General
JERG-2-215A

DESIGN STANDARD
SOLAR ARRAY PADDLE

Mar

27,2013

RevA

Japan Aerospace Exploration Agency

JERG-2-215A

This is an English translation of JERG-2-215A. Whenever there is anything ambiguous in thisdocument, the original document (the Japanese version) shall be used to clarify the intent of therequirement.

Disclaimer
The information contained herein is for general informational purposes only. JAXA makes nowarranty, express or implied, including as to the accuracy, usefulness or timeliness of any
information herein.

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

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

JERG-2-215A-Contents1. Scope ············································································································ 1
2. Applicable documents ························································································ 1
3. Terminology, Definition and Abbreviation ································································ 1
4. General requirements ·························································································1
5. Subsystem design and analysis ··············································································3
5.1

Electric design and analysis ..................................................................................................... 3

5.1.1

Power generation design ....................................................................................................... 3

5.1.1.1

Array circuit component design ........................................................................................ 4

5.1.1.2

Power generation analysis ................................................................................................. 5

5.1.1.3

Array circuit configuration (layout) design ..................................................................... 10

5.1.1.4

Design of interface between PDL and paddle drive mechanism ..................................... 15

5.1.1.5

Shunt interface design ..................................................................................................... 16

5.1.2

Operation mode and power consumption design ................................................................ 18

5.1.3

Electromagnetic compatibility design................................................................................. 18

5.1.4

Signal design....................................................................................................................... 18

5.1.4.1

Suspended capacity of signal wiring ............................................................................... 18

5.1.4.2

Consideration of influence of noise from the power line associated with shunt
operation .......................................................................................................................... 18

5.2

Structural design and analysis ................................................................................................ 19

5.2.1

Panel substrate design ......................................................................................................... 22

5.2.2

Boom and yoke design........................................................................................................ 24

5.2.3

Design of release section .................................................................................................... 24

5.2.4

Structural analysis ............................................................................................................... 26

5.2.4.1

Analysis of structure during storage ................................................................................ 26

5.2.4.2

Structural design and analysis after paddle deployment .................................................. 27

5.2.5

Mass characteristic analysis ................................................................................................ 30

5.3

Mechanism design and analysis ............................................................................................. 30

5.3.1

Expansion mechanism design ............................................................................................. 33

5.3.2

Synchronization mechanism design.................................................................................... 33ii

JERG-2-215A

5.3.3

Deployment design and analysis ......................................................................................... 34

5.4

Thermal design and analysis .................................................................................................. 36

5.4.1

Thermal design ................................................................................................................... 36

5.4.2

Thermal analysis ................................................................................................................. 39

5.5

Paddle drive mechanism design ............................................................................................. 41

5.5.1

Electric design .................................................................................................................... 41

5.5.2

Machine design ................................................................................................................... 43

5.5.3

Thermal design ................................................................................................................... 45

5.5.4

Reviewing environmental resistance .................................................................................. 45

5.5.5

Clarifying test plan and test items ....................................................................................... 47

5.6

Paddle drive control design and analysis ............................................................................... 47

5.7

Space environment design ..................................................................................................... 49

5.7.1

Radiation design and analysis ............................................................................................. 49

5.7.2

Ultraviolet (UV) countermeasure design and analysis ....................................................... 52

5.7.3

Atomic oxygen (AO) countermeasure design and analysis ................................................ 53

5.7.4

Debris/micro meteoroids countermeasure design and analysis .......................................... 53

5.8

Anti-charging/discharging countermeasure design and analysis ........................................... 54

5.9

Installation design .................................................................................................................. 54

5.10 Definition of interfaces .......................................................................................................... 56
5.10.1 Definition of external interfaces ......................................................................................... 56
5.10.2 Interface demarcation ......................................................................................................... 56
5.10.3 Interface conditions............................................................................................................. 56
5.10.3.1 Electrical interface conditions ......................................................................................... 56
5.10.3.1.1

Power supply system interfaces ................................................................................. 57

5.10.3.1.2

Telemetry/command interfaces .................................................................................. 57

5.10.3.1.3

Attitude control system interfaces .............................................................................. 58

5.10.3.1.4

Other electrical interfaces........................................................................................... 59

5.10.3.2 Mechanical interface condition ....................................................................................... 59
5.10.3.3 Thermal interface condition............................................................................................. 59
5.10.3.4 Dynamics interface conditions ........................................................................................ 60

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5.10.3.5 Visual field interface ....................................................................................................... 61
5.10.3.6 Movable section envelope interfaces ............................................................................... 61
5.10.3.7 Interfaces with ground support equipment ...................................................................... 61
6. Verification··································································································· 62
6.1

Deployment/verification plan ................................................................................................ 62

6.2

Manufacturing and testing ..................................................................................................... 65

Appendix I Definition of name ........................................................................................................... 69iv

JERG-2-215A

1. Scope
This design standard is specified to be applied to the lightweight, rigid-type solar
array wing system of the spacecraft developed by the Japan Aerospace Exploration
Agency (JAXA). The content of this design standard specifies the key requirements
to be applied to the design and development of the lightweight, rigid-type solar
array wing system.

2. Applicable documents
JERG-2-141

Space Environment Standard

JERG-2-143

Radiation Resistance Design Standard

JERG-2-144

Space Debris Protective Design Standard

JERG-2-152

Disturbance Control Standard

JERG-2-200

Electric Design Standard

JERG-2-211

Charging/Discharging Design Standard

JERG-2-212

Wire Derating Design Standard

JERG-2-213

Insulation Design Standard

JERG-2-214

Power Supply System Design Standard

JERG-2-310

Thermal Control System Design Standard

3. Terminology, Definition and Abbreviation
Shown in Appendix I.

4. General requirements
4.1 General
The spacecraft solar array wing system converts solar light into electricalenergy, and most of the functional portions are placed outside the spacecraft’s
structure. The system must generate electricity required by the spacecraft overthe entire mission period while being exposed to “thermal and mechanicalenvironment during launch and ascent” from launch to spacecraft separation
and a “hostile space environment” after separation for an extended period of
time.

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4.2 Composition and components
A lightweight, rigid-type solar array wing is a solar array wing having an
aluminum honeycomb sandwich structure covered by CFRP as the substrate
and a solar cell panel with a solar array affixed as the main component.
Figure 4-1 shows an example of the general composition of a lightweight,
rigid-type solar array wing system.
Because the necessity of each component differs depending on the mission
requirements and power generation requirements of the spacecraft, thenecessary function configuration shall be specified in accordance with themission requirements of each spacecraft in the course of development. For alarge solar array wing system requiring a large amount of power, it is difficult
in some cases to verify the entire paddle system due to restrictions of theground test equipment. In the verification of this development, in addition to
performing a series of verifications on each component, it is necessary to createa proper development plan and verification plan according to the solar array
wing system configuration and check the validity of the design and thesoundness of the flight products.
Solar array wing system

Solar array wing

Solar cell panel

Solar array
Substrate

Boom and yoke

Holding and releasing
mechanism

Paddle deployment synchronization mechanism

Holding mechanism

Preload (hold-down)Bolt, wire, etc.

Releasing mechanism

Bolt cutter, explosive device, etc.

Deployment mechanism

Synchronization mechanism

Paddle drive mechanism

Deployment hinge, deployment spring
Deployment status switch
Latch mechanism
Damper, etc.

Tension adjustment mechanism, anti-vibration mechanism
Synchronization wire, pulley etc.

Slip ring
Actuator

Step motor
Harmonic drive, etc.

Others• Solar tracking sensor• Accelerometer• Temperature sensor etc.

Figure 4-1 Example of lightweight, rigid-type solar array wing system configuration

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JERG-2-215A

5. Subsystem design and analysis
5.1 Electric design and analysis
This section describes the design points to note in solar array wing electric design
and analysis.
5.1.1

Power generation design

In power generation design, the power required by the system shall be satisfied
throughout the mission period with particular attention paid to the following design
items and design contents.
(1) Selection of solar cell type and review of array circuit configuration (number ofseries-parallel circuits)(2) In reviewing the number of series-parallel circuits, consideration shall be given
to the influence of series resistance, temperature characteristics, deterioration
due to influence of space environment such as radiation and cover glass,adhesive loss and the like in performing the power generation analysis.
Figure 5-1 shows an abbreviated flow of power generation analysis.

Rigidity requirement• At launch storage• On-orbit paddle deployment

Power generation
requirements through mission
periodIs it necessary to review
the design preconditions
including system
adjustment?

Power generation analysis
System structure I/F(retention point/envelope
range)

Paddle/panel
structural design

Calculation of number ofrequire series parallels(calculation of necessary
surface area)

Characteristics of array circuit components assumed to be used
On-orbit radiation environment (radiation deterioration prediction)

Cell temperature based on thermal analysis

Implementation of
power analysis
iteration inaccordance withreview of panel size
and the like

Panel size holddown position

Array circuit configuration
design (layout design)

Shunt I/F

Implementation of power analysis
iteration in accordance withreview of array circuitcomponents

Are the paddle design standard
requirements and system I/F
requirements satisfied?

[Determining power generation design]• Determining array circuit components• Number of series-parallel circuits, string
configuration• Array circuit layout

Design condition output of“PDM interface design”
and “shunt interface
design”

Figure 5-1 Abbreviated flow of power generation design

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5.1.1.1

Array circuit component design

The following contents shall be clarified in designing the array circuit components.
(1) Solar cell
The solar cell appearance and following specifications shall be clarified. When adedicated charge array is provided, the specifications shall be clarified.
(a) Cell type(b) Output parameter (short circuit current Isc, open voltage Voc, maximum
power current Imp, maximum power voltage Vmp)
For the maximum power current Imp, the average minimum load current (minimum
average) in the bare cell shall be clarified in consideration of the cell grade.
(c) Power generation efficiency in BOL and EOL(d) Cell size (length × width × thickness)(e) Average mass

(2) Cover glass
The cover glass appearance and the following specifications shall be clarified.
(a) Glass type (model name)(b) Substrate material(c) Coating(d) Glass size (length × width × thickness)(e) Average mass(f) Solar light transmittance and UV deterioration rate(3) Interconnector and bus bar
The interconnector, bus bar appearance, shape and the following specifications
shall be clarified. In manufacturing the interconnector and bus bar, design
instructions shall be given so that no burrs are left in the edge.
(a) Material(b) Thickness(c) Average mass

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JERG-2-215A(4) Bypass diode for shadow countermeasures
The appearance of a bypass diode for shadow countermeasures and the following
specifications shall be clarified.
(a) Forward voltage Vf(b) Reverse current Ir(c) Shape(d) Average mass
(5) CIC (Cover Integrated solar Cell)
Out of the above-mentioned solar cell circuit components, solar cells with theinterconnector welded and the cover glass adhered are generally called CIC. The
following specifications shall be clarified with respect to such assemblies.
(a) Interconnector weld strength (45 degree peel strength)(b) Surface characteristics under CIC state (α, ε)(c) Output grade classification under CIC state

5.1.1.2

Power generation analysis

The power generation specified point shall be the paddle drive mechanism output
terminal. The power generation at the interface voltage with the system shall beanalyzed.
Based on the result of thermal analysis and radiation analysis, the required
number of series-parallel circuits shall be designed after indicating the V-I curvecharacteristics of the entire solar cell array circuit satisfying the power required by
the system throughout the mission period. In addition, the trend of power
generation between BOL and EOL in each season (spring, summer, fall and winter)shall be clarified.
In power generation analysis, the solar cell output parameters shall be considered
properly as a precondition for analysis in accordance with the development phase.
•

Design phase: Consideration shall be given to the solar cell output distribution
indicated in the solar cell application data sheet and the like and the output
grade distribution in CIC state accordingly.

•

Manufacturing and test phase: Reflection of array circuit output characteristics
measured in an irradiation test.

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JERG-2-215A

In addition to the above-mentioned analysis preconditions, the following analysis
parameters shall be clarified before performing power generation analysis.

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JERG-2-215A(1) Solar light strength
The details of AM0 solar light strength (solar constant) at one astronomical unit
(1 AU = average distance between the earth and the sun), which shall be inaccordance with JERG-2-141 Space Environment Standard, are given below as
reference.
[1] Solar light strength = 1336 W/m2@1AU[2] Uncertainty of solar light strength (including variation due to the number ofsunspots and measurement errors) = ±10 W/m2
[3] The approximate values of seasonal variation of solar radiation strength due
to changes in the distance between the sun and the earth (excluding
uncertainty of the above-mentioned [2]) are shown in Table 5-1. The error(1σ) in the table is 0.6 W/m2.
Thus, the solar light strength used in power generation analysis shall beroughly expressed in the formula below:
Solar light strength max/min = (1366±10 W/m2) × (seasonal variation of thesun-earth distance change +3.36%/-3.22%)
Table 5-1 Approximate value of seasonal variation of solar radiation strength
MonthSolar radiation strength W/m2
January 3 (perihelion)
1412
February
1406
March
1391
April
1368
May
1345
June
1329
July 4 (aphelion)
1322
August1326
September
1342
October1363
November
1387
December
1405(2) Solar light incident angle
In the case of a static earth-orbiting spacecraft, a solar light incident angle of±23.4 degrees shall be taken into consideration. As for the polar-orbit spacecraft,the cant angle shall be taken into consideration at the cant paddle at the local
time of descending node passage.
(3) Paddle tracking angle error
Consideration shall be given to the tracking error estimated as the base of7

JERG-2-215A

analysis.
(4) Solar cell temperature
Consideration shall be given to the temperature prediction error marginestimated with respect to the orbital nominal prediction temperature obtained
from the thermal analysis result.
See the technical data for the estimated temperature prediction error margin
track record of earth-orbiting spacecraft.
The temperature prediction error margin corresponds to the uncertainty ofdesign and equal to <maximum prediction temperature range> - <designprediction temperature range>. Normally, a temperature prediction error
margin of 10°C or more is used. In the case of use in a similar environment withan orbital track record, a proper temperature margin shall be specified based on
the technical judgment of the project(5) Preservation ratio of solar cell output parameters in accordance with radiation
exposure(6) Solar cell each output parameter temperature coefficient
Consideration shall be given to the temperature coefficient in accordance withradiation exposure.
(7) Cover loss and ultraviolet radiation deterioration rate(8) Line drop (harness resistance)
Consideration shall be given to the isolation diode, interconnector resistanceand their temperature characteristics.
(9) Measurement error of paddle V-I characteristics
Consideration shall be given to the light source setting error, illumination
distribution error and measurement error as error factors of V-I characteristics
measurement using the solar simulator.
(10) Yaw steering and positional error
Consideration shall be given to the attitude angle profile in accordance with theyaw steering and the positional error with respect to the profile.
(11) Paddle having a reflecting plate or condensing plate

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JERG-2-215A

For a paddle having a reflecting plate or condensing plate, consideration shall begiven to the reflection and condensing efficiency.
(12) Shadow influence
When the spacecraft structure or loaded objects cast a shadow on the paddle cellsurface, power analysis shall be conducted in consideration of the influence ofthat shadow.

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JERG-2-215A

When adopting a solar cell having an IBF (Integrated Bypass diode Function)function, care shall be taken not to allow reverse bias calorification due to the
influence of shadow (especially in the solar cell range where a shadow is cast) tocause a high-temperature state exceeding the allowable range of the solar celland substrate.
Additionally, consideration shall be given to the following factors as uncertain
errors/margins.
(a) Solar cell circuit random failure
About 5% of the power generation or one-array circuit, whichever is larger, is
recommended as the random failure margin due to space debris.
Refer to the technical data for the track record of power generation variation inearth-orbiting spacecraft.
(b) Loss due to ESD
The details shall be in accordance with JERG-2-211Charging/Discharging
Design Standard and the influence of ESD (Electrical Static Discharge) shall bereflected in the power generation analysis based on the following review.
i) The discharge occurrence frequency during the mission period shall beanalyzed.
ii) The power generation deterioration rate shall be grasped by conducting a
coupon panel test as needed.
iii) The power generation deterioration rate shall be analyzed from theabove-mentioned two conditions.
5.1.1.3

Array circuit configuration (layout) design

After determining the required area as the panel based on the number ofseries-parallel circuits obtained from the power generation analysis, an arraylayout shall be designed in consideration of the following points.
(1) The number of parallel circuits in one-array circuit
Array circuit (parallel array circuit family corresponding to one shunt)configuration in the range satisfying I/F with the power supply system shunt
element. The one-array circuit current upper limit for avoiding sustaining
discharge shall be in accordance with the requirements specified in JERG-2-211
Charging/Discharging Design Standard or the following (2) isolation diode shallbe installed for every series circuit.

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JERG-2-215A(2) an isolation diode for every series circuit in the array circuit
In view of failure separation and avoidance of sustaining discharge,
consideration shall be given to the installation of an isolation diode for every
series circuit.

11

JERG-2-215A(3) CIC output grade grouping
After assembling the CIC, obtain the output characteristics for each CIC,
perform grouping for each output level and apply to the assembly of the CIC
module so that no specific CIC will be the output-limiting cell.

(4) CIC layout position
In view of efficient power generation, in consideration of wiring loss and shunt
I/F upper limit current, the CIC module layout shall be optimized in accordancewith the output grade.
(5) Considering laydown around holding point
While adjusting with the holding point position during storage required fromthe structural design, CIC laydown shall be performed around the position inconsideration of the following viewpoints.
(a) Avoid CIC laydown to places where shadow is cast steadily due to a bolt
catcher and the like.
(b) Ensure a distance for avoiding occurrence of sustaining discharge if thearray high potential portion is close to the hold-down section.
(6) Considering potential difference between adjacent cells
In designing the solar cell array layout, a potential difference between different
array circuits or between folded adjacent cells shall be laid out to meet therequirements specified in JERG-2-211 Charging/Discharging Design Standard
for avoiding sustaining discharge. When evaluating the potential difference
between other array circuits, consideration shall be given to the potential
difference due to shunt operation.
(7) Controlling magnetic moment
To control magnetic moment which develops on the solar cell panel, takeaccount of a design that can cancel the magnetic moment developed in the solar
cell panel in consideration of the following points regarding cell layout and

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JERG-2-215A

harness routing.
(a) In the layout, the polarity of the solar cell array circuit in the solar cell panel
shall be line-symmetric to the greatest extent possible.

13

JERG-2-215A(b) In the harness layout from each circuit of the solar cell array to the connector
which interfaces with the paddle drive unit, occurrence of the magnetic field
associated with the current loop shall be minimized.
(8) Setting the bypass diode for shadow countermeasures
If a shadow of the structure mounted in the spacecraft is cast on the solar cell panel surface
when the spacecraft orbits in a general attitude or attitude is lost, the shaded cell will be in areverse bias state. The worst case is when only one cell in one series circuit is totally shaded.
A bypass diode for shadow countermeasures must be installed to prevent a breakdown
without exceeding the reverse bias withstand voltage of the cell. Because the general siliconcell has a reverse withstand voltage of about 30 volts, it is desirable to install a bypass diodefor shadow countermeasures for every 30 series circuits. As the reverse withstand voltage
has temperature dependence, the number of insertion steps shall be set after grasping thetemperature characteristics and unevenness of the solar cells to be used in consideration ofproper derating (50% or less recommended). Because the III-V compound-based solar cell
has a low reverse withstand voltage in general, the bypass diode for shadow
countermeasures shall be installed on a cell-by-cell basis. In selecting bypass diode parts forshadow countermeasures, a model which is capable of passing electric current for one series
circuit in consideration of derating (50% or less recommended) shall be chosen. Some cells
are integrated with a bypass function (IBF; Integrated Bypass Function). When IBF cells areused, it is possible to simplify the panel integration because the independent shadow diodecan be omitted.
(9) Charge array layout
For a solar array wing having a charge array, consideration shall be given to the following
points with respect to the layout of the charge array.
(a) Consideration to risk of power reduction due to ESD
Based on charging analysis, consideration shall be given to the array layout to thelocation which minimizes the risk of reduction in power generated by the charge arraydue to ESD.
(b) Minimizing influence of shadow cast by satellite body
Consideration shall be given to the array layout to the position which minimizes the
reduction in power generation due to a shadow of the satellite body cast on the charge
array.

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JERG-2-215A

5.1.1.4

Design of interface between PDL and paddle drive mechanism

As an electrical interface design of a PDL-to-paddle drive mechanism, consideration
shall be given to the following design points with respect to:1) Paddle drive mechanism slip ring design;2) Array circuit occurrence voltage and slip ring withstand voltage design;
3) Processing of spare ring; and
4) Design notes for ground line.
(1) Paddle drive mechanism slip ring design
Based on the design of the power generation analysis and telemetry (such astemperature sensor line and angle detection signal line)/command and design of theheater line and ground line, electric design of the slip ring (ring current capacity
and the number of rings) required in the paddle drive mechanism shall be
performed. In selecting the slip ring current capacity, consideration shall be given
to derating (50% or less recommended).
In determining the current capacity of the hot-side ring for electric power, the ring
capacity selection condition differs depending on whether the shunt is mounted tothe solar cell array side (so-called yoke mounted) from the paddle drive mechanism
or to the spacecraft structure side. When the shunt is mounted to the structure side,
it shall be necessary to select the hot-side slip ring for electric power having thecurrent capacity corresponding to the one-array circuit connected to the shunt
element. When the yoke is mounted, the large-capacity slip ring can be used in therange where both HOT and RTN rings can be selected.
In selecting the slip ring for electric power capacity selection, consideration shall begiven to derating with respect to the worst case including abnormal attitude andthe like (conditions under which generation current is maximized (BOL, solar lightvertical incidence, high temperature condition and array circuit shunt state)).
Capacitive surge current due to charging/discharging of nongrounded insulatorssuch as solar cell array cover glass may pass through the slip ring, harness andconnector (including connector pin). Although a surge current due to discharge isabout 100 A at the maximum and it continues for about several tens of μ seconds(depending on insulator capacity and charged potential), it is desirable that thiscapacitive surge be considered in selecting the capacity of the slip ring and the like.

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JERG-2-215A(2) Array circuit occurrence voltage and slip ring withstand voltage
For the withstand voltage between adjacent slip rings and the withstand voltage
between brushes, consideration in design shall be given to the creepage distanceand spatial distance having a proper discharge tolerance with respect to the
maximum voltage (PDM output terminal voltage in electric power source (EPS)
system interface) generated by the solar cell array and other signal voltages. In
designing the withstand voltage distance, consideration shall be given to influenceof abrasion powder due to sliding between the slip ring and the brush. Open mode
failures related to the harness and connector shall be always reviewed with respect
to the worst-case voltage.
(3) Processing of spare ringIn view of charging and discharging (see JERG-2-211Charging/Discharging DesignStandard for details), the spare ring shall be grounded so as not to leave any
suspended conductive materials. When performing grounding, it is desirable toinsert resistance to the empty ring in which a signal capable of passing a large
current is laid out to the adjacent ring.
(4) Design precautions to ground line
When employing a design on the premise that the substrate ground fault of thesolar cell array is permitted without mounting the breeder resistance (ground line
resistance), the number of slip rings and ground line wire current capacity which
are properly within the allowable range with respect to the estimated maximum
ground fault current shall be designed.
5.1.1.5

Shunt interface design

In accordance with the power generation analysis, it shall be clarified that thedesign satisfies the maximum allowable current with respect to the shunt circuit.
As the electrical interface design between PDL and shunt, the following design
consideration shall be given to the current and voltage.
(1) Current interface
For the current capacity design of the one-array circuit, the allowable current inlight of derating shall not be exceeded in the worst case including attitude
abnormalities (conditions under which the generation current is maximized (BOL,
solar light vertical incidence, high temperature condition and array circuit shuntstate)).

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JERG-2-215A(2) Voltage interface
For the design where the shunt is mounted to the solar cell array side (so-calledyoke mounted) from the paddle drive mechanism, when brush chattering is likely tooccur due to an environment where the equipment undergoes a physical shock,consistency with the shunt element withstand voltage shall be checked with respect
to the worst case voltage (BOL, low-temperature condition and array circuit openstate) as array circuit occurrence voltage. In case of inconsistency, it shall beclarified that measures be taken with respect to the power supply system shunt
design.

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JERG-2-215A

5.1.2

Operation mode and power consumption design

The power consumption of the paddle system subsystem in each operation mode shall be
specified. Mainly, a paddle drive mechanism driving power and damper heater power
shall be included.
5.1.3

Electromagnetic compatibility design

Design requirements for grounding system design, bonding and electromagnetic
interference limit shall be entered.
5.1.4

Signal design

5.1.4.1

Suspended capacity of signal wiring

In telemetry design of the solar array wing, PA (passive analog) telemetry may be
used for deployment status signal and panel temperature. In the solar array wing,
the time constant related to wiring stray capacitance will be large due to the
influence of these lengthened signal wiring routes and temperature(low-temperature environment). This is the large solar array wing ground test and it
is difficult to check consistency with the spacecraft system in simulation of the
low-temperature environment in an End-to-End test. Thus, consideration shall begiven in design and analysis to consistency between the data sample timing of thedata handling system and the time constant of the above-mentioned PA telemetry
signal (reduction in stray capacitance, checking of sample timing interface
consistency, design corresponding by active analog or deployment status design todecrease the time constant after deployment).
5.1.4.2

Consideration of influence of noise from the power line associated with shunt
operation

The signal lines (such as temperature telemetry) of the solar array wing aresusceptible to noise from the power line associated with shunt operation. In large
solar array wing ground test, it is difficult in this part to check the consistency withthe spacecraft system in an End-to-End test. Thus, consideration in design shall beapplied to the signal lines such as adoption of shield lines or separation from thepower line in the line routing in the range free from troubles in deployment
reliability as necessary. If it is difficult to deal with noise, the signal lines susceptible
to noise shall be designed with a function only for a telemetry monitor and not be
used for on-board control.

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5.2 Structural design and analysis
This section describes the design points to note in solar array wing electric design
and analysis.

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JERG-2-215A

The design items to be checked by carrying out structural design and analysis shallbe as follows. The structural design flow is shown in Figure 5-2.
• In the solar array wing storage state, the structure shall have the necessary
strength to withstand the load conditions provided from the spacecraft systemand satisfy the rigidity requirements for the solar array wing system. In
structural design and analysis, consideration shall be given to acceleration atlaunch, acceleration during a sine wave vibration test, cumulative fatigue,
acoustic environment, shock environment and heat load between the structureand solar array wing. When another subsystem device is mounted to the solar
array wing system, a proper interface shall be set to the mechanicalenvironment of the section where the device is mounted. Consideration shall begiven to load concentration due to uneven rigidity in the system structure sidein the solar array wing mounting section.
• During solar array wing deployment operation, the structure shall have aproper strength against shocks from the deployment latch.
• In the solar array wing deployed state, the structure shall have strength totolerate external disturbance caused by spacecraft orbit control and attitudecontrol and rigidity to satisfy interface requirements with the attitude control
system. Interface requirements with the attitude control system shall besatisfied with respect to the orbital disturbance in which the solar array wings
can be a source of vibration such as thermal snap.

20

JERG-2-215A

Structural design flow

[Design precondition setting]• Mass, mass characteristic• Load condition (quasi-static acceleration, sine wave
vibration, sound, heat)• Rigidity requirement• Layout restrictions (surface area, hold-down position) etc.

[Review of paddle configuration]• Shape (during storage and at deployment)• Component strength• Component rigidity• Installation design, etc.

AS NEEDEDAS NEEDED[Material property data acquisition]
Young's modulus, Poisson's ratio,
coefficient of linear deployment,
strength, space environment tolerance(For new materials)[Structure characteristics dataacquisition]• Hinge rigidity, strength• Paddle drive mechanism rigidity[Selecting structural analysis case]• Paddle storage state
Natural frequency, quasi-static acceleration, sine wave
vibration, heat distortion (as needed)• Paddle deployment state
Natural frequency, quasi-static acceleration, heat distortion,
heat-induced dynamics[Building structural mathematical model][Structural analysis]

Is it within design precondition?
Are rigidity requirements and strength
requirements satisfied?
Are orbital disturbance requirements satisfied?

[Determining structural design]
Figure 5-2 Structural design flow21

Is it necessary toreview the designpreconditions
including system
adjustment?

JERG-2-215A

5.2.1

Panel substrate design

The substrate is a panel mounting the solar cell array which is a major part of thesolar array wing.
The design items related to a substrate shall be as follows:• A substrate is a major structural element of the solar array wing and usually has
a CFRP face sheet honeycomb panel structure to realize lightweight and highrigidity. Because the CFRP face sheet on the substrate surface is a conductive
material, a polyimide film must be attached to the surface mounting the solar cellarray to prevent a short circuit. Before starting laydown, the polyimide film shallbe checked for insulation resistance.
• The honeycomb core shall be checked to confirm whether perforations areproperly open before using. When using t10 (10 mm thick) or less, while noting
the influence of absence of perforations.
• To prevent the solar cell surface from reaching a high temperature, emissivity
shall be improved by affixing some heat control material or increasing the surface
roughness.
• The substrate hold-down shall be surface-treated to prevent the joint surface fromadhering in orbit or in use with the surface-treated pad pinched. Care shall betaken to prevent the hold-downs from coming into contact with each other during
deployment due to the linear deployment difference between the spacecraft
structure and the paddle, thereby hindering deployment.
• Because the substrate is often exposed to a severe temperature environment,
when selecting the material to be adhered to CFRP, the difference of linear
deployment coefficients shall be as small as possible. In using adhesive, proper
thickness shall be controlled. Care shall be given so as not to develop cracking
under a thermal environment. Materials having a glass transition point within
the operation temperature range shall be evaluated using the property values at
the maximum and minimum temperatures.
• All metal materials to be mounted to the substrate shall be electrically coupled tothe conduction reference point (substrate ground point) which is electricallycoupled to the satellite. By so doing, care shall be taken so that no potential
difference is generated. When the honeycomb core is coupled, care shall be taken
so that electrical coupling between honeycomb cores is ensured.
• Consideration shall be given so that no damage occurs due to the difference ofheat distortion between the substrate and the spacecraft structure when thesubstrate is mounted to the spacecraft.
• Contact prevention measures shall be taken with portions where displacement islarge under a vibration environment such as substrate angle and center sections.

22

JERG-2-215A

Consideration shall be given so that no joints are adhered in orbit by performing
solid lubrication or ceramics film treatment on contact-prevention measuresportions so that deployment is not hindered.

23

JERG-2-215A

5.2.2

Boom and yoke design

The boom and yoke shall have a function to deploy a solar cell panel from thespacecraft body to a proper distance in orbit and maintain the panel. The design ofboom and yoke shall be determined in consideration of the range of shadow
generated by spacecraft-loaded objects, mass characteristic requirements,
spacecraft body structural element layout (hard point), solar array wing
synchronization mechanism layout and harness layout.
The design items related to boom and yoke shall be as follows:• The length of boom and yoke elements shall be determined in accordance with therequirements for a spacecraft system (analysis of shadow formed by objects loadedin the spacecraft, mass characteristic analysis).
• Check the distance between the boom/yoke elements and solar cell panel. Element
design shall be done so that no boom/yoke elements will come into contact withthe panel even when the elements are vibrated under a vibration environment.
• A harness wiring shall be located in a place out of direct solar light where possible
in consideration that harness temperature is likely to increase in portions close to
the spacecraft structure due to reflected heat of the structure. If it is inevitable toinstall wiring under direct solar light, shielding shall be provided by MLI or SLI.
• Consideration in design shall be given to adhesion in the hold-down section as in
the case with the substrate.
• Consideration in design shall be given to bonding to CFRP elements as in the casewith the substrate.
• The boom and yoke shall be electrically coupled to all the metal materials to be
mounted in the boom and yoke.
5.2.3

Design of release section

The release section shall have a function to fix and hold objects such a solar cellpanel and boom and yoke with bolts at launch of the spacecraft and release theparts in orbit with the releasing parts including explosive devices incorporated.
The design items related to release section shall be as follows:• During releasing operation, it is desirable that the release section will come into
contact with no other structural elements regardless of the locus of the preloadbolt. In case of contact, it shall be fully checked that snagging will not hinder
deployment.
• To prevent the preload bolt from moving backward after releasing the holding
state, preventive measures shall be taken such as providing an unreturning
spring material.
• Consideration in design shall be given to adhesion in the hold-down section as in24

JERG-2-215A

the case with the substrate.
• In consideration of conditions of installation to the system structure, the template
for locating the holding console shall be prepared as needed.

25

JERG-2-215A• When installing an explosive device, the explosive device shall be located easily.
There shall be easy access to the explosive device.
• The coupling interface with the system side EED harness shall be clarified. Check
the presence/absence of MLI on the EED harness side. Check for interference withthe deployment locus of the paddle structural elements.
• When installing the MLI to the holding console, an interface with the system
structure side MLI shall be clarified with regards to aspects such as thepresence/absence of the tape.
• The release section shall be electrically bonded to the explosive device to beinstalled using wires and the like.
• For shock release, a proper interface with the system structure shall be specified.
When conducting a verification test at the subsystem level, the configuration suchas jig composition shall conform to the actual machine.
5.2.4

Structural analysis

5.2.4.1

Analysis of structure during storage• Analysis of characteristic values during storage shall be conducted to check thatthe rigidity requirements to the solar array wing are satisfied.
The boundary conditions with the system structure shall be clarified (completely fixed ingeneral).
• Acceleration load strength analysis at launch and strength analysis during thesine wave vibration test shall be conducted to check that the strength
requirements for the solar array wing are satisfied.
Examples of strength assessment shall be as follows: In the case of evaluating the strength ofan actual machine, the test result of a model simulating the actual machine shall be used
where possible.
- Material fracture (Tsai-Wu)- Face sheet interlayer shear- Skin buckling (intra cell buckling (*1) or wrinkling (*2))- Insert joint evaluation (such as opening and slip)(*1) Fracture mode in which the honeycomb panel face sheet buckles in therange within the honeycomb core cell.
(*2) Fracture mode in which the honeycomb panel face sheet buckles acrossthe wall surface of the honeycomb core cell core due to a short wavelengthand the panel is damaged as the load is not transferred.

26

JERG-2-215A

5.2.4.2

Structural design and analysis after paddle deployment(1) Flexible structure mode after deployment
The subsystem side of the solar array wing system shall analyze the flexible structure of
the solar array wing system after completion of normal deployment in orbit by using
finite element methods for structural analysis (such as finite element method,
differential equation numerical solution and differential equation analysis solution).
Modelization and analysis shall be carried out on the entire solar array wing system.
When the system has a paddle drive mechanism, a paddle drive mechanism shall be
included. In analysis, fix-free boundary conditions in which the spacecraft bodyinstallation section is assumed to be a completely rigid interface surface shall be applied.
The natural frequencies from the lowest order mode to the waveband specified incoordination with the spacecraft system side in this analysis and the corresponding
zero-order and primary coupling coefficient vectors and the mode-damping ratio shall be
indicated. These flexible structure parameters and corresponding allowable errors shallbe specified and provided to the spacecraft system. It shall be checked that theseflexible structure parameters and corresponding allowable errors are within theallowable range of the flexible structure parameters specified by the spacecraft system.
The spacecraft system side shall specify the flexible structure parameters under fix-freeboundary conditions provided from the subsystem side to the spacecraft model andobtain the flexible structure parameters under fix-free boundary conditions, thereby
performing analysis of spacecraft dynamics, attitude control and attitude stability using
the model. Consideration shall be given to the flexibility of the installation section in
the spacecraft body side. The flexible structure parameters from the spacecraft systemside and corresponding allowable errors shall be specified. It shall be required to specify
not only the minimum natural frequency but also the allowable waveband of the
low-order natural frequency in accordance with the need for attitude control andstability.
The above-mentioned analysis, regulations and evaluation shall be carried out on major
deployment abnormality modes as needed.
Verification of the flexible structure mode after deployment shall be carried out at leastby building the entire mode through buildup of rigidity measurement results ofindividual parts or similar parts. In actual machine verification at the subsystem level,
the rigidity in a deployed state shall be statically measured to check the natural

27

JERG-2-215A

frequency of the low-order flexible structure mode in a deployed state.

28

JERG-2-215A(2) Orbital heat distortion analysis
The subsystem side shall perform analysis of orbital temperature variations in eachportion after reaching a thermally stationary cycle state with respect to the solar array
wing system which completed deployment in orbit by using the finite element methods
for thermal analysis (such as finite element method, differential equation numericalsolution and differential equation analysis solution). Consideration shall be given to
thermal influence of the spacecraft body and rotation of the paddle (spacecraft body) inanalysis. Any temperature change in both sides of the paddle shall be modeled. The
obtained temperature distribution profile shall be entered to the finite element methods
for structural analysis (such as finite element method, differential equation numericalsolution and differential equation analysis solution) to obtain the orbital deformation
profile of the solar array wing system in a non-controlled attitude state (quasi-static
heat distortion profile and dynamic response profile) and the installation point torque.
The transient responses consisting of the quasi-static heat distortion variations anddynamic responses between equilibrium state under irradiation and equilibrium statein the shade which occur immediately after the start and end of an eclipse shall becalled “heat-induced dynamics.”(3) Strength analysis
Strength analysis shall be carried out on each section of the solar array wing systemwith respect to the following load inputs using the finite element methods for structural
analysis (such as finite element method, differential equation numerical solution anddifferential equation analysis solution) to check that M.S. is positive.
a) Acceleration, angular acceleration input to the solar array wing system installation
section at thruster injection and application of large torque: Consideration shall begiven to the barycenter offset resulting from deformation due to heat distortion in
the solar array wing system as well as inertia.
b) Deformation of the solar array wing system due to thermal factors during orbit(quasi-static heat distortion profile and dynamic response profile) based on analysisin (2).
c) Coupled vibration of paddle drive mechanism and solar array wing and spacecraft
body associated with paddle driving.
For b) and c), the fatigue tolerance to the cycle applied during mission period shall bechecked. It shall be necessary to check the tolerance of not only the mechanism but also
the structural elements such as substrate, electric parts such as solar cells and theircomposite articles.

29

JERG-2-215A(4) Heat-induced dynamics analysis
The subsystem side shall perform heat-induced dynamics analysis under fix-freeboundary conditions based on the result of transient response between equilibrium
state under irradiation and equilibrium state in the shade in heat distortion analysis in
the above-mentioned (2). Analysis shall be carried out in accordance with the finite
element method, motion equation numerical simulation and analysis solution. It shallbe necessary to provide the spacecraft system side with a paddle deformation profile,installation point reaction torque and barycenter variation obtained from analysis.
The spacecraft system side shall carry out heat-induced dynamics analysis including
contribution of the spacecraft body, other flexible structures and attitude control system
based on the analysis of the subsystem side, thereby evaluating the influence ofdisturbance affecting attitude control and attitude stabilization.
5.2.5

Mass characteristic analysis• Mass evaluation shall be carried out based on the measured values of the solar
array wing to check that solar array wing mass requirements are satisfied.
• Mass characteristic analysis in a storage state shall be carried out to check that thesolar array wing mass characteristic requirements are satisfied. Tuning of theanalysis model shall be carried out based on the measured values of mass
characteristic test to calculate the mass characteristic values with respect to all thecoordinate axes.
• Mass characteristic analysis in a deployed state in orbit to check that the solar
array wing mass characteristic requirements are satisfied. Tuning of the analysis
model shall be carried out based on the measured values of a mass characteristic
test to calculate the mass characteristic values with respect to all the coordinate
axes.
5.3 Mechanism design and analysis
Solar array wing mechanism parts (deployment mechanism, synchronization
mechanism) shall have a function to develop the solar cell panel in orbit and maintainthe deployment state over the entire mission period. The parts shall have the strength
to withstand external disturbance due to spacecraft orbit control and attitude control
and rigidity to satisfy interface requirements with the attitude control system.

30

JERG-2-215A

This section describes the design points to note in solar array wing electric design andanalysis.
Figure 5-3 shows the mechanism design flow.

31

JERG-2-215A

Mechanism design flow

[Design precondition setting]• Expansion time
• Torque margin• Expansion envelope range• On-orbit temperature• Structure interface load etc.

[Review of paddle configuration]• Shape (during storage and at deployment)• Expansion torque, deployment inhibition torque• Synchronization wire tension, play amount• Damper damping rate
• Micro switch resistance• Installation design, etc.

Is it necessary toreview the designpreconditions
including system
adjustment?

As needed

As needed[Expansion analysis dataacquisition]• Hinge deployment torque• Hinge sliding torque• Harness resistance torque• Damping rate

[Expansion strength data acquisition]• Hinge strength• Paddle drive mechanism strength[Selecting deployment action analysis case]• High-temperature case
• Low-temperature case
• Ground test prediction
Analysis shall be carried out on the deployment profile,
deployment time and latch shock level of each above-mentioned
case.

[Building deployment action analysis
model][Expansion action analysis]

Is it within design precondition?
Are deployment time and deployment envelope
requirements satisfied?
Are strength requirements and disturbancerequirements at deployment satisfied?

[Determining mechanism design]
Figure 5-3 Mechanism design flow

32

JERG-2-215A

5.3.1

Expansion mechanism design

The deployment mechanism shall deploy the boom, yoke and solar cell panel released inorbit at a constant speed and maintain the deployment state.
The design items related to deployment mechanism shall be as follows:• The sliding section at deployment of each hinge shall be surface-treated such as
solid lubrication to prevent adhesion.
• The bearing section shall be designed so that the clearance of the outer ring/innerring is ensured under a low-temperature environment in consideration of thetemperature range to be loaded.
• The hinge of one side inside the hinge line shall be provided with free play in thethrust direction and designed to absorb displacement difference due to linear
deployment difference between the panels sandwiching the hinge line.
• Calculation result of deployment torque margin derived from the following formula
shall be 1 or more. However, if the requirements cannot be satisfied, judgment
shall be made based on the project.
deployment torque margin = deployment torque/resistance torque -1(resistance torque = harness resistance torque + hinge sliding torque)• Harness resistance torque shall be designed based on the measured value using aharness equivalent to that of the actual machine where possible.
• Hinge sliding torque shall be designed based on the measured value using thehinge equivalent to that of the actual machine where possible.
• The telemetry to confirm the completion of deployment of the hinge shall be
detected and output by the micro switch provided in the hinge. When the micro
switch is installed, connect a different resistance for every hinge and design themechanism so that you can check which hinge completed deployment bymonitoring the resistance value.
• The mechanism shall be designed so that the result MS of deployment latch shock
evaluation will be positive. The damper shall be installed to relieve shock at thetime of latch deployment where appropriate.
5.3.2

Synchronization mechanism design

When multiple components such as a boom, yoke and solar cell panels in the solar array
wing are coupled by hinges and it is likely that a collision will occur between
components and with the spacecraft body at deployment action, it shall be necessary to
install a synchronization mechanism to synchronize the solar array wing with the
deployment angle of the hinges.
• The deployment synchronization mechanism shall be designed to provide thesynchronization wires with a proper tension in consideration of on-orbit
temperature variations at paddle deployment action.
• The synchronization wire section including caulking parts shall be checked inadvance for soundness by applying proof load.

33

JERG-2-215A• The mechanism shall have a structure not to develop slack due tonon-synchronization at deployment action to the extent possible. The deployment
action envelope of the synchronization wires shall be free from protrusion andheight difference so that no snagging occurs in case of contact.
• The synchronization wires installed in the boom and yoke are likely to causechanges in amplitude during a vibration test due to the free play of the wires.
Steady rests shall be installed at intervals so that no contact with the panels
occurs.
5.3.3

Deployment design and analysis

The subsystem side shall analyze the deployment behavior of the solar array wings inorbit by multi-body dynamics numerical simulations (such as multi-body dynamics
numerical simulation software, differential equation numerical solution, differential
equation analysis solution) to obtain the deployment configuration profile, deployment
angle profile, deployment signal profile, deployment time, latch shock torque andinstallation point reaction torque profile. Analysis shall be carried out at the moment of
releasing in the case of passive deployment by the spring and the like and immediately
before motor drive by the motor and the like as for active deployment. For the boundary
conditions of paddle fixed end, be sure to carry out analysis under not only fix-free
conditions but also free-free conditions including the satellite body capable of rotating
the space freely. (Free-free conditions shall be the minimum required conditions.) The
satellite body for analysis under free-free conditions may have a rigid body when otherflexible structures are not deployed. The subsystem side at least shall carry out analysis
under fix-free conditions. The subsystem side or the satellite system side or both sides
shall carry out analysis under free-free conditions. However, the satellite system side
shall bear the ultimate responsibility. The subsystem side shall interface the conditions
required in analysis under free-free conditions to the satellite system side. For fix-freeor free-free conditions, analysis shall be carried out at least on two cases including
assumed highest and lowest temperature cases. Among analysis parameters shall behinge torque, hinge deployment torque, harness resistance torque, damper resistancetorque, deployment synchronization mechanism tension, temperature, hinge offset andthe like. The values verified based on the results of element prototype tests shall be
used including characteristic changes due to temperature. It shall be required to carryout analysis of worst-case combinations of parameters which envelop the variation
range of analysis parameters to check the deployment reliability. Based on the analysis

34

JERG-2-215A

results, check whether the deployment time, latch shock torque, M.S. at latch shock anddeployment configuration envelope range satisfy the requirements. In addition, thesatellite system side shall check whether the attitude rate at deployment and attitude
angle at deployment satisfy the requirements.

35

JERG-2-215A

The subsystem side shall verify the deployment of the solar array wings in a ground
deployment test and model the deployment behavior of the solar array wings in the
ground deployment test in consideration of the contribution of aerodynamic force andfrictional force as well as varied boundary conditions (under fix-free conditions in aground deployment test in general), thereby performing analysis by multi-bodydynamics numerical simulation as in the case with on-orbit behavior. Tune the analysis
parameters of ground deployment to conform to the ground test results and reflect thetuning results in the on-orbit deployment analysis.
As needed, set the conditions in simulation of fluctuations in harness resistance torqueout of the assumed range, failures of damper and deployment synchronization
mechanism and conduct on-orbit deployment analysis to check for robustness during
deployment.
5.4 Thermal design and analysis
An example of thermal design flow in solar array wings is Figure 5-4.
5.4.1

Thermal design

In accordance with the equipment allowable temperature of parts and components ofthe solar array wing system, it shall be necessary to carry out thermal design of thesystem and clarify temperature control requirements with respect to the spacecraft
system. In accordance with JERG-2-310 Thermal Control System Design Standard andthe thermal design standard specified in the project, design policies regarding design
uncertainty (uncertainty of errors and tolerances) and design margin (allowance forunexpected occurrence) shall be specified.
[1] Precondition• Internal calorific value (consider the exothermic density as for equipment with high calorificvalue)• External heat environment (during test, transportation, storage, launch, on-orbit operation
mode and abnormal attitude)• Thermal interface conditions provided from the system side
• Equipment allowable temperature (characteristic values of parts and components) and rate ofallowable temperature change• Requirements of operating life and reliability• Power generation requirements (the conversion efficiency of the solar cell is decreased as the36

JERG-2-215A

temperature increases)[2] Thermal design37

JERG-2-215A• Basic structure of the paddle shall be determined in consideration of the following points:- Shape- Surface area
- Thermo-optical characteristics of cover glass and panel back side
- Installation method- Surface treatment- Temperature regulation point- Thermophysical property• Selecting thermal control materials (parts and materials)
Selection shall be performed in consideration of thermophysical properties (such as
thermo-optical characteristic, heat conduction, heat capacity), reliability, environmental
resistance (such as radiation, atomic oxygen, charging/discharging), process (quality,
reproducibility), outgas, cleanness, performance deterioration due to contamination, flighttrack record, availability, useful operating life and workability.
• Conductivity of thermal control materials and grounding• Optimization of heat radiation area and heating capacity (power)• Calculation of high and low worst-case predicted temperatures• Temperature measurement points and temperature sensors• Adhesive strength, outgas, deterioration (aging, temperature)• Vibration at launch and strength and separation under a reduced pressure• Influence of adhesion of objects separated from the spacecraft structure and their shadow[3] Test temperature• Qualification test: Temperature range higher than the maximum predicted temperature range
by 5°C or more and not exceeding each equipment allowable temperature (parts andcomponents)• Acceptance test: Temperature range outside the maximum predicted temperature range andnot exceeding the acceptance test temperature range of each equipment allowable temperature(parts and component)• Proto-flight test: Same as qualification test

38

JERG-2-215A

5.4.2

Thermal analysis• In accordance with the thermal interface conditions, a detailed thermal math model of solar arraywings shall be created. While improving accuracy by verification and co-relation, design review ofsolar array wings, on-orbit temperature prediction, thermal equilibrium test evaluation, design
validation and on-orbit evaluation shall be carried out.

• As thermal analysis cases, the worst case (including solar light strength, thermal interface
conditions and thermo-optical characteristics) shall be selected for both the high-temperature andlow-temperature cases of solar array wings. As for the high-temperature case, consideration shallbe given to full shunt. The high-temperature worst cases of harness bundles and connectors shallbe considered in consideration of self-heating.

• As needed, an interface thermal math model for system thermal analysis shall be created. The
validity shall be checked through testing and analysis.

[Reference: JERG-2-310 Thermal Control System Design Standard]• Predicted design temperature range: Nominal design temperature in consideration of seasonalvariations and aging deterioration• Maximum predicted temperature range: Temperature in consideration of uncertainty such as
difference between actual machine and thermal math model, analysis tools, design parameters,
manufacturing tolerance and measurement errors.

<design uncertainty> = <maximum predicted temperature range> - <predicted design temperature
range> Example: 10°C or more<design margin> = <equipment allowable temperature> - <maximum predicted temperature range>
Example: 5°C or more

39

JERG-2-215A

Thermal design flow

As needed[Thermal balance test][Design precondition setting]• External thermal environment• Thermal interface condition• Power generation requirements• Allowable temperature, etc.

[Review of paddle configuration]• Shape (during storage and at
deployment)• Cover glass characteristics• Panel backside characteristics• Thermal control material selection• Harness bundle• Installation design, etc.

[Building thermal math model][Selecting thermal analysis case]• High-temperature worst case
• Low temperature worst case
• Harness bundle high-temperature
worst case etc.

[Thermal analysis]

Within the allowabletemperature range?
Is the temperature margin
assured?

[Power generation
analysis]

Are power generation
requirements
satisfied?

[Determining thermal design]

Figure 5-4 Thermal design flow40

Is it necessary toreview the designpreconditions
including system
adjustment?
As needed[Thermooptical characteristic dataacquisition]• For new materials• When no data after
deterioration is available

JERG-2-215A

5.5 Paddle drive mechanism design

The functions of paddle drive mechanism shall be as follows:(1) Rotates the solar array wings around the 1 axis to direct the solar arraywings to the sun.
(2) Transfers generated power of the solar array wings to the spacecraft body.
(3) Transfers electric signals between the solar array wing and spacecraft body.
(4) Outputs drive shaft rotation position signals by the angle detector.
(5) Outputs drive shaft reference position signals by the reference point sensor(in accordance with system requirements).

5.5.1

Electric design(1) Actuator
The actuator shall be clarified after performing I/F adjustment with the attitudecontrol system/drive circuit with respect to the specifications of the following items.
(a) Drive characteristics• Excitation circuit method• Excitation sequence• Winding impedance/inductance• Drive voltage/current• Drive phase counter electromotive voltage• Electric power consumption, calorific value• Pulse rate
• Step angle(2) Slip ring assembly
The slip ring assembly shall be clarified after performing I/F adjustment with the power supply
system and attitude control system with respect to the specifications of the following items.

(a) Ring configuration: Rated current value and number of power ring (RTN), powerring (HOT) and signal ring41

JERG-2-215A(b) Maximum working current value: 1/2 or less of rated current value. It is desired
to consider the capacitive surge.
(c) Maximum working voltage: Maximum applied voltage to each slip ring(d) Connector: Type, number, pin assignment(e) Empty ring: Empty rings (reserve rings) which are not to be used and notconnected to anything shall be grounded in the system.
(f) Static resistance value Static resistance shall not increase at the stop. (However,temporary increase under a ground environment (when the value returns to thenormal value due to rotation) shall be allowable.)(g) Dynamic resistance The variable width of resistance value during rotation shallbe regulated.
(h) Short circuit prevention The barrier section shall have a proper height andwidth to prevent a short circuit between adjacent rings caused by the brushexceeding the barrier due to vibration, shock and the like. Interior of the slip
ring case shall be short-circuit proof.
(i) Countermeasures for discharge The distance between adjacent rings anddistance between brushes shall be designed properly to prevent discharge
between slip rings.
(3) Angle detector
The angle detector shall be clarified after performing I/F adjustment with the attitude control
system/drive circuit with respect to the specifications of the following items.
• Machine angle: (Definition of relationship between mechanical origin and angle detector
origin in particular)• Angle/output characteristic: (Including angle detector unit accuracy)• Input/output impedance• Input voltage• Number of working slip rings• Connector pin assignment(4) Selecting maximum working current value of harnesses, connectors and connector
pins• As in the selection of slip ring maximum working current value, the maximum current value
for the connector and connector pin used in PDM shall be 1/2 or less of the rated current
value.

42

JERG-2-215A

However, the total conduction current of the connector shall be designed within theallowable temperature range in accordance with MIL-W-5088-based derating guidelines
indicated in the connector application data sheet and, ultimately, thermal analysis.
• The current value of harnesses shall be in accordance with JERG-2-212 Wire Derating
Design Standard while the current value of externally exposed harnesses (outside ofspacecraft structure, shaft section harnesses and the like) shall be specified properly inaccordance with the thermal analysis requirements in 5.4.2.

5.5.2

Machine design(1) Review of structure characteristic
After clarifying the I/F conditions with the solar array wings and spacecraft structure,
characteristic value analysis and strength analysis shall be carried out on the paddle drive
mechanism to check whether the rigidity requirements and strength requirements are satisfied.
Major evaluation items shall be as follows:• Allowable load (translational force, bend, torsion)• Allowable temperature (cabinet/output shaft temperature difference)• Equivalent rigidity (bend, torsion)(2) Reviewing drive characteristics
For the paddle drive mechanism, the following items shall be reviewed and clarified withrespect to drive characteristics• Rotation direction• Step angle• Backlash• Rotation angle accuracy (maximum accumulated errors)• Output shaft generating torque and torque margin (definition shall be clarified)• Operating life analysis/evaluation(3) The holding torque of the paddle drive mechanism in the stand-by mode and holdmode shall satisfy the requirements for paddle drive mechanism.
(4) Torque margin
The torque margin M shall be defined as in the following formula and the requirement value
shall be 1 or more.

43

JERG-2-215A

M=T0/(TI+TD)-1

44

JERG-2-215A

T O:

Output torque

TI:

Torque to accelerate the solar array wings

T D:

Paddle drive mechanism internal drag torqueOutput torque TO can be obtained by step motor output torque x deceleration ratio.
Internal drag torque can be obtained with:- Harmonic drive ® starting torque- Harmonic drive ® efficiency,- Friction torque of drive module bearing, spur gear efficiency- Friction torque of slip ring assembly, and- Friction torque of drive shaft bearing.
Because these values and step motor output torque vary depending on the temperature, averification test shall be conducted on newly adopted articles in advance.

5.5.3

Thermal design

For the paddle drive mechanism, the following items shall be reviewed and clarified
with respect to thermal design.
• Equipment calorific value• Paddle drive mechanism installation (contacting surface area)• Installation surface exothermic density• Surface treatment of installation surface and other than installation surface• Component temperature analysis and heat resistance evaluation(For the power wire harnesses in particular, it is desirable to carry out review and analysis
in consideration of the thermal input from an external heat environment as well as
self-heating due to energization.)• Thermal design attribution as equipment• Paddle drive mechanism I/F temperature (installation surface, harness connector I/F
point)

5.5.4

Reviewing environmental resistance

Check for conformity to the Space Environment Standard JERG-2-141. (It is desirable
in particular to evaluate the externally exposed sections of the paddle drive
mechanism (including harnesses) for tolerance to a complex environment such as UV,
AO, thermal cycle and radiation.)45

JERG-2-215A

46

JERG-2-215A

5.5.5

Clarifying test plan and test items

For the paddle drive mechanism, the test plan including test items shall be clarified.
5.6 Paddle drive control design and analysis(1) Paddle solar tracking control

Paddle solar tracking control shall be generally carried out by the spacecraft system (attitude control
system in many cases) with the use of the angle detection sensors (such as paddle solar tracking
angle sensor and paddle drive mechanism rotation angle sensor) and the drive actuator mounted in
the solar array wing system. The subsystem side of the solar array wing system shall output the angle
detection sensor measurement information with a specified accuracy and rotate the drive actuatorwith an accuracy specified according to the input signals. The sensor measurement information andactuator drive model shall be provided to the spacecraft system side. Alignment information shall be
included in these sensor models and actuator models.

The spacecraft system side (attitude control system in many cases) shall design a paddle solar
tracking control system based on the sensor models and actuator models provided from the solar
array wing system to perform analysis of tracking control function and performance (such as control
sequence, control accuracy, and disturbance).

(2) Drive mode design

Generally, the subsystem side of the solar array wing system shall design and specify paddle drive
mechanism modes including standby mode, transition mode, hold mode, through mode, and clockmode as drive modes of the solar array wing system. On the other hand, the spacecraft system side(in many cases, attitude control system performing paddle solar tracking control) shall design andspecify paddle sun acquisition standby mode, paddle sun search mode, solar tracking mode, array
trim mode, automatic array trim mode, constant clock rate mode, and manual array trim mode (withduplication) as drive modes for macro action. The subsystem side of the solar array wing system andthe spacecraft system side shall clarify the correspondence and combination of drive modes andinterface each other. Information on the driving rules, action and sequence related to the drive modes
shall be reflected in the tracking control function/performance analysis and paddle drive disturbance47

JERG-2-215A

analysis of the modes.

(3) Failure detection identification reconfiguration (FDIR)

The spacecraft system side (in many cases, attitude control system performing paddle solar tracking
control) shall design a failure detection identification reconfiguration (FDIR) function of paddle
solar tracking control. The judgment conditions of abnormality detection, action sequence after
detection, and restoration procedure in operation shall be clarified and provided to the subsystemside of the solar array wing system. The subsystem side of the solar array wing system shall checkthese for inconsistency with the design of the subsystem side of the solar array wing system.

(4) Paddle drive disturbanceThe subsystem side of the solar array wing system having the paddle drive mechanism shall create acoupled dynamics model consisting of the step motor, harmonic drive, gear, solar array wing flexible
structure mode, and spacecraft body and analyze the drive disturbance, vibration profile, reaction
torque, and barycenter variations when driven in accordance with the tracking control drive signals
provided from the spacecraft system by numerical simulation and provide the analysis results as wellas the dynamics model to the spacecraft system. The subsystem side shall check that the analysis
results are within the range specified by the spacecraft system.

The spacecraft system side shall incorporate the paddle drive disturbance model provided from thesubsystem side into the spacecraft dynamics model including other flexible structures and conduct
attitude/orientation disturbance analysis by numerical simulation. Consideration shall be given to theflexibility of the spacecraft installation section. Check for conformity to the attitude/orientation
accuracy requirements in accordance with the results of attitude/orientation disturbance analysis.
Specify the allowable disturbance amount according to the waveband with respect to the subsystemside of the solar array wing system.

See the disturbance control standard (JERG-2-152) and disturbance control manual(JERG-2-152-HB001) for the standard and method of disturbance control.

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5.7 Space environment design
5.7.1

Radiation design and analysis

Details of radiation environment, radiation dose and radiation deterioration are described in Space
Environment Standard JERG-2-141 and Radiation-proof Design Standard JERG-2-143. An
overview is as follows:•

The radiation environment and radiation dose which the spacecraft encounters shall be
predicted on a mission-by-mission basis using the radiation environment model specified by
Space Environment Standard JERG-2-141. Because the confidence level which must be
applied to each mission period is specified with respect to the prediction of solar proton, it
shall be necessary to observe Space Environment Standard JERG-2-141.

•

As the prediction of radiation dose involves substantial uncertainty, a proper margin shall be
provided when reflecting the result in the design.

•

It is difficult to determine the margin quantitatively but important items shall be reviewed andevaluated properly for an assured and efficient radiation-proof design.

•

It shall be verified whether radiation-induced deterioration can be allowed with respect to themission period by testing if possible.

•

Omission of testing may be reviewed when using materials that have a track record of use in
the orbit and period to be operated.

In particular, the influence of radiation exposure in orbit of the solar array wings shall be reflected in
the design in the following two main points.

(1)

Cell electric performance deterioration due to the solar cells exposed to radiation(2)

Physical deterioration of employed materials due to radiation exposure

For the above-mentioned (1), the deterioration degree of solar cells shall be quantified in accordancewith the predicted radiation environment conditions and reflected in the electric design (determine
the initial generated power in anticipation of deterioration). Among solar cell deterioration
prediction methods are relative damage coefficient (RDC) method and displacement damage dose(Dd) method. Because the details of specific deterioration prediction procedures based on RDC
method and Dd method are described in Radiation-proof Design Standard Handbook
JERG-2-143-HB001, an overview and flow chart of the two methods are described here.

[Relative damage coefficient (Relative Damage Coefficient: RDC) method]49

JERG-2-215A[1] Radiation test shall be conducted on electrons and protons of the target solar cells. Or alreadyacquired data sheets shall be prepared.
[2] The energy spectrum distributions of electrons and protons fluence during mission period shallbe obtained in accordance with the radiation environment prediction code.
[3] The deterioration curve of solar cells with respect to a 1 MeV electron beam shall be preparedfrom experience and data sheets.
[4] RDC with respect to electrons and protons of target solar cells (without shielding materials,
radiation vertical incidence) shall be calculated. In addition, a 10 MeV proton/1 MeV electron
beam conversion coefficient shall be derived.
[5] RDC with respect to electrons and protons of target solar cells shall be converted under theconditions of radiation omnidirectional incidence with the shielding materials.
[6] With the use of the RDC after conversion and the particle energy distribution of electrons andprotons, the equivalent fluence of a 1 MeV electron beam shall be calculated and derived forelectrons while the equivalent fluence of 10 MeV protons for protons. With the use of 10 MeV
proton/1 MeV electron beam conversion coefficients, the equivalent 10 MeV proton fluence
shall be converted to the 1 MeV electron beam fluence and added to the equivalent 1 MeV
electron beam fluence of the former electron beam. Thus, the deterioration of cells due toelectrons and protons flying with various energy/fluence can be converted to one indicator ofdeterioration with respect to the electron beam of 1 MeV.
[7] In the deterioration curve with respect to 1 MeV electron beam, the cell output at the end ofoperating life can be estimated by calculating the deterioration amount of solar cell output(preservation ratio) from the obtained equivalent 1 MeV electron beam fluence.
[1]

[2]

Implement a radiation exposure test on target solar
cells (however, without cover glass) and acquire
the deterioration curve (horizontal axis: fluence).

Determine the energy spectrum of
incidence radiation particle (electrons,
protons) count in accordance with theenvironment prediction code.

[4]

[5]

Derive the relative damage
coefficient (RDC) and 1 MeV
electron / 10 MeV proton
conversion coefficients.

Convert the relative damage coefficientunder the condition of omnidirectional
incidence with the cover glass (having adesired thickness).

[3]

[6]

Acquire the deterioration curve with respect to 1
MeV electron beam.

Calculate the equivalent 1 MeV
electron beam fluence during mission
period.

[7]

Estimate the deterioration amount ofsolar cells at the end of mission(EOL).

Figure 5.7-1 Procedure for predicting the solar cell deterioration amount by relative damage
coefficient (RDC) method50

JERG-2-215A[Displacement damage dose (Displacement Damage Dose: Dd) method][1] Derive the particle (proton/electron beam) energy dependence of non-ionizing energy (NIEL)
value of solar cell materials (such as Si, GaAs).
[2] Indicate the horizontal axis (electron beam or proton beam fluence) of deterioration
characteristic data of solar cells by using an indicator called displacement damage dose (Ddvalue = NIEL value × fluence).
[3] Calculate the radiation dose during mission period by using the radiation environment model.
[4] Calculate the radiation spectrum after transmission of shielding materials (such as cover glass
and CFRP skin).
[5] Calculate all the Dd values during mission period from radiation dose proceeding into the solar
cells and NIEL values.
[6] Estimate the cell output at the end of operating life in accordance with the cell deterioration data
and all Dd values during mission period.
[1]

[3]

Acquire and derive non-ionizing energy
loss (NIEL) values of constituent elements
in target solar cells.

Determine the energy spectrum of incidence
radiation particle (electrons, protons) count inaccordance with the environment prediction
code.
[4]

[2]

Implement a radiation exposure test on solar
cells (however, without cover glass) andacquire the deterioration curve (horizontal axis:
displacement damage dose Dd).

Convert the incidence radiation particle countenergy spectrum under the condition ofomnidirectional incidence with the cover glass(having a desired thickness).
[5]
Calculate the displacement damage amount during mission period.

[6]
Estimate the deterioration amount of solar
cells at the end of mission (EOL).

Figure 5.7-2 Procedure for predicting solar cell deterioration amount by displacement damage dose(Dd) method

For the above-mentioned (2), calculate the total ionizing dose and estimate the radiation
deterioration in characteristics of applicable parts and materials based on the radiation resistancedata of parts and materials. See Radiation-proof Design Standard JERG-2-143 and Radiation-proof
Design Standard Handbook JERG-2-143-HB001 for details of the total ionizing dose.

•

Consideration shall be given to the radiation deterioration evaluation of thermo-optical andmechanical characteristics and radiation resistance of the materials employed in solar array

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wings (such as harnesses, adhesive, cover glass, thermal control materials).
•

It is desirable basically to select materials excellent in radiation resistance or materials having
a flight track record.

•

For example, wires covered with radiation-crosslinked ETFE excellent in radiation resistanceincluding exposed sections are often used as cables for transmitting generated power, which isa basic function of solar array wings. The wires have a considerable track record in domestic
spacecraft and overseas commercial spacecraft.

•

When the radiation-crosslinked ETFE must be used under a high radiation environment andcracks may be developed in the covering, measures shall be taken to prevent a discharge andshort circuit with peripheral elements even if the internal electric wires are exposed.

•

When characteristic deterioration in parts and materials is within an allowable range in viewof usage, the application is judged to be proper.

•

If characteristic deterioration exceeds the allowable range, it shall be necessary to takemeasures such as shielding.

•

Addition of shielding materials causes a considerable impact to spacecraft design in terms ofweight and the like. Because there are still many uncertain elements including space radiation
environment in radiation analysis, a certain degree of margin must be incorporated in design.

•

If radiation resistance characteristics of parts and material are not known, deterioration data
must be acquired by conducting a radiation test.

•

Because the influence of radiation on materials may vary depending on the radiation type,energy and dose rate, attention shall be paid to this point when evaluating the data.

5.7.2

Ultraviolet (UV) countermeasure design and analysis

Among influences of ultraviolet on materials employed in solar array wings are ultraviolet
deterioration in CIC generated power (See 5.1.1.1 (2) and 5.1.1.2 (7)), characteristic changes inorganic material surfaces such as harness films. (In particular, the solar light absorbing ratio α
increases due to browning of white harness films and the increase tendency has a temperaturedependence.) In temperature analysis of on-orbit harnesses (in particular harness bundles), care shallbe taken to prevent deviation from the harness’ allowable working temperature in consideration ofthese deterioration factors. It shall be necessary, as needed, to take measures such as harness routing
to surfaces out of direct exposure of ultraviolet rays and installation of light shielding plates orprotection by SLI and the like.
Refer to the Collection of Wire Derating Test Data (JERG-2-212-TM001) for details of
ultraviolet-induced characteristic changes in surfaces of organic materials such as harnesses.

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5.7.3

Atomic oxygen (AO) countermeasure design and analysis

Atomic oxygen exists in low orbit depending on the orbit height where earth-orbiting spacecraft
aviate. This causes erosion in the materials employed in the solar array wings, resulting insignificantly high erosion yields of metals such as silver and osmium and organic materials such aspolyimide. As materials employed in the solar array wings, silver is used in solar cell electrodes,
interconnectors and bus bars of solar cell array circuits while polyimide in panel substrate surface
and SLI/MLI. The paddle shall be designed so that functional maintenance of the above-mentioned
materials will not be affected by atomic oxygen erosion depending on the spacecraft orbit altitude
and mission operation period. In erosion yield analysis, consideration shall be given to the incidence
direction of AO. Among countermeasures are adoption of a proper material thickness for functional
maintenance or coating by materials having a low reaction rate (such as gold plating for silver and
ITO, SiO2 or germanium coating for polyimide).
See the Radiation-proof Design Handbook JERG-2-143-HB001 for details of atomic oxygen
distribution density and reaction rate of each material with respect to orbit altitude.
5.7.4

Debris/micro meteoroids countermeasure design and analysis

Because the details of possibility of collision of debris/micro meteoroids, collision influence andcountermeasure design are described in the Space Debris Protective Design Standard (JERG-2-144),
countermeasure design is described here in view of improving the reduction in generated power.
The solar array wing is a piece of externally installed equipment which is directly exposed to outerspace. It is difficult to take protective measures against debris/micro meteoroids for the solar cellarray circuit, in particular, due to its functional action. Thus, it is desirable to consider the following
countermeasures in view of improving the reduction in generated power.
(1) Reduction of influence by means of separation of power harness HOT/RTN lines and increasein the number of power harness routes in the system EMC allowable range.
(2) Installation of the power harnesses closely contacted with the structure elements to the extent
possible for improving the shielding effect. (However, proper stress relief must be
maintained.) When using both harness installation and thermal control materials installation,
proper consideration shall be given to thermal design and thermal control materials
grounding.
(3) Array circuit ground fault prevention by adoption of bleeder resistance53

JERG-2-215A

5.8 Anti-charging/discharging countermeasure design and analysis
The details of anti-charging/discharging countermeasure design and analysis are specified in
Charging/Discharging Design Standard (JERG-2-211) and shall be in conformity with the applicable
design requirements (JERG-2-211 5.4 “Solar Cell Panel”), analysis requirements (JERG-2-211
6.2.2.4 “Computer Aided Model and Analysis Results”) and test requirements (JERG-2-211 6.4“Solar Cell Panel”). Except for panel surface side (array circuit side) at present, there are no standard
design indicators equivalent to JERG-2-211, Section 9 Appendix-II “Test Results of Inter-solar Cell
Distance and Discharge” with respect to the panel back side, paddle drive mechanism and otherspots. Thus, it shall be necessary to identify critical spots at which self-lasting secondary discharge
attributable to primary discharge may occur and judge the necessity of testing in line with theregulations specified in JERG-2-211 6.4.1.3 “In the case no testing is required.”
5.9 Installation design
Cautions in solar array wing installation design are described here.
(1) Precautions for installation for improving reliability and survivability
Table 5-1 shows the precautions for installation for improving reliability and survivability.
Installation is carried out for improving reliability and survivability. Failure to pay attention
during installation to the points listed in the table may result in failure or contradiction to
design intent.

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Table 5-1 Precautions for installation design
Major role
Target parts or
installation

Sustaining
discharge
control

Groundfault
current
control

Secondaryfailureavoidance(1) Grouting between cell gaps



N/A

N/A

(2) Bleeder resistance

N/A



N/A

(3) Blocking or isolation diode



N/A



(4) Bypass diode (panel back side
installation is target)

N/A

N/A

N/A

(5) MTC/D-sub connector

N/A

N/A

N/A

(6) 8W8P connector

N/A

N/A

N/A

(7) Panel ground line harness
capacity and ground line
slipping capacity

N/A

N/A



Notes for installation design[1] When performing grouting between solar cells to control secondary
discharge, proper deforming shall be performed during mixing to leaveno voids.
[2] Baking shall be performed properly to prevent grouting materials frombecoming a contamination source in orbit.
[1] To avoid contact with discharge plasma, the resistance shall be moldedby RTV or covered with insulating (heat shrinkable) tube without
exposing the electrodes. If it is difficult to perform installation due tostress relief, consideration shall be given to prevent discharge by
providing a sufficient distance and the like.
[2] Resistance values shall be selected properly in accordance with theamount of on-orbit influent electrons. Selection shall be done inconsideration of normal rated power of resistance in case of ground
fault.
[3] In case of multiple circuit ground faults concurrently with shunt ON
action, the multiple currents exceeding the shunt element rating may
flow in the same shunt element. Thus, when installing a bleeder
resistance, resistance shall be installed between every panel, yoke andboom to the extent possible.
[4] In the case of the paddle mounting the shunt yoke, a ground fault of
power harness (HOT) with respect to the boom/yoke is single failuremode of bus short circuit. Thus, it is desirable to install a bleeder
resistance to the boom/yoke.
[5] When installing a bleeder resistance, deployment synchronization
mechanism (synchronization wire) may become an electricallyconducting path between panels. To function the original design
intention for the above-mentioned [3], insulation between the panel
hinge and the panel shall be assured. Care shall be given to prevent thesynchronization wire from coming into contact with the substrate or
steady strain, which may cause in contact conduction between the panel
hinge and panel.
[1] To avoid contact with discharge plasma, the resistance shall be moldedby RTV or covered with insulating (heat shrinkable) tube without
exposing the electrodes. If it is difficult to perform installation due tostress relief, consideration shall be given to prevent discharge by
providing a sufficient distance and the like.
[2] Because a poor soldering in component electrodes and power harnesses
may directly cause power drop, it is desirable to check for proper
thermal cycle tolerance and install a blocking diode serving as aredundant route.
[1] To avoid contact with discharge plasma, the resistance shall be moldedby RTV or covered with insulating (heat shrinkable) tube without
exposing the electrodes. If it is difficult to perform installation due tostress relief, consideration shall be given to prevent discharge by
providing a sufficient distance and the like.
[1] In accordance with the installation example of Insulation DesignStandard (JERG-2-213), the connector back face (harness connectionside) shall be molded by RTV or covered with insulating (heatshrinkable) tube without exposing the electrodes to avoid contact with
discharge plasma.
[1] In accordance with the installation example of Insulation DesignStandard (JERG-2-213), the connector back face (harness connectionside) shall be molded by RTV without exposing the electrodes to avoid
contact with discharge plasma.
[1] When an array circuit ground fault is permitted in the design without
installing a bleeder resistance, current capacity shall be selected inconsideration of avoidance of secondary failures caused by current
concentration to the left line due to multiple ground faults.

(2) Precautions for installing wire harnesses
When wire harnesses are routed and fixed on the panel surface and the like, consideration
shall be given to the following points.
• In consideration of thermal shrinkage of wires with respect to the environment temperature
range, wire stress release between fixed points shall be provided.
• Consideration shall be given to the minimum bend polarization rate in bending wires such as

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when passing through panel through-hole. Care shall be given to susceptibility of film cracks
at bend sections under a complex space environment (radiation and thermal cycle).
• When routing is performed near sharp edges such as substrate edges or when harness films
may be worn under a vibration environment or near movable sections such as deployment
action, covering of sharp edges and protection covering of harness shall be carried out.
• When fixing a harness bundle to the boom/yoke section using a tie wrap, consideration shallbe given to a proper restraint force and harness film covering in harness bundle restraining
sections to avoid cold flow of harness films.
• If noise causes a critical influence on the functions of signal lines, the provision of routes
separated from power lines or the use of shield lines.
• Harness routing in boom/yoke sections shall be performed so that development of a magnetic
field associated with a current loop will be minimized in principle. However, in view ofsurvivability requirements, robustness requirements or noise immunity requirements, when acurrent loop is developed due to multiple routes, adjustment shall be made with the systemwith respect to the allowance and specified in ICD and the like.

5.10 Definition of interfaces

5.10.1 Definition of external interfaces
Interfaces (electrical, thermal and mechanical) between a paddle subsystem and other subsystems
shall be clearly illustrated.

5.10.2 Interface demarcation
Electrical, mechanical and thermal interface demarcations shall be clearly illustrated.

5.10.3 Interface conditions
In the interface conditions defined in 5.10.1, conditions at least for the items listed below shall beorganized and clarified.

5.10.3.1 Electrical interface conditions
As electrical interfaces, conditions at least for the items listed below shall be organized and clarified
for the power supply system, telemetry/command system, attitude control system and otherinterfaces defined in 5.1.

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5.10.3.1.1

Power supply system interfaces

It shall be necessary to organize and clarify the interface conditions related to the power
supply-related interfaces between a paddle drive mechanism and power supply system whichtransfer the power generated by the paddle subsystem, voltage/current interfaces with the shuntcircuit, primary power source interfaces required by the paddle system such as damper heater, andODC pass interfaces related to explosive device ignition power. For shunt action procedures of thearray circuit, if there are no special system requirements, make correspondence between the shuntlevel in the power supply system and the shunt action array circuit understandable and clarify themeasures in ICD.
Other interfaces shall be in conformity to the regulations specified in Power Supply System DesignStandard (JERG-2-214).

5.10.3.1.2

Telemetry/command interfaces

All general requirements for telemetry/command shall be in accordance with Electric DesignStandard (JERG-2-200). Details of TeleCommand item setting standard and electrical signal
interfaces shall be in accordance with the regulations specified in Examples of Signal Interface(JERG-2-200-TM001).
Recommended telemetry items required for a solar array wing system are shown in Table 5-2.

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Table 5-2 Recommended telemetry items
Target component
Remarks
Solar cell panel• It is desirable to directly measure the solar
cell temperature but it is difficult to install atemperature sensor on the panel surface side.
Thus, install a temperature sensor on thepanel backside and grasp the solar celltemperature based on the thermal mathmodels and thermal vacuum test result.
• It is ideally desirable to measure each solar
cell panel. However, when there are
restrictions on the telemetry count and thenumber of slip rings, it is desirable at leastthat the farthest and nearest panels of thespacecraft structure be measured.
Boom/yoke temperature Boom, yoke
• It is desirable to perform measurement in theproximity of the harness bundle installed.
Harness temperature
Boom/yoke section• It is desirable to perform measurement at thefitting harness bundle
center of the harnesses bundled in theboom/yoke section.
Paddle deployment status Solar array wing
• It is desirable to monitor the deployment
status of each hinge by using the micro
switch and resistance.
Damper temperature
Damper
Paddle drive unit
Paddle drive unit
• In the paddle drive unit, it is desirable totemperature
measure the temperature difference between
the inner ring and outer ring of the bearing.
• Shaft temperature• Flange temperature
Paddle drive mechanism Paddle drive unit
• The reference point sensor is not a mere
angle signal
piece of telemetry but a signal used for thecontrol system. Thus, it shall be designed in• Potentiometerconsideration of the interfaces with the• Segment switch• Reference point sensor
attitude control system.
Solar tracking error angle Solar tracking sensor • The solar tracking sensor is not a mere piecesignalof telemetry but a signal used for the control
system. Thus, it shall be designed inconsideration of the interfaces with theattitude control system.
Shunt temperature
Shunt
Applied to the spacecraft mounting the shunt
yoke.
Shunt drive signal
Shunt
Telemetry item
Solar cell panel
temperature

5.10.3.1.3

Attitude control system interfaces

As solar tracking interfaces, the interfaces related to paddle drive mechanism drive and the interface
conditions related to paddle drive mechanism angle detection shall be organized and clarified.

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5.10.3.1.4

Other electrical interfaces

When the solar tracking sensor, deployment monitor (DM) and accelerometer are installed to thesolar array wing, the electrical interface conditions with the applicable equipment shall be clarified.
5.10.3.2 Mechanical interface condition
The panel size and envelope area in solar array wing storage state shall be specified.
The mechanical interface conditions related to holding and installation between solar array wing andpaddle drive mechanism, between paddle drive mechanism and structure, and between PDL andstructure shall be clarified. When the solar tracking sensor, deployment monitor (DM) andaccelerometer are installed to the solar array wing, the mechanical interface conditions related toapplicable holding and installation shall be clarified.
5.10.3.3 Thermal interface condition
The thermal interfaces (including calculation formulas of conductive heat exchange and radiant heatexchange amount) conditions related to holding and installation sections between solar array wing
and paddle drive mechanism, between paddle drive mechanism and structure, and between PDL andstructure shall be clarified. When the solar tracking sensor, deployment monitor (DM) andaccelerometer are installed to the solar array wing, the thermal interface conditions related toapplicable holding and installation sections shall be clarified.
For thermal analysis input conditions, it shall be necessary to clarify the interface conditions(combination of the solar light incidence conditions from launch to each on-orbit operation modeand the orbit thermal input constants (solar light strength, albedo constant and earth infraredintensity)) including attitude conditions at the time of attitude abnormality and specify the conditionsin application documents such as thermal design standard.
For the above-mentioned solar light incidence conditions, care shall be given to the influence ofreflected light from the system structure thermal control materials (such as OSR and Ag-Teflon).

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5.10.3.4 Dynamics interface conditions

For design, analysis and verification in 5.2.4.2, 5.3.4, and 5.6, the following dynamics interface
conditions shall be specified and interfaced with the spacecraft system. As the solar array wing
system is coupled mutually with the spacecraft body and other subsystems, the dynamics will change
the behavior. Thus, a torque interface with the paddle system end fixed rigid is not proper because it
is impossible to express the motion of the spacecraft and paddle system by the torque interface.
Therefore, the spacecraft system shall perform interfacing using a dynamics model which allows for
dynamics analysis of all machines.

(1) Paddle mass characteristic• Storage state, and mass, barycenter position, inertia moment, product of inertia, tolerance andallowable range and regulations coordinate system under a deployment state (after normal
deployment and, as needed, at deployment abnormality)(2) Paddle external disturbance characteristic• Surface characteristics and surface area for solar radiation pressure calculation• Residual magnetic moment(3) Paddle flexible structure characteristics• Natural frequency, zero-order coupling coefficient vector, primary coupling coefficient vector,
mode damping ratio, tolerance and allowable range, regulations coordinate system (after normal
deployment, and, as needed, at deployment abnormality)(4) Paddle heat distortion deformation• Deformation models, deformation profile, installation point torque profile, quasi-static
deformation (warpage) attributable to heat during orbit(5) Paddle heat-induced dynamics• Heat-induced dynamics models, deformation profile, installation point torque profile, barycenter
variations profile(6) Paddle deployment dynamics

60

JERG-2-215A• Deployment configuration profile, deployment angle profile, deployment signal profile,
deployment time, latch shock torque, installation point torque profile, attitude angle profile,
attitude rate profile (at least two temperature conditions including high-temperature andlow-temperature conditions)(7) Paddle solar tracking control• Angle detection sensor models• Drive actuator models(8) Paddle drive disturbance• Step motor, harmonic drive, gear, paddle flexible structure mode, drive profile or driving rules
parameter• Drive disturbance, vibration profile, installation point torque profile, barycenter variations
5.10.3.5 Visual field interface
Install a solar tracking sensor. When there are requirements of visual field interference, visual fieldinterfaces shall be specified in ICD and the like.
5.10.3.6 Movable section envelope interfaces
The action envelope (range in which no mechanical interference occurs with spacecraft system)
during deployment action of the solar array wings and on-orbit solar tracking rotation shall be
specified in ICD and the like.
5.10.3.7 Interfaces with ground support equipment
It shall be necessary to clarify a set of ground support equipment for manufacturing, assembly, test,
maintenance and inspection of the solar array wing system and define electrical and mechanicalinterfaces.

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6. Verification
It shall be verified that a solar array wing system is designed properly and therequirements related to generated power and the like are satisfied. Verification
shall be carried out in the PFM/FM component test, intrasystem combined test, andsystem test in the design analysis and manufacturing/test stages based on the
results of manufacturing and testing of EM and the like up to the design stage.
6.1 Deployment/verification plan
In devising a deployment plan, perform technology analysis of the requirement
specifications based on similarity and analysis whether element test anddeployment test on the components shown in Figure 4-1 are necessary or not.
Device a deployment/verification plan in which manufacturing and testing of thenecessary element test models and deployment test models are clarified. Organize
the plan as a requirement confirmation matrix and an environment test matrix fordesign and analysis requirements specified in 5. Table 6-1 shows an example ofengineering models with consideration given up to system test.
Table 6-1 Example of definition of engineering models

PDM

Solar array wing

Item

Model name
Remarks (objective)
Charging/discharging coupon
Used for discharge tolerance evaluation
panel
Coupon panel
Used for checking for thermal cycle life performance(Other element models)
Deployment hinge section model Acquisition of deployment element parameters (suchas torque and rigidity)
Holding/releasing mechanism
Shock level confirmation
section model
etc.
Structure dummy
Used for system structure model
Thermal dummy
Used for system thermal model
EM
Used for total design verification
PFM/FM
Used for flight quality verification
Operating life test model
Used for checking movable sections such as slip rings
and actuators for operating life
Structure dummy
Used for system structure model
Thermal dummy
Used for system thermal model
EM
Used for paddle combination test and system electric
model
PFM/FM
Used for flight quality verification

For the models shown in Table 6-1, precautions for manufacturing models shall be
as follows.
(1) Coupon panel
There have been fractures of solar cells and interconnector welds under a thermal vacuum
environment in a past project. When changing substrates, solar cells or interconnectors, it
shall be necessary to perform verification in a coupon panel test and the like.

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JERG-2-215A

Thus, coupon panels shall be manufactured with materials and design and manufacturing
processes similar to those of flight articles in consideration of the following points.
• In addition to solar cells, parts installed to flight articles (including resistance, diodes,
connectors, thermal control materials and, as needed, brackets such as hold down pads)
shall be simulated.
• It is desirable to include a solar cell replacement simulation.
• Surface film joints (flight articles having joints) shall be simulated.
• CFRP face sheet layer number (including doubler boundary section) shall be simulated.
See Solar Cell Panel Quality Assurance Handbook (JERG-2-216-HB001) for requirements
for test and evaluation. When adopting an array circuit isolation diode, it is desirable toperform a cycle test with coupon panels with energization state in separate diode section
assessed.
(2) Charging/discharging coupon panel• For intercell gap, the minimum gap shall be simulated within the manufacturing tolerance
standard range.
For items other than those above, simulation shall be reflected as in the case with coupon
panels. See Charging/Discharging Design Standard (JERG-2-211) for requirements for test
and evaluation.
(3) Structure dummy
After being assembled in the spacecraft system structure model, the structure dummy shallbe supplied to verify the structural design of the system and check the structure interface
design of the spacecraft system and solar array wing system for validity.
The structure dummy shall be manufactured in principle according to the following
requirements. Unless otherwise specified, the following requirements shall be applied toboth the solar array wings and paddle drive mechanism.
• The exterior shape and dimensions shall be the same with PFM in principle.
• The installation hole position, installation hole diameter (installation fastener size), andplate thickness of installation section shall be same with PFM including tolerance.
• The mass and moments of inertia shall be same with the PFM nominal value. (Allowableerrors shall be adjusted with the system.)• The mass center position shall be same with the PFM nominal value. (Allowable errors
shall be adjusted with the system.)• Shall be same with the main mode. Shall have the same strength with PFM. Normally,
notching other than the acceleration equivalent to mass center shall not be allowable.
• The natural frequency of the main mode shall be same with PFM. (Allowable errors shallbe adjusted with the system.)63

JERG-2-215A• PFM occurrence shock shall be simulated as needed.
(Applied only to the solar array wing structure dummy.)64

JERG-2-215A(4) Thermal dummy•

Normally, the main objective of system thermal vacuum test is to check thethermal control state under operation state after solar array wing deployment.
It is ideal that solar array wings are supplied in deployment state to system
thermal vacuum test. However, it is extremely difficult to install solar arraywings in deployment state due to equipment restrictions. Thus, a solar array
wing thermal dummy shall be manufactured so that the dimension and shape(give consideration so that the form coefficient conforms as much as possible),heat capacity and thermo-optical characteristic are equal to those of the solar
array wing flight articles in line with the evaluation purpose and evaluation
items of system test where possible.
However, solar array wings are exempted from evaluation in a system thermal
vacuum test because the solar array wings are insulated from thespacecraft/PDM and independently thermal-controlled. For solar array wings
having a small thermal bonding with the spacecraft structure due to the use of
yoke and the like, it is allowed in some cases to perform testing with only thesolar array wing installation interface hold-down installed to the structure
without using the solar array wing thermal dummy. This is because of
circumstances related to test configuration such as the power line connections
from the solar array simulation power source to the PDM installed in thespacecraft structure.

•

If it is difficult to use the flight article paddle drive mechanism in a system
thermal vacuum test, it shall be necessary to manufacture a thermal dummy ofthe paddle drive mechanism having the dimension and shape, heat capacity,heat conduction (installation surface thermal conductance), thermo-opticalcharacteristic, and calorific value equivalent to those of flight articles to theextent possible without interfering with the evaluation purpose and evaluation
items of the system test and then to supply the dummy to the system thermal
vacuum test.

6.2 Manufacturing and testing
Figure 6-1 shows an example of manufacturing/test flow of proto-flight in solar array wing system.
Precautions (reflection of nonconformance cases) for each process of manufacturing/test flow arelisted in the notes in the figure and 6.2 (1). Review the test implementation items andimplementation procedure properly before verifying the design and analysis requirements in
Paragraph 5. The verification items for requirements confirmation matrix and environment test
matrix shall be satisfied.

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JERG-2-215A

Recommended check items in in-process process are shown in Table 6-2. When checking the solar
array wings for flight quality soundness, important items which cannot be checked in a complete
state shall be checked for soundness during in-process operation by setting proper criteria.

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JERG-2-215A

Table 6-2 In process recommended check items
Target category
Solar cell panel Substrate

CIC peel strength
Deployment
Hinge
synchronization
mechanism

Holding andreleasingmechanism

Rotary damper
Preload bolt

Target unit
Actual machine

Check item
Polyimide surface insulation resistanceinspection
Dimple developmental rate, dimple depth
Parallel sample Flatwise strength test
3-point or 4-point bend test
Condition setting Checking interconnector for welding
samplestrength
Actual machine Sliding torque/deployment torque at
deployment action
Rigidity at deployment
Synchronization wire (tightening portion)
proof load tolerance
Actual machine Damping rate
Actual machine Proof load tolerance(1) Precautions in manufacturing/test flow (reflection of nonconformance cases)<Solar array wing> In-process operation• Defective appearance of substrate/CIC due to handling (such as flaws, hollows on substrate andcracks in cover glass)• Re-melting of solder in MTC connector heat shrinkable tube operation
 Proto-flight test
• Defective appearance of substrate/CIC due to handling (such as flaws, hollows on substrate andcracks in cover glass)• Incomplete deployment due to poor maintenance of deployment supporting device• Damage in mechanism components due to poor notch level setting• Deviation from standard due to shock developed by explosive device• Temperature control standard value deviation in thermal vacuum test
<Paddle drive mechanism> In-process operation• Increased resistance value of slip rings and noise generation due to insufficient run-in• Peeling in incompletely plated slip ring section and sliver developed by intrusion of foreign
matter into brush Proto-flight test
• Standard deviation due to temperature change at reference point sensor trigger point• Defective appearance caused by handling (flaws on cabinet and connectors)• Temperature control standard value deviation in thermal vacuum test

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JERG-2-215A

Solar array wing
CIC/module
manufacturing

Paddle drive mechanism
Substrate, boom, yoke andmechanism section manufacturing

Drive module
assembly

Solar cell laydown

PDM assembly

Run-in of paddle drive
mechanism (in particularslip ring assembly) shallbe carried out in
Completion inspection in-process operation.

In-process operation

Panel unit initial performance test
• Electric performance test(including irradiation test)

Thermal vacuum test
(paddle, yoke, harness unit,
mechanism section)

Note: Thermal vacuum
test shall be carried out
as proto-flight test.

Panel unit final performance test
• Electric performance test(including irradiation test)

Initial function
performance test

Hinge/harness torque measurement

Mechanical environment test

Hinge rigidity test

Random vibration test
Sine wave vibration test

Paddle assembly (integration)

Function performance test after
mechanical environment• Conduction/insulation inspection• Visual inspection• Dimension measurement• Mass measurement

Thermal vacuum test

Final function
performance test

Paddle initial function test

Proto-flight test

Slip ring assembly• Deployment test
• Electric performance test(including irradiation test)

Final visual inspection

Mechanical environment test
• Sine wave vibration test
• Acoustic test
Paddle final function test
• Mass characteristic test
• Deployment test
• Electric performance test(including irradiation test)• Wing rigidity test
• Final visual inspection

Paddle-PDM combined subsystem test

System test (related to paddle)• Electric confirmation before system test (DC integration)• Sine wave vibration test
• Acoustic test
• Deployment (or release)• Shock test
• Electric confirmation after system test
• Visual inspection

Verification test after system test
• Deployment test
• Electric performance test (including irradiation test)• Visual inspection

Figure 6-1 Manufacturing/test flow

68

• For electrical confirmation before and after
system test, 5.10.3.1.2 Conduction check oftelemetry/command interfaces and array circuits(simplified irradiation test) shall be carried out as
END-to-END confirmation operation to the extent
possible.
• In a deployment test, power shall be supplied to
explosive devices from the spacecraft’s power
supply system.

JERG-2-215A

Appendix I Definition of name• Definition of name
The name of components in the solar array wing system used in the design standard document
shall be defined as follows:
Solar array: Electric circuits in solar array wing

CIC
moduleconnection unit
to 1 shunt
element = arraycircuit

1 series
circuit(Solar cell) panel

Substrate; Structure attachingarray circuits(Solar cell) paddle

Yoke

Boom

(Solar cell) paddle subsystem[1] Solar cell: Solar cell element itself. Simply called cell or bear cell.
[2] CIC (Cover-glass Integrated Cell or Connector Integrated Solar Cell): Solar cell installed withinterconnector and cover glass.
[3] CIC module: Unit where CIC is connected to arbitrary series-parallel circuit unit withinterconnector[4] Array circuit: Parallel-connected unit corresponding to 1 shunt element[5] One series circuit: One series circuit in array circuit.
[6] Substrate: Structure attaching array circuits[7] Solar cell panel: Unit where array circuits are assembled to one substrate[8] Boom: Beam element which is a part of structure coupling the paddle drive mechanism edge
and solar cell panel edge.
The solar array wing system has no boom in some cases.
[9] Yoke: Y-bend element which is a part of structure coupling the paddle drive mechanism edge
and solar cell panel edge and attached to the solar cell panel side. The solar array wing system
has no yoke in some cases.
[10] Hinge: Mechanism which rotates two structures relatively around itself[11] Solar array wing: Unit in which solar cell panels, boom and yoke are assembled.
Also called as PAD or SAP.

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JERG-2-215A[12] Solar array: Generic name for power generation circuits of solar array wing
[13] PDM (Paddle Drive Mechanism): Paddle drive mechanism. Also called SADM (Solar Array
Drive Mechanism).
[14] Solar array wing system: Entire subsystem including solar array wing and paddle drive
mechanism. Also called SPS or PDL.

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