4.2.1 Configuration and components
A schematic example of a spacecraft power subsystem configuration is shown in Figure
4.2-1. This document covers those parts of the instrumentation subsystem which are related
to power and those parts of a solar array wing or load equipment which are related to theinterface requirements and defines them as a power subsystem. The power subsystem
performance requirements required in this document are applied at the power output
interface point. For the solar paddle subsystem, refer to JERG-2-215.
Power generation(solar array wing)

Power control(PCU)

Power distribution(PDCU)

Power storage(BAT; energy storage)

Other power
distribution (ODC, etc.)

Primary power bus
Load equipment

Other power bus
Load equipment
Instrumentation
related to power

Coverage of thisDesign Standard

Power output
interface point

Load equipment
input interface
point

Figure 4.2-1 Schematic diagram showing example of power subsystem configuration
As shown in Figure 4.2-1, a power subsystem is made up mainly of the following functions:(1) Power control function (including surplus power control, bus voltage control andbattery charge and discharge control)(2) Power storage function(3) Power distribution function (including power supply to pyrotechnics and voltage
conversion for other power buses*Note 1)(4) Protection related to the above, i.e., failure and abnormality limiting function (this is
implemented as a function in the configuration of the three functions above)
Note 1: Other power buses refer to the power buses of voltages used in the past (28V, 50V,etc.).

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JERG-2-214

For a specific power subsystem architecture (selection of bus voltage,
regulated/non-regulated bus type, multiple bus type, etc., design of hardware and software
components, implementation techniques (circuit type), etc.) to implement the functions
above, the selection should be made in light of the requirements and restrictions at the
spacecraft mission level. As a selection example, the following may be included.

• Power control function(Solar array power generation and voltage control type)a.

Series control typeb.

Shunt control type(Digital sequential shunt type and analog partial shunt type)c.

Peak power tracking control type

(Battery discharge type)a.

Floating typeb.

Step-up or step-down converter regulation type

(Battery charge type)a.

Charge array typeb.

Charge regulator type

• Power storage function(Selection of power storage element)a.

Nickel-cadmium batteryb.

Nickel-hydrogen batteryc.

Lithium ion battery• Power distribution function
Includes the following functions:b.

Over-current protection functionb.

Multiple bus isolation and coupling functionc.

Bus changeover function in case of the backbone distribution typee.

Others10

JERG-2-214- Related considerations If the peak power tracking type is selected out of the power control function class, it is difficult
to keep track of the operating point on the solar array V-I curve. Hence, it is difficult to predict
and manage the maximum possible power generation in response to on-orbit degradation and in
response to using voltage and current telemetry of the solar array itself. For a spacecraft which
requires prediction and management of possible on-orbit solar array power generation, it isrecommended to take prediction and management measures by mounting a monitor cell or using
other means.

4.2.2

Operation mode and transition conditions

The operation modes of the selected power subsystem and their transition conditions should
be clarified, and the relationship with the operation state of components and its confirmation
method, etc. should be clarified by telemetry data and status to make clear the specificationof the characteristics, etc. described below. As a related consideration, an operation mode
example is shown below.

- Related considerations The definition of operation modes of a power subsystem should be described to make clear
the specification of the characteristics, etc. described below. An operation mode example isshown below.
• During the launch
Charge by ground equipment, change over from ground power to battery, battery
off/automatic on during cold launch, etc.
• When in a transfer orbit
Pyrotechnic ignition, wing deployment, orbit or attitude change, etc.
• When in a stationary orbit
Charge (constant current/constant voltage/voltage control/trickle charge), shunt operation,
reconditioning/battery off, discharge in a penumbra or during an eclipse, etc.
• When protection function is in operation
Load abnormality, UVC, battery temperature abnormality, etc.
When using a special term, abbreviation, etc., its definition should be shown.

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4.2.3

Definition of interface

Define the interfaces of the power subsystem with the spacecraft system, its related
subsystems, etc. and specify their requirements, restrictions, common design standard,
information about interface FMEA, etc. In the spacecraft system, these are organized into
ICD (interface control document/drawing) or other documents and managed collectively.

(1)

a.

Electrical interface

Power supply
For power supplying bus power lines (including redundant lines), specify the diagramof the interface with every piece of equipment (including the presence/absence of arelay, fuse and current limiter) and the mean and peak of supplied power to each pieceof equipment (including the causal factors, period, duration, etc. of the peak). For the
power consumption of a component of the power subsystem, separately specify it
together with similar specifications.
Of the power supply characteristics, specify the input end voltage on the load side
besides the bus voltage reference point. Besides the on-orbit supply characteristics,
specify the transient at power-on from zero volts, which cannot be avoided in a groundtest, and at power-off at an emergency, etc.
For electromagnetic compatibility and power control design analysis, the power input
impedance on the load side should be presented.

- Related considerations The specifications for power supply are design requirements for the power subsystemand power distribution, and they specify the upper limit in the design stage and aretaken as a reference to confirm that the subsequent power supply level updates in thedevelopment stage are within the upper limit. In connection with this, the supply(consumed) power serves as source data of a required power profile. As such and as anallocation item of the system, it is desirable to separately specify and manage it using atable, clearly describe the grounds for the setting (prediction/measurement, etc) in theremarks field, and update the specified value as well as review the description as thedevelopment stage proceeds.

12

JERG-2-214

For the specification of supply characteristics, pay attention to matching with theload-side interface besides the power supply bus output characteristics. For this purpose,
set this specification considering the line drop of the system harness. Separately specify
the conditions at power-on or power-off during a ground test to prevent equipment
damage, status change or other similar situation which could occur at
power-on/power-off if designed only with always on on-orbit conditions. In the designof AGE, ensure that an excessive inrush current, which however will not occur during
actual operation, is not applied to the power subsystem equipment.
b.

Telemetry/command
For telemetry, select items for which it is possible to monitor the health check offunctional operation, performance and equipment operation and the supply power flow
(power control, power storage and load power supply), and all telemetry specifications(analog/digital distinction, sampling rate, accuracy, etc.) and their content should beclarified. For a command to set a power subsystem operation, verify telemetry should
be issued by which the result of its execution (change of operation states) can be known.
For especially important items, ensure that a redundant telemetry or a cross check withother items is possible.
For a command, clarify the command specifications (discrete/serial distinction,
parameter setting accuracy, etc.) and its content (the functional operation set by thecommand). Exclude any single command that can damage the power subsystem orbring it to an unrecoverable state.
For the operation state of an automatic judgment function (over-voltage andover-current protection functions, etc.) of the power subsystem, pay attention to atelemetry output by which operation can be confirmed and a reset function triggered by
a command.
For the details of telemetry/command item setting standards and electrical interfaces,
conform to the specifications of the Signal Interface Example (JERG-2-200-TM001),
etc. Telemetry/command is a system allocation item, so it should agree with thespecifications for it.
Telemetry items whose output is recommended for the power subsystem aregiven below.

13

JERG-2-214• All function settings, operation statuses and on/off status of the power subsystem• Main bus voltage (including secondary bus voltage)• Bus voltage error amplifier output signal• Overall load supply current (should have a resolution that is sufficient to keep trackof the load behavior. If necessary, divide into multiple load supply groups and obtainthe load supply current of each. The same applies to the secondary power bus
current.)• Shunt current and shunt drive signal• Overall generated current of solar array (if the current on the RTN side is to be
detected, the structure short-circuit current can be calculated from the load supply
current, shunt current and battery charge current values.)• Battery voltage (individual battery cell voltages or unit group voltage as the needarises) and battery charge and discharge voltages• In case of a Ni-H2 battery, the internal pressure of at least one battery cell per battery
assembly
The temperature at a reference point of each component and, for a battery assembly, thetemperature at a representative point in the assembly (main redundant configuration isdesirable)c.

Control signal
For an interface related to control signals between the power subsystem and anothersubsystem, specify its functions, boundaries, electrical characteristics, sending andreceiving circuits of the signal, comprehensive signal lines and operations, etc. In some
types of design such as integrated design, an algorithm processing component
belonging to another subsystem performs control related to a power subsystem function
based on power subsystem telemetry. For these control signals as well, specify theirlines and operations. In this case, operation control takes place not by individual control
signals but by on-board commands in some types of design. However, clearly specify
those control signals as a function also in the control signal category.
The control signals include the following, for example,• LLM signal• Spacecraft separation signal• Solar array wing deployment signal

14

JERG-2-214• Battery charge voltage control (V/I control)Note 2• Battery reconditioning controlNote 2• Battery high temperature protectionNote 2• Heater on/off controlNote 2

Note 2: For the functions formerly implemented by hardware of the power subsystembut now, through integration, etc., implemented by processing of computer
software or other subsystem, make sure there are no omissions.

- Related considerations Control of related and coordinated signals of these multiple subsystems takes place atthe system level. For a design change to one subsystem, in particular, it should bepossible to judge the effect on the involved subsystems, comprehensive function
ensuring, etc. quickly and appropriately. For this purpose, the related matters should bespecified also in the specification, etc. of the involved subsystems from acomprehensive point of view.
For the simulated signals needed when testing these control signals, consider theirgeneration method, etc. in advance from the design stage.

d.

Pyrotechnic function

Specify the all power lines (including redundant lines) from the ignition power supply
source to the pyrotechnics including the boundary to the power subsystem andrelationship with the spacecraft grounding.
Specify the major specifications of the power source, pyrotechnics, relays, anti-static
resistances, etc., the battery and ODC input voltages, ignition current and supply
duration, and the sequence of operations up to ignition.

e.

Grounding

The return of the power supply bus line should be in principle grounded to thespacecraft structure at a single point. For the details of the requirements for grounding,conform to the Electrical Design Standard (JERG-2-200).

15JERG-2-214- Related considerations Grounding should be designed to exclude a place of indeterminate potential withrespect to the spacecraft structure and not to let a stray current flow to the spacecraft
structure. It is desirable to place the single grounding point so that it is possible toverify that there is no return current. Defective grounding can cause a malfunction, etc.
due to electrostatic charge (electrostatic discharge in the worst case). The effects ofelectrostatic charge and discharge are sometimes examined by an ESD test as a part ofan EMC test.
f.

AGE
For all items of the spacecraft power subsystem with which the AGE has an interface,specify the diagram, content, specifications (voltage and current ranges, signal pattern,
accuracy, etc.) of each component, separately for umbilical cable and for test cable. For
the return line of each item, show its relationship with the grounding of AGE and withthe spacecraft structure grounding in a systematic way. In addition, show their status
and pattern of use during each ground support test and the launch and their treatment
while they are not used.
For AGE involved in power supply, provide a treatment to prevent an excessive inrush
current, excessive voltage application due to failure or an operational mistake, reverse
flow of battery power, etc. and use a design that does not damage the spacecraft. If
voltage remote sensing is to take place in supply voltage control, a fail-safe design isnecessary in case an abnormality occurs in that connection. If on-board equipmentchanges its status, etc. at power-on or power-off, it should not be dealt with on the
ground side or manual operations but its cause should be searched for and the result
should be reflected in equipment design.
- Related considerations For the interface with AGE, consider not only the items necessary for power supply andpower subsystem state setting (for example, battery on/off) during a test or launch but
also the items to make it possible to ensure the comprehensive performance of thepower subsystem. It is desirable to clearly describe the relationship of each item withthe functional configuration of the power subsystem. For battery-on/off, the design ismade so that the state of all batteries can be set by one command because of a
restriction of umbilical cable. Pay attention to ensure that internal failure of thespacecraft receiver of that command will not become critical (battery fixed to the off
state, etc.).

16

JERG-2-214

For grounding, fully know the power lines of test equipment and range equipment andtake measures for measured data error prevention, clarification of the power supply
return, electrical shock prevention, lightning protection, etc. For power supply in therange (at the launch point in particular), consider the effects of umbilical line
impedance.
For a short-circuit or other failure on the AGE side (including its cable), a protective
resistor or other consideration is made for the signal line. However, this is difficult on
the power supply line, so it is necessary to deal with it on the ground side. In this case,avoid dealing with it by using an operation procedure but reflect it in the device design.
If a diode is used for reverse current prevention, etc., pay due attention to its heatgeneration.
g.

Inside the power subsystem
Referring to the configuration diagram of the power system, specify all interface items
between components, including the intervening harnesses and slip rings, and thecorrespondence between shunt operations and arrays through a slip ring (prescribedshunt sequence and shunt operations corresponding to shunt drive elements), etc.
Specify detailed specifications in a management document, etc.
- Related considerations For the solar array and battery, clearly describe the reference points of output power, etc.
and their voltage and current ranges. In solar array output control, etc., the control
element and control signal processor belong to different components in many cases. In
such cases, it is necessary to specify that interface from the aspect of comprehensive
operations of the control subsystem. As for the items involved in the connection, etc. ofthe solar array and spacecraft, there are items that are difficult to verify by a ground test.
For design error prevention, therefore, it is necessary to clarify wire connections and, inparticular, the polarity and relationship between return and ground, and the like.

(2) Mechanical interface
For a power subsystem component, specify its mounting position on the spacecraft,including the coordinate axes.
For the mechanical interface between each component and spacecraft structure, show
its outer dimensions, the dimensions and reference hole positions of the attachment
portion, mass, center-of-mass position, materials used and surface treatment, roughness

17

JERG-2-214

and flatness of the mounting surface, and mounting method (bolts, tightening torque,etc.).
Identify the connectors of each component and clarify their compatibility with theircounterpart of the spacecraft system. The mass and dimensions of a component are
specified along with its physical properties, and this specification must agree with thesystem allocation.
(3) Thermal interface
For each component, specify the heat capacity, amounts of heat absorption andgeneration, radiating surface absorptance, amount of heat flux and thermalmathematical model (as necessary), and heater power. As for the amounts of heatabsorption and generation, amount of heat flux and heater power, specify them for each
operation state.
Show the allowable temperature range and allowable temperature change rate of eachcomponent (during storage and transportation, on-orbit operation, at power-on/off, andin non-operating state). For a thermally critical component (battery and the like),
specially mention the reasons and requirements for it and show the specifications forthe measure to deal with it (active temperature control, etc.).
Clarify the temperature reference point that is basic to the interface. Obtain thetemperature at the reference point through telemetry.
For the battery charge operation as a range operation with the spacecraft mounted, pay
attention to the possibility of taking an interface which can ensure a sufficient amountof charge for the heat rejection environment during the charge.
In particular, a Ni-H2 battery generates a large amount of heat at the end of charge anddue to self-discharge. To sustain a sufficient charge, therefore, it is desirable to keep theambient temperature environment as low as possible. In some cases, however, it isdifficult to keep a sufficient temperature environment for charging a Ni-H2 battery in a
mounted state with the air conditioning capacity of the range facility including thefairing air conditioning. Therefore, pay due attention to thermal interface design forrange charge from the spacecraft system design stage.

(4) Functional interface
Enumerate the items that functionally interface (coordinate/interfere, etc.) with anothersubsystem and, if the functional requirement is a requirement from the powersubsystem, specify its content and specifications.

18

JERG-2-214- Related considerations As a functional interface, the following items, for example, may be cited.

a. Matters with the attitude control subsystem related to a pyrotechnic ignition signal
or abnormality
b. Matters related to the supply of bus voltage and (sensor, heater, etc.) power otherthan pyrotechnic ignition power.
c. Matters related to the spacecraft separation switch
d. Battery charge voltage control (V/T control)Note 2
e. Battery reconditioning controlNote 2
f. Battery high temperature protectionNote 2
g. Heater on/off controlNote 2

Note 2: For the functions formerly implemented by hardware of the power subsystembut now, through integration, etc., implemented by processing of computer
software or other subsystem, make sure there are no omissions.

(5) Work interface

For each of the components, specify the item name, fabrication number, serial number
and other identification for configuration management. As shipment preparation,
designate the delivery configuration and specify the history management,
handling-related documents and transport and packing method.
In addition, specify the special work in spacecraft integration or a test, support content
related to legal procedures, etc. of high-pressure gas, etc. and the points to consider
with regards to work safety.

- Related considerations For a heavy component, it is necessary to make a design considering not only the actual
on-orbit operation state but also the workability on the ground from the viewpoint ofweight and dimensions.

19

JERG-2-214

4.2.4

Common design standard

In the design of a component of the power subsystem, the following common standard
should be complied with.
(1)

Electrical design
The power subsystem and each component should be designed to implement therequirements for the electrical interfaces and performance as simply and as efficiently
as possible.
The power supply and telemetry/command interfaces should conform to the electricalinterface conditions. Design the interface so as to prevent failure propagation.
In response to the presentation of a power input impedance from the load side, verify
the electromagnetic compatibility and power control feasibility as the power subsystem.
Set the connector pin arrangement considering the electromagnetic compatibility,
redundancy, etc. The use of a fuse and the management of magnetic characteristics
should be done according to the Electrical Design Standard (JERG-2-200).
- Related considerations When a fuse has to be used because short-circuit protection measures other than a fuse
are not practical as a short-circuit protection means (for example, short-circuit
protection measures for an electrolytic capacitor, power transistor, or other similar
devices and for a power processing unit), requirements in principle are appended as forthe conformance to the criteria set about fuse type, derating, redundancy, etc. and as fora means to judge from outside the equipment whether the installed fuse is good or bad.
As for magnetic management, etc., pay attention to the following in the design of apower subsystem.
a. The power line should be of a twisted pair construction.
b. Ground the return of the power supply bus at a single point in principle (do not use
the structure for a return). If, however, a structural body is used as a power return
line, do it according to the requirements of the Electrical Design Standard(JERG-2-200) and lay wire along the spacecraft structure to minimize the current
loop and harness inductance.
c. For devices that use magnetism (transformer, inductor, etc.), select a construction,etc. of low flux leakage.

20

JERG-2-214(2)Mechanical design
Design the construction of each component based on implementing high-density
packing and aiming at compactness and lightness. Considering the prescribed
environmental conditions (including not only the mechanical environment but also
thermal stress, etc.) and the characteristic change or degradation of the materials used
due to this, make a design to satisfy the strength requirements and rigidity requirements
for the prescribed load conditions. For a pressure vessel, design it to satisfy theprescribed proof pressure and burst pressure requirements.
The attaching method, etc. should conform to the mechanical interface. For large heavyequipment or a panel or other similar equipment for which an interference is expectedwith the system structural characteristics or attitude control subsystem, present a
structural model and perform a coupling analysis, etc. with the system.
As an example, a Ni-H2 battery is sometimes designed in such a configuration that thebattery cells are arranged individually using a battery cell bracket (sleeve) from thepoint of view of ensuring a heat path for each battery cell. In this case, the vibration
mode becomes complex and the stress on the harness between battery cells becomeslarge. If needed as a result of a coupling analysis with the system, ensure there issufficient overall rigidity as a battery assembly by having a mechanical design such as
mutually connecting the sleeves or otherwise reinforcing them.

(3)

Thermal design
Design each component while ensuring that the temperature of a device, etc. used falls
within the specified allowable range and minimizing the heat path for a spot of highheat generation in particular.
And, carry out design while considering the characteristic change and degradation of
the materials, etc. used due to the prescribed environmental conditions and heat.
Pay attention to prevent the thermal conductivity, etc. from declining due to themechanical environment.
The attaching method, surface roughness, surface treatment, etc. related to thermal
design should conform to the mechanical interface.
For a component with a large temperature gradient across its attaching surface, amodularized power subsystem, a panel, or other equipment exposed to the outer space

21

JERG-2-214

environment, present a thermal model and perform a coupling analysis, etc. with thesystem.
The harness bundles including power wires should satisfy the specifications of
JERG-2-212 Wire Derating Design Standard.
- Related considerations In the thermal design, heaters are used as a means of active temperature control;
however, a thermal design involving much heater power even at EOL is not desirable inview of power budget. Ensure that the maximum generated current value at BOL will
not deviate the shunt element’s allowable temperatures. From the viewpoint of heatdissipation surface area design, consider the change in shunt current value and thermal
control material characteristics from BOL to EOL and consider a combination of design
conditions so as not to overdesign.

(4)Electromagnetic compatibility design
Design each component to suppress electromagnetic noise generation and keep it as low
as possible and not be influenced by external electromagnetic noise as much as possible.
For static electricity prevention, etc., provide the heat control thermal blanket, etc.
besides the equipment chassis with bonding according to JERG-2-211. If an insulating
material is used for surface treatment, obtain the agreement of the system side.
The requirements of electromagnetic compatibility are specified concerning thefollowing items:
a. Electrical bonding and insulation resistanceb. Conducted noise (in a steady state or during a transient)c. Conducted susceptibility (ditto)
d. Radiation noise (electric and magnetic fields)
e. Radiated susceptibility (ditto)
The test method is specified in the electromagnetic compatibility document. As acondition, select an operation mode which forms the worst case when conducting thetest. (As an example, employ the upper limit of bus voltage if voltage is intended andthe lower limit of bus voltage if current is intended. Conduct the test also in operation
modes which generate periodical noise or transient noise.)22

JERG-2-214

In a test on a component on the receiving side of a power supply line, use thedesignated power supply side impedance model.
- Related considerations The power supply characteristics of the power subsystem are parameters which
determine the limit level of an electromagnetically compatible power line’s conductedsusceptibility. On the other hand, the requirements for a power line’s conductedinterference noise are parameters which determine the power supply characteristics ofthe power subsystem. Therefore, the design of the power subsystem must proceed whileconsidering the electromagnetic compatibility requirements. In a generalelectromagnetic compatibility design, a specification related to power supply voltage
transients sometimes requires an application, etc. of a voltage twice as high as thesteady-state voltage. However, this should be specified considering the causes, etc. forthe occurrence. The applied level in the test should be consistent with the electricalinterface conditions including the allowable setting tolerance of test conditions.
When using a power supply side impedance model, the power supply capacity of thetransient inrush current, etc. at power-on are influenced by the power supply
capacitance connected to that input side. In the case of the power subsystem, it is also
necessary to verify that the bus line transient voltage occurring at a load variation orload failure is kept within the applied voltage level in the steady-state disturbance andabnormality disturbance requirements.
(5)

Environmental conditions
For the environment of each component, determine the following items in each stage
from testing, storage, transportation, launch to orbit insertion, and on-orbit operation.
These are specified as a reference for environmental resistance design.
a. Temperature (should conform to the thermal interface, except for on ground)b. Pressure (from ground surface to high vacuum state in outer space)c. Humidity (On ground: Generally 60% or more in relative humidity)d. Cleanliness (On ground: Generally ISO Class 8 (Class 100 thousand) or higher)e. Acoustics (applied to exposed area)f. Acceleration (including the steady acceleration component and low-frequency
vibration acceleration component)
g. Vibration (sinusoidal and random: level depends on the mounting position.)h. Impact (impact associated mainly with spacecraft separation and pyrotechnic
ignition: level depends on the mounting position)

23

JERG-2-214

i. Radiations (from launch to on-orbit operation)
j. Exposed area environment (excluding acoustics)
For an object exposed outside the spacecraft, an environment under the influence ofthermal radiation from the fairing, free molecular flow heating, thermal radiation andjet heating from the rocket, etc., atomic oxygen, meteoroids/debris, solar radiation,
terrestrial radiation, and plasma and thruster plume gas around the orbit is specified.
The environment of each component should be determined concerning the following
items for each stage from testing, storage, transportation, launch to orbit insertion andon-orbit operation and these should be specified as a reference for environmental
resistance design.
- Related considerations The environmental conditions differ for qualification/acceptance tests and are set with adesign margin included. For example, temperature is set as expected maximum
temperature ±5°C and vibration or impact is set as expected maximum level plus 3
decibels. However, it is necessary to set them appropriately considering the margin of
the expected condition, design impact on the equipment side, avoidance of unnecessarystress application on the flight model (e.g., on the battery, etc.), and others.
(6)

Safety design
When developing a power subsystem (fabrication, testing, transportation, unpacking,range work, etc.), analyze the level (catastrophic, serious, critical) of risk items(hazardous substance, explosion, pointed object, heavy object, heat generation, UV ray,electric injury, static electricity, corrosion, rupture, short-circuit, etc.) and their
possibility of occurrence and take effective measures (change of materials, etc. used,thoroughgoing danger marking, complete provision of a work environment, fixtures,etc., installation of protective and alarm devices, preparation of emergency procedures,and others) to ensure the safety of the personnel and hardware.

(7)

Reliability design
Each component of the power subsystem should be designed to ensure that the powersubsystem will satisfy the prescribed performance requirements throughout the life
period with the reliability specified by the system allocation. By FMEA associated with
the inside and interface, analyze the failure mode, propagation effects and criticality andmake a design to ensure that a failure of a single part in particular will not become24

JERG-2-214

critical including its propagation. Examine the related troubles that occurred in previous
developments or preceding research to know their causes and take recurrence
prevention measures.
A failure of the power subsystem influences normal power supply. Because it is difficult
to ensure redundancy in the subsystem or equipment, a redundant design (majority
decision, series parallel, etc.) of the circuit or device level is employed. In this case,consider a means to verify that the redundancy was not lost during fabrication or
testing.
In the design, perform a worst case analysis considering the characteristic variance of adevice or the like, its (electrical and thermal) operation conditions, changes due toenvironmental conditions (vibration, impact, radiation, etc.) malfunctions due to contact
chattering, etc. of a relay or slip ring, redundancy-free state, and others. Keep thenumber of adjusting points in the process to the minimum necessary in order to ensurefabrication reproducibility and prevent deviations from the design.
Damage to a device, etc. due to radiation includes that due to accumulation of exposure
dose (total dose effect) and that caused by the incidence of one heavy ion particle(single event effect). In the power subsystem, attention is paid mainly to the total dosesuch as characteristic changes of semiconductor devices. In the signal processor, etc.,
however, it is also necessary to examine the resistance to single events
The setting of the worst case should be consistent with the reliability prediction model.
Perform a sneak circuit analysis to prevent an unintended current flow, status change,
etc. from occurring at power-on or -off, changeover, and connector connection or
disconnection concerning the power supply line and return from power supply (inparticular, multiple) buses, ground power, etc.
Clarify the selection criteria of the devices, materials, and processes used. When
applying a device or material, perform an analysis on the electrically, thermally,mechanically, radiation-, high vacuum- or otherwise induced stresses in it. For a device,in addition, do appropriate derating concerning power, voltage, current, and temperatureaccording to the applicable standards. Employ a proven design, device and material fora part that is not in conflict with the development objective. If the device used, material
used, or the like necessary for the design lacks its application data, conduct aqualification test, etc. separately.
Select measurement data so that the normal operation of the power subsystem can be
known by a test, etc. and set an evaluation criterion for the trend, etc. of the data.
Consider redundancy also in data acquisition and make a design considering the casewhere data measurement becomes partly impossible and also a substitution with otherdata in such a case.
Evaluate the accumulated fatigue of a flight item by ground tests to confirm that the25

JERG-2-214

fatigue does not influence the durability necessary when in actual use. For a battery,
relay, connector or other limited life item subject to wear and aging degradation and an
item with a restricted number of times of operation, set and manage restrictions on thetime and the number of times of operation to the launch. For a battery, specify its
storage method before being mounted and maintenance method after mounted.
- Related considerations In a redundant design of the power subsystem, standby and changeover is not allowed
in most parts. Therefore, even if a trouble occurs in a part of the redundant line on the
ground, its confirmation is difficult in many cases. As a measure and at the equipment
proper level, a signal to force a part of redundancy to become inoperative is applied
from outside and a means to verify that the remaining part is normal is taken. Or, a test
connector is used to measure characteristics (voltage drop of a diode, capacitance of acapacitor, etc.). Yet another method is to provide a signal to monitor the operation ofeach individual part as an output and monitor those signals. However, these methods
are generally difficult to conduct at the system test level. As an alternative means, trend
assessment of measured data is utilized. In trend assessment, it is important to verify
not only that the data lie within the specification but also that the data havereproducibility and change as predicted. So, pay attention to subtle data variations as
well. A relevant item is telemetry data. Besides this, for important items, provide a test
connector output and pay attention to the steady-state value and transient waveform
although under the restraints of the test configuration.
(8)

Maintainability design
In the period from fabrication and installation in a spacecraft to launch, there may be acase where a component requires inspection, repair, replacement, etc. Allow for such anoccasion, that is, make a design to ensure that the work can be done easily and has little
influence on other pieces of equipment or systems.
- Related considerations Concerning the power subsystem, battery replacement must be considered. A flightbattery is a limited life article and its maintenance after being mounted on a spacecraft
is troublesome. So, it is desirable to mount it just before the launch within the realm of
possibility and store it separately until then. Therefore, in spacecraft structural design,
work schedule setting, fixture preparation, etc., be conscious of item replacement. If the26

JERG-2-214

power distribution section has an upstream fuse for short-circuit fault isolation, pay
attention to the connector accessibility to confirm fuse integrity.
4.3

Power supply capacity
The power subsystem should have a power supply capacity that satisfies the required power
profile throughout the entire mission period of the spacecraft. The following specifications
are an item of basic performance of the spacecraft system and should be consistent with theapplicable requirements of a specification document, etc.

4.3.1

Required power profile

Specify the require power profile throughout the entire mission period as follows: In the caseof multiple buses, specify for each bus.
(1)

Define the reference point of supply power to a load.

(2)

Define all operation modes from the spacecraft launch to the end of mission (including
contingencies, etc.)(3)

For each power-allocated load item (irrespective of its bus voltage but including a load
which operates continuously), calculate the power consumption in each of its operation
modes. In the power consumption, include the load power and the loss by the harness
from the supply power reference point.

(4)

Calculate the supply power in each operation mode.

(5)

Along the operation plan from the spacecraft launch to the end of mission, show therequired power change with time (required power profile) at every constant period (forexample, from launch to orbit insertion and then at every revolution in orbit).
- Related considerations For clauses (2) to (4), it is desirable to organize items as shown in Table 4.3-1.
The supply (consumed) power serving as the source data for the required power profile
is generally updated as the development stage progresses. Therefore, the required power
profile is usually specified as the upper limit of the level required in a certain period of
the development stage. It is desirable to separately track the update history using a table,clearly describe the grounds for the update in the remarks field, and review thedescriptions together with the specified value update with the progress of thedevelopment stage. However, the acceptance or rejection of an update is examined27

JERG-2-214

based on a power budget analysis, etc. performed at an appropriate time.
For the supply (consumed) power in a required power profile, the transient peak poweris generally ignored. For example, the pyrotechnic ignition power is ignored. In general,an upper limit (including a margin) is specified, and its positioning (development target
value/expected worst value of a test model/measured worst value) and setting margindepend on the management policy of system development and the progress of thedevelopment stage.
For the operation modes associated with the required power profile, consider thepredictable worst case from the viewpoint of power budget.
Table 4.3-1 Example of supply power table by operation mode
(Unit: W)
Operation mode

In the sunshine
5

6

7

8

60

60

60

60

60

125

125

125

125

125

125

25

25

25

25

25

25

25

77

77

77

77

77

48

48

48

Thermal control subsystem

184

184

184

184

184

114

114

114

Total power consumption of bus
lines

471

471

471

471

471

372

372

372

Mission equipment A

429

132

0

0

124

429

0

124

Mission equipment B

200

300

300

100

300

200

300

300

Mission equipment C

50

0

0

100

200

50

0

200

Mission equipment D

50

0

120

120

100

50

120

100

Total power consumption of mission
equipment

729

432

420

320

724

729

420

724

Total power consumption of bus andmission

1200

903

891

791 1195 1101

792 1096

Harness loss (2%)

24

18

18

16

16

Total supply power

1224

921

909

807 1219 1123

4.3.2

1

2

3

TT&C subsystem

60

60

60

Attitude and orbit control
subsystem

125

125

Power and paddle subsystem

25

Secondary propulsion
subsystem

During shade

4

24

22

22

808 1118

Required size of power source

The required size of a solar array and battery is specified according to the characteristics in
Sections 4.5.1 and 4.5.2. The characteristics of this power source should conform to therequired power profiles of all operation modes and the depth of discharge should be kept
within the allowable value for the battery. Judge conformity with the required power profile

28

JERG-2-214

by power budget analysis and present the specifications of that power budget analysis model.
- Related considerations In the power budget analysis, establish and maintain a power budget based on the peak
power value and an energy budget based on the mean power value. Take the following items
into consideration.
Distance from spacecraft to sun
Duration of sunshine and eclipse• Aspect angle of sun• Orientation accuracy• Effects of environmental temperature and degradation• Matters associated with reliability and safety (effects of failure)
Based on the final power budget analysis in the design stage performed on therequired power profiles of all operation modes, the required size of the solar array andbattery is evaluated as follows:• Required size of solar array
Show the required power generation profile during the entire mission period.
• Required size of battery
Specify the required discharge energy profile during the entire mission period. Or, set thestored energy of the battery and show the required depth-of-discharge profile of thatbattery.
Applying an appropriate margin for power budget analysis to these values, thecharacteristic requirements for a solar array and battery are determined and they become
specifications for the required size from then on. For the appropriate margin, each project
should determine an appropriate margin in each design phase. Margin examples include
the following:29

JERG-2-214

Design phase
PDR
CDR
PSR

Spacecraft configuration
New bus design and new payload/mission design

15%

10%

6%

In the case where large-scale design changes are made to an
existing bus or payload design

10%

8%

6%

5%

8%

4%

5%

5%

3%

In the case where minor changes are made to an existing bus or
payload design
Existing bus design and existing payload/mission design

As the available power, supply of the required power should be possible even when one
solar array circuit or one battery cell fails at the end of the specified life of the powersubsystem.
Incidentally, once characteristic requirements for solar array and battery hardware are
specified, there is no scope for making a change after that. In the stage of determining
these specifications, it is common practice to proceed with a solar array and battery
design simultaneously with a power budget analysis for the required power profile.
On this occasion, pay attention to the following:a. The allowable depth of discharge of a battery is generally specified for the maximum
level in the stationary discharge energy profile. If a large transient depth of discharge
is expected, identify that operation mode and specify an allowable value separately.
b. If a battery discharge may occur in the sunshine, it is desirable that a battery charge
completes before an eclipse begins.
c. It is desirable that the battery charge power can be supplied at the upper-limit level ofcharge current set by charge control. For a LEO spacecraft, there may be a case,depending on operational requirements for the mission equipment, where it isdifficult to ensure the upper-limit charge current. In this case as well, it is desirable
that a battery charge completes before an eclipse begins.
d. If it is often the case that a charge is not complete in the sunshine, the operation plan
ideally should ensure that the depth of discharge does not become large during the
succeeding eclipse. Conformity with the required power profile and power source
size is evaluated by power budge analysis at an appropriate time as the development
stage proceeds. Present the specifications for the power budget analysis model (seethe example of Appendix I) used for these analyses together with the required power

30

JERG-2-214

source size specification and, in the same way as the data associated with therequired power profile, use the values updated with the progress of the development
stage. The required solar array and battery sizes specified here are the lower limit
values that satisfy the requirements, and the margin, redundancy, etc. on the
hardware side are considered in each design.

4.4

Power supply characteristics
The power supply characteristics of the power subsystem should be consistent with theelectrical interface conditions related to the spacecraft. Here, only the characteristics of thebus voltage generated by the power subsystem itself are considered. Take the reference pointat the power distribution output end (power subsystem output end). Unless otherwise
specified, use all the spacecraft load conditions and environmental conditions as specified
conditions.

4.4.1

Bus voltage

Show the steady-state voltage range for bus voltage. In the case of the regulated bus type, it
is desirable to select from among 28 V, 50 V and 100 V. For a selection guide, refer to theElectrical Design Standard (JERG-2-200). In the case of the non-regulated bus type,appropriately divide and show the range considering the operation from the launch tostationary orbit insertion and the operation modes while there is sunshine/an eclipse andin-between transient period on a stationary orbit.
Example: 50-V non-regulated bus: 33.0 to 52.0 V
4.4.2

Voltage stability

In the case of regulated bus type, show the bus voltage stability considering the steady state
and a load transient. The stability is desirably defined including the time of a load transient;however, it may be defined separately in Section 4.4.5. Specify the time of a load transient,
for example, as the case of a 50% load change.

31

JERG-2-214

In the case of non-regulated bus type, leave this out of the scope of application by showing
the bus voltage range according to Section 4.4.1. If a stable operation by a shunt is involved
while there is sunshine; however, it is effective for understanding the operation to show thevoltage stability in the sunshine.
Example: 100-V regulated bus: 100.0 to 103.0 V while there is sunshine
When shaded: 97.0 to 100.0 V
4.4.3

Ripple and spikeShow the upper limit of peak-to-peak value of periodic voltage variations superimposed on
the bus voltage. Specify it in the time domain and take the frequency band of themeasurement system to be 50 MHz and above. As the specified value, show a maximum
considering the operation modes while there is sunshine/an eclipse and a transient.
In addition to the above-stated ripple voltage, show the upper limit of spike voltage also in
the time domain.
As the ripple voltage specified here, only the voltage generated by the power subsystemitself and appearing at the power subsystem output end is considered.
For the definition of ripple and spike, refer to Figure 4.4-1.

Spike

Ripple• Ripple: Periodic voltage variations superimposed on
the power bus• Spike: A variation in the form of a spike generated bya switching power source but not forming anenergy source.

Figure 4.4-1 Definition of ripple and spike(Example)
Regulated bus: Ripple voltage = 0.5 Vp-p or less, Spike voltage = 2.0 Vp-p or
less
Non-regulated bus: Ripple voltage = 0.2 Vp-p or less, Spike voltage = 0.5
Vp-p or lessNote: A setting examination example for the above-stated specification isshown in Table 4.4-1.

32

JERG-2-214

If seen as a spacecraft system, the ripple voltage on the bus voltage will contain not only theripple generated by the power subsystem itself as described above but also the reboundcomponent consisting of the superimposed conduction interference noise current of eachload and applied to the power subsystem. Therefore, the ripple specification as a spacecraft
system will be set at the level coming from the power subsystem as described above plus therebound component from each load.

Table 4.4-1 Characteristics of ripple and spike
Bus type

Regulated bus
RippleNon-regulated bus
RippleSchematic diagram
Spike

Ripplevoltage

SpikevoltageDuringchargeDuringdischarge
DuringchargeDuringdischarge

Spike

0.2Vp-p

0.2Vp-p

0.5Vp-p0.1 Vp-p or less

0.5Vp-p

0.5Vp-p

2.0Vp-pNone• For each bus type, the respective maximum values of ripple voltage and spike voltage are shownwith a circle.
• The specification for ripple voltage and spike voltage will be set based on the maximum values
above.

4.4.4

Output impedanceShow the upper limit of output impedance of the power subsystem proper in the frequency
range from 1 Hz to 100 kHz. Appropriately divide and show the specification considering
the operation modes while there is sunshine/an eclipse.
4.4.5

Transient characteristics

Considering the power subsystem transient operation involved in peak power supply to aload, power-on, power-of, paddle deployment, abnormality inside the power subsystem,
load abnormality/restoration, etc., show the bus voltage transient with time at the time of a

33

JERG-2-214

load change and abnormality. Appropriately divide and show the specification considering
the operation modes while there is sunshine/an eclipse and in-between transient (penumbra)
period. Specify the time of a load change, for example, as the case of a 50% load change.
Determine it considering the actual loads and coordinating with the system.
If, in the above-stated case, the spacecraft system has a pulse power load, it is necessary toconsider it as a load involving a transient operation. As a load abnormality/restoration, forexample, the case of a fuse blown and then restored is assumed.

Sunshine voltage

Steady-statevoltage
Shade voltage

Restoration timeafter eclipse

Occurrence of load change(b) Transient characteristics involved inload change(a) After eclipse (non-regulated bus)
Steady-statebus voltage

Steady-statebus voltage

Steady-statebus voltage(c) Transient characteristics involved inload abnormality/restoration

Steady-statebus voltage(d) Transient characteristics involved inwheel stop
Note: Consider this if the wheel side does
not have a bleeder resistance or otherregenerative power absorption means.

Figure 4.4.2 Types of transient characteristics- Related considerations -34

JERG-2-214

As the reference point, designate a point at which the power supply characteristics can beevaluated. Figures 4.4-3(a) and (b) show methods of power distribution. Because the power
distribution characteristics of a power subsystem take the reference point at the powersubsystem output end, it is necessary to consider the contribution of the system harness tothe characteristics at the load input end. In the case of the backbone method in Figure
4.4-3(b), separately specify using an electrical interface, etc. for a dedicated unit end or loadend.
Specifications for ripple are specified in Section 4.4.3 as the level generated by the powersubsystem itself; however, at the spacecraft system level, they are specified as containing
also the rebound component consisting of the conduction interference noise current of eachload superimposed and applied to the power subsystem but not considering only the voltage
generated by the power subsystem itself. The actual output impedance to convert (rebound)
an interference noise current into a ripple may be related not only with the output impedanceof the power subsystem proper but also with the connected load input filter. Therefore, it isnecessary to keep track of each load input part in management of electrical interfaces orelectromagnetic compatibility.
A transient depends on the peak power of the load, amount and rate of current change at
power-on/-off, assumption of failure modes, etc. Therefore, these values and assumptions
must be specified by electrical interfaces, etc.
The specifications for power supply characteristics are premised on the actual operation
state; however, they sometimes include the transient at ground power start-up, shut-off, etc.
in a ground test. (See Clause a of 4.2.2, (1).)

Power
subsystem

Instrumentation subsystem(System harness)

Load 1Power
subsystem

Load 1

Load 3

Load 5

Dedicated
unit 1

Dedicated
unit 2

Dedicated
unit 3

Load 2

Load 4

Load 2Power
supply bus

Load 3

Power supply
bus

Load 4
Load 5(a) Radial Type(b) Backbone Type

Figure 4.4-3 Methods of power distribution

35

JERG-2-214

4.5

Power subsystem
The power subsystem consists of the interface with a solar paddle subsystem having thefunction of power generation, a battery having the function of energy storage, andequipment having the function of power control, distribution, protection, etc. A specificconfiguration and these pieces of equipment, etc. are designed based on the examination of
spacecraft system and its actual performance. Here, major components of a powersubsystem are outlined, and their characteristics, related technology and methods, specificcondition of use, considerations in design, etc. are described.

4.5.1

Interface with solar paddle subsystem

In the design of the interface with a solar paddle subsystem, clearly specify the following
items.
(1) Interface reference point
Clearly specify an interface reference point to specify the power generation of
the solar paddle subsystem. In many cases, a system with a paddle drive
mechanism (PDM or SADM) takes the paddle drive mechanism output end as
an interface reference point and a system with no paddle drive mechanism
takes the solar array output connector end as such a point.
(2) Maximum voltage and maximum current at interface reference point
Specify the upper limit current value considering the power subsystem shunt
element specifications and the maximum on-orbit current output conditions(time, direction of the sun, temperature, etc.) of an array circuit.
(3) Round-trip line resistance between interface reference point and PCU(4) Shunt operation sequence
Clearly specify the correspondence between the shunt operation sequence, set
in the system design, of the array circuits on the solar array and the operation
sequence of the power subsystem shunt circuits. If the analog partial shuntmethod is used, specify the shunt tap voltage range.

As another interface the power subsystem has with the solar paddle subsystem, it has aninterface with the pyrotechnic. For details, refer to Power Distribution (Section 4.5.4).

4.5.2

Battery

The battery cell capacity, charge and discharge voltage characteristics, number of batteries,

36

JERG-2-214

etc. should be designed considering the conditions of use (depth of discharge, temperature,etc.), characteristics of the cell (including its life), margin for the life characteristics,
redundancy, etc. and should be designed to satisfy the requirements in 4.3 Power Supply
Capacity and be compatible with charge and discharge control in 4.5.3 Power Control.

(1) Configuration
Show the specifications of the cell used (type, nominal capacity of the cell, etc.),
number of cells connected in series, electrical diagram of the battery (including the
installed sensors and heaters), insulation performance of the cell case, allowabletemperature range and allowable internal temperature difference of the battery,allowable depth of discharge for each operation mode, etc. (during a contingency, arequired discharge energy within a limited time may have to be set), and the number ofbatteries.
(2) Ground initial characteristics
Specify the charge and discharge current, time, temperature and other conditions andshow the requirements for initial charge and discharge characteristics and, if necessary,for short-cycle characteristics. If there is an intermediate tap or the like for apyrotechnic, show its required characteristics.

(3) Life prediction
Show a long-term characteristic change prediction in the state not considering the(batter level/cell level) redundancy under the on-orbit actual conditions of use.

(4) Sealing/pressure-proof characteristics.
A battery cell functions using electrochemical reactions, and chemical changes of theelectrode materials and electrolyte associated with the reaction proceed inside.
Therefore, the structure of a battery cell or battery must prevent leakage of the internalmaterials and endure increasing internal gas pressures. These structural endurance
requirements specify the initial characteristics and long-term characteristics associatedwith corrosion resistance, pressure cycle characteristics, etc.

(5) Battery redundancy
Considering a single failure of a battery, the battery should be designed to have a
redundant function so that the battery function will not be lost.

37

JERG-2-214(6) Precautions for preservation and use of battery(a) The spacecraft design should ensure that the battery can be removed from thespacecraft and replaced without influencing other equipment, etc.
(b) A battery cell or battery preserved for a long time should be brought to the ambient
temperature state gradually. In addition, perform the low-current conditioning cycle
according to the manufacturer’s specifications to bring it to the nominal performancestate.
(c) As the minimum conditions for the use and preservation of a battery cell and battery,comply with the precise preservation and reconditioning requirements for thefollowing items.

• Longest ground preservation period (before and after reconditioning if applicable)• Longest unused period in the state of no special start-up cycle taking place.
• Maximum battery temperature and its duration in the pre-launch stage and operation
stage.
• Battery maintenance in the integration stage and pre-launch stage including a launchdelay time.
• Requirements for preservation procedure, preservation temperature, and battery celldischarge• Humidity and packing during preservation• Reconditioning procedure after preservation• Preservation procedure, preservation temperature, charge state, whether or not toshort-circuit the battery cells individually, details of trickle charge, and regular
maintenance(d) To prevent damage and performance degradation over the entire remaining life after
that, the flight battery shall not be used for ground operations as far as possible.
(e) Make sure before the launch that the requirements for preservation, handling, andoperation (for example, longest allowable time of being at the upper-limit
temperature, exactness of a maintenance activity schedule, etc.) are satisfied.

(7) Battery safety requirements
Almost all batteries mounted on and used in a spacecraft can pose a risk if not
managed appropriately. In most batteries, for example, a very large current can flow

38

JERG-2-214

when they are short-circuited. If a battery cell is abused, the internal pressure canbecome excessively high and the contents may gush out; in an extreme case, even an
explosion can occur. Moreover, the emitted electrolyte, reactants and/or reaction
products of the battery cell are corrosive, inflammable and poisonous, so they canjeopardize all personnel present nearby in addition to adjacent devices. Therefore,
perform safety verification for each one. The safety requirements for a lithium ion
battery are described below. For conventional Ni-H2 and Ni-Cd batteries, refer to thespecifications of JMR-001 and JMR002.
(a) Safety requirements for battery cells
The battery cell should be provided with a rupture part to avoid a severe rupture.

The battery cell should be provided with the rupture part at two places to haveredundancy.

The battery case should have a safety margin of three times the ordinary working
pressure.

It should be provided with a rupture part that surely operates at a lower pressure
than the case withstand pressure. This should be demonstrated by a test.

The battery cell should have external short-circuit protection function.



The cell should have the safety for a forced discharge. This should bedemonstrated by a test.

(b) BatteryIn the battery design, each cell should be provided with a shunt function toprotect against an over-charge.

The design should be such that when battery cells are assembled into a battery,the cell safety valve operation is not hindered.

(c) Battery transportation and handling toolsFor battery transportation, pack the batteries securely to prevent an external
short-circuit.

The battery transportation packing should be proven as to the following by
carrying out a drop test from 1.2 m. 1) The batteries inside and battery cells are
not damaged. 2) The batteries inside and battery cells do not experience anexternal short-circuit. 3) The batteries inside and battery cells are not thrown out.

When handling the battery proper, provide a strong protective gear for a drop,

39

JERG-2-214

etc. This protective gear should have a design to prevent an external short-circuit
and a battery cell’s forced out of the battery even when dropped from 1.2 m.
(d) Others(1) If a battery cell group is connected in parallel, an effective means for reverse
current prevention should be provided.
(2) Control should take place to prevent the upper-limit operation temperature frombeing exceeded.
(3) Conduct a test to verify that the battery cells and battery function safety even at
the upper-limit operation temperature. Also consider the heat generation ofbattery cells due to external heat input and internal loss.
(4) Identify the processes important for safety. These important processes are
subject to an examination by JAXA.
4.5.3

Power control

Power control of the power subsystem consists mainly in output control of the solar arrayand charge and discharge control of the battery.
Each control should be designed to be compatible with the solar array and battery
characteristics and satisfy the requirements of 4.4 Power Supply Characteristics.
(1) Control system
Show the diagram and control law of all functions involved in power control. For solar
array output control and battery charge and discharge control, clarify the operating range
and priority of control, and coordination management not to allow control operations tooverlap each other detrimentally.
In the case of a shunt-type 50-V non-regulated bus and charge control by a battery charge
controller, the operating range and priority of control are, with increasing bus voltage, inthe order of bus voltage control by charge control (bus voltage width about 1 V), a gap(arbitrary) and bus voltage control by shunting (bus voltage width about 1 V) as
standard.
In the case of Multibus, show the functional diagram and control law for eachindependent operation. If there is a coordinated operation of multiple buses, show theiroperation in the coordinated state separately. For the parallel operation of multiple buses
and multiple batteries, specify the allowable level of their current share.

40

JERG-2-214

For a single failure inside the power subsystem and an abnormality on the load side,
designate the place of the failure and failure mode and specify the bus voltage transient
involved in the failure.
- Related considerations In a coordinated operation, provide a monitoring means to keep track of its state anddesign to ensure that there is no detrimental interference or alternating intermittent
operation.
The abnormality bus voltage transient should be consistent with the specifications ofSection 4.4.5.
(2) Operating point analysis
Clarify the behavior of the operating point of all voltages (bus voltage, control voltage,
and their monitoring voltage, etc.) in voltage control.
- Related considerations In the operating point analysis of voltage, clarify the operating point of the solar array,bus voltage and battery voltage. In the analysis, it is necessary to consider BOL, EOL,
shade/sunshine transient, load characteristics and their variations, etc.
If the lock-up phenomenon (the battery discharges regardless of whether the load power
can be supplied by the solar array) is expected to occur, take steps to identify and, if
necessary, prevent that.
(3) Control diagram
For all operating and non-operating regions of power control, show a control diagram(equivalent circuit, simulation model, etc.), by which the bus voltage stability, output
impedance, control stability, and transient response characteristics can be evaluated, andthe effective frequency region of each.
For a feedback loop, ensure a gain and phase margin of about 5 to 15 dB and 25 to 50
degrees or more, respectively.
An appropriate gain crossover frequency is several hundred to several ten kHz or so.
- Related considerations In the control diagram, clarify the calculation grounds for the characteristic value of eachcontrol element.

41

JERG-2-214

Have a design so that the control performance will not be impaired by temperature,radiation degradation, load conditions, etc.
When evaluating the effects of load conditions, it may be necessary to know the input
impedance of all pieces of equipment on the spacecraft.
The gain and frequency band width of feedback loop control should at such a level that asufficient margin can be ensured for the requirements for bus voltage stability,
impedance and transient response characteristics, battery charge and discharge current
and temperature. The design should be such that an appropriate gain/phase margin can beensured and, even in the worst case, control stability can be ensured.
Have a design so that, for each control, not only the static stability in a control mode but
also the transient response characteristics at a transfer within a control mode and between
control modes satisfy the transient specifications.
The modeling on which a control diagram is predicated generally needs approximation,etc. of nonlinear characteristics into a linear form. So, pay attention to its validity limit.
For a simulation model, use a high frequency equivalent circuit as close to the actual
circuit as possible in the model, by using parasitic parameters not present on the circuit
diagram or otherwise.
(4) Characteristics of each power control
In some design, a control function consists of multiple elements (pieces of equipment).
In such a case, it becomes necessary to specify interface boundaries and specify thecharacteristics individually. Here, the characteristics by control function are described.
Specify the loss by (efficiency of) power control taken in as a parameter of the powerbudge analysis mode, individually in the relevant part.

a.

Solar array output control
The solar array output control should have a processing capacity compatible with thesolar array characteristics. Specify this in terms of the power for that processing (meanand peak current and voltage). Specify the monitoring method to keep track of theoperation state of solar array output control.

b.

Charge control
Specify the set values for the control parameters (voltage, current, temperature, etc.) to
be compatible with the battery characteristics.

42

JERG-2-214

And, specify the operations of the protection function (charge suspension at antemperature abnormality, etc.) for heater control (if built in) and charge of the battery.
Pay attention also to the transient of bus voltage and battery charge current at a chargesuspension and resumption due to the action of a protection function.
c.

Discharge control
Specify the requirements for discharge power (peak and abnormality values, in addition
to mean value).
Specify the operations of reconditioning function and protection function (UVC,
isolation of a short-circuit on the load side, etc.) for discharge of the battery.
- Related considerations For solar array output control, consider that the output characteristics are varied on ashort-term basis by temperature, sunlight incident angle, etc. and on a long-term basisby radiation degradation, etc. In addition, an environmental resistance at the same levelof the solar panel is required for the solar paddle and other control elements attachedoutside the spacecraft. If direct measurement of solar array power generation is difficult,a monitoring means to keep track of the operation state of output control is utilized as
an indirect evaluation means in some cases.
For battery charge and discharge control, it is desirable, considering the long-termchange in the battery characteristics, to provide multiple control levels (V/T curve,
over-voltage protection, over-heat protection, UVC level, etc.).
In general, a fuse blowing capability to isolate a short-circuit on the load side is also
required in discharge control.

4.5.4
4.5.4.1

Power distribution
Power distribution pattern(1) Bus configuration
If broadly divided, there are architectures of single bus configuration and multibus
configuration. Because each has its own merits and demerits, consider an appropriate bus
configuration based on the system requirements.
(2) Wiring pattern
As the power distribution wiring pattern for a component made of main/subordinate
components (including a component of an internally redundant configuration), do not use
a wired OR connection (so-called π connection) at the connector but employ individual
power distribution. If a protective circuit (fault isolation function) of a circuit-breaker orfuse is to be provided on the power distribution side, install it on each power distribution

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JERG-2-214

line.
(3) Selection of wire and wiring method used for power distribution
For the selection of wire and wiring method used for power distribution, conform to
Section 4.6.1 of JERG-2-200 Electrical Design Standard.
(4) Supply power monitor
With an understanding of keeping track of the state of power supply to the load side
during a system test and on-orbit operation, consider an appropriate supply powermonitoring function (resolution and system).
4.5.4.2 Operating temperature of power distribution harness
For a harness of bundled wires engaged in power distribution, perform a thermal analysis
considering the effects of heat from mounted peripheral equipment, etc. and the heatgeneration by current flowing in the harness itself, and confirm that the operating
temperature satisfies the specifications of JERG-2-212 Wire Derating Design Standard.
4.5.4.3 Inrush current/outrush current
When an inrush current/outrush current flows at power-on/off by normal operation ofequipment on the load side or at an operation of a protective circuit (on both the power
distribution side and load side), the protective circuit on an unintended distribution line
should not malfunction.
4.5.4.4 Specification of power distribution specifications
As the output of a power distribution design based on the requirements of Sections 4.5.4.1 to4.5.4.3, specify the following specifications concerning each item of power distribution by
the bus voltage and in the form of a power distribution management table.
a. Supply pattern (normally on, on/off by a command, on/off by a designated sequence,etc.)
b. Supply current (mean/peak/profile mean in the operation mode)c. Voltage drop from the power supply bus to each power supply end
d. Harness type and voltage drop from each power supply end to the corresponding input
end on the load sidee. On/off relay switch type (predicted number of operations, related command number,installed place)f. Telemetry number related to keeping track of power supply state (voltage/current/status
and there measuring positions)g. Fuse or power-off function provided or not (rating and other specifications, and placed
position)
Also for power distribution by other than bus voltage (by adapting a prefabricatedcomponent or otherwise), specify the power supply characteristics within the necessary
scope and according to Section 3.3 and specify items similar to the above. If, as withpyrotechnic ignition power, a supply sequence function is incidental to the power subsystem,
specify its details, etc.
- Related considerations Because of an inappropriate design or with the progress of the design stage, a related
specification or the like may have to be changed. When examining the impact of such achange, End-to-End management concerning power supply is performed. Therefore, even if
the above-mentioned related specification is an item not classified as a function of the powersubsystem, the specification should not be divided or omitted.

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For a platform spacecraft*Note 3 bus, power distribution to unitized subsystems, etc. isemployed. In such a configuration, the functions of power supply and distribution to theequipment in a unit are dispersed to each unit. For Clauses “a” to “g” above, however,specify and manage them as belonging to the power subsystem.
These specifications are sometimes collectively incorporated in a related document ofelectrical design or interface management. The specifications desirably use a table in which
they are used in common with Clause (1), a of Section 3.2.2.
For pyrotechnic ignition power distribution, consider Clauses (1) to (7) and specify andmanage it using an actual system diagram including the pyrotechnic harness and in the same
was as above.
(1) For sure ignition, give redundancy to the ignition system (ignition circuit, wire,
pyrotechnic). Provide the automatic ignition sequence, etc. with a backup of groundcommand.
(2) Letting the sequence to ignition have three stages (enable/disable, ignition), monitor the
status.
(3) In the sequence, incorporate the spacecraft/rocket separation signal, the attitude control
subsystem’s pre-operation completion signal, etc. in a circuit.
(4) For malfunction prevention, simplify the system, clarify the relationship, etc. withspacecraft grounding, and connect an antistatic resistor to the wire to the pyrotechnic.
(5) Because the heating wire to fire the powder in the pyrotechnic (bridgewire) cannot fulfill
the firing function if the supply current or time is insufficient, know the transient
response set the supply time of the power supply source.
On the other hand, an unexpected contingency (release becoming impossible,
short-circuit to the structure, etc.) occurs if an excessive current flows. Therefore,
incorporate a current limiting function (should have the function to prevent an excessive
current and the function to prevent continuous voltage application) so that therecommended firing current of the pyrotechnic flows.
(6) A relay is generally used for powering on, etc. a pyrotechnic. However, contact reversal
or chattering due to vibration or impact poses the risk of causing serious troubles fromsystem malfunction. Therefore, evaluate the environmental resistance of the relay,
pyrotechnic controller, etc. in advance and take measures.
(7) In the case where the number of operations is limited, verification using actual
equipment is difficult. Therefore, implement a test plan to verify the system performanceafter sufficiently examining it.
*Note 3: Refers to the spacecraft type of ADEOS and ADEOS-II.
4.5.5 Protection
4.5.5.1 Overview
The protection function has the roll of localizing failure or abnormality, which may have aserious adverse effect on power supply, to avoid a critical situation as a system or a situation
affecting mission accomplishment.

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JERG-2-214

For the power subsystem, uninterrupted operation is required, so it is necessary to have adesign on the premise that the primary functions cannot be interrupted as fail-safe.
4.5.5.2(1)

Protection in power subsystem
Redundancy(a) The power subsystem should not have a single failure point that can be assumed by
failure analysis.
If a single failure point subsists, identify that single failure point and show thegrounds for the subsistence. The grounds for this include the need for a technicalevaluation of occurrence possibility and, if any, an examination of a means that canremove the single failure point and measures to reduce the probability or latenteffect of failure occurrence.
(b) For an abnormality or failure of a device due to an open circuit or short circuit, takemeasures such as series-parallel connection, circuit and equipment redundancy, etc.
Pay attention to the failure rate of the OR circuit, change-over circuit, etc. toachieve a redundant configuration.
(c) No two protection functions supporting an essential function in the converter or
regulator should be performed in the same hybrid or monolithic IC; they should use
independent reference voltages and drive power sources.
This does not apply to a hybrid IC which accommodates function groups physicallyisolated to prevent failure propagation.
(d) The primary power source should be recoverable when the abnormal condition (forexample, abnormal attitude, etc.) is removed, including the case of loss of
secondary power.

(2)

Insulation
Insulation should be in accordance with the Insulation Design Standard (JERG-2-213).
The matters to be specifically noted as a power subsystem are described here.
(a) In power distribution from the primary power bus, there are parts where
over-current protection cannot be implemented (include the bus bar, harness,connector, and PCB). Protect these parts by double insulation to the point of thefirst protective device (fuse, current breaker or current limiter, or reverse current
prevention diode)(3)

Over-current protection(a) Install a protective circuit as close to the power source as possible to minimize theprimary power bus length.

46

JERG-2-214(b) If a fuse is used to protect the distribution wire of the primary power bus, the ratingof this fuse shall be selected considering the maximum current in the load path
including the transient state (of inrush current, etc.). In addition, the power source
should be able to supply such a current that can blow the fuse by the time when thebus voltage reaches the abnormality lower limit.
(c) If a fuse is used to protect the distribution line of the primary power bus, the powersubsystem should be designed so that, when the load is short-circuited, this fuse canbe blown within the specified time determined by the electrical design standard ofeach project.
(d) If a fuse is used to protect the primary power bus, an open circuit of the fuse can be
confirmed from outside the component. It is desirable that the fuse is accessible andcan be replaced even when the spacecraft is integrated.
(e) If a fuse is to be installed in the subsequent stage of a DC/DC converter, consider
the melting characteristics of that fuse to ensure that the DC/DC converter has theability to blow the fuse for certain.
(4)

Battery protection(a) A battery cell should have protection function so that a short circuit, open circuit orother abnormality of it does not affect other battery cells or primary power bus.
(b) A battery should have protection function such as suspending the charge to preventthe battery from ending up critical failure if an over-temperature occurs in thebattery.

4.5.5.3

Protection in cooperation with other subsystem

For the protection functions related to power supply shown in (1) to (5) and in cooperation
with other subsystem, specify the specifications of Clauses a to f in the form of a protection
function summary table. An example of protection function summary table is shown in
Table 4.5.5-1. For a complex function, specify it using a system diagram, state transition
diagram, etc.
(1)

Malfunction of attitude control subsystem, tracking subsystem, etc.

(2)

Abnormality of operating point, etc. of solar array(3)Over-discharge of battery(4)

Load short circuit of battery(5)

Abnormality of multiple cooperating busesa.

Operating conditions (detection level and response) and operation confirmation
method (monitoring, etc.)47

JERG-2-214b.

State transition and recovery procedurec.

Self-diagnosis method of protection functiond.

Protection function and malfunction measures (reset, override, etc.)
Override refers to setting of ENA/DIS, forced protection cancellation, etc.

e.

Normality and operation confirmation method by ground test, etc.

f.

Related (protection for others, etc.) function and interference prevention measures

48

JERG-2-214

Table 4.5.5-1 Protection function summary table

a. Operating conditions (detection level
and response) and operation
confirmation method (monitoring,etc.)

Input power source dynamic variations
simulating attitude control subsystem,
tracking subsystem, etc.
UVC monitor, bus voltage, and input
power(2) Abnormality of solar panel operating (3) Battery over-discharge
point, etc.
Prevention of lock-up by load variations Operation confirmation of LVC
Transition to survival mode
Confirmation of survival power
Confirmation of abnormal load
disconnection function
Solar battery current monitor by load
Simulation of battery over-charge by load
variations
variation
Generated voltage monitor
Output voltage, current monitor, battery
temperature monitor, and bus voltage
monitorb. Status transition and recovery
procedure

Normal power  UVC mode  survival
power  power recovered  normal
power

Normal power  Maximum load mode
Normal power  UVC mode  Load
 Power drop  Normal power  Light disconnected  survival power  power
load power  Normal power
recovered  normal powerc. Self-diagnosis method of protection
functionTransition to survival mode due to input
power decrease by abnormality of
attitude control subsystem, tracking
subsystem, etc. and power supply after
restoration to normality from thisabnormality.
Automatic restoration

Solar battery lock-up (lock-up does notoccur) by load variations and restoration
function by light load

Over-discharge the battery byshort-circuiting the battery (loadvariation) and then disconnect the
short-circuited part and confirm thepower supply resume function.

Automatic restoration (or reset)

Automatic restoration (or reset)

Automatic restoration (or reset)

Characteristic confirmation by input
power source dynamic variations which
simulate the attitude control subsystem,
tracking subsystem, etc.
It is difficult to cause a malfunction of
the attitude control subsystem, tracking
subsystem, etc. actually at the satellite
level.
It is difficult to cause a malfunction of
the attitude control subsystem, tracking
subsystem, etc. actually at the satellite
level; however, operation confirmation of
the power subsystem is relatively easy bytaking this event as an effect on the
power subsystem and simulating the
malfunction of the attitude control
subsystem, tracking subsystem, etc. by
reducing the test equipment input power.

Confirmation of solar battery lock-up by
load variation (confirmation of nolock-up occurring) and confirmation ofrecovery function by light load
Impossible with the solar battery
connected.

Simulate a battery over-discharge byvarying the load and disconnect the
short-circuited part, and confirm thepower supply resume function after that.
This is very dangerous operation
confirmation and difficult at the satellite
level.

Not allowed

Operation can be confirmed bydisconnecting the solar battery andsimulating the solar panel operation pointwith the test equipment input power.

Although it is very dangerous operation
confirmation and difficult at the satellite
level, it is possible to confirm a battery
over-discharge by reducing the test
equipment input power as a battery
over-discharge.

Verification functiond. Protection function and malfunction
measures (reset, override, etc.)
e. System test allowed or not

f. Risk in case of performing as asystem test, and the reason forimpossibility
g. Alternative confirmation method(1) Malfunction of attitude control andtracking subsystems
Survival mode detection (UVC function)
Power recovery function
survival power

49(4) Battery load short-circuit
Confirmation of overload function

Confirmation of power supply shut-off
function and power recovery
characteristic function by overload
characteristics
Output voltage, current monitor, battery
temperature monitor, and bus voltage
monitorNormal power  overload mode  Loaddisconnected (UVC mode  survival
power)  power recovered  normal
power
Disconnection of short-circuited part andpower supply resume function

It is impossible to form a short circuit
actually. This is utterly impossible
especially if a fuse is used for battery
output protection.
For an over-current prevention circuit
using an electronic circuit,
implementation may be possible byissuing a simulated signal after
over-current sensing.

JERG-2-214

4.5.5.4

Protection against ground equipment abnormalities

It is not practical to take measures on the spacecraft side against ground equipment
abnormalities. Therefore, verification of spacecraft should be performed using sufficiently
verified ground test equipment.

50

JERG-2-214

5. Design verification
5.1

GeneralVerify that the power subsystem is appropriately designed and the requirements for power
supply are satisfied.
The verification takes place in a design review based on fabrication and test of EM, etc. tothe design stage, a component test of PM/FM by the fabrication and test stage, acombination test of the power subsystem, a spacecraft system test, and a board of review
after these tests. The requirements, plan, organizational structure, method, etc. for the boardof review and tests are specified in the following documents.
• Reliability Program Common Specification and Quality Assurance Program Standard• Power Subsystem and Spacecraft System Specification• Test Requirements for Spacecraft
The requirements for the power subsystem shown under Design Standard of Section 4
should be organized into a confirmation matrix showing the classification of confirmation
methods (inspection, analysis and similarity, test) in the quality assurance clauses of thePower Subsystem and Spacecraft System Specification. According to this, for a test in
addition to analysis, its purpose, specifications of a specimen, configuration, method,decision criterion, timing of execution, organizational structure, etc. should be specified in aplan, etc. The verification results should be organized into a design review material,
post-test review material, pre-delivery review material, etc. and their conformance to therequirements should be confirmed.
Table 5.1-1 shows the basic policy of verification for design standard items in the form of aconfirmation matrix for use as a check list, etc. of a design review. In the table, Inspection,
Analysis (including similarity), and Test show the classification of verification methods. For
the tests related to the power subsystem, there are levels and items as follows:• Component proper test (electrical performance, temperature, mechanical environmental,
electromagnetic compatibility, thermal vacuum and thermal cycle, etc.)• Power subsystem combination test (electrical performance, temperature, etc.)
For a unitized*Note 4 power subsystem, the object of the combination test takes the flight form
and mechanical, thermal vacuum, etc. tests are performed.
• System integration and system test
(Electrical performance, mechanical environmental, electromagnetic compatibility, thermal
vacuum and thermal balance, etc.)
Besides these, a charge and discharge cycle test is conducted for the purpose of confirming
the battery life. In addition, the execution contents, conditions, etc. are different in theclassification of model (development, qualification and acceptance) tests according to thedevelopment stage.
Comments on these test classes, etc. are shown in Table 5.5-1.
Here, points to consider about the tests, analyses, etc. recommended to be done as the powersubsystem are described.
* Note 4: Refers to the spacecraft type of ADEOS and ADEOS-II.

51

JERG-2-214

Table 5.1-1 (1/3) Design verification matrix

4.2
4.2.1

4.2.2
(1)
a
b
c
d

e
f
g
(2)

(3)

(4)

(5)

Definitions of and restrictions on goods
items
Configuration and components• Configuration diagram/components(equipment)• Operation modes/transition conditions
Definition of interface
Electrical interface
Power supply (power consumption)
Telemetry/command• List/calibration curve
Control signal
Pyrotechnic ignition power• Diagram/sequence• Grounding (single grounding/diagram)
AGE
Umbilical/test line
Inside the power subsystem• Between spacecraft and paddle, etc.
Mechanical interface• Mass/dimensions/attachment/connector, etc.
Thermal interface• Heat generation/thermal
characteristics/temperature range, etc.
Functional interface• Cooperation/interference with othersubsystem
Work interface• Identification/shipment
preparation/work support

4.2.3
Common design standard(1) Electrical design(2) Mechanical design• Strength/rigidity/proof and failurepressures, etc.
(3) Thermal design(4) Electromagnetic compatibility(5) Environmental conditions
a Temperatureb Pressurec Humidity
d Cleanlinesse Acoustic
f Accelerationg Vibrationh Impact
I Radiation
j Exposed part environment• Heating during launch• Atomic oxygen• Meteoroid/debris• Sunlight/albedo/earth radiation• Orbit peripheral plasma• Thruster plume

Test

Analysis

Design requirements

Inspection

Design verification method
Remarks

Check actual articles, drawings, etc.


































Systematically confirm in a system test.
Confirm pin arrangement before connection.
Confirm connection in integration process.
(Pay attention to paddle I/F confirmation method.)Confirm characteristics, etc. of electrical performance
tests.

Confirm by system integration.

Confirm by thermal vacuum and thermal balance tests.












Pay attention to configuration.

Check actual articles, documents, and work history.












Confirm by design review.



Requirements and restrictions



Application criteria
Measures taken in design











Results of preliminary trial fabrication / test
Validity of design analysis
Results of individual qualification test

Inspection items and test method







52

Test items and test methodProperty assessment data
Management and maintenance items

JERG-2-214

Table 5.1-1 (2/3) Design verification matrix(6) Safety design(7) Reliability design• Reliability• Trouble (recurrence prevention)measures• FMEA• (Critical) single failure measures• Confirmation of redundant design• Worst-case analysis(Variance, temperature, radiation, etc.)• Sneak circuit analysis• Selection of device/ material/ process• Stress analysis• Derating• Individual qualification test
• Trend assessment• Limited life/operations articles• Ground maintenance of battery





Test

Analysis

Design requirements

Inspection

Design verification method

Remarks

Confirm by design review.













Requirements and restrictions









Application criteria



Results of preliminary trial fabrication/test









Measures taken in design

Validity of design analysisResults of individual qualification test
Inspection items and inspection method
Test items and test method
Property assessment data

(8) Maintainability design
Battery exchange procedure
4.3

Power supply capacity
4.3.1 Required power profile• Operation mode/supply capacity bymode
4.3.2 Required size of power source• Power budget analysis

4.4







4.5.2
(1)
(2)
(3)
(4)
(5)
(6)
(7)

Compare and evaluate analysis and test results.
Component unit test







Bus voltage
Stability
RippleOutput impedanceTransient characteristics

Properties of major components
4.5.1 Interface with solar paddle subsystem• Interface reference point• Maximum voltage and current
specifications at reference point• Line resistance between reference
point and PCU• Shunt operation sequence
Battery
Configuration
Ground initial characteristics
Life prediction• Related actual data
Sealing/pressure-proof characteristics
Battery redundancy
Precautions for storage and use of battery
Safety requirements for batteries

Check validity by design review.

Update data according to test results, etc.

Check power budget by power analysis.

Power supply characteristics• Reference point/measurement method• Power subsystem operation
mode/simulated power source• Trend assessment
4.4.1
4.4.2
4.4.3
4.4.4
4.4.5

4.5

Management and maintenance items















Power subsystem combination test
System test

Inspect by drawings, etc.
Check paddle subsystem irradiation test data.









Confirmation by drawings, ICD, etc. and the operation on
the simulated solar array power in a system test.
Inspect by drawings, etc.


Do quality confirmation, operation simulation test, etc.















In-process test, etc.
Check actual articles, documents, and work history.

53

JERG-2-214

Table 5.1-1 (3/3) Design verification matrix

4.5.3 Power control(1) Control system• Functional diagram and control law• Cooperation management(2) Operation point analysis(3) Control diagram(4) Characteristics of each power control
a Solar array output control• Compatibility with solar arraycharacteristics• Operation (power generation)
monitorb Charge control

4.5.5 Protection• Protection function summary table

Test

Remarks
Compare and evaluate analysis and test results.




















Development (partial trial fabrication, etc.) testComponent unit testPower subsystem combination test
(Pay attention to solar paddle I/F.)
Confirmation by a system test using simulated solar
array power.
Ditto
Check by battery combination test and system test
(thermal vacuum).

• Compatibility with battery
characteristics• Battery protection function
c Discharge control• Required discharge power• Battery protection (maintenance)
function
4.5.4 Power distribution• Power distribution management
table

Analysis

Design requirements

Inspection

Design verification method
















Integration → System test











Pay attention to prior examination of test method.

54

JERG-2-214

5.2

Definitions of and restrictions of goods items

5.2.1 Configuration and components
The configuration and components of the power subsystem should be verified by inspecting
goods items, drawings, etc.
All operation modes of the power subsystem should be verified by confirming the functions
and operation by a test.
When confirming an operation mode transition, confirm by telemetry data, status, etc. thatthe operation mode transition takes place under the desired transition conditions.
Operation confirmation is desirably done for all modes by the time of completion ofsubsystem tests and under the condition of the full configuration and actual operation
simulated within the realm of possibility.
Related considerations
In some of the tests involving the solar panel (paddle), on-orbit simulation is difficult
because of restrictions on the test equipment, etc. Simulated solar array power is used as
AGE that simulates panel output; however, this is not the full configuration as a powersubsystem. Operation confirmation involving UVC or protection function will also use asimulated signal, etc. It is necessary to make a confirmation for all modes (including theprotection function, etc.) without any omission in the end, considering these restrictions andusing analysis, etc.
5.2.2

Definition of interface

The requirements for the definition of an interface should be confirmed at each stage of thedesign review, unit test of elements, and combination test and should be verified by a system
integration test and system test in the end.
(1)

Electrical interface

a. Power supply
It should be confirmed that the prescribed power supply characteristics are ensured andthe supply current (level/waveform) satisfies the specifications and, in addition, when
integrated, the similarity to the unit state should be confirmed. Confirm that thetransient at power-on or power-off in a ground test is also within the specification.
Consider normality confirmation of a line and redundant function part (connector, wire,etc.) which use a parallel (redundant) fuse.
Related considerations

55

JERG-2-214

The power consumption of the power subsystem is divided into a part depending on
power control operation and a steady part and is generally defined as the steady part.
From the viewpoint of evaluating the parameters in a power budget analysis model, it isdesirable to measure, as source data for example, the (input) current at the interface
point between the solar array and a power control element and the (output) current at
the interface point between the power control element and a power distribution element
in a typical operation mode, etc.
b. Telemetry/command
In a unit test, it should be confirmed using a dedicated test and adjustment equipment,
etc. and by the specified method that the timing, input and output characteristics, etc.
are in accordance with the desired design.
At the system level, confirmation of all telemetry/command list items should be madeby a functional test, etc.
However, for other than the functional test items related to LLM/UVC, protection
function, pyrotechnic ignition power, etc., a confirmation should be made in therelevant test as the need arises.
The source data (calibration curve) involved in engineering value conversion oftelemetry data should be obtained in each EM and PFM unit test and be confirmed tomeet the accuracy requirements before use in the system EM and system PFM tests.
c. Control signal
As a unit, all functions related to a control signal should be confirmed by a dedicatedtest and adjustment equipment using the simulation circuit, etc. defined in the interface.
At the system level, an End-to-End functional confirmation should be made within therealm of possibility for all control signal items.
(The considerations for simulated signal generation are described in the relevant clausesof Chapter 3.)
d. Pyrotechnic ignition function
In the unit test, a functional test of the whole system (including redundancies) should be
performed using dedicated test equipment to confirm that the input voltage, ignition
current, supply duration and operation sequence to ignition are in accordance with thedesired design.
Also in a functional test on a system EM or system PFM not using a pyrotechnic, aconfirmation should be made for the same items stated above, and an End-to-End
functional confirmation should be made in the full configuration within the realm of
possibility.
In the case of a design with a limited number of times of operation, verification of thedesign and fabrication should be performed by confirming sequence-based conduction,
insulation, etc.
e. Grounding56

JERG-2-214

Confirmation related to grounding and bonding should be made in system integration,and it should be confirmed by individual measurements that the specified resistancevalues are satisfied. A single-point grounding of the primary bus return is usuallyconfirmed in the power subsystem if it is to be connected within the power subsystem.
If it is not connected within the power subsystem, a confirmation should be madeduring system integration.
The comprehensive validity of a grounding system should be verified by confirming thenon-occurrence of loop formation on the grounding system diagram, conforming no
return current present at the single grounding point and by a system EMC test, etc.
Evaluation of a grounding design to prevent the effect of static charge and discharge
should be performed and verified by an EMC test.
f. AGEFor all umbilical cable/test cable items, confirm that the requirements of a prescribed
specification, etc. are satisfied, before connecting the device to the spacecraft, etc. andusing it. In particular, pay attention to the following:• Test results of the (overvoltage and overcurrent) protection of a power supply line,
etc. for prevention of damage to the spacecraft, etc.
• Fool-proof design to prevent damage to the spacecraft, etc. due to a mistake insetting, etc. by test personnel or operator.
For connection of ground test equipment, pay attention to the grounding system on the
ground side and spacecraft side. For umbilical cable items, in particular, confirm thecompatibility with the lightning measures of the range. Confirm the risk prevention
measures from the viewpoint of work safety.
g. Inside the power subsystem
As a unit, confirmation of the conformance to the specification takes place, and aconsiderable part is confirmed by power subsystem combination tests. On this occasion,
also combine the simulated solar array power and other AGE used in the system test.
The data thus obtained becomes source data for evaluation of system test results.
For an interface between the solar array and power subsystem (spacecraft), its operation
confirmation with current, etc. actually flowing is difficult even if they are connected.
Therefore, the (conduction and insulation) normality of a connection will be confirmed
from a connector pin, etc. On this occasion, pay attention to the review of drawings, etc.
and clarification, etc. of instructions in advance because a trouble may be brought about
in orbit if there are errors in drawings, etc. which form a criterion for confirmation.
There are practical examples of operation confirmation using a simulated lamp, etc.,
and it is desirable to examine some operation confirmation method from the design
stage.
For an interface between the solar array and power subsystem (spacecraft), its operation
confirmation with current, etc. actually flowing is difficult even if they are connected.
(2)

Mechanical interface

57

JERG-2-214

As a unit, conformance to ICD, etc. should be inspected.
In the end, verify the conformance in the system integration process of assembling it in
the spacecraft main body.
(3)

Thermal interface
Confirmation of thermal interface validity at the unit level should be made byconfirming the grounds, etc. for numerical value calculation. At the system level, verify
by confirming the reference point temperature by a system thermal balance test.
For a thermally critical component, model validity evaluation should be madeaccording to the instructions of the thermal model and related test results.

(4)Functional interface
Verification of linked items is performed at the place of the relevant function, and its
conditions and method are specified in a related document, etc. Therefore, confirm that
there is no omission in the related documents of the two and check the related item of
the power subsystem, for example, the protection specifications for an abnormality
signal, the relevant power distribution specifications for pyrotechnic ignition power,
and the specifications of a signal receiving function, etc. for an isolation switch.
Verification of interfering items is performed by a review of analysis results andconfirmation of the data, etc. used for the analysis.

(5)Work interface
On receipt, inspect to ensure that the identifications, history/instruction manual, packing
method, etc. are as specified.

5.2.3 Common design standard
For the items required by the Design Standard, their compatibility is proven by a review ofthe design and analysis results, etc. except for the specifications to be confirmed by aninspection or test. Here, the validity of the analysis technique itself is also subject to thereview. For an item which is related to the design policy and whose assessment criterion in
the review is not clear, that is, simply stated, for example, efficient, high density packaging,and small size and light weight, consider a comparison with past records or similar cases
and the current technical level, etc.
Items to be confirmed mainly by inspection or test are described below.
(1)

Electrical design
For a connector pin arrangement, confirm by the drawing, etc. For a fuse, review thevalidity of its type, conditions of use, etc. as well as checking the traceability of its
usability after it is installed and the usability results.
For magnetic characteristics, besides conducting a review, make a magnetic
characteristic measurement at the component or system level though this depends on
the system.

58

JERG-2-214(2)Mechanical design
The validity of a design is evaluated and confirmed mainly by (vibration, acoustic,
impact, acceleration, etc.) mechanical environmental tests at the unit level and, for apressure vessel, by a proof pressure test and destructive test. For large, heavyequipment, perform a validity evaluation of its numerical model by a test, etc. on its
structural model and perform an evaluation and confirmation of the compatibility withthe system by coupling analysis(3)

Thermal design
The validity of a design is evaluated and confirmed mainly by (thermal vacuum,
thermal balance, etc.) thermal environmental tests at the unit level. As the order ofprecedence, thermal environmental tests should be performed as a general rule after
mechanical environmental tests to confirm no characteristic degradation caused by themechanical environment. For equipment needing a thermal model, perform a validity
evaluation of a numerical model by thermal balance test, etc. on the thermal model, etc.
and evaluation and confirmation of the compatibility with the system by coupling
analysis, etc.
For a harness bundle not satisfying the requirements of the Wire Derating DesignStandard (JERG-2-212), it should be subjected to a harness temperature evaluation test
using a thermal model and temperature evaluation and confirmation (acquisition of
harness temperature data) by a system thermal vacuum test.

(4)

Electromagnetic compatibility design
The validity of the design should be verified by the unit-level and system-level
electromagnetic compatibility tests specified in the electromagnetic compatibility
management document. For the system-level electromagnetic compatibility test,
employ the test conditions which simulate the on-orbit operation within the realm of
possibility.
From the viewpoint of the Related Considerations described below, it is necessary toalways confirm the bus voltage behavior over the entire period of the unit-level andsystem-level tests, so it is desirable on this occasion to pay attention to the varying
waveform not only by telemetry data of limited information content but also by test
connector outputs, etc.
Related considerations
The power supply characteristics are closely related to electromagnetic compatibility of
conductivity and the operating conditions of each load is related to the power supply
characteristics. On the other hand, keep in mind that the output impedance of a powersubsystem depends on the operating region of power control of the power subsystem.

(5)

Environmental conditions: Ditto, except for the representation below.
The environmental resistance related to the requirements for mechanical and thermal
design should be verified by an environmental test according to the conditions andmethod specified in related standard, test specification, etc. Each item is described
below.

59

JERG-2-214a.

Temperature and pressure
Should be evaluated and confirmed by a thermal vacuum test. The power subsystem isequipment operating from the launch and the pressure decreases from the ground levelto the on-orbit level. Make a confirmation that no trouble occurs due to a corona
discharge, etc. during this period.
Prior to a component thermal vacuum test, a temperature test using a thermostatic ovenis sometimes conducted at the card level or module level to evaluate and confirm theeffects of temperature alone. By this temperature test, evaluate and confirm the validityof the design (to determine the comprehensive characteristics with devices oftemperature characteristic variations reduced by a negative feedback circuit, etc.) to
hold down the effects of temperature to the lowest possible level and the effects, etc. of
temperature characteristic variance of the devices. In particular, for the operation of aprotection circuit and other functions whose operation is difficult to confirm by a test at
the component level, verify by this test.

b.

Humidity and cleanlinessNo particular test is conducted concerning these conditions.

c.

Acoustic, acceleration, vibration (sinusoidal and random), and impact
The power subsystem is equipment operating from the launch. The environmental test
applied in the Environmental Resistance Standard should be conducted in its operatingstate of operation simulation as a general rule. Do inspection, performance test, etc.
before and after these tests to confirm that no degradation has occurred.
Related considerations• The acoustic test is applied to a panel and other exposed equipment, etc.
• ･ The acceleration test is intended for verification concerning the strength andrigidity requirements; however, the acceleration-induced load conditions seldom
become critical. In the power subsystem, there is no case where a test is actuallyconducted, except for the battery element, (to evaluate the effects of acceleration on
the electrolyte). Of the effects on the electrolyte, the directional steady acceleration
component is more important than the vibration acceleration component consideredin the load conditions.
By vibration and impact tests, confirm, besides the environmental resistance, forequipment having relays, slip rings or other movable parts built in that there is noabnormality in the power subsystem due to chattering, contact reversal, etc.

d.

Radiation and exposed part environment (excluding acoustic)
As for radiation resistance, verify the design validity at the equipment level by the
analysis results based on existing data at the device and material levels.
Also for evaluation of the effects of the exposed part environment, verification by a testat the equipment level is difficult in many cases. So, conduct an effect evaluation by
analysis for each influential element (effects of thermal radiation from the fairing, free60

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molecular flow heating, heating by thermal radiation and jet from the rocket, etc.,
atomic oxygen, meteoroid/debris, sunlight radiation, terrestrial radiation, plasma andthruster plume gas around the orbit, etc.) and, if necessary, verify the validity of thedesign by an element test.
(6)

Safety design
Review the safety analysis results and check that the measures indicated by the results
are reflected appropriately in the design, work environment, fixture, relevant equipment,
procedure, etc.

(7)

Reliability design
The main work is a review of reliability, FMEA and analysis results; however, themeasures against a single failure are subject to confirmation redundancies and,depending on their content, to an in-process inspection, etc. For redundancies, theintegrity of all should be confirmed before the system delivery as a general rule.
However, for a redundant design, etc. at the circuit level whose direct confirmation in
the system is difficult, make an indirect confirmation by trend assessment or the like.
Worst-case analysis is subject to a review; however, an evaluation using circuit analysis
software or trial fabrication model is also made at the design stage. Devices, materials,
process selection, stress analysis, and derating are subject to a review, and for thedevices, materials and processes, whether they are implemented and managed
according to the selection is subject to the in-process inspection. In an individual
qualification test, review the test plan (method and decision criteria) and confirm thatthe expected results can be obtained. In a trend assessment, do a preliminary review of
the intended items and a confirmation of the assessment results in each test. For alimited life/operations article, make a preliminary review of the intended items andmanage the cumulative time/number of times in each test. For ground maintenance ofbatteries, carry out a review of the validity of its method, an inspection of maintenancecondition, and a performance test after the maintenance (before mounting).

(8)

Maintainability design
Confirm in the design review that actions and preparations (procedures, etc.) as per thedesign are ready. The workability of battery replacement should be confirmed in thesystem EM phase or system PFM.

5.3

Power supply capacity
Verification of power supply capacity is to verify that the power supply capacity satisfies therequired power profile during the entire mission period, and this is done by confirming thepower supply capacity required from power budget analysis of the system.
Verification of power supply capacity is done in the development phase and by thefollowing approach.
• From conceptual stage to design stage: Review of design results• Fabrication and test stage: Review of fabrication and inspection results of eachcomponent

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5.3.1 Required power profile
Based on the examination and test results, etc. in each stage, update the table of supply
power by operation modes and the operation plan to keep the profile up to date and confirm
that the design requirements are satisfied.
In the fabrication and test stage, the source data of the table of power supply by operation
mode is updated with the measured values of power consumption of the component.
Moreover, in a system test, or a simulation test and thermal balance test of the operation
modes in particular, the supply power and required power profile by operation mode are
updated in a comprehensive way. However, there are limitations to a simulation of operation
modes because of restrictions, etc. on a ground test, so the power consumption of themission subsystem, ion engine, attitude control subsystem, thermal control heater, etc.
cannot be set in a ground test as assumed on orbit. Thus, make an appropriate correction tothe obtained data by analysis.
5.3.2 Required size of power source
The hardware characteristics for the required size of a solar array and battery are verified in
4.5.1 and 4.5.2. The specifications for required size are set by power budget analysis to alevel satisfying the final required power profile in the design stage. Therefore, if the
hardware fabrication and tests proceed as designed and it is confirmed that the required
power profile and the requirements for a solar array and battery are satisfied, this establishes
verification that the power subsystem can supply power to achieve the spacecraft mission.
However, the source data (power consumption, model variables, etc.) will be updated withthe progress of the development; therefore, it is necessary to appropriately confirm that thepower budget is kept in balance. Therefore, manage the power budget analysis model and its
source data continuously in and after the design stage of development.
Related considerations
A policy, etc. of coping with the case of severe operation as to power budget is shown
below.
In such a case, when examining the depth of discharge or stored energy recovery state of abattery, it is desirable, considering the limitations to power budget analysis, to conduct anoperation simulation test (see Clause (3) of 4.5.2) on the battery to obtain data for that.
(1)

Initial stage of launch
In the case of a paddle-shaped solar panel, the power budget is generally severe in the
period from launch to paddle deployment and sun orientation. As a measure, scaling up
the panel and battery is conceivable, although this results in a mass increase. It isconceivable to increase the depth of discharge of the battery in this period; however,what to do first is thorough power saving of the entire spacecraft in this period (byreducing the power consumption of equipment, etc., putting unnecessary equipment in
the non-operating state, or otherwise).

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JERG-2-214(2)Contingencies
At an abnormality covered by the protection function, prediction of the depth ofdischarge of a battery takes place also by power budge analysis. In such a situation,
there are limitations to the prediction of panel power generation, load condition,
duration, etc. So, thorough energy saving of the entire spacecraft should be pursued as
far as possible.

(3)

In case where balancing the budget of one orbital revolution is difficult
While there is sunshine, a battery discharge does not continue in general and is
negligible in most cases. For a LEO spacecraft, stored energy recovery during one
orbital revolution (charge completion while there is sunshine) is difficult in some cases.
If this state continues, the stored energy in the battery will continue to decrease, so asole measure is to enlarge the solar array size except for power saving. If this state does
not continue and the stored energy is recovered after several orbital revolutions, adesign subject to this may be judged to be valid (considering the effect on the life) Even
for an LEO spacecraft, a significant battery discharge takes place even in the sunshine if
there is necessity for power supply to an ion engine or other large load.

5.4

Power supply characteristics
The requirements related to power supply characteristics are verified by a design review
based on analysis, etc. and by a system test, and the element unit tests and power subsystem
combination tests are done for the purpose of judging that a power subsystem component
may be subjected to a system test. Here, paying attention to a system test as final
verification, overall points to consider are described.

(1)

Reference point/measurement method
In a system test, monitoring items are restricted, and a precondition is that the methodof monitoring the characteristics subject to verification has a sufficient accuracy.
Therefore, a review and confirmation of reference points, voltage monitoring method,etc. should be done in advance. Because detailed evaluation is difficult with telemetry
data alone, provide a hardware line monitor terminal, etc. using a test connector.
Because the power subsystem is a subsystem operating at all times, AGE that cancontinuously monitor and record acquired data from the power start-up at the beginningof a test to the power shut-off at the end of the test should be prepared. Besides voltage,current monitoring is desirably possible for its evaluation, etc. (by clip-on or otherwise).

(2)Power subsystem operation modes/simulated power source
Conduct a power supply characteristic verification test for all operation modes of thepower subsystem in 3.2.1 and under the conditions in combinations including the
expected maximum and minimum of the solar array output characteristics, battery's
state of charge or depth of discharge, and supply (load) power. For this purpose, prepare63

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a solar array simulator and battery simulator that can set the expected solar array output
characteristics and a battery’s charge and discharge (voltage) characteristics. Thesesimulated power sources are desirably those which were used for a unit test, etc. For thebattery, conduct a series of tests on an actual one, besides using the simulator. For thesolar array, it is desirable to conduct a test for the purpose of confirming theconnections, etc., using a simulated sunlight source, lamp, etc. For the load, designate
in advance the conditions to evaluate the supply power characteristics based on thetable of supply power by spacecraft operation mode. On this occasion, it is necessary to
examine and prepare in advance the heat rejection measures of large power equipment
and thermal control heaters during a ground test.
(3)

Trend assessment
The verification is to confirm that the specifications concerning power supply
characteristics are satisfied. However, the specification itself assumes the situation
where an abnormality of a redundant part, or a trouble of a battery cell, etc. had
occurred.
Therefore, it is impossible in some cases to judge the inside to be normal even within
the specification. Thus, the test should designate bus voltage or other power supply
characteristic as a trend assessment item and evaluate the reproducibility ofcharacteristic behavior in all test processes.

5.4.1

Bus voltage
The specifications in 3.4.1 assume the state of actual operation. So, confirm in thepertinent operation state that the bus voltage data remains/varies as designed and there
is no abnormality in the trend.

5.4.2

Voltage stability
In a state where the active bus voltage stabilization function operates, measure the bus
voltage under the most favorable/adverse operating conditions to confirm that the
specification is satisfied and there is no abnormality in the trend.

5.4.3

Ripple and spike
Measure the ripple and spike in each operating mode to confirm that the respective
specifications are satisfied and there is no abnormality in the trend. On this occasion,
besides the ripple component generated by the power subsystem, the reflected ripple
component from the load side is also intended. So, obtain source data in advance toevaluate these components. It is not practical to make an evaluation in all test
conditions. Set the condition of the solar array, battery, and load (for example, acondition for large ripple) in advance and evaluate the data of that case in detail. To
measure ripple and spike, a band of 500 MHz or more is required, so it is impossible toevaluate them with telemetry data.

5.4.4

Output impedance

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This specification intends the output impedance of the power subsystem proper and is aconfirmation item in a power subsystem combination test. So, it is deleted from an
confirmation item of a system test. However, it is also possible to incorporate the powersubsystem at the time of system integration and conduct a system test in theconfiguration with no load connected.
Stability, ripple, and transient characteristics can serve for indirect evaluation andconfirmation of output impedance.
5.4.5

5.5

Transient characteristics
Measure the bus voltage transient at the transient operations (paddle deployment, load
connection and disconnection, peak power supply, sunshine/eclipse and in-between
transient, etc.) imposed on the power subsystem according to the spacecraft operation toconfirm that the corresponding specification is satisfied and there is no abnormality inthe trend. On this occasion, set certain conditions in advance as with ripple and evaluate
the data of that case in detail. The transient at a load abnormality or an internal
abnormality of the power system is generally deleted from a confirmation item of asystem test as with output impedance (if otherwise, it is necessary to prepare a
dedicated line for the system harnesses and test connectors) and is taken to be aconfirmation item of a power subsystem combination test and unit test.
Power subsystem

5.5.1 Interface with solar paddle subsystem
Because of restrictions of AGE, it is very difficult in general for the solar paddle subsystem
to verify by a test the consistency of the power subsystem interface design in a configurationin combination in the system. Therefore, verify the satisfaction of the requirements by the
results of an analysis designed based on existing data, results of a unit test and inspection of
the two subsystems, consistency confirmation of the two subsystems’ ICD and otherdrawings, operation confirmation on the simulated solar array power (solar array simulator),etc.
A verification example of the interface design requirements in Section 4.5.1 is shown below.
(1) Interface reference point
Consistency confirmation of two subsystem’s ICD(2) Maximum voltage and maximum current at interface reference point
Confirmation of solar paddle subsystem irradiation test results and analysis.
(3) Round-trip line resistance between interface reference point and PCU
Confirmation of power subsystem (system harnesses of power lines) unit inspection
results.
(4) Shunt operation sequence
Two subsystems’ ICD consistency confirmation and confirmation of shunt operation
sequence by operation on simulated solar array power.

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5.5.2 Battery
For the requirements for a battery, perform a verification by unit tests, system tests, etc.,
except for the life-related characteristics. For capacity and some other characteristics,
however, evaluate the trend in voltage, etc. because these cannot generally be evaluated by asystem test. Life-related verification is made by a battery and battery cell design review
based on related past records and preceding test results, confirmation of fabrication and test
results, and (if specified in the battery specification document) quality confirmation test
results of shipment lots.
(1)

Configuration
Besides ordinary inspection of appearance, dimensions, mass, etc., confirm theinsulation between the battery cell case and battery cell holding structure. Additionally,in relation to insulation, measure the voltage between the case and battery cell terminals(positive/case and case/negative) to confirm it to be proper.
For the temperature difference in the battery, confirm in a thermal vacuum test that it is
kept under the allowable value when under specified charge and discharge conditions.

(2)

Ground initial characteristics
Under the specified conditions of charge and discharge current, time, temperature, etc.,
confirm by a test that the requirements for initial charge and discharge characteristics
are satisfied. And, confirm that there is no abnormal behavior, etc. in the characteristics
related to the short-cycle characteristics required as a characteristic trend.
For example, reproducibility exists in the characteristic results of an initial test and final
test under the same conditions.
The characteristics of a battery cell may be influenced by its history. Therefore, keep
the conditions before and after the test as constant as possible. If the test is suspended,
give the same treatment as for storage.
In a unit test, measure the voltage of each battery cell to evaluate the matching quality
and its trend of the characteristics of the battery cells making up the battery.

(3)Life prediction
Review the results of design analysis (similarity) and preliminary evaluation to confirm
that the life requirement is satisfied.
Confirmation that the battery cell used has the same characteristics as the battery cell
subjected to the preliminary evaluation is made by evaluating the battery cell’sfabrication process data, delivery test results, and (if required in the battery cell
specification document, etc.) lot quality confirmation test results.
When conducting an operation simulation test for preliminary evaluation, make aspacecraft actual operation plan and a test plan to simulate charge control to evaluate
the validity of the entire operation of the battery. In particular, it is desirable not to limit
to charge and discharge operation in a stationary orbit but to incorporate the charge at
the launch (the temperature condition is generally severe), transient operation in a

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transient orbit (the depth of discharge tends to go deep until paddle deployment),
contingencies (confirm required energy supply at an abnormality/evaluate the validityof battery charge control during steady-state operation/estimate the restoration period)and confirmation of the effect of battery cell refresh (*1) or reconditioning (*2).
In power budget analysis, for an operation, etc. in which the budget balance of one
orbital revolution cannot be ensured, there are limits to analysis of the battery behavior,
which therefore should be evaluated by an operation simulation test.
*1: Battery cell refresh
A lithium ion battery has no saturation voltage (the voltage is not constant atcompletion of charge) unlike conventional NiCd and NiH2 batteries. Therefore, if an
imbalance of the state of charge occurs among the battery cells connected in series,
this variance remains unchanged and the series-connected battery will be restrictedby this varying state of charge of the battery cells. Battery cell refresh is a function
to recover the battery capacity by equalizing the battery cell voltage, and theequalization takes place by a bypass circuit connected to, and in series with, eachbattery cell to make the state of charge even among the battery cells.
*2: Reconditioning
NiCd and NiH2 batteries have a property (memory effect) in which, if charge anddischarge are repeated at a level before reaching a complete discharge, it becomes
impossible for discharging to take place exceeding that amount of discharge eventhough there is a margin of capacity. Reconditioning is a function to prevent this andrecover the amount of discharge by periodically deepening the amount of discharge
to activate the electrodes.
(4)

Sealing and pressure-proof characteristics
If possible with the battery, conduct a leak test of the electrolyte and an air-tightness
evaluation trace (He) gas. Confirm related data at a delivery test and process data
related to the sealing/pressure proof characteristics. If a leak test is conducted on thebattery, it is desirable to employ the same conditions as for the battery cells.
The pressure-proof characteristics should be verified with a battery cell proper.

(5)

Battery redundancy
In a battery-redundant design, etc., verification should be made on all redundancies as a
general rule to confirm the integrity of the batteries. If, however, a trouble, etc. of a
specific battery cell cannot be verified, the verification should be made by analysis.

(6)

Safety verification

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For a lithium ion battery, do the following.
Safety verification of a battery should take place, as basic verification, by managing allprocesses from battery cell fabrication to battery test completion so that they are done
appropriately. Because a direct safety test on a payload article can damage that article,
conduct such a test on a lot sample of the battery cell. Because a safety test of a battery
cell can endanger the operating personnel and neighboring equipment, conduct the test
while paying due attention to each risk.
(a) Safety of battery cell
As a safety test of a battery cell, conduct an external short-circuit test and overcharge
test on each type of battery cell for safety verification. For the safety valve (rupturepart) of the battery cell, do a test for each production lot of the rupture part. By thisrupture test, confirm the design operation pressure.
(b) Safety of battery
For the protection function for battery overcharge, make an operation confirmation of
the protection circuit not only in its proper state but also in the its incorporated state in abattery.
(c) Transportation safety
The battery transportation test is conducted on the dedicated battery packing and is todrop it from 1.2 m. The following requirements should be satisfied.
[1] The battery and battery cells inside are not damaged.
[2] The battery and battery cells inside does not form an external short-circuit.
[3] The battery and battery cells are not forced out.
(d) Parallel connection protection
For the protection circuit for parallel connection of batteries, make a confirmation, orverification by analysis, of its operation.
(e) Others
For the means provided for battery protection, do a test in the respective subsystems(temperature control and voltage control).

5.5.3 Power control
Design related to power control is verified in a comprehensive way in 5.4 Power Supply
Characteristics. However, there are some items whose verification is difficult to make with asystem test. Here described is an overview of the overall confirmation items in the unit test
and combination test performed before the system test and the evaluation items in thesystem test.

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5.5.3.1 Unit test
Confirm that the characteristics of each element are within the required specification value
or design specification value. In particular, a comparison between the measured andpredicted effect of the variations of device characteristics in a temperature test and aconfirmation of characteristic reproducibility of EM/FM under the influence of variance ofdevice characteristics are effective for validity evaluation of the design. In spite of device
characteristic degradation, etc. due to radiations, etc., the required specifications have to besatisfied, and verification for that is made by a review of design analysis results.
Because the simulated solar array power source, battery simulator, external start-up powersource, etc. are used also in combination tests and system tests, make a plan so that theeffect of the difference in test configuration will not pose a problem in subsequent tests.
Redundant design confirmation at the circuit level, transient evaluation at an internal single
failure or load abnormality, etc. are items for which it is difficult to do a system test. Thus,
do a unit test for all such items.
The electromagnetic compatibility test becomes complete by a unit test alone in some cases(especially for FM), so a characteristic comparison with the EM, etc. whose compatibility
with the system was confirmed is important.
For data belonging to a trend evaluation item in a system test, obtain source data for theevaluation, and obtain calibration curves for all telemetry data.

5.5.3.2 Combination test
Although there is a difference from a system test because no system harness or actual load isconnected, confirm by a combination test that the comprehensive operation (power control,in particular) of the power subsystem is in accordance with the design, the required power
supply characteristics can be ensured, and a transition to a system test may be made. On thisoccasion, make an evaluation that the results of a unit test of each element are reproduced
and do management of the trend items and calibration curves.
In a combination test, confirm the monitoring method of comprehensive operation andevaluate the data behavior of voltage, current, status, etc. for all operation modes of thepower subsystem in 4.2.1 and at power-on and power-off to obtain source data forevaluating the power supply characteristics (see 5.4, (2)) in a system test.
In a temperature test, confirm that the condition setting of equipment in combination is inaccordance with the prediction of actual operation.
5.5.3.3 System test
Mainly make a confirmation of the power supply characteristics in 5.4, and evaluate thecomprehensive operation of the power subsystem based on the combination test results.

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Control system
By comprehensive operation confirmation by a combination test and system test,
confirm that solar array output control, charge control, and discharge control operate
systematically, the operation range, priority and coordination management of control
agree with design, and the functions related to operation modes are operating normally.
For parallel operation of multiple buses or multiple batteries, confirm that current shareis appropriate.

(2)

Operation point analysis
Along with verification of the control system, evaluate the behavior of operation
monitoring data of bus voltage, solar array output current, load current, etc. to confirm
that the operation point of each control operation agrees with the design prediction.
For lock-up and other operating point abnormalities, confirm the identification method
and measures (automatic restoration) by load reduction, etc.

(3)

Control diagram
Confirm that the power supply characteristics by a unit test, combination test andsystem test agree with the characteristics predicted from the control diagram. Although
the prediction and measurement (of dynamic characteristics in particular) may notnecessarily agree depending on the characteristics of the simulated solar array powersource (including the test cable), load, etc., evaluate that the basic behavior does notdiffer from the prediction to confirm the validity of the control diagram. Also confirm
that an abnormal shift of the operation point does not occur due to load variations, solar
array output variations, a shift of the state of charge of the battery or other disturbances,and unstable vibrations, latch-up, etc. of control do not occur due to inappropriate
damping characteristics, parasitic oscillation, etc.
Because it is difficult to measure and simulate the dynamic characteristics of a solar
array, it is common practice to design so that the dynamic characteristics do not
influence the control characteristics However, if the degree of this influence poses a
problem for ensuring the control stability, etc., it is necessary to make an evaluation, etc.
of an analysis model for a partial panel. For the simulated solar array power source, usea power source having a small output capacitance, a higher speed of response thanpower control, and frequency characteristics close to the actual solar array.

(4)

Characteristics of each power control
The control characteristics are evaluated by a unit test, combination test and system test,
and design verification is performed for the items of (1) to (4). Here, points to consider
related to the test of each control element are described.

a.

Solar array output control
Prior to a test, confirm that the simulated solar array power source incorporates thesolar array output (V-I characteristics) prediction and can simulate the characteristic

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change from BOL to EOL in particular and the output ripple, etc. before paddle
deployment. The solar array output characteristics change greatly due to the rapid
temperature rise after an eclipse, and it is desirable that this situation can also be
simulated.
The monitoring means to keep track of the operation condition of output control should
be confirmed as a monitoring method of comprehensive operation of the powersubsystem. Besides this, if direct measurement of power generation of the solar array isdifficult, obtain and evaluate related data as an indirect measurement means of power
generation.
If the operation of the control element of solar array output control is switching (digital
shunt), shielding of the noise associated in principle with the operation cannot be
accomplished as with general equipment and, in addition, it is difficult to use an actual
solar array. Therefore, when conducting an electromagnetic compatibility test on otherequipment while solar array output control is taking place, it is necessary to examine
the test method or analysis means in advance.
b.

Charge control
When testing the charge control function of a battery, it takes a considerable time if an
actual battery is used. Therefore, use a test power source and, by making adjustable thesimulated signals battery voltage, temperature, etc. by a battery simulator and test
adjustment device, make a confirmation of the V/T level, change-over of charge current,
constant current/constant voltage charge control, bus voltage control function, heater
control, protection function, etc. The battery simulator is convenient for characteristic
evaluation, etc. of equipment and is used in a combination test and system test in somecases, besides a unit test. However, a simulator (including harnesses) does notcompletely simulate the characteristics of a battery. For a combination test before asystem test which uses an actual battery, it is desirable to plan to also connect an actual
battery.
For a spacecraft needing to confirm the efficiency (loss) in the power budget, measure
the power conversion efficiency paying attention especially to constant-current charge
and bus voltage control operations, in the above-stated charge characteristic
confirmation.
On this occasion, note that the efficiency varies with temperature, battery voltage andcharge current.

c.

Discharge control
As with charge control, make a confirmation of the discharge function and protection
function (UVC, etc.) using a battery simulator.
The blowing capability of the fuse at a load abnormality should be confirmed byanalysis or a test using a battery.
When confirming the efficiency (loss), note that it varies with temperature, battery
voltage and discharge current, as with charge control.

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5.5.4 Power distribution
The requirements for power distribution should be verified along with the confirmation of
power supply of the electrical interface of 2.2, (1) under System Test. Arrange the results ina power distribution management table and confirm that the requirements are satisfied.
In a unit test, besides confirming the distribution function (conduction, voltage drop,
turn-on/off, and monitoring), make a confirmation of the normality of built-in protection
functions (fuse and shut-off function) which are difficult to confirm by a system test.
Design involving the operating conditions of a fuse (current, time, etc.) should be verified
by evaluations up to the design stage.
For pyrotechnic ignition power supply, in the case where the number of operations is limited,
what to do with FM is, in some cases, only to confirm all functions such as conduction,
insulation and an ignition sequence using small current. In such a case and in an EM test,
etc., a function and performance confirmation using actual current (actual pyrotechnic)
should be made as verification of the design.
5.5.5

Protection

5.5.5.1 Overview
The power subsystem is required to operate at all times and its fault tolerant design is madeon the premise that its major functions cannot be stopped as fail-safe. Therefore, it isnecessary to make verification considering that verification of its protection function is noteasy.
In particular, when simulating the occurrence of an abnormality, it is necessary to take care
not to impose an unnecessary overstress on the devices, etc.
5.5.5.2 Protection in power subsystem(1) Redundancy(a) For a single-point failure, make design verification using FMECA, etc.
(b) For an abnormality or failure of a device, etc. due to an open-circuit or short-circuit,make design verification using FMEA at the device level and functional FMEA.
In addition, if a module or equipment takes a redundant configuration, conduct a test
assuming that one module or piece of equipment failed.
(c) If a protection function is to be performed in the same hybrid IC, make design
verification by the circuit diagrams, manufacturing drawings, etc. of the hybrid IC toensure that no chain of failures will occur.
(d) When the solar battery (simulated power source) starts output, confirm that it operates
without needing a special command.
(2) Insulation(a) Perform insulation measurements in units of components. Determine the applied
voltage considering the function of the component.

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JERG-2-214(3) Overcurrent protection(a) Using interface FMEA, etc., confirm that the primary power bus is protected by aprotection circuit.
(b) If a fuse is used to protect a primary power bus distribution wire, show the fuse’s rating
and, by analysis, the maximum current in the load path, including transient states (dueto in-rush current, etc.), and show that the rating of the fuse is appropriate. Perform
analysis of fuse blowing to verify that the bus voltage can satisfy the lower limit at anabnormality.
(c) If a fuse is used to protect a primary power bus distribution line, verify by analysis thatthe power subsystem can blow this fuse when the load is short-circuited within thespecified time determined in the Electrical Design Standard of each project.
(d) If a fuse is used to protect the primary power bus, confirm at the time of systemdelivery and at the component level that there is no open-circuit in the fuse. After
integration of the spacecraft, confirm as the need arises that there is no open-circuit in
the fuse.
(e) If a fuse is installed in the subsequent stage of a DC/DC converter, consider themelting characteristics of that fuse and verify by a unit test of the DC/DC converter
that the DC/DC converter has the ability to blow the fuse for certain.
(4) Battery protection(a) Perform a test or analysis which assumes that one battery failed.
(b) If an over-temperature occurs in the battery, charge suspension or other protection
function will operate. Verify this by a test. If a test battery is not available, simulate thebattery over-temperature state by a simulation signal to avoid an overload on thebattery.
5.5.5.3 Protection in cooperation with other subsystems
For all protection functions related to power supply, confirm them according to theoperation confirmation method by a ground test in the protection function summary tablespecified in Section 4.5.5.3. For confirmation of a function, make confirmation, etc. of
simulation signals and statuses to establish end-to-end verification from detection toprotection action in the realm of possibility.
Confirmation of a protection function about an item requiring cooperation with othersubsystem can only be performed systematically by a system test. Thus, specify the points intime of confirmation in advance and arrange the evaluation results in a summary table toconfirm at the end of the system test that all items have been tested and evaluated.
5.5.5.4 Protection for abnormality of ground equipment
Ground test equipment should be subjected to sufficient validation before used forverification of a spacecraft.

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Appendix-I Design Overview (as reference materials)
Of the electricity subsystem (power subsystem plus solar paddle subsystem) of a spacecraft,
this appendix mainly describes an overview of design of the power subsystem.
I.1

GeneralThe electricity subsystem of a spacecraft has the role of supplying required electrical powerat the prescribed voltage and other various characteristics without interruption over theentire mission period.
For this purpose, the electricity subsystem has the following functions:[Solar paddle subsystem](1) Power generation[Power subsystem](2) Storage of electrical energy(3) Power control(4) Power distribution(5) Protection: Limitation of failure or abnormality
Figure 4.2-1 in the text shows an overview of the electricity subsystem configuration. The
electricity subsystem has, besides the basic functions shown in the figure, atelemetry/command processing function.
Generation of electrical power is achieved by a solar paddle and storage of electrical energy
by a battery, and the size of these power sources determines the power supply capacity ofthe electricity subsystem. The power supply capacity must satisfy the required power profile
over the entire mission period of the spacecraft.
Power control consists of the functions of solar array output control and battery
management (charge and discharge control), and these determine the power supply
characteristics (voltage and other various characteristics). The power supply characteristics
must satisfy the electrical interface conditions related to power supply to the spacecraft
system.
Power distribution is the function of distributing power to each load in the spacecraft, and its
output end forms an interface boundary of power supply and is taken as a reference point of
power distribution characteristics.
The power subsystem has a power supply bus to cooperate the functions of the solar paddle,
battery and power control, and power is supplied from this bus. And, the voltage at thereference point of power supply characteristics is called primary bus voltage. As a powersubsystem configuration, the power supply bus consists of one line (single-bus) or multiplelines (multiple-bus), and in some multiple-bus configurations, the lines do not necessarily
have the same structure or characteristics and the bus voltage is set to multiple levels. For
power supply, besides power to the equipment, a secondary power bus is needed for
propulsion motors, paddle and antenna deployment, etc. and a pyrotechnic ignition powerbus is provided to supply power there from the battery through a dedicated distribution line.
Protection is a function to keep normal power supply. It suppresses propagation of failure or74

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abnormality not only inside the power subsystem but also of other subsystems, and shares
functions and coordinates the operation with other related subsystems.
The electricity subsystem occupies a large system share of mass, etc. and is closely related
to the (electrical, structural, attitude control, etc.) design of the whole system.
Development proceeds at the levels corresponding to the development stages of conceptual
design, basic design, detailed design, and maintenance design/fabrication and test. In the
meanwhile, a review is performed on the results of preceding study, investigation of current
technology, analysis, trial fabrication, tests, etc., and validity confirmation, revision andchange of the design are made based on the technical feasibility of the development target,
development time, risk, cost, etc. For a design change, it is necessary to consider its effects,
etc. on others. On and after the basic design stage, therefore, it is reviewed and treated basedon the configuration program plan.
Described below is an overview of major system considerations, major specifications of thepower subsystem (amount of power supply and power supply characteristics) and major
components when proceeding with power subsystem design up to the detailed design stage.
Figure I.1-1 classifies the necessary considerations in power subsystem design into System
Level Examination Items, Subsystem Level Examination Items and Major Component
Level Examination Items and shows their relationship.

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1.2.3 Required power profile

Major specifications of
power subsystem

System considerations

I.2.1/ I.2.4 Orbit/life requirement
Sunshine duration
Eclipse duration
Transient orbit
Stationary orbit
Life requirement• Mission period• Number of battery charge
and discharge cycle

Subsystem power consumption(mean/peak)Margin(existing/refurbished/developed)
Operation plan, modes, prediction ofheater power at abnormality (thermal
design)Eclipse operation of mission equipment
Launch (hot launch/cold launch)
Launch window analysis

Bus voltage
Single-bus/multiplebus
Power supply
methodMaximum supply power

Required storage current capacity, allowable depthof discharge

Conditions of use
Battery cell used
Assembly structure
Battery configuration
Energy density

Temperature control
Operation simulation
evaluation
Reconditioning
Life prediction
Ground maintenance1.2.5 Design restrictions (4.2.3-/5.2.2-3 in the test)

1.2.2 Satellite shape
Attitude stabilization method(triaxial/spin)
Sunlight incident angle to panel
Paddle tracking method(fixed/uniaxial/biaxial)
Paddle placement (shadow, interference,
flexibility and plume)
Paddle deployment sequence

I.3 Power supply capacity (4.3/5.3 in text)
Discharge power

I.5.1 Battery (4.5.2/5.5.2 in text)
Major components

Primary power bus(Section 4.2 in text)

Common specification and other requirements
Electrical, mechanical and thermal interface management (including
external interfaces)
Reliability, quality control and device and material process standards
Environmental resistance, test, maintainability and safety
System allocation
Mass, mass characteristics, reliability and power consumption
Telemetry and command items

Note: Connecting lines to show
relation type

Required power
generation

Bus voltage
Line drop
Stability (regulated/non-regulated) Power transmission loss
RippleImpedance
Output impedanceTransient characteristics

Directly related
Commonly related
Related by control
Related by system
harnesses

Charge power

See Solar Paddle Design Standard (JERG-2-215)
Environmental conditions
Solar cell used
Panel structure
Array circuit configuration
Specific power (per unit mass)

Deployment mechanism
Ground evaluation of initial
outputDegradation prediction
Handling
Interface between spacecraft

I.5.2-4 Power control, distribution and protection (4.5.3-5/5.5.3-5 in text)
Solar array output control method (ground test of control characteristics)
Battery charge and discharge control method (compatibility with battery)Distribution method (single-/multiple-bus, load, turn-on/off, monitoring)
Protection method (intended events, detection, emergency measures,
restoration)

Figure I .1-1 Overview of power supply design76

1.4 Power supply characteristics (4.4/4.5 in text)

I 5.6 Others
System harness
Pyrotechnic
ignition circuit
Umbilical
Test cableJERG-2-214

I.2

System considerations
The power subsystem of a spacecraft needs a design suitable for the mission requirements
for the spacecraft from the viewpoint of mass, efficiency, reliability, etc. These mission
requirements are basic conditions to determine the spacecraft shape (including the attitudeconditions), required power profile and life requirement.
Design examination of a power subsystem is closely related to system design concerning [1]orbit, [2] spacecraft shape and [3] required power profile and proceed in pace with thedesign of these.
To implement a system, [4] design restrictions are set on its subsystems and components,
and the design proceeds in compliance with these.
An overview of the basic conditions of [1] to [4] above is described below.
This design overview does not cover the power subsystem of a spacecraft for which
replenishment or recovery is intended. Therefore, such considerations specific to a system
are not described individually. With such a spacecraft, most operations are transient innature, and the required power and time, attitude conditions (extent of sunlight utilization),
power transfer to or from other systems, operations to avoid abnormality propagation(collision, impact outside the planned area, etc.), required level of redundancy (triple
redundancy, etc.), etc. in the mission period have a large influence on the design of thepower subsystem.

I.2.1 Orbit
An orbit determines eclipse conditions (time of occurrence and duration), sunshine duration,and sunlight incident angle to the orbit plane and dictates the environmental conditions of
the spacecraft main body and solar paddle and the charge and discharge conditions, etc. of
the battery.
Typical examples of orbits are the geostationary earth orbit (GEO) and a low-altitude earth
orbit (LEO). On the GEO, an eclipse of about 70 minutes at the longest occurs only in
20-odd days before and after the vernal and autumnal equinoxes, so more than 20 hours canbe taken for a charge and the state of full charge is kept for most of the mission period. On
the other hand, on an LEO, an eclipse of about 30 minutes occurs at every revolution aroundthe earth in most cases, and the charge time is one hour or so, and the number of eclipse
occurrences a year reaches about 5,000 or 6,000 in some cases.
The solar panel temperature changes rapidly in the transient period between an eclipse andsunshine. On the GEO, for example, the temperature falls from about 50°C during sunshine
to about minus 200°C during an eclipse.
In addition, the power generation output of a solar array degrades due to radiations, etc. inouter space, however, this situation depends on the orbit. An orbit is divided into astationary orbit for fulfilling a mission and a transient orbit up to a transition to thestationary orbit. For GEO, power supply on a transient orbit tends to depend on the battery.
I.2.2 Spacecraft shape
In general, the shape of a spacecraft is related to its attitude stabilization method, that is, thespin method takes a cylindrical form and the triaxial method takes a box-like form with adeployed paddle. And, the solar array is attached on the main body cylinder in the case of
the spin method and on the deployed paddle in the case of the triaxial method. Therefore,
the sunlight incident angle to the panel is determined by the spacecraft attitude in the case of77

JERG-2-214

the spin method. In the case of triaxial method, the placement of the solar paddle andtracking method are involved. The tracking method is generally uniaxial, however, thebiaxial or fixed method is also used. For a paddle-shaped solar panel, power is transmitted
to the spacecraft main body through the slip ring of the solar paddle drive mechanism.
In the case of the spin method, the dimensions of the panel generally coincide with those of
the spacecraft main body. Thus, the power generation requirement is restricted by the size ofthe spacecraft. On the other hand, few restrictions are imposed on the triaxial method
because of a deployed paddle. The placement, tracking method and shape of a paddle aredetermined by a comprehensive examination of the power generation requirement,
efficiency of the solar cell used (consider the temperature characteristics and degradation
characteristics as well), field interference with the sensors and antennas, shadow of thespacecraft main body, interference and contamination from the propulsion subsystem,
attitude disturbances, etc. In this process, the required dimensions, etc. of the solar paddle
are also determined.
For a deployed paddle, power supply depends on the battery in the period from launch to
deployment. Therefore, it is necessary to set the deployment sequence and the required
power profile during this period to maintain the depth of discharge of the battery within theallowable value.
I.2.3 Required power profile
The required power profile is determined by the component power consumption of eachsubsystem and the operation mode based on the spacecraft operation plan, etc. If therequired power changes in time many times, it is expressed as a time series of operations.
The power consumption intends mean power and peak power, contains a margin according
to the records of a subsystem, component, etc. (newly developed/refurbished/new), and is
updated with the measured value, etc. with the progress of the development stage.
For such an update of source data, the power consumption of the on-board equipment can bebased on the measured value of a test. However, for example, the heater power, etc. inthermal design of the spacecraft system is difficult to keep track of accurately even in a(thermal balance, etc.) test. Thus, such a quantity will have to rely on prediction eventhough the accuracy is improved from the design stage. In addition, at EOL when the powerbudget is severe, it is necessary to predict the required heater power (smaller than at BOL ingeneral) considering the deterioration of surface characteristics of the thermal blanket.
Therefore, for a system in which source data, like the thermal control heater power, relying
on a predicted value occupies a large proportion of the required power profile, how to deal
with the margin of a prediction becomes challenge.
Set the operation mode for the entire mission period, that is, for the launch stage, transition
stage and steady-state operation, and consider contingencies (abnormality of attitude, etc.)
as well. In the triaxial method, the solar paddle is deployed by the time of transition to astationary orbit, and the power supply until the deployment tends to depend on the battery.
Include such a transient operation mode as well.
The eclipse operation of mission equipment influences battery size determination.
Of the launch types, cold launch does not put the spacecraft in operation until the rocket
separation, so it reduces the burden on the battery. However, hot launch which put the
spacecraft in operation using the battery before the launch is generally employed. The
requirements concerning electricity in the launch stage are considered in launch window
analysis.

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I.2.4 Life requirement
Design and consider to satisfy the battery life (on-board life, charge and discharge cycle life,
ground storage life, etc.) required by the spacecraft mission requirements. Other life
requirement examinations should be in accordance with Clause (7) Reliability Design in Section
4.2.4 Common Design Standard of the text.
I.2.5 Design restrictions
For development of a spacecraft system, common restrictions (requirements) to comply with areestablished and set.
These matters are indispensable for proceeding the development on a full scale and are
documented in the form of a common specification or system allocation. However,determination of details consumes a considerable time for coordination, etc. between the systemand subsystem parties concerned.
The power subsystem and its constituent components are designed and fabricated in compliancewith these matters, and the design validity is demonstrated by a test or analysis.
(1)

Common specifications, etc.
The following common specifications, etc. are established in accordance to thearchitecture (example) of specifications and standards of each project.
a. Electrical Design Common Specification (Standard)b. Mechanical Design Common Specification (Standard)c. Thermal Design Common Specification (Standard)d. Electromagnetic Compatibility Common Specification (Standard)e. Environmental Resistance Design Standard
f. Interface Management Specification
g. Test Common Specificationh. Device, Material and Process Standard
i. Maintenance Program Standard
j. System Safety Standard
k. Reliability Program Common Specificationl. Quality Assurance Program Standard
The power subsystem has interfaces with other subsystems for power supply throughsystem harnesses, and the power supply characteristics are incorporated in the common
specifications for electrical design and electromagnetic compatibility. The grounding linesof the power lines of the spacecraft system are managed by the common specification andgrounding is taken at one point in the spacecraft structure.
The constituent elements (components) to implement the required function of the powersubsystem are electrically connected by the system harnesses in general. Therefore,
although management of the relevant harness takes place in the system, it is necessary toincorporate the characteristics of the harness in the design of the power subsystem.

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Power subsystem components involves various types of heat generation and absorption
when they do an energy conversion, storage and control. Therefore, for maintaining theirtemperature within a range to ensure the prescribed performance and characteristics, thethermal interface with the spacecraft system is an important design and management item.
The power subsystem protection function plays the role of preventing anabnormality/failure of some subsystems (including the power subsystem) from propagating
throughout the spacecraft through the power bus as far as possible. Assumption, preventivemeans, etc. of intended abnormalities/failure are designed and managed from the viewpointof the spacecraft system, and function sharing, etc. are set.
During a ground test or launch, the power subsystem receives charge power for the battery
from AGE and power, etc. from a simulated solar power source through a dedicated line (ofumbilical cable, etc.). This interface must be appropriately designed and managed
considering its purpose.
(2)

System allocation (design budget)
System allocation is an important management item for the success of a mission, and it isset concerning mass/mass characteristics/power/reliability/telemetry and command items,
etc.
The target values should be appropriately allocated to each subsystem/component andmanaged based on the system design. The maintenance state of system allocation isreported in progress management.
When changing a set value, examine all influences on others. For example, increasing a
power-related set value will change the required power profile, which in turn makes it
necessary to reevaluate the compatibility with the power supply capacity of the powersubsystem. If incompatible, an increase in mass, a change to the operation plan, etc. may be
involved.

I.3 Power supply capacity
I.3.1 Overview
The mean value of power supply capacity of a power subsystem is roughly determined by thescale of the solar array and battery and from the following relation.
[Power supply capacity during sunshine] ≈ [Solar array output power] - [Battery charge
power][Power supply capacity during an eclipse] ≈ [Battery discharge power]
Here, the charge power is necessary for recovering the discharge power of the battery. If energy
recovery does not take place appropriately, the stored energy of the battery gradually decreases
and it becomes impossible to ensure the power supply capacity during an eclipse, possibly
affecting the spacecraft operation. The discharge power of a battery is restricted by the total
stored energy of the battery {= (stored energy per battery unit) × (number of battery units)},
allowable depth of discharge and discharge time.

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When in sunshine as well, the peak value of power supply capacity is given as the sum of thesolar array power generation and (allowable) discharge power
From these relations, a rough value of power supply capacity is determined by the size (power
generation, stored energy, etc.) of the power sources (solar array and battery). The fundamental
requirement for power supply capacity is to satisfy the required power profile over the entire
mission period. Power budget analysis is performed to evaluate the compatibility with thisrequirement.
I.3.2 Power budget analysis
In power budget analysis, evaluate the compatibility between the required power profile and thepower supply capacity of the power subsystem, that is, verify that the stored energy of thebattery is appropriately maintained with respect to the required power profile.
Figure 2.3-1 shows an example of a power subsystem model for power budget analysis(example of main array type). In the analysis, the parameters to consider include:
PS: Solar array output powerηS: Transmission efficiency of generated power
CS: Instantaneous stored current capacity
ηC: Charge control efficiency
of battery
ηB: Energy efficiency of battery charge anddischarge
PE: Power consumption of powerηD: Discharge control efficiency
subsystem
PL: Load powerηL: Transmission efficiency of power supply
N: Number of battery power lines
Solar array

Load

Power
subsystem

ChargeDischarge

Battery

Transition equation for instantaneous stored current capacity CS (Δt: Time intervalof calculation)
In theBattery charge power = [(PS • ηS)-(PL/ηL)-PE] •ηCsunshine= Limit value by charge control: in case where the value ofthe above equation is larger than the limit valueCs (t+Δt) = Cs(t) + ηs • (Battery charge power) • Δt/V_bat(chg)Note: In case where no discharge is taking place
DuringBattery discharge power = [((PL/ηL) + PE]/ηD)/N]an
Cs(t+Δt) = Cs(t) – (Battery discharge power) • Δt/V_bat(dischg)
eclipseNote: V_bat(dischg) denotes the mean battery discharge voltage.
Figure I.3-1 Example of power budget analysis model81

JERG-2-214

The definitions of these parameters are described below.
PS:

Solar array power generation Because solar batteries degrade by radiation, etc., thepower generation requirement at the beginning of life (BOL) is set from the power
generation requirement at the end of life (EOL) based on a degradation prediction anddetermines the size of the solar array. PS varies due to panel temperature, sunlightincident angle, shadow from the spacecraft body, etc., besides long-term degradation.
The surplus power generation of PS is treated by output control of the solar array.

ηS:

Transmission efficiency of generated power Set considering power transmission loss
due to the system harnesses, etc. from the solar array to the power subsystem.
Value corresponding to power conversion efficiency, etc. of charge control
Value corresponding to power conversion efficiency, etc. of discharge control
Amount of discharged current (Ah) divided by the minimum amount of charge current(Ah) necessary for that recovery. In particular, the charge energy after completion of acharge becomes heat in the battery, so it does not contribute to an increase in stored
energy.
The upper limit of CS is the total amount of stored electricity CB determined by battery
capacity, discharge voltage and the number of battery units. This CB determines thesize of the battery.

ηC:
ηD:
ηB:

CS:

PE:
ηL:

PL:

Is the instantaneous stored current capacity (Ah) and can be called remaining current
capacity (Ah). There is no data that directly represents this value and it is an analyticalparameter. The depth of discharge is given in percentage of the expression (CB CS)/CB) from the CS at the end of discharge. The allowable depth of discharge is set
considering the file requirements for the battery. In relation to life and depth ofdischarge, operation at a nearly 100% depth of discharge is not common considering an
insufficiency of solar array power generation if an attitude abnormality, etc. occurs
immediately after an eclipse.
Steady-state power consumption of power subsystem itself.
Efficiency of power supply to load. Set considering the power transmission loss due tothe system harnesses, etc. from the power subsystem output end to each subsystem orcomponent.
Load power. Specified by the required power profile and varies with spacecraft
operation modes, etc.

Power budget analysis is usually performed by a power subsystem simulation program built
up on a computer. In a power budget analysis using a computer program, make time series ofPL, PS, etc. and calculate the instantaneous stored current capacity CS (Ah) or the depth ofdischarge. And, if this value is kept within the allowable value, the power supply capacity is
judged to be compatible with the required power profile. The simulation of a powersubsystem in Figure Ⅰ.3-1 is the simplest one and hardly considers the voltage-current

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characteristics of the solar array, bus voltage behavior, change and discharge characteristicsof the battery, the effect of temperature on these, and the like. To deal with this, set andupdate the parameters to appropriate average values by hardware test results and evaluation
repetition. In another method for this purpose, a power subsystem operation analysis program
incorporating the hardware characteristics as much as possible is applied for the powerbudget analysis considering the following items.
••••

V-I characteristics of solar array (consider temperature characteristics, radiation
degradation, etc. as well.)Electrical characteristics including state of charge of battery (consider temperature
characteristics, aging, etc. as well.
Electrical characteristics of each piece of equipment of power subsystem (solar array
output control, charge and discharge control, etc.)
Load characteristics (constant power, constant current, resistance)

For example, in a constant-voltage charge, which is one of the battery charge methods, thecharge current decreases exponentially and this change in time depends of the characteristicsof the battery. In addition, to analyze the change of stored current capacity CS (Ah) in time,
which is a key parameter for power budget analysis, it is necessary to consider the charge anddischarge energy efficiencies as well, besides the correlation between this charge control andbattery characteristics. However, modeling of the charge characteristics, etc. of a battery for
operation analysis involves many problems, and improvements through accumulation of
application experience or otherwise are required. In addition, a comprehensive operation
analysis program is conceivable for the spacecraft system, including the powersubsystem-related operations, etc. of the thermal control subsystem. However, to maintainand develop such a program effectively, a repeated improvement of modeling throughapplication experience is required. Considering this, a long-term policy to use the same typeof spacecraft system a considerable number of times is a precondition.
For reference, an example of power budget analysis is shown in Figure I.3-2.

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Elapsed time(minutes)

Interval(minutes)

Item

Acquisitionof earth

PaddlepresetShadeStart of sun-tracking of
paddleEnd of paddle stretch

Into shade

Start of paddle stretch

End of boom deployment

Satellite separation
End of rate dump

Configuration
at launch

Rate dumping

Attitude control

Paddle

Mode

Event

Boom deployment

Battery mode transition
LaunchShade

Triaxial stabilization
establishedYawacquisition

Paddle stretch

Hold

Sun trackingPowerconsumption

Power
consumption

Figure I.3–2 Example of power budget analysis (sun-synchronous sub-recurrent orbit)

I.3.3 Related tasks in each development stage
Setting of required power supply capacity, namely, of the required size of the solar array andbattery, is an important item in spacecraft system design. Power budget analysis is a means toevaluate power supply capacity and its compatibility. Here, tasks related to the analysis in eachdevelopment stage are described.
(1)

Conceptual design stage
In this stage, a rough framework of the system to achieve the mission is examined. The
greater part of the system considerations related to the power subsystem are indeterminate
and are subject to trade-offs. For the required power profile, orbit conditions and power84

JERG-2-214

consumption level are assumed. As power supply capacity, rough sizes of the solar arrayand battery are subjected to a case study considering the records, current state of technology,
results of previous studies, etc. In this stage, a rough value of the mean power supply
capacity described in I.2.1 Overview is discussed, but a full-scale power budget analysis is
not performed. On this occasion, besides development period and risk and cost, the requiredmass, dimensions, etc. associated with power supply capacity become evaluation criteriawith respect to performance.
For reference, for the mean on-orbit load power PL, take the allowable depth of discharge as
DODA, the eclipse duration as TE, then the required total stored current capacity CB (Ah) ofa battery is as follows:
CB={[(PL/ηL)+PE]/ηD } • (TE/DODA)}/V_bat (dischg)
Taking the sunshine duration as TS, the required mean charge power PC is:
PC={([(PL/ηL)+PE]/ηD)/V_bat(dischg)} • {(TE/TS)/ηB} • V_bat (chg)From these, a rough value of the required power generation PS of the solar array becomes
as follows:
PS={[(PL/ηL)+PE+(PC/ηC)]/ηS
From PC and CB above, the required size of the solar array and battery is roughly set
considering characteristic degradation and redundancy.
(2)

Basic design stage
In this stage, the baseline of the system is determined based on trial fabrication tests of trial
models, etc. For the power consumption of each component or subsystem, which acts as aload, rough values exist but are indeterminate. Nonetheless, a power budge analysis withrespect to a required power profile is conducted based on such an indeterminate power
consumption (containing margins) and operation plan, to set the required value of power
supply capacity.

(3)

Detailed design stage
In this stage, through fabrication and test of a development modes, etc., the power
consumption of a component settles with a certain error range and the required power
profile containing an operation plan become definite.
In parallel with this design progress, power budget analysis is performed at appropriate
times and the compatibility between the previously set power supply capacity and required
power profile is evaluated. If not compatible, the required size of the solar array and battery,allocation of power consumption, operation plan, etc. are revised. However, the freedom ofrevision is restricted as the design proceeds. In particular, the required size of the solar
array and battery exerts a large impact, etc. on the mass, and in the case of a panel, on the85

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system shape, etc. In general, therefore, a large revision is difficult. The parameters of
power budget analysis are updated to reflect the progress of the design, etc. of the relevant
part.

(4)

Maintenance design/fabrication and test stage
In this stage, the specifications for each individual component have been determined, andevaluation of management of their feasibility are performed. To reflect the characteristics ofactual hardware, the source data (required power profile, analysis parameters, etc.) involvedin power budget analysis are updated. The power budget analysis in this stage has the
purpose of evaluate the impact of these updates. A power budget-related margin isincorporated in the design of each component in the design stage, and these margins areaggregated in this stage as a system margin. In this stage, if the required power supply
capacity is not compatible with the required power profile, the only measure that can betaken is a revision of the operation plan in ordinary cases.

(5)

Evaluation stage
Major tasks related to power supply capacity in the stage toward hardware development are
as follows. The power budge analysis used in this process is utilized also in the subsequent
launch, tracking control and evaluation stages and forms a part of database of the actual
operation plan. On this occasion, the final parameters in the fabrication and test stage areused for analysis and, in addition, by updating them or otherwise, application records ofanalysis are accumulated and utilized for the next development.

I.4 Power supply characteristics
The power supply characteristics of a power subsystem include bus voltage, stability, ripple,
output impedance and transient characteristics. The reference point of these characteristics aretaken to be the output end of power distribution.
I.4.1 Bus voltage
For the bus voltage, about 30 VDC has been used conventionally. This is partly because thesupply power was small but mainly because the design platform, including applied devices, wasthe 28 VDC power system of aircraft, etc.
With the increasing size and power usage of spacecraft, it is possible indeed to rely on theconventional bus voltage. However, to suppress a mass increase of the power supply harness(including the wiring on the solar panel), etc., the bus voltage was raised. In several-kW
spacecraft, a voltage of about 50 V is already used. For 10 kW or so, 100 V or higher is thought
to be efficient.
When raising the voltage, the following items must be examined.

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Restrictions of applied devices
Restrictions of devices are a problem involving the design platform. So, attention mustbe paid to that the problem involves the design not only of the power subsystem but
also of all subsystems supplied with power. In particular, with respect to withstandvoltage and switching durability, power semiconductors, relays and capacitors may be
used (a measure by using them in parallel is possible but causes an increase of mass andvolume). Their availability and applicability must be confirmed in advance (consider
alternatives, etc. if circumstances require).

(2)

Paschen’s dielectric breakdown law and interaction with outer space plasma
environment
The breakdown is caused by the presence of residual gases or an outer space plasma
environment (of coexisting electrons and positive ions of 100 keV or less) which iselectrically conductive. With the interaction, in particular, a problem occurs in a low
orbit where the plasma density is high.

(3)

SafetySafety is a matter related mainly to an electric shock of workers. First, make therelevant work (assembly, test, range work, etc.) conform to the safety standard. Then, it
is necessary to pay attention to reflecting the results into the design of equipment(including AGE).
As a voltage for power supply, besides direct current, the alternating current of aircraft,
etc. using ground commercial power or engine generator output is also conceivable. In aspacecraft, however, alternating current is rarely used because the spacecraft powersource is a solar array and battery and their output is direct current. Alternating current
has the following advantages: a voltage change and insulation between primary andsecondary sides are easy with a transformer, direct drive of a rotating machine ispossible, and short-circuit breaking is easy because of the existence of zero current
point. For alternating current, however, reverse current prevention by a diode is
impossible unlike direct current. Because of power factor (reactive power) due to phasedifference between voltage and current and waveform distortion as an effect of harness
impedance at the AC frequency, and the like, the specifications and function sharing for
power supply interfaces become complex. In addition, when link the AC outputs ofmultiple power sources, phase synchronization is necessary. Thus, there are many
challenges.
High frequencies are indispensable for size and weight reduction, and as a waveform,
sinusoidal waves are desirable with respect to electromagnetic compatibility, etc.
However, a considerable amount of development efforts are involved to solve theproblems above and realize the advantages of alternating current.

I.4.2 Regulated bus/non-regulated bus
The power subsystem types are broadly divided with respect to the variation width of bus
voltage into regulated bus and non-regulated bus.
On a regulated bus, the variation width of bus voltage is narrow (for example, nominal±1%). In each subsystem receiving power, no active regulation (line regulation) function or
performance for input variations is required in general.

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On the other hand, a non-regulated bus has a wide bus voltage variation width and requires
active line regulation function and performance in each subsystem.
With respect to quality of supply power, regulated bus is superior. To implement a regulatedbus, however, the power subsystem must add and improve (load) regulation functions andperformance compared to non-regulated bus. Therefore, the difference between regulated
bus/non-regulated bus types can be restated as centralizing the necessary regulation to thepower subsystem or distributing it to each subsystem, and this choice of type affects thedesign of the entire spacecraft system.
At the choice, design examination becomes necessary with respect to regulation mass/loss(heat generation), reliability (fault propagation, redundancy, etc.), etc. Consideration mustbe paid also to the situation of technology to implement the regulation (in particular,
switching regulator circuit technology and applicable related devices).
With a regulator using a regulated bus, a design remaining intact not only for the predictablepeak power but also for failure of a load is necessary.
For bus voltage variations, the voltage drop due to spacecraft system harness must also betaken into consideration.

Powersupply bus

Load 1

Variation

Load 2PowersourceLoad N

I.4.3 Relation with electromagnetic compatibility
Power supply characteristics are closely related to the specifications for conductivity of theelectromagnetic compatibility design. Figure I.4-1 simplifies the power supply configurationfrom the power supply bus to each load and shows the relationship between noise voltage δ
VBUS in the bus voltage and conductive noise current δ IN of a load, output impedance ZOUT
of the power supply bus, and noise voltage δ VGEN generated in the power subsystem. The
term converted by ZOUT into a voltage is called the rebound component and ZOUT is called
common impedance.

Spacecraft structure grounding

Figure I.4-1 Relationship between bus voltage noise component and noise generated onload side88

JERG-2-214

For each load, the noise voltage at its input end becomes the allowable required level of its
conducted susceptibility and is the sum of δVBUS and Zi×δIi, taking the impedance of thesystem harness as Zi. Therefore, to hold down the noise voltage, it is necessary, besides
reducing the noise generated in the power subsystem, to hold down the output impedance
ZOUT and the noise current of each load. On this occasion, it is a basic policy forelectromagnetic compatibility design to make a design balanced all over the system without
imposing severe requirements on one part, and the power subsystem design proceeds along
this policy.
For δIi related to the rebound component, if these cannot be treated as the sum of noises anda beat of frequency fluctuations (low frequency component, in particular) occurs, its
reduction becomes troublesome. Thus, management of, for example, switching frequency of
the DC/DC converter on the load side is taken into consideration.
Grounding of the power line of the spacecraft system is taken at a point on the spacecraft
structure from the power supply bus (on the return side in general). Grounding is taken
elsewhere than the places where the potential is indeterminate with respect to the spacecraft
structure and in such a way that no stray current flows in the spacecraft structure. The
potential of the spacecraft structure itself depends on the environment. Changes of structure
potential become a common-mode noise with respect to the power supply bus (by making
the grounding at one point, DC-like influence is suppressed).
With regards to this potential, pay attention to the following points.
a.

GroundDuring integration or a test of the system, connect the spacecraft structure to thegrounding of the workshop or test site.
When connecting a simulated solar battery or external power serving as a power source
during a test, the power supply line on the equipment side must be connected to thegrounding of the workshop or test site by the grounding of the spacecraft structure toprevent a current from flowing on the grounding line. If the connection to such
equipment is disconnected from the spacecraft, the relevant line’s potential becomesindeterminate. If joined to the rocket in the range, systematically manage therelationship between the grounding of the rocket structure and the grounding, etc. of arange facility.

b.

In orbit
The potential is determined by the interaction with the plasma environment in outerspace. In the state where there is no sunlight incidence, the spacecraft is negatively
charged because of the velocity difference between ions and electrons, so it prevents an
inflow of electrons but collects ions and reaches a negative potential where the current
between the spacecraft and plasma becomes zero (this depends on plasma temperature;
several volts in a low orbit and 10 kV or over during a magnetic storm in a stationary
orbit). On the other hand, if sunlight is incident, the potential changes toward positiveby the photoemissive effect. In a spacecraft in stationary orbit, this interaction makes apotential difference of about several kV between the structure and an insulated part andinduces an electrostatic discharge.
The structure potential is related to the plasma environment which depends on the orbit
altitude, etc. as well as on the shape, surface characteristics, etc. When docking
spacecraft, the potential difference between them becomes a factor to cause a discharge.
At an electrical connection of a solar array exposed to plasma, a bias in proportion to
the bus voltage with respect to the plasma is superimposed on the potential. In a low
orbit where the plasma density is high, the resultant potential causes, besides a leak89

JERG-2-214

current (power loss), a discharge due to local concentration of electric fields.
I.5 Major components
A power subsystem consists of a battery serving the function of energy storage and equipment
serving the function of power control, distribution, protection, etc. A specific configuration,equipment, etc. are designed based on an examination of the spacecraft system, past records, etc.
Here, an overview of major components of the power subsystem, characteristics, related
technology and method, particular use conditions, and consideration in the design are described.
I.5.1 Battery
A battery consists of battery cells assembled and connected in series to obtain the necessary
voltage. It is charged when in sunshine using the output power of a solar array and discharges
power during an eclipse or other period when the solar array output is insufficient to serve thefunction of supplying the spacecraft.
The capacity of the battery (capacity of cells used), number of cells connected in series, and thenumber of battery units are determined considering the service conditions (life, depth ofdischarge, temperature, discharge voltage, etc.), redundancy, required discharge energy, and thelike. As an important performance index of a battery, stored energy per unit mass (energy
density, mass efficiency: Wh/kg) or stored energy per unit volume (energy density, volume
efficiency: Wh/liter) is used.
The battery cell used, battery structure, matching necessary for the constituent cells of a battery,service conditions, and redundant design concerning a battery are described below.
(1)

Battery cell
As the battery cell used for spacecraft, the nickel-cadmium (Ni-Cd) cell andnickel-hydrogen (Ni-H2) cell are common. The Ni-Cd cell is a secondary cell using nickelhydroxide for the positive electrode and cadmium hydroxide for the negative electrode andhas a long track record of use.
The Ni-H2 cell is same as the Ni-Cd cell for the positive electrode but uses a hydrogencatalyst electrode for the negative electrode and is a secondary cell using hydrogen gas asan active material. Because it does not use a cadmium electrode which is easily degraded,
it is good at cycle life characteristics and has the advantage of being able to take the depthof discharge deep. However, because its external shape is a pressure vessel, it is poor atvolume efficiency and requires a large installation area. Usually, the hydrogen gas reaches
a gauge pressure of 1 MPa or over, so the cell is subject to the High Pressure Gas Safety
Act.
On spacecraft, cells with a capacity of several Ah to 100 Ah or so are used. For the Ni-Cd
cell, a rectangular type is usual although a cylindrical type with small capacity is available,
and 50 Ah or less is usual.
The Ni-H2 cell is not suitable for small capacities because of the pressure vessel shape, etc.
and 35 Ah or more is usual.
Besides these two types of battery cells, another nickel-hydrogen cell (Ni-MH cell; to90

JERG-2-214

distinguish from the Ni-H2 cell, it is sometimes called a high-pressure nickel hydrogencell) is also used. There are expectations that a lithium ion cell will be a future battery cellfor spacecraft.
The lithium ion cell is a generic term for battery cells made of a combination of a positive
and negative electrode that can absorb and release lithium ions not associated with thesolution or precipitation reaction of lithium. Compared to the above-mentioned cells, It has
two or more times as high a mass efficiency. In the lithium ion cell developed in Japan, thepositive electrode used lithium cobaltate (LiCoO2) and the negative electrode uses a
carbon material (C). Development of new electrode materials is now under way and afurther performance improvement can be expected.
(2)Structure
A battery consists of battery cells assembled and connected in series to obtain thenecessary voltage. In structure, a heater for battery temperature control and connectors, etc.
for data transmission of power, cell voltage and temperature are attached. The cell holding
method differs depending on the battery type. For the Ni-Cd battery, the cells are tightened
and held using an end plate and tie rods. Heat release takes place by the cell support rib
placed between cells, and heat is released to the mounting surface from there. For theNi-H2 battery, it is held using a base plate and sleeve and heat release takes place also by
the sleeve. For a lithium battery in an extended elliptical cylindrical shape, shell frame
structure in which the cells are fixed to an structural chassis by a thin metal cell-holding
plate is employed in some cases. A structural conceptual diagram is shown in Figure I.5-1
in comparison with the stack structure used for Ni-Cd batteries.
An appearance of a Ni-Cd battery is shown in Figure I.5-2, and an appearance of an Ni-H2
battery in Figure I.5-3. In structural design, etc., the following matter is examined.
a. Insulation between the cell case and cell-holding structure(If insulation quality is not ensured, the case suffers from electrolytic corrosion.)b. Bonding, fixing and placing method of cell, heater, etc.
c. Heat release pathd. Wiring design
e. Mechanical strength enduring launch environment, cell’s internal pressure, etc.
f. Method of fixing to spacecraft and preparingg. Ground handling ease
In view of handling ease, etc., for some batteries with a large capacity, a unit is assembled
with half the necessary number of series cells and one combination battery unit is made up
of two such units electrically connected in series.

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NutThrough bolt

Cell

Inverted T
section chassis

Basic blockconnecting bolt

End plate

Cell

Shell frame structural chassis
Cell-holding plate

Basic block (4-cell module)
Battery module(a) Stack structure (conventional)(b) Shell frame structure

Figure I.5-1 Conceptual diagram of shell frame structure

Figure I.5-2 Appearance of Ni-Cd battery

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Figure I.5-3 Appearance of Ni-H2 battery(3) Matching of battery constituent cells
The capacity of a battery is determined by the minimum of the capacities of the constituent
cells. Therefore, if there is even one cell with insufficient capacity or a defective cell in thebattery, the characteristics of the entire cell will become poor. Besides this, the variance of
charge characteristics of the constituent cells (voltage and internal pressure) also form afactor that hinders appropriate charge, etc.
When assembling a battery, therefore, it is necessary to line up constituent cells with nocharacteristic variance or as little variance as possible (matching). For that, it is necessary to
select cells based on the capacity data, charge characteristic data, etc. at the cell receiving
test.
From the viewpoint above, it is necessary to have cells of the same product (type) number. In
addition, from the viewpoint that the lot-to-lot variance and long-term characteristics dependon the fabrication process, cells are generally selected from cells of the same lot.
(4) Conditions of use
Because a battery cell functions using electrochemical changes (oxidation and reduction
reactions) of the internal materials, its characteristics tend to be influenced by its service
conditions and history. Thus, the following service conditions must be given consideration.
From this viewpoint, to ensure life, a battery used for a test for a long time is not used in aflight, and another flight article is prepared separately in general cases.

93

JERG-2-214a.

Temperature conditions
The characteristics of a battery cell have a strong dependency on temperature. Besides the
charge and discharge voltage, discharge capacity, charge acceptability, the charge anddischarge cycle life also differ with temperature. In general, the higher the temperature,
the lower the charge voltage and discharge capacity and the shorter the life tends to be. On
the other hand, at low temperature, charge and discharge reactions, etc. are hindered.
Therefore, it is necessary to consider keeping the battery temperature in an appropriate
temperature range.
The recommended range of battery temperature differs with the cell type. In general, it is 0
to 20°C for a Ni-Cd cell, -10 to 10°C for a Ni-H2 cell, and 10 to 25°C for a lithium ioncell. Outside this range, a characteristic degradation may be brought about. Thus, makesufficient examination concerning the service temperature range.
Concerning temperature control, the charge and discharge reactions, etc. of a battery cellare associated with complex heat generation and absorption. In ordinary cases, a thermal
shield or the like is placed to improve the heat release performance of thebattery-mounting part and suppress the influence on other equipment, etc. To avoid an
extremely cold state of the battery, temperature measurement and temperature control takeplace to make the heater operate as the need arises.
The temperature difference between multiple batteries or units of batteries is desirably
suppressed as low as possible, and it is necessary to decide the mounting place consideringsuch a thermal viewpoint besides the electrical and mechanical designs and maintainability.

b.

Charge conditions
The charge current level, which forms the basis of charge conditions, depends on the orbit.
In GEO, the eclipse duration is at most 72 minutes and the sunshine (charge) duration is
20 hours or more. Therefore, the charge current may be the 20-hour rate or so. The charge
current at the 20-hour rate is such a value that when charged for 20 hours, the stored
electricity reaches the nominal capacity of the battery. In an LEO, on the other hand,because the orbital period is about 100 minutes and the sunshine duration is about 60
minutes, a rapid charge is required to recover the amount of discharge. For this charge
current, to charge more electricity than the depth of discharge during sunshine, forexample, for a depth of discharge of 25%, about the 3-hour rate or higher is set to charge
one-third the electricity of the nominal capacity for one hour.
For the current set by the orbit conditions, the current value of full charge is generally
employed. An overcharge with this charge current causes the voltage, temperature andinternal pressure to rise rapidly and impose a stress on the battery. Thus, reduce the charge
current using the V/T curve, etc. In the case where the charge current is changed over to alower level, this charge is sometimes called a trickle charge. During the total sunshine
period in GEO, a trickle charge takes place to make up for self-discharge and keep thestate of charge. In LEO, a further charge by a trickle charge is expected in some designs.
The current adjustment as a charge condition is set as a function of charge control. The
temperature condition as a charge condition takes place by temperature control.

c.

Discharge conditions(i)
DOD: Depth of discharge

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In general, the shallower the depth of discharge of the battery, the longer the battery life
tends to be. Therefore, it is necessary to set the allowable depth of discharge and manage
the amount of discharge of the battery. When setting the allowable depth of discharge, itis necessary to duly consider the operation conditions and mission period of thespacecraft, and life characteristics, etc. of the battery cell.
(ii) Overdischarge prevention measures
When using (discharging) the battery, make sure to set the discharge under-voltage to
avoid an overdischarge. In particular, if discharging completely, an overdischarge of acell must be avoided in the usage to a deep depth of discharge. Thus, It is necessary tomonitor the discharge voltage of each cell. If a Ni-Cd cell is reverse charged, there is therisk that an unrecoverable capacity reduction or a rapid increase of cell internal pressure
may destroy the battery case (explosion, in the worst case). If a lithium ion cell isdischarged to lower than the appropriate voltage, (even if not reverse charged), an
unrecoverable short-circuit fault can occur.
It is necessary to pay due attention to overdischarge prevention measures, not only
during on-orbit actual operation, but also during a ground test in which complete
discharge treatment takes place (management is apt to become sparse because of thetasks before and after the test). In general, from the experience of consumer apparatuseswith a built-in secondary battery, a lack of understanding of battery overdischarge isseen in test personnel, etc. An overdischarge of a battery with several ten parallel cells is
a very dangerous event, and this must be made known and understood thoroughly in thetest procedure, etc. Overdischarge prevention (under-voltage control function) is
considered as a protection function of discharge control. If the function of measuring
each cell voltage, etc. to prevent an overdischarge is provided, pay attention to thedischarge current (very small in general) to the measurement circuit. When leaving thebattery in an open-circuit state for a long time on the ground, the connections of themeasuring system can become a factor that induces a difference in the state of charge
between cells. Disconnection or other treatment is necessary.
d.

e.

Storage condition
When storing a battery for a long time, treatment before storage differs greatly for the
Ni-Cd cell, Ni-H2 cell and other aqueous batteries and for the lithium ion battery. For the
Ni-Cd battery, etc., all cells are discharged completely before storage while the lithium ion
battery is not discharged completely for the purpose of avoiding the overdischarge shownin Clause c and stored in a certain level of the state of charge considering self-discharge,
etc.
Store either battery cell at a low temperature in a dry atmosphere so as not to bring about
corrosion, etc. of the terminals and cell case. For the short-circuit/open-circuit treatment
during cell storage, comply with the recommended condition of the cell manufacturer.
Storage in a charged state can cause characteristic degradation due to self-discharge orother change of internal materials. When using a battery which has been stored for a longtime, a rapid increase of cell voltage or internal pressure must be avoided. For this purpose,
it is common practice to do an activation by discharging after charging at a low rate lower
than the 20-hour rate or otherwise.
Maintenance conditions (on orbit)

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In the case of the Ni-Cd cell, Ni-H2 cell and other aqueous battery, if keeping the battery ina charged state for a long time or repeating a charge and discharge, a decrease of thedischarge voltage of the internal materials not yet discharged (memory effect) occurs.
Therefore, when the battery is used or discharged to a deep depth, the discharge voltage islower than the ordinary voltage. Because this phenomenon can be removed by a very small
current of the 50- to 100-hour rate or so, do this before the launch of the spacecraft andbefore and after the stationary spacecraft enters an eclipse. This operation is called
reconditioning. For a low-orbit spacecraft, reconditioning for a long time involves asuspension, etc. of mission operation or temporary imposition of a large load on otherbatteries. So, do not do reconditioning unless it is necessary. Because reconditioning
involves a deep discharge, it is ordinarily included as a related function of discharge
control.
As a reconditioning method, the battery is discharged simply to the prescribed battery
voltage. Besides this, in one method, each cell voltage is monitored to avoid an
overdischarge or overcharge. In another method, a resistor is connected to each cell and thebattery is discharged to a considerably low voltage.
The lithium ion battery does not have a memory effect in general and does not need theabove-stated reconditioning. In some cases, however, the stored electricity variance
between battery cells is leveled by bypassing the charge current in the necessary amount tothe circuit connected in parallel to a cell. To suppress the degradation due to long sunshine
duration, a charge takes place at a low temperature or a low voltage.
(5) Redundancy design
The battery cell is taken as a limited-life item because its characteristics degrade. Even now,
its design and process management rely on technology that is not adequately understood.
Therefore, redundancy design for mechanical failure is required in general. In some cases inthe early period of space development, many batteries were mounted considering multiple
occurrences of battery failure. Now, because of the accumulation of records and improvementof design and process management, failure of multiple batteries is not assumed though thisdepends on the required reliability. A short-circuit fault of a battery cell merely causes a
decrease of battery charge voltage while an open-circuit fault of one battery cell causes
battery failure. Thus, redundancy design depends in part on the assumed failure mode of thebattery cell.
As a design example, besides unit redundancy which is disadvantageous with respect to mass,
the following element level (to cope with open-circuit/short-circuit) is available.
These are a contrivance to keep the charge and discharge functions of the battery even when abattery cell fails by using other normal battery cells as far as possible.
a.

b.

Use of cell bypass relay (effective only in short-circuit mode)
In a parallel charge of multiple batteries, to match the discharge voltage of normal
batteries to the discharge voltage of a battery with element failure, some battery cells of
the normal battery are bypassed.
Use of cell bypass diode96

JERG-2-214c.

d.

A charge and discharge bypass diode is connected to every cell in parallel. The powersubsystem is configured in advance so that the power supply bus can respond(configuration change) on each battery.
Use of cell short-circuit relay or switch (responds open-circuit mode)
In this configuration, a relay or a switch is connected instead of the bypass diode in
Clause b.
Discharge control on each battery (effective only in short-circuit mode)
By discharge control, each battery is discharged irrespective of the difference indischarge voltage.
These contrivances assume random failure of battery cells and are not effective forcharacteristic degradation common to the manufacturing lot of the element.

(6) Destructive failure mode of battery
The battery accumulates energy. If a phenomenon to release the energy for a short time isinduced, failure doing catastrophic damage to the periphery of the battery will result.
However, such a mode is prevented for an element and a battery appropriately designed andsubject to sufficient process management by using the battery appropriately and providing aprotection function for emergencies. Defenses against the released energy or other measures
are not generally taken on the spacecraft.
For the cell failure prevention measure in (5), it is necessary to examine in advance the effectof a (short-circuit or other) failure of that measure itself on the cell.
Concerning such battery hazard, the mode, causes, effects, measures and their confirmation
methods are reviewed.
I.5.3 Power control
Power control of the power subsystem include solar array output control, battery charge
control and battery discharge control. The operation region of these control functions isusually designated by the (bus) voltage of the power supply bus as shown in Figure I.5-4 andthere the coordination of the control functions is kept appropriate. Then, the specifications for
supply of required load power and for bus voltage of the power supply characteristics are
satisfied.

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Solar array output
control

Solar array output control
Dead band between solar array
control and charge control

Charge control

Dead band between charge
control and discharge control
Discharge control
Lower limit of bus voltage

Abscissa corresponds totreatment power.

Figure I.5-4 Operation region of control

For the bus voltage, if it falls, discharge control keeps it higher than the lower limit, and if itrises, charge control charges the battery first. Then, the battery charge power is held there
because it keeps the charge conditions appropriately (corresponding to the region in broken
line in Figure I.5-4). If the bus voltage rises further, solar array output control keeps it lower
the upper limit.
The operation point of the bus is determined by the intersection of the characteristics of thesolar array, charge and discharge control, load, etc. As an unusual case, a (lock-up) state canoccur in which discharge control operates within the charge control region. To leave this state,
the load power is decreased or other means are taken.
The bus voltage variation range is specified by the upper limit and lower limit. A regulated
bus corresponds to the case where this region is ±1% or so of the nominal voltage. On theother hand, for a non-regulated bus, the upper limit is generally specified above the voltage tocharge the battery and treat the maximum surplus power, and the lower limit is specified by
the battery discharge voltage. The region is ±20% or so of the mean of the upper and lower
limits.
The lock-up state poses a problem especially in the case of a non-regulated bus with a widebus voltage range. Coordination of control functions is achieved by issuing, from the bus
voltage detection processor, a signal to operate the control function corresponding to theprescribed operation region. Failure of this processor influences the entire power control andcauses abnormal variations of bus voltage. In redundancy design to prevent the grave effectof this single failure, a standby method or simple parallel method cannot be employed. Thus,
the majority-decision method is applied.
An overview, type, etc. of each power control is described below.

98

JERG-2-214(1) Solar array output control
The voltage-current characteristics of a solar array is as shown in Figure I.5-5, and its output
current (power) is determined by the operation point of the solar array. When simply
connecting the panel and load, the operation voltage is determined by the intersection of theload (voltage-current) characteristics and solar array characteristics., and the power supply(bus) voltage will vary due to load power variations, panel temperature variations, etc. As ameasure for this, suppressing the bus voltage variations is a function of solar array output
control. That is, by adjusting or otherwise handling the power generation of the solar array,load power (power of the required power profile, power consumption of the power subsystem,
charge power, etc.) is supplied from the solar array and the bus voltage is kept within therange specified by the power supply characteristics.
Current

A⇒B: Decrease of solar cell temperatureA⇒C: Decrease of incident sunlight
Radiation degradation

Current

Figure I.5-5 Solar array current – voltage-current characteristics
The adjustment methods are divided as follows: A schematic configuration of it is shown in
Figure I.5-6.
Shunt type

Full shunt

Panel output
control

Partial shunt
Series typeOf the shunt type, in full shunt, the operation point voltage of the solar array corresponds tothe bus voltage. Although it cannot adjust the power generation of the panel itself, it diverges(shunts) surplus power to the control element, supplies power to the load, and keeps the bus
voltage within the specified range. This control element operates corresponding to a relief
valve in a pressure line.
In partial shunt, a control element is connected to the intermediate tap of the solar array. It
makes adjustable the surplus power diverging to the control element and thereby adjusts thepower generation of the solar array itself. On this occasion, the operation voltage of the upper
panel is determined by the load current while the operation voltage of the lower panel is madeadjustable by the control element.

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Full shunt

Load

Solar
array

Power supply
Controlbelement

Load

Solar
array

Controlelement

Power supplybSeriesLoad

Intermediate
tap

Partial shunt

Controlelement

Upper part

Solar array

Lower part

Power supplyb

Figure I.5-6 Types of solar array output control
The series type inserts a control element in series between the solar array and load. It keeps
the bus voltage within the specified range, makes adjustable the operation point voltage of thepanel and adjusts the power generation. This control element operates corresponding to athrottle valve in a pressure line.
For the shunt type, adjustment is treatment of surplus of solar array power generation, andsolar array delivers the maximum power with the control element in a non-operating state.
The characteristics do not appear in the budget analysis model in Figure I.3-1. On the other
hand, for the series type, when the solar array output is at the maximum, the treated power ofthe control element also becomes maximum and the effect has an influence on ηS in FigureI.3-1.
a.

Features, etc. of type
In solar array output control, adjustment (treatment of surplus power, etc.) of power
generation of the solar array is associated with a considerable amount of heat generation,
and a smaller amount of such heat is more suitable for the system from the viewpoint ofthermal design. From this viewpoint, features, measures, etc. of each type are described
below.
(i)

Features of shunt type
For this type, the larger the load power, the smaller the surplus power is treated. In
partial shunt, the configuration, etc. becomes complex compared to a full shunt in thatan intermediate tap is used; however, because the heat generation is small, it is suitable
for spacecraft that use a large amount of power. The operation of the control element isbroadly divided into analog and switching (see Figure I.5-7), and switching is superiorin view of heat generation; however, it needs attention to electromagnetic compatibility,

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Load

Controlelement

Solar arrayetc. As a method to reduce the power generation, there is a configuration in which thecircuit in the solar array and connection of the control element are further divided andoperated in sequence (sequential operation) (Figure I.5-8), and it is suitable for controlof large power; however, the interface becomes complex.

Current balance: ILOAD = Iarray - ISHUNT

Analog type

: Continuously adjusts ISHUNT (surplus) to keep balance.

Switching type

: Turns ON/OFF the control element and adjust the ratio D of that
ON time to make ISHUNT in the balance equation above equal to D
× Iarray or so. Actually, a function to smooth the voltage variations
due to switching is needed.

Load

Controlelement 1

Reversecurrent
prevention

Controlelement 2

Arraycircuit 2

Arraycircuit 1

Figure I.5-7 Operation of control element (in case of full shunt)

The control element appropriate for the load is operated in sequence. For example, under alight load, only control element 1 is operated and all others are put in the ON state so as
not to supply the output of the corresponding array circuits to the load.
On the other hand, under a large load, only element N is operated and all others are put in
the OFF state so as to supply the output of the corresponding panel to the load. As theoperation of a control element, either the analog type or the switching type may be
employed.

Figure I.5-8 Sequential operation of control elements (in case of full shunt)

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JERG-2-214(ii) Features of series type
For this type, all of the power generated by the solar array applied to the load passes
through the control element, the loss (heat) by the control element must be reduced.
Therefore, the operation of the control element is switching. An advantage is that the bus
voltage is not equal to the operation point voltage of the panel, so the interface with thesolar array has flexibility and it is also possible to add a function to track the maximum
output of the panel, thus making up for the loss by the control element.
As for selection of type, the shunt type can be applied for both GEO and LEO while theseries type is effective for LEO because the share of charge power in the solar array
power generation is large. In GEO, the greater part of power generation of the solar
array is load power, and the loss by the control element of the series type reduces thepower supply capacity to the load, so the advantages of this type (e.g., the operation
point of the panel can be made adjustable) cannot be realized.
b.

Solar array power generation monitoring
Figure I.5-9 shows an example of a power subsystem telemetry arrangement (full shunttype) for solar array power generation monitoring. To evaluate on-orbit solar array
performance (maximum generation capacity), a method to put the control element in the
non-operating state and increase the load or change the sunlight incident angle orotherwise handle it is necessary. These are difficult to employ in part from the viewpointof operation and design. However, to evaluate the integrity of the solar array, thenecessary minimum integrity evaluation is possible by the telemetry arrangement shownin Figure I.5-10 as necessary telemetry as a power subsystem. To monitor the generated
current of the solar array, the following two methods are available.
(i) Evaluation by the total of currents on the HOT side
Evaluate as Iarray=ILOAD+ISHUNT+Ibat_chg. To evaluate in terms of power, make anevaluation in a simple way by multiplying the bus voltage or multiplying the correctedpaddle interface reference point voltage.
(ii) Evaluation by the total array RTN current
Evaluate as Iarray=Iarray_RTN. To evaluate in terms of power, make an evaluation in a
simple way by multiplying the bus voltage or multiplying the corrected paddle
interface reference point voltage.
Evaluation is possible by above-mentioned [1] alone; however, together with [2],identification of an open-circuit or short-circuit is possible when an array circuit fails.

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Solar paddle
subsystem

Chargecontrol

Discharge
control

Load

Solar array

Control element

Bus voltage
control pointPowersubsystem

Current
TLM

Voltage
TLM

Figure I.5-9 Example of solar array power generation monitoring (full shunt type)c.

Simplification of output control
There is a method in which the solar array output is not adjusted finely but the output
characteristics are used as they are. For the series type, it is conceivable to divide thesolar array as shown in Figure I.5-8 and use relays for the control elements in series andturn them on and off. In these methods, the power supply characteristics become coarse
and the battery has the role of a shunt. Thus, they cannot simply be employed; however,they are effective for system simplification.
In some cases, the required capacity of the control element can be reduced by using arelay, etc. For a paddle-shaped panel, it is conceivable to change the sunlight incident
angle by the paddle tracking subsystem to do output control of the solar battery. However,this poses a problem concerning response and operation coordination between the attitudecontrol and power subsystems (the propagation range of malfunction becomes larger), sothey can hardly be said to be an ordinary design.

(2) Charge control
Charge control has the function of keeping the battery charge conditions appropriate. In theappropriate range, it actively control the bus voltage by this operation using all the solar array
power generation, excluding the load power, for charge (continuous line in Figure I.5-4). The
appropriate range of charge conditions is set by charge current, voltage, temperature and thelike so as not to impose an excessive stress on the battery. In general, charge control functions
so that the charge current and voltage will not be inappropriate (upper limit) value (broken
line in Figure I.5-4). As a type for that, the two types of constant current-constant current andconstant current-constant voltage are commonly used.
If sensing the battery RTN side (A6 in Figure I.5-9) as a telemetry detection point used forbattery charge control, pay due attention to, and examine, the design so that charge control
takes place normally during an operation associated with a battery discharge during sunshine.

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Charge type(i) Constant current-constant current
First, charge taking the upper limit of charge current as the full charge current value. The
charge current at this time is usually the full charge current value though this depends on
the panel power generation and load power. When the charge proceeds and the battery
reaches the specified state (when the voltage rises and reaches a set level), lower theupper limit of charge current below the full charge current value to suppress the stresson the battery and continue the charge. Charging with a lowered level is sometimes
called trickle charge or supplementary charge and is generally aimed at maintenance ofthe state of charge.
In a total sunshine period in GEO, the state of charge of the battery is maintained for along time by trickle charge. Instead of a trickle charge, it is also possible to maintain thestate of charge of the battery with a constant voltage charge as shown in (ii).
(ii) Constant current-constant voltage
First, a charge takes place taking the upper limit of charge current as the full charge
current value in the same way. However, after the battery has reached the specified state,charge continues with the battery voltage kept constant. The charge current during thisconstant voltage charge usually decreases exponentially. That is, since the voltage isheld at the constant voltage value Vc, the current (Ic) decreases when the stored
electricity (Q: Ah) increases. Thus, assuming the following proportional relation:Ic ≈ - K×Q
Here, because of Q = I c×t, after all, the charge current becomes as follows:
dIc/dt ≈ - K×Ic ∴ Ic ≈ Ico×exp (-Kt)
Since the charge current changes with time like this, it is sometimes called a taper
charge.
Thus, during a charge, when the battery reaches the specified state, the charge current isreduced to suppress the stress. Reaching the specified state is judged generally byconsidering that the battery voltage has reached a set level, and that level is set taking
into consideration the characteristics and service conditions of the battery. With too low
a level, the amount of discharge is not restored by a charge sufficiently, and with toohigh a level, unnecessary stresses are imposed on the battery. It is necessary to pay
attention especially to the aging of the battery characteristics. It is common practice toprovide multiple levels and select from them during operation. In case of doing
temperature compensation, these levels are sometimes called a V/T(voltage-temperature) curve. As a technique to judge that the charge has reached thespecified state, besides the technique by V/T, the following is available.

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JERG-2-214[1] Introduce measurement of the amount of electricity (time integral of current) andjudge whether the amount of charge electricity has reached a specified value withrespect to the amount of discharge electricity.
[2] Detect that the element temperature rises due to the heat generation associatedwith charge completion or that the charge voltage decreases by this temperaturerise. With a stationary spacecraft, ground operation of confirming this voltagechange (called Head Poking because the value goes down after rising in the
charge) and issuing a command to make a current change-over is available.
[3] Monitor a physical phenomenon (for example, change in internal pressure or
generation of oxygen, etc.) in which the state of charge inside the battery cell isreflected.
Note: The lithium ion cell does not cause the water-splitting reaction or other side
reactions seen in the Ni-Cd cell, Ni-H2 cell or other aqueous batteries.
Therefore, full charge judgment in [2] and [3] is impossible.
In relation to the effect of temperature on charge of an aqueous battery, when charge isalmost complete, charge power changes into heat, and this tends to raise the battery
temperature if heat rejection is poor. The temperature change during charge poses a
problem especially during constant voltage charge. That is, because of the temperature
characteristics of the battery, even if the voltage is held constant, if the temperature rises,
the charge current increases. Therefore, the tendency of a taper charge in which thecurrent increases with an increase in the amount of stored electricity is lost, and thetemperature rise and charge current rise enter a positive feedback relation, and fall into astate of thermal runaway. The charge conditions are not set merely by the charge current
determined by the orbit and depth of discharge. This setting requires sufficient knowledge
about the battery characteristics and temperature environment.
The lithium ion battery causes no exothermic reaction if charged in an appropriate voltage
range. Therefore, it never falls into a state of thermal runway.
b.

Control type
The types are broadly divided into the shunt type and series type, as with the types of solar
array output control.
For the shunt type, the control element of full shunt in Figure I.5-6 takes a configuration inwhich the charge circuit and battery are in series. In the series method, the panel output
control in the series method doubles as charge control and takes a configuration in which
the battery is contained in the load in series in Figure I.5-6. A design in which charge
control is superimposed on panel output control is the shunt type at the same time; however,this integration of control tends to coarsen the charge control capability of the battery andreduce the flexibility.
As operation of the control element, the selection is made from the viewpoint of loss
considering the level of treated power, etc. In LEO where charge power is large, switchingis indispensable for reducing the loss.

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For charge control of multiple batteries, it is common practice to provide each battery withdedicated charge and discharge functions to keep the charge conditions appropriate
individually. When doing charge control of multiple batteries connected in parallel, it isnecessary to pay due attention to the battery temperature uniformity, etc. to ensure
equalization of charge currents. Separation of a battery exhibiting a low-voltage
abnormality (charge current becomes excessive) is essential.
Charge control includes the function of detecting an abnormal temperature and suspending
the charge to operate the battery at an appropriate temperature (power subsystem protection
function).
(3) Discharge control
Discharge control has the function of keeping the power supply characteristics within
specification using the stored energy in the battery when the required power profile cannot
be maintained with the power generation of the solar array. In one type, it is common
practice to insert the relevant function between the battery output and power supply bus.
Figure I.5-10 shows the most common arrangement of charge and discharge control using
charge control of the shunt type. The control element is associated with the type of regulated
bus/non-regulated bus. In the case of a regulated bus (to raise the battery charge voltage
above the lower limit of bus voltage), it has an active power conversion function. In the caseof a non-regulated bus, it is made up only of a passive reverse current prevention diode andon/off relay. And, in the case of an active function to treat large power, switching is
indispensable for operation of the control element to reduce the loss.
Power supply bus
Chargecontrol

Discharge
control

Battery

Figure I.5-10 Arrangement of charge and discharge control (in case of shunt type)For discharge control of multiple batteries, on/off and other functions are installed
individually, and power is discharged in parallel in general. In this case, the variance ofdischarge current is suppressed by equalizing the resistance, etc. of the discharge lines
because the discharge voltage of the battery is relatively stable. On the other hand, somedesigns provide each battery with an active power conversion function to equalize thedischarge current.
To prevent an overdischarge, etc. of the battery, discharge control has an under-voltage
control function (power subsystem protection function), which monitors the battery
discharge voltage and, when the specified level is reached, issues an LLM signal to change
the spacecraft operation over to low load.

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I.5.3 Power distribution
Power distribution is the function to supply power from the power subsystem to each subsystem,component and the like with the prescribed characteristics. In the case of a spacecraft, the powertransmission line, which is the output of the power subsystem, is taken as a part of the system
harness. Here, without regard to this division, an overview of the types, bus configuration,
switchgear, monitoring function, etc. is described.
(1) Distribution typeThe distribution types include the radial type and backbone type in Figure I.5-11 (viewpoint of
redundancy omitted). Besides these, there is also the ring type and others; however they are notcommon on spacecraft. Major features of each type in Figure I.5-11 are described below.
a.

Radial typeIn the power subsystem, there is a bus, which dictates the power supply characteristics andsupplies power to each load through a dedicated harness. Therefore, its design is easy and it
is excellent with respect to supply characteristics, sharing of fault protection measures, etc.
However, it makes the system harness complex and is poor in terms of extensibility. On
spacecraft, this type is common conventionally.

b.

Backbone type
The power supply bus branches out to the system harness and the power supply interface tothe load is implemented by dedicated Vnits. Therefore, in this configuration, most of the
distribution function is distributed to the dedicated units. (The reference point becomesambiguous.) The power supply characteristics are inferior to the radial type if a voltage
regulation function is not given. However, the system harness is organized and response toextensions is easy.
Instrument
subsystemPower
subsystem(system harness)

Load 1
Load 2Power
supply bus

Load 3
Load 4
Load 5(a) Radial type107

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Powersubsystem

Powersupply bus

Load 1

Load 3

Load 5

Dedicated
unit 1

Dedicated
unit 2

Dedicated
unit 3

Load 2

Load 4(b) Backbone
Figure I.5-11 Types of power distribution
For a modularized spacecraft, the loads of the radial type correspond to subsystems, etc.,
and power supply to each component takes place radically by the dedicated units. An
hierarchical structure is sometimes taken.
In a redundancy, it is common practice to provide a dedicated harness from the power
supply bus. If the redundancy is incorporated in one chassis, the interface part with thepower subsystem is made common.
(2) Bus configuration
As a configuration of a power subsystem, (in the case of a single-bus), the power supply bus
combines the solar array, battery and power control function and consists of one line, and (in thecase of multiple-bus) it consists of multiple lines. A single-bus configuration is commonly usedbecause it is simple in configuration while a multiple-bus configuration is sometimes employedbecause the features are associated with the following distribution function.
a. Basic equipment of a spacecraft is supplied with power from multiple bus lines (diode-or
connection, etc.). Therefore, in the event of a failure of one line, loss of function can beprevented.
b. Different bus voltages, etc. suitable for the load can be supplied.
If multiple buses having the same configuration and characteristics are used, linked
operation of the buses is sometime employed for accommodation,etc. of power supply
capacity among the buses.
In this case, also in a state of not being linked, it is necessary to pay attention to theallocated power share and required power profile of each bus so as to prevent thedifference in service conditions of the battery of each bus from becoming large.
A multiple-bus configuration is sometime employed to implement buttery redundancy inunits of batteries by connecting one battery to multiple buses.

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JERG-2-214(3) Switching function
The switching function is generally implemented using a relay with a break or by using amechanical contact or a semiconductor device. In the case of a relay, chattering or contact
reversal can occur due to vibrations or impact. The design must consider measures for that (asto occurrence of such a phenomenon, select a device with no possibility, or implement a
method to exclude an occurrence or to prevent the effect at an occurrence). However, it issuitable for switching a power line because the power loss is small. Semiconductor
switchgear has superior characteristics to those of a relay in terms of response,electromagnetic compatibility, reliability, operation easiness, etc.; however, a power loss
occurs.
The service conditions of a relay and load characteristics in particular influence the life of thecontact. In the transient period of current switching, an electrostatic discharge occurs
inevitably between the pair of approaching contacts (the electron emission by an electric field
becomes a trigger for a temperature rise of the contact material, which promotes evaporationof the metal and ionization of the environmental gases, and occurrence of an arc plasma).
While the discharge conditions are satisfied, the discharge continues to promote contact
ablation and deformation
The duration of a discharge has a proportional relationship with the power being conducted.
Special care is required when handling a large-power load or inductive load. For discharge
suppression, it is effective to reduce the increasing rate of contact voltage versus contact
opening speed and, as a measure to take, a snubber circuit (which forms a transient route forthe current at contact opening) is connected to the relay in parallel.
In some cases, the relay is used only for a switch-over of connections and turning-on/off ofequipment is done separately.
If the number of times of opening/closing poses a problem, it is common practice to use aswitchgear using semiconductor devices.
(4) Monitoring function
To monitor the power supply condition, supply current, etc. are monitored. This also is
included in the distribution function; however, monitoring takes place not on each load but on
total output current, etc. in general. When evaluating the operating condition, etc. of manyloads by monitoring the total output, there are restrictions as with monitoring of solar array
power generation. So, it is necessary to accumulate and evaluate sufficiently detailed related
data by ground tests, etc. in advance. Especially for equipment in which it is necessary toknow current consumption data to know its operating condition, for example, mission
equipment experiencing large load variations during operation, it is desirable, if possible, toindividually include current data, etc. as a telemetry item.
In relation to the switchgear function, the on/off status is included in the distribution function
in some cases. The sharing of these functions should belong to the design policy at the systemlevel and specified in the design specification, etc. of the system.
(5) Power supply other than primary power bus
The distribution function of a power subsystem covers the power supply at the bus voltage.
Besides this, as power that the power subsystem distributes, special power, for example,
pyrotechnic ignition power for paddle or antenna deployment, is included in some cases. For

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that, the battery output is directly used in general, and a dedicated distribution function isprovided. Another special case is equipment or a subsystem that does not require manyelectric or electronic circuits, for example, power supply to the heaters and sensors to thethermal control subsystem, propulsion subsystem, etc. By using existing components, ameans to supply power other than the bus voltage (a specific stable voltage in case of a
non-regulated bus) is used in some cases.
I.5.4 Protection
The protection function has the role of localizing a failure or abnormality that causes serious
effects on power supply to avoid a critical situation as a system or a situation hindering
mission achievement. The phenomena it covers are defined as a contingency, etc. based on
FMEA in reliability design. For abnormalities or failure of a device due to an open-circuit or
short-circuit, measures such as a series-parallel connection and device or equipment
redundancy are considered in the design. The power subsystem is required to operate at all
times and its fault tolerant design is made on the premise that its major functions cannot be
stopped (for example, like the engine of an automobile) as fail-safe. Therefore, its fault
tolerant design is made by incorporating the following measures.
•
•
•
•
•
•
•
•
•
•
•
•

Multi-connection of array circuit and redundant connection between array circuits
Reverse current prevention diode on each one-series circuit of array circuit
Redundancy in the number of slip rings/brushes and harnesses
Battery cell bypass circuit, unit redundancy and isolation relay of battery
Reverse current prevention diode and reverse current detection/change-overmechanism of battery
Series-parallel connection of diode, transistor, etc.
(Parallel coupling) redundancy and current limiting circuit of power supplies forinternal circuit
Majority vote redundancy method for bus voltage monitoring of power control
subsystem
Short-circuit detection/circuit isolation relay or fuse
Reinforced insulation of power wiring, etc.
(Parallel coupling) redundancy of power control section, etc.
Cross strap coupling of battery output and discharge control section

Therefore, the phenomena covered by the protection function are mainly non-criticalabnormalities such as an open-circuit and short-circuit and malfunction of a subsystem due tosuch an abnormality. Detection and emergency measures (isolation) of these are the role ofthe protection function.
Conventionally, restoration (correction) has been done by a command from the ground ingeneral. A ground monitor is necessary to supplement the protection function and for
operation evaluation and abnormality indication detection by trend analysis of important data
and for contingency measures. Currently, to respond to these on board, a protection function,etc. having intelligent capabilities is being applied. For example, if multiple abnormality
detection inputs are available for emergency measures of the protection function, recoding
what has happened in detail is useful for ground work (establishing an action plan) for which
limited data is available.

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Table 3.5.5-1 in the text shows an example of the phenomena covered by the protection
function and their detection and isolation methods, etc. Of these phenomena, those which
require action considering the degree of influence and frequency of occurrence by FMEA, etc.
are objects of the protection function.
The circuit, etc. for the required original function should not be made complex by adding aprotection function. Some emergency measures do not necessarily require to be automateddepending on the spacecraft and some depend on the situation of the abnormality. Restoration(correction) should be performed based on the survey and investigation of the causes of theabnormality. As a correction, besides a change-over to a redundant line, changing the powersubsystem configuration (reconfiguration) is also conceivable. For emergency measuresinvolving power supply, such as a transition to a light-load mode, increase or decrease in load,
shut-off or isolation of a load, flowing of a fuse, and linking of multiple buses, it is necessary
to do interface coordination, function sharing, etc. with a related subsystem, etc.
As a feature of the protection function, although it is in operation at all times, its operation
itself does not emerge at times other than an emergency. So, its malfunction or failure has agrave influence on the system and, in addition, its confirmation test is not easy to conduct. In
the design of such a protection function, matters to consider are as follows:a.

Status monitoring
If the protection function has a status to detect an abnormality, at an occurrence of an
emergency measures-requiring state, it must be confirmed whether it is due to an operationof the protection function or not, and monitoring takes place for that. If multiple protection
functions do one emergency measure, the status of each protection function must be
monitored. The operation history, etc. of the protection function may become necessary forpost analysis, etc.

b.

Override
Refers to making it possible to disable or cancel the protection function from the ground at
its malfunction, abnormality, etc. However, pay attention to the fact that a decision between
normality and abnormality of the protection function itself can be made from other data forover-temperature protection, etc. in many cases while it is not easy for overcurrent shut-off,
etc.

c.

Reset
As a premise of a protection function, set it so that it can operate again at re-closing(restoration). If the protection function has hysteretic characteristics as to detection level,
do not re-close until it enters an operable state again.

d.

Use of fuse
It is useful for isolation of an internal local short-circuit, etc. of equipment. Design andselect it so that it will not melt during a normal state. At an emergency, the powersubsystem should be able to deliver the melting current without being damaged. Use of afuse should be based on a electrical design standard, etc.

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Detection level and response
Ensure that the protection function will not operate under a normal transient situation. If
degradation cannot be avoided during the life period, provide multiple levels and make achange-over.

f.

Coordination of protection
Ensure that multiple protection functions do not operate in succession (like dominoes). A
transient state due to operation of one protection function can cause another protection
function to operate in some cases. As a conceivable example, an instantaneous battery
voltage drop due to a short-circuit may be judged as an over-discharge of the battery; whenisolating a short-circuit on the load side, the circuit-breaker of the power subsystem also
operates to shut off power supply to other load; during a transient period of a transition toan emergency measures-requiring state, it may be judged to be an abnormality of the bus
linkage or may influence the operation of another subsystem. Prevent these situations andensure coordination between protection functions. After emergency measures are taken,
contingency analysis, etc. is performed to prevent the spacecraft from going into a worse
state.

g.

Confirmation test
Examine a method to confirm by a ground test, etc. in advance that the protection function
will operate normally. When adding an intelligent capability the protection function, aself-diagnosis function, etc. is indispensable.

I.5.5

OthersAlthough not classified as an element of a power subsystem, system harness, pyrotechnic
control, umbilical cable, test cable, etc. are closely related.
(1) System harness
Power of the power subsystem is supplied to subsystems and equipment through a system
harness. The voltage drop, etc. associated with this power transmission has an influence on the
power supply characteristics at the load’s receiving end. And, the power loss due to a voltage
drop has an influence on power supply capacity. For example, a 1 V voltage drop for a 1 kW
load’s receiving end voltage of 50 V means a system harness resistance of 50 mΩ, supply end
voltage of 51 V and a loss of 20 W (2% of load power), that is, the power the power subsystem
must supply is 1020 W.
In general, the power supply characteristics at a load’s receiving end is an electrical interface
management item; however, their details are closely related with major specifications of thepower subsystem.
A system harness consists of many cables bundled and formed in a shape that is easy to install at
a prescribed position in the spacecraft. It is taken as one of the components for hardware
integration of a spacecraft and is designed based on the Electrical Design Standard.
Besides the system harness, hardware integration includes, electrical elements, etc. involved inbonding, shielding, antistatic measures, etc. Selection of connectors, harness cables, service112

JERG-2-214

conditions (arrangement, current derating, etc.) are specified in the Electrical Design Standard.
A harness is classified into categories for power supply, signal (high level/low level),pyrotechnic, high frequency, etc. Isolate different categories as far as possible to avoid
electromagnetic interference. An especially critical harness line is fabricated with shielding.
A system harness test consists of confirming the prescribed connector pin connections at bothends, insulation from other cables, withstand voltage, etc. However, because may cables arebundled, the confirmation is not easy.
Measurement, confirmation, etc. of line resistance also takes place on this occasion.
If a polarity reversal of a power line or a confusion with another cable is overlooked, equipment
damage results at the initial current application. At an integration, therefore, follow a proceduresuch as confirming the voltage again before connecting to equipment. At a ground test,
meticulous inspection is necessary for a line whose operation cannot be confirmed in fullconfiguration. The environmental resistance of a system harness is considered in design.
Besides this, in its single state, a vacuum test is conducted, and at the system level, thenormality, etc. of installation and contact are confirmed using vibrations, impact, etc.
(2) Pyrotechnic ignition circuit
For paddle or antenna deployment and AKM ignition, etc., pyrotechnic (EED: electro explosive
device) is used. Failure or non-operation of this circuit brings about a critical situation for thesystem and forms a grave hindrance to achievement of the mission. In the ignition circuit design,
therefore, pay attention to the following items.
a.

b.
c.
d.

e.

f.

For certain ignition, make the ignition line (ignition circuit, wiring, pyrotechnic)
redundant.
Provide the automatic ignition sequence, etc. with a backup of ground commands.
Make a sequence to ignition as a 3-stage sequence (enable/disable, ignition) andmonitor its status.
In the sequence, incorporate a spacecraft/rocket separation signal, attitude control
subsystem’s pre-operation completion signal, etc. in a circuit.
To prevent malfunction, simplify the line, clarify the relationship, etc. with thespacecraft grounding, and connect an antistatic resistance to the wire to thepyrotechnic.
Because the heating wire (bridgewire) to fire the powder in the pyrotechnic cannot
fulfill the firing function if the supply current or time is insufficient, know thetransient response of the power source and set the supply time. On the other hand, an
unexpected contingency (release becoming impossible, short-circuit to the structure,etc.) occurs if an excessive current flows. Thus, insert a current-limiting resistance or
use a constant-current power supply circuit to ensure the recommended firing currentof the pyrotechnic,
To turn on, etc. the pyrotechnic power, a relay or semiconductor device is usually used.
However, contact reversal or chattering due to vibration or impact poses the risk of
causing serious troubles from a system malfunction. Therefore, evaluate theenvironmental resistance of the relay, pyrotechnic power controller, etc. and takemeasures in advance.

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If a relay is used, that line is limited as to the number of times of operation andverification using an actual piece of equipment is difficult. Therefore, pay special
attention to the test plan to confirm the performance of the system, and carry it out.

(3) Umbilical cable
From assembly to launch of a spacecraft/rocket, power supply, state monitoring and battery
charge take place at an umbilical cable. As to the power subsystem, the following items areincluded under an umbilical cable.
a.

b.

c.

Insert a reverse current prevention function to ensure that the spacecraft is notdamaged inside by a short-circuit outside the external power lines (for power supply inthe range and for battery charge). Restrict the supply power to the necessary amountand take protection, etc. measures for failures or operational mistakes so that noexcessive power or current is delivered to external power sources.
Insert a protection resistance, etc. to ensure that the spacecraft is not damaged insideby an external short-circuit of the monitor lines of bus voltage, battery voltage, relay
status, battery temperature, etc. Pay attention also to the noise immunity of groundmonitoring lines (in particular, CMR: common mode rejection to electromagnetic
interference of ground lines) and treatment of connection ends not yet connected.
Battery ON/OFF relay drive line
Treat in the same way as a range power supply line.

Besides this, an umbilical cable includes propellant-related pressure and temperature lines,
etc. For an umbilical cable, it is necessary to make a design so that, when a connector ispulled out, no effect (status reversal, etc.) will be induced inside the spacecraft.
Know the lightning protection measures of the range and give consideration so that themeasures for the umbilical cable can be taken promptly. In range work, a prior confirmationis conducted by paying attention to the above.
(4) Test cable, etc.
In a ground test, it is difficult to operate the power subsystem with the power generated by
the solar array, and a simulated solar array is generally used. Test cables include, besides thisconnection line of a spacecraft and simulated power source, various types of simulation
signal lines such as for hardware line monitors, sensors, etc. of spacecraft state, redundantfunction confirmation line, etc. This test cable line operates in a more severe environment as
to installation and handling situation, etc. than an umbilical cable. The same measures arenecessary to ensure that the spacecraft will not be damaged by its abnormality (short-circuit,etc.). When using a reverse current prevention diode, consider the measures for its heatgeneration. The spacecraft’s receiving side of a simulated power source is the connection partof a solar paddle subsystem output. It is not fixed in many cases (wire, etc. of a slip ring on
the triaxial type) and its treatment and fixing method require care. In addition, it is necessary
to provide the simulated power source with measures so as not to deliver an excessive
voltage.

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Because a test cable and test equipment are not included in the flight model, their design
review and other checks are apt to be insufficient. However, they are connected to the flightmodel indeed, so their inspection, etc. before use should be duly performed.

115