PACE Solar Array Panels Specification - Amazon AWS

Effective Date: December 18, 2018

Expiration Date: December 18, 2023

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400-FORM-0002 (4/16/2014)

PACE-PWR-SPEC-0062, Revision A

Plankton, Aerosol, Cloud, ocean Ecosystem (PACE),

Code 427

PACE Solar Array Panels Specification

Goddard Space Flight Center

Greenbelt, Maryland

National Aeronautics and

Space Administration

GSFC PACE CMO

12/18/2018

Released

https://ipdtdms.gsfc.nasa.gov/

PACE Solar Array Panels Spec PACE-PWR-SPEC-0062, Revision A


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Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

400-FORM-0002 (4/16/2014)

PACE Spacecraft Solar Array Panels Specification

Signature/Approval Page

Prepared By:

John Lyons

Reviewed By:

Beth Weinstein

Melyane Ortiz-acosta

Trevin Dear

Dave Sohl

Approved By:

Andre’ Dress

Electronic Signatures available online at: https://ipdtdms.gsfc.nasa.gov/





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400-FORM-0002 (4/16/2014)

Preface

This document is under Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) Mission

configuration control. Changes to this document require prior approval of the PACE

Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be

submitted to the PACE Configuration Management Office (CMO), along with supportive

material justifying the proposed change. Changes to this document will be made by complete

revision.




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400-FORM-0002 (4/16/2014)

Change History Log

Revision Effective

Date

Description of Changes

(Reference the CCR & Approval Date)

Revision - 11/16/2018 Baseline Release following the approval of PACE-CCR-0459

Revision A 12/18/2018 Updated following the approval of PACE-CCR-0489




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400-FORM-0002 (4/16/2014)

Table of TBDs/TBRs/TBSs

Action Item

No.

Location Summary Individual/

Organization

Actionee




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400-FORM-0002 (4/16/2014)

Table of Contents

1.0 Introduction ........................................................................................................... 1 1.1 General Information .......................................................................................... 1 1.2 Scope ................................................................................................................ 1

2.0 Applicable Documents .......................................................................................... 2

3.0 Contract Description ............................................................................................. 4 3.1 Hardware Description ....................................................................................... 4

3.1.1 Flight Solar Array Panels ........................................................................... 4 3.1.2 Qualification Coupons ................................................................................ 4 3.1.3 Insulated Substrates .................................................................................. 4

4.0 Functional/Performance Requirements ................................................................ 5 4.1 Solar Array Panels Flight Unit Functional/Performance Requirements ............. 5

4.1.1 Test Condition Power................................................................................. 5 4.1.1.1 Flight Panels .......................................................................................... 5

4.1.1.2 Qualification Coupons ............................................................................ 5 4.1.2 Power at Highest Predicted Operating Temperature ................................. 5

4.1.3 Power at Highest Predicted Temperature .................................................. 5 4.1.4 End-of-Life Power ...................................................................................... 5 4.1.5 Solar Cell and Bypass Diode ..................................................................... 6

4.1.5.1 Solar Cell Mechanical ............................................................................. 6 4.1.5.2 Solar Cell Layout .................................................................................... 6

4.1.5.3 Solar Cell Power .................................................................................... 6 4.1.5.4 Limit to Solar Cell Shadowing ................................................................ 6

4.1.6 Solar Cell Cover ......................................................................................... 6

4.1.6.1 Cover Material and Thickness ................................................................ 6

4.1.6.2 Cover Orientation ................................................................................... 6 4.1.6.3 Antireflective/Indium Tin Oxide Coating (AR/ITO) .................................. 7

4.1.7 Blocking Diodes and Terminal Boards ....................................................... 7

4.1.8 Wire and Wire Layout ................................................................................ 7 4.1.8.1 Coarse Sun Sensor (CSS) Wiring .......................................................... 8 4.1.8.2 Hinge Damper Heater and Thermistor Wiring ........................................ 8

4.1.8.3 Monitor Solar Cell ................................................................................... 9 4.1.9 Connector Wiring and Connector Type ...................................................... 9 4.1.10 Platinum Resistor Thermometers (PRTs) ................................................ 14

4.1.11 Insulated Substrate .................................................................................. 14 4.1.11.1 Substrate Insulation Resistance ........................................................... 22 4.1.11.2 Substrate Grounding ............................................................................ 22

4.1.12 Panel Performance in Thermal Vacuum Environment ............................. 23 4.1.13 Panel Performance in Depressurization Environment ............................. 23 4.1.14 Allowable Degradation Due to Charged Particle Radiation ...................... 23 4.1.15 Allowable Degradation Due to Humidity................................................... 23

4.2 Resource Allocations ...................................................................................... 23 4.2.1 Mass Allocation ........................................................................................ 23




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4.3 Power .............................................................................................................. 23 4.3.1 Solar Array Panels Power Wire Redundancy .......................................... 23

4.4 Electrical Grounding ........................................................................................ 23 4.4.1 Primary Power DC Isolation ..................................................................... 23

4.4.2 Mechanical Contact Resistance ............................................................... 24 4.4.3 Connector DC Resistance ....................................................................... 24

4.5 Signal And Data Interfaces ............................................................................. 24 4.5.1 NA ............................................................................................................ 24 4.5.2 Passive Analog Telemetry ....................................................................... 24

4.5.2.1 Platinum Resistor Thermometer (PRT) ................................................ 24 4.5.2.2 PRT Performance................................................................................. 24

5.0 Physical Requirements ....................................................................................... 25 5.1 Interface Documentation ................................................................................. 25

5.2 Mass Properties .............................................................................................. 25 5.2.1 NA ............................................................................................................ 25

5.3 Physical Envelope ........................................................................................... 25 6.0 Environmental Requirements ............................................................................. 26

6.1 NA ................................................................................................................... 26

6.2 NA ................................................................................................................... 26 6.3 NA ................................................................................................................... 26

6.4 NA ................................................................................................................... 26 6.5 NA ................................................................................................................... 26 6.6 NA ................................................................................................................... 26

6.7 NA ................................................................................................................... 26 6.8 Pressure.......................................................................................................... 26

6.8.1 Operating Pressure Range ...................................................................... 26 6.8.2 NA ............................................................................................................ 26

6.9 NA ................................................................................................................... 26 6.10 Ground Environments ..................................................................................... 26

6.11 Thermal Requirements ................................................................................... 26 6.11.1 Flight Interface Design Temperature Limits ............................................. 26

6.12 Charged Particle Radiation Requirements ...................................................... 27 6.12.1 Definitions ................................................................................................ 27

6.12.2 Total Ionizing Dose .................................................................................. 27 6.12.2.1 Minimum TID Tolerance for EEE Parts ................................................ 27 6.12.2.2 NA ........................................................................................................ 33

6.12.3 NA ............................................................................................................ 33 6.12.4 NA ............................................................................................................ 33

6.12.4.1 NA ........................................................................................................ 33 6.12.4.2 NA ........................................................................................................ 33

6.12.4.3 NA ........................................................................................................ 33 6.12.5 Charging Environment ............................................................................. 33

6.13 Atomic Oxygen Fluence .................................................................................. 33 6.13.1 Atomic Oxygen Analysis .......................................................................... 33




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400-FORM-0002 (4/16/2014)

6.13.2 Atomic Oxygen Testing ............................................................................ 33 7.0 Cleanliness ......................................................................................................... 34

7.1 Surface Contamination ................................................................................... 34 7.1.1 Surface Contamination Levels at Delivery ............................................... 34

7.1.1.1 Particulate Contamination .................................................................... 34 7.1.1.2 Molecular Contamination ...................................................................... 34

7.1.2 Surface Contamination Generation .......................................................... 34 7.1.2.1 Particulate Generation .......................................................................... 34 7.1.2.2 Molecular Generation ........................................................................... 35

7.2 Electrostatic Cleanliness ................................................................................. 35 7.2.1 Conductive Surface Ground Path ............................................................ 35 7.2.2 Conductive Surface Resistivity ................................................................ 36 7.2.3 Exposed Harness Specific Requirements ................................................ 36

7.3 Magnetic Cleanliness ...................................................................................... 36 8.0 Design & Construction Requirements ................................................................. 37

8.1 Parts, Materials & Processes (PMP) ............................................................... 37 8.1.1 EEE Parts ................................................................................................ 37 8.1.2 Materials .................................................................................................. 37

8.1.2.1 Material Conductivity ............................................................................ 37 8.1.2.2 Material Limitations for Debris Casualty Area ...................................... 37

8.2 Electrical ......................................................................................................... 38 8.2.1 Test Sensors ............................................................................................ 38 8.2.2 Interface Requirements ............................................................................ 38

8.2.2.1 Connector Selection ............................................................................. 38 8.2.2.2 Signal Segregation ............................................................................... 38

8.2.2.3 Test and Flight Signal Isolation ............................................................ 38 8.2.2.4 Test Interfaces ..................................................................................... 39

8.2.3 Mitigation of Internal Charging ................................................................. 39 8.2.3.1 NA ........................................................................................................ 39

8.2.3.2 Floating Conductors ............................................................................. 39 8.2.3.3 Dielectric Structures ............................................................................. 39

8.3 Safety .............................................................................................................. 39 8.4 NA ................................................................................................................... 40

8.5 Identification and Marking ............................................................................... 40 8.6 Workmanship .................................................................................................. 40

8.6.1 Workmanship Standards .......................................................................... 40 8.6.2 Connector ................................................................................................ 40

8.6.2.1 GSE Cable Connectors ........................................................................ 40

8.6.2.2 Prevention of Connector Mismating ..................................................... 40 8.6.2.3 NA ........................................................................................................ 41

8.6.2.4 Connector Identification ........................................................................ 41 8.6.2.5 Protection of Unused Test Connectors................................................. 41 8.6.2.6 Connector Savers................................................................................. 41

8.7 Reliability and Mission Lifetime ....................................................................... 41




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400-FORM-0002 (4/16/2014)

8.7.1 Mission Life .............................................................................................. 41 8.8 Ground Handling ............................................................................................. 41

8.8.1 Ground Support Equipment (GSE) Design .............................................. 41 8.8.2 NA ............................................................................................................ 42

8.8.3 NA ............................................................................................................ 42 8.8.4 NA ............................................................................................................ 42 8.8.5 NA ............................................................................................................ 42 8.8.6 NA ............................................................................................................ 42 8.8.7 GSE Cleanliness ...................................................................................... 42

8.8.8 GSE Bakeout ........................................................................................... 42 8.8.9 Test Harness ........................................................................................... 42

9.0 Mechanical Design Requirements ...................................................................... 43 10.0 Logistics ............................................................................................................. 44

10.1 NA ................................................................................................................... 44 10.2 Ground Support Equipment ............................................................................ 44

10.3 Transportation Equipment ............................................................................... 44 11.0 Verification Requirements .................................................................................. 46

11.1 Verification Methods ....................................................................................... 46

11.1.1 Inspection ................................................................................................ 46 11.1.2 Analysis ................................................................................................... 46

11.1.3 Test .......................................................................................................... 46 11.2 Inspection Requirements ................................................................................ 46

11.2.1 Visual Inspection ...................................................................................... 46

11.2.2 Physical Measurement............................................................................. 47 11.2.3 Documentation Search ............................................................................ 47

11.3 Analysis Requirements ................................................................................... 47 11.4 Test Requirements .......................................................................................... 47

11.4.1 Definitions ................................................................................................ 47 11.4.2 NA ............................................................................................................ 48

11.4.3 Test Tolerances ....................................................................................... 48 11.4.4 Test Restrictions ...................................................................................... 48

11.4.4.1 Failure During Tests ............................................................................. 48 11.4.4.2 Modification of Hardware ...................................................................... 49

11.4.4.3 External Adjustment ............................................................................. 49 11.4.4.4 Re-Test Requirements ......................................................................... 49

11.5 Required Tests ................................................................................................ 49 11.5.1 NA ............................................................................................................ 50 11.5.2 NA ............................................................................................................ 50

11.5.3 NA ............................................................................................................ 50 11.5.4 NA ............................................................................................................ 50

11.5.5 NA ............................................................................................................ 50 11.5.6 NA ............................................................................................................ 50 11.5.7 NA ............................................................................................................ 50 11.5.8 NA ............................................................................................................ 50




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400-FORM-0002 (4/16/2014)

11.5.9 NA ............................................................................................................ 50 11.5.10 Thermal Vacuum Bake-out ...................................................................... 50 11.5.11 Thermal Vacuum Test.............................................................................. 53

11.5.11.1 Thermal Vacuum Test Parameters ................................................... 53

11.5.12 NA ............................................................................................................ 55 11.5.13 NA ............................................................................................................ 55 11.5.14 NA ............................................................................................................ 55 11.5.15 NA ............................................................................................................ 55 11.5.16 Solar Cell and Bypass Diode Qualification Tests ..................................... 55

11.5.17 Solar Array Panel Qualification Tests ...................................................... 55 11.5.17.1 Solar Array Panel Life Cycle Coupon Tests ..................................... 56

11.5.18 Flight Solar Array Panel Tests ................................................................. 57 11.5.19 Test Condition Power Verification ............................................................ 57

11.5.20 Bypass Diode Functionality Verification ................................................... 57 11.5.21 Substrate Insulation Resistance Verification ............................................ 58

11.5.22 Solar Cell Mechanical Verification ........................................................... 58 11.5.23 Cover Orientation Verification .................................................................. 59 11.5.24 Cover Grounding Verification ................................................................... 59

11.5.25 Flight Connector Type Verification ........................................................... 59 11.5.26 Platinum Resistance Thermometer Type Verification .............................. 59

11.5.27 Platinum Resistance Thermometer Performance Verification .................. 59 11.5.28 Parts and Assembly Layout Verification................................................... 60 11.5.29 Mission Life Verification ........................................................................... 60

11.5.30 Shelf Life Verification ............................................................................... 60 11.5.31 Substrate Ground Verification .................................................................. 60

11.5.32 Cleanliness Verification............................................................................ 60 Appendix A Abbreviations and Acronyms .................................................................... 62




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400-FORM-0002 (4/16/2014)

List of Figures Figure Page

Figure 1-1. PACE Spacecraft ......................................................................................... 1 Figure 4-1. PACE Solar Array Conceptual Diagram ....................................................... 8 Figure 4-2. Panel 3 to Panel 2 Electrical Interface ....................................................... 11

Figure 4-3. Panel 2 to Panel 1 Electrical Interface ....................................................... 12 Figure 4-4. Panel 1 to SADA Electrical Interface .......................................................... 13 Figure 4-5. Panel 1 Front and Side View ...................................................................... 15 Figure 4-6. Panel 1 Back View (KOZ = “Keep-Out Zone”) ........................................... 16 Figure 4-7. Panel 2 Front and Side View ...................................................................... 17

Figure 4-8. Panel 2 Back View (KOZ = “Keep-Out Zone”) ........................................... 18 Figure 4-9 Panel 3 Front and Side View........................................................................ 19

Figure 4-10. Panel 3 Back View (KOZ = “Keep-Out Zone”) ......................................... 20 Figure 4-11. Qualification Coupon Front and Side View............................................... 21

Figure 4-12. Qualification Coupon Back View (KOZ = “Keep-Out Zone”) .................... 22 Figure 6-1. Total Ionizing Dose-Depth Curve (includes x2 margin) .............................. 28

List of Tables Table Page

Table 2-1. Applicable Documents .................................................................................. 2 Table 6-1. Solar Array Panel Temperature Environment .............................................. 27 Table 6-2. Dose (including x2 margin) as a Function of Shielding ............................... 29

Table 6-3. Trapped Electron Spectrum ........................................................................ 30

Table 6-4. Trapped Proton Spectrum ........................................................................... 31 Table 6-5. Solar Proton Spectrum ................................................................................ 32 Table 8-1. Limited materials for debris casualty area ................................................... 37

Table 11-1. Test Tolerances ........................................................................................ 48




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400-FORM-0002 (4/16/2014)

1.0 INTRODUCTION

1.1 GENERAL INFORMATION

The Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) mission is a strategic climate continuity

mission that will extend the high quality ocean ecological, ocean biogeochemical, cloud, and

aerosol particle data records begun by NASA in the 1990s. The mission will be capable of

collecting radiometric and polarimetric measurements of the ocean and atmosphere, from which

these biological, biogeochemical, and physical properties will be determined. PACE data

products will not only add to existing critical climate and Earth system records, but also answer

new and emerging advanced science questions related to Earth’s changing climate.

An artist’s conception of the PACE spacecraft is presented in Figure 1-1.

Figure 1-1. PACE Spacecraft

1.2 SCOPE

This specification describes the electrical, mechanical, environmental, and verification testing

requirements for space-qualified Solar Array Panels that will provide electric power for the

NASA Goddard Space Flight Center (GSFC) PACE Mission.






400-FORM-0002 (4/16/2014)

2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall

apply to the fabrication and to the electrical, mechanical, and environmental requirements of the

Solar Array Panels to the extent specified herein. In the event of conflict between this

specification and any referenced document, this specification will govern, with the exception of

the PACE Solar Array Panels Statement of Work (PACE-PWR-SOW-0025), in which case the

Statement of Work takes precedence.

The following is a list of the applicable specifications and publications.

Table 2-1. Applicable Documents

Document Number Title

AIAA S-111A-2014 Qualification and Quality Requirements for Space

Solar Cells

AIAA S-112A-2013 Qualification and Quality Requirements for

Electrical Components on Space Solar Panels

PACE-PWR-SOW-0025 PACE Solar Array Panels Statement of Work

NFPA 70 National Fire Protection Association National

Electric Code

NASA-STD-5001B Structural Design And Test Factors Of Safety For

Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements

NASA-STD-6016 Standard Materials and Processes Requirements for

Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC1246E Product Cleanliness Levels And Contamination

Control Program

ASTM E-595-07 Standard Test Method for Total Mass Loss and

Collected Volatile Condensable Materials from

Outgassing in a Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and

Aluminum Alloys

AMS 2488 Anodic Treatment - Titanium and Titanium Alloys

Solution pH 13 Or Higher

MIL-A-8625F Anodic Coatings for Aluminum and Aluminum

Alloys

EEE-INST-002 Instructions for EEE Parts Selection, Screening,

Qualification, and Derating

DOD-HDBK-83575 General Handbook for Space Vehicle Wiring

Harness Design and Testing

GSFC-STD-7000A General Environmental Verification Standard

(GEVS)






400-FORM-0002 (4/16/2014)

Document Number Title

NASA-STD-5019A Fracture Control Requirements for Spaceflight

Hardware

NASA-STD-5020 Requirements for Threaded Fastening Systems in

Spaceflight Hardware

NASA-STD-5017A Design and Development Requirements for

Mechanisms

FAA AC 20-71 Federal Aviation Administration Advisory Circular

(AC) 20-71, “Dual Locking Devices on Fasteners".

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning,

General Practices for

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for

Stress Corrosion Cracking Resistance in Sodium

Chloride Environments

AIAA S-111-2005 Qualification and Quality Requirements for Space

Solar Cells






400-FORM-0002 (4/16/2014)

3.0 CONTRACT DESCRIPTION

3.1 HARDWARE DESCRIPTION

3.1.1 Flight Solar Array Panels

Each flight solar array panel is one-third of a deployable solar array for the PACE spacecraft.

The PACE solar array will convert solar energy to electrical power for the spacecraft.

The contract includes the population of insulated qualification and flight substrates with solar

cells and associated components and testing of the completed solar array panels.

3.1.2 Qualification Coupons

The Qualification Coupons are small solar array panels that are representative of the Flight Solar

Array Panels in every respect except size. They are to be used for tests to qualify the PACE Solar

Array Panel design.

3.1.3 Insulated Substrates

GSFC will provide insulated substrates for the qualification coupons and flight solar array

panels. Insulated Substrates are unpopulated panels. Solar cells, harnessing, etc. are not included

at this level of assembly. The Insulated Substrates will be made of composite facesheet with

aluminum honeycomb core with a layer of insulation on top of the front side facesheet.






400-FORM-0002 (4/16/2014)

4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

The Solar Array Panels shall be designed to withstand the operational and non-operational

environments specified in the following section without degradation to mission goals and

performance requirements.

4.1 SOLAR ARRAY PANELS FLIGHT UNIT

FUNCTIONAL/PERFORMANCE REQUIREMENTS

4.1.1 Test Condition Power

4.1.1.1 Flight Panels

Under simulated Air Mass Zero (AM0) illumination at 28 degrees Celsius (º C), normal

incidence, the sum of all the solar cell string power output on all three flight panels, taken at the

test connectors, shall exceed 3,459.0 watts (W) at a load voltage of 46.8 volts (V).

4.1.1.2 Qualification Coupons

a. The qualification coupon(s) shall have at least one example of the type of circuit on the

flight panels.

b. The qualification coupon output from each of these circuits shall be proportional to the

panel power requirement of Section 4.1.1.1, Flight Panels.

4.1.2 Power at Highest Predicted Operating Temperature

The contractor shall extrapolate the current-voltage curves of the flight panels and qualification

coupon(s) in accordance with the SOW, PACE-PWR-SOW-0025.

4.1.3 Power at Highest Predicted Temperature

The contractor shall extrapolate the current-voltage curves of the flight panels and qualification

coupon(s) to 1.405 AM0, 115°C at BOL and to 1.033 AM0, 93°C at BOL in accordance with the

SOW, PACE-PWR-SOW-0025.

4.1.4 End-of-Life Power

The contractor shall predict the panel End of Life (EOL) I-V curve in accordance with the SOW,

PACE-PWR-SOW-0025.






400-FORM-0002 (4/16/2014)

4.1.5 Solar Cell and Bypass Diode

4.1.5.1 Solar Cell Mechanical

No cell on a panel shall have a crack, visible at 7x or less magnification.

4.1.5.2 Solar Cell Layout

a. The Contractor shall propose the solar cell layout. The solar cell layout of the three solar

array panels shall be divided into 24 parallel segments, each having an equal number of

solar cell strings in parallel, except for one segment, which may have a smaller number of

strings if the total power generated by the sum of all the parallel segments is sufficient to

meet the test condition power requirement in 4.1.1. At BOL, 1.05 AM0, 115°C, the short-

circuit current of any segment shall not exceed 3.8 amperes.

Flight and qualification panel dimensions including stay-out zones are included in

Figures 4-5 through 4-12.

b. To the extent practicable, the Contractor shall not divide individual segments between

panels.

4.1.5.3 Solar Cell Power

At 28°C and Air Mass Zero, the PACE solar cells shall produce sufficient power to meet the test

condition power specified in Section 4.1.1.

4.1.5.4 Limit to Solar Cell Shadowing

No cell shall experience more than 3 percent degradation over the life of the mission in

maximum power output as a result of the cell being repeatedly shadowed.

4.1.6 Solar Cell Cover

4.1.6.1 Cover Material and Thickness

Each coverglass shall be 100 µm (nominal) thick cerium dioxide doped glass (CMG or

equivalent) and shall cover 100 percent of the active area of each solar cell.

4.1.6.2 Cover Orientation

The contractor shall orient the cover using either an etch symbol or a stain.






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4.1.6.3 Antireflective/Indium Tin Oxide Coating (AR/ITO)

The contractor shall use covers that are coated with AR/ITO having a nominal resistivity of 109

ohms per square with the coating on each cover grounded to the solar cell front contact pad.

4.1.7 Blocking Diodes and Terminal Boards

a. The Contractor shall design the terminal (sometimes called diode) boards.

b. The Contractor shall insulate the terminal boards from and bond them to the rear face

sheet with an adhesive or hardware proposed by the Contractor and approved by the

GSFC.

c. Each terminal board shall contain parallel redundant blocking diodes which connect in

series with the solar cell strings.

d. The contractor shall conformally coat the terminal boards prior to delivery to GSFC.

4.1.8 Wire and Wire Layout

Figure 4-1 is a conceptual diagram depicting components that are to be wired to the panel

connectors.

a. For the Panel 3 to Panel 2 power and signal harnesses, the contractor shall provide 90-cm

pigtails from the outboard edge of Panel 2. These will be formed into shape at GSFC and

mated to the Panel 3 inboard connectors in accordance with the pin assignments to be

defined by the contractor.

b. Similarly, for the Panel 2 to Panel 1 power and signal harnesses, the contractor shall

provide 90-cm pigtails from the outboard edge of Panel 1. These will be formed into

shape at GSFC and mated to the Panel 2 inboard connectors in accordance with the pin

assignments to be defined by the contractor.

c. The contractor shall protect wire wherever abrasion may be a problem.

d. The contractor shall use stress relief between wire tie points to avoid strains, particularly

on the solar cell string terminations.

e. The contractor shall address how it will stake wire.






400-FORM-0002 (4/16/2014)

Figure 4-1. PACE Solar Array Conceptual Diagram

4.1.8.1 Coarse Sun Sensor (CSS) Wiring

a. The contractor shall route wiring for the Coarse Sun Sensors (CSSs) to the panel

connectors on the backs of the panels using GFE twisted, shielded wire.

b. The Contractor shall ground the shield to the connector shell.

c. NASA/GSFC will attach the CSSs after the panels have been delivered. To provide for

this, the Contractor shall terminate the CSS wiring in a 30 cm pigtail at the CSS locations

on the outboard panel.

4.1.8.2 Hinge Damper Heater and Thermistor Wiring

a. The contractor shall route wiring for the hinge damper heaters to the panel connectors on

the backs of the panels using twisted wire.

Panel 3 Panel 2 Panel 1

Segment 1

Segment 8

SubstrateGround

SubstrateGround

SubstrateGround

SubstrateGround

SubstrateGround

SubstrateGround

Segment 9

Segment 16

Segment 17

Segment 24

Pos. Plus Rtn. Wire Pair

Twisted, Shielded Pair

Segment Incl. Terminal Boards, Blocking Diodes, Test Connectors

Ground Wire

1. All pos. wires are run adjacent to their returns to the extent practicable.2. PRT, CSS, and Hinge Pot wires twisted and shielded. Shield grounded to connector shell.

Connector

3. This conceptual diagram does not show physical locations of components.4. Most segments not depicted. Their wires would run in a manner similar to those shown.

Coarse Sun Sensor (CSS)

PRT

PRTPRT

PRT

Hinge Damper w/Heater

PRT

PRT

SAD

A In

terface

Monitor Solar Cell

Twisted, Shielded Trio

Thermistor Thermistor






400-FORM-0002 (4/16/2014)

b. The contractor shall route wiring for the hinge damper thermistors to the panel

connectors on the backs of the panels using twisted, shielded wire.

c. NASA/GSFC will install the hinges with dampers and heaters after the panels have been

delivered. To provide for this, the contractor shall terminate the hinge damper heater

wiring in a 30 cm pigtail at the hinge damper locations shown in Figures 4-5 through 4-8,

on Panels 1 and 2.

d. NASA/GSFC will install thermistors on the hinge dampers after the panels have been

delivered. To provide for this, the contractor shall terminate the hinge damper thermistor

wiring in a 30 cm pigtail at the hinge damper locations shown in Figures 4-5 through 4-8,

on Panels 1 and 2.

4.1.8.3 Monitor Solar Cell

The contractor shall place a single monitor solar cell on Panel 1. The contractor shall wire the

sensors to the panel connector using twisted, shielded wire.

4.1.9 Connector Wiring and Connector Type

Allowable power, signal, and test connector locations are shown in Figures 4-5 through 4-10.

The contractor shall attach the connectors to the substrates using click bonds, part number

CB4000C08CRA8.

a. Government-furnished Sommer DW-747 connectors shall be used for power and signals.

Power and signal wires shall be routed to separate connectors. The contractor shall

define the connector pin assignments. Power and signal connector pin functions are given

in Figures 4-2 through 4-4.

b. Test connectors shall be Glenair Series 79 Micro-Crimp connectors which the Contractor

must fix to each panel to test each individual solar cell string as specified in paragraph i

of this section.

c. A test connector on one panel shall not be in same position as the flight and test

connector from another panel when the panels are stowed unless approved by the

NASA/GSFC COR. This is to insure adequate dynamic clearance.

d. For power, the contractor shall parallel the strings at a terminal board or boards for each

segment.

e. The contractor shall use two parallel-redundant JAN TXV 1N5811 blocking diodes for

each solar cell string on the terminal board or boards.

f. The cathodes (output) of the diodes shall be paralleled and the circuit returns shall be

paralleled to form each of the segments.






400-FORM-0002 (4/16/2014)

g. The contractor shall use wire meeting MIL-W-22759/44.

h. From each segment to the panel power connectors, the contractor shall provide two

AWG #20 pair wiring; that is two positive wires and two return wires.

i. From the anode side of the blocking diodes for each solar cell string to the panel test

connectors, the contractor shall provide one AWG #20 wire. From the return side of each

solar cell string to the panel test connectors, the contractor shall provide one AWG #20

wire; that is one positive wire and one return wire.

j. From each hinge damper heater to the panel power connectors, the contractor shall

provide one AWG #20 pair wiring; that is one positive wire and one return wire.

k. From each PRT, thermistor, and CSS, to the panel signal connectors, the contractor shall

provide one AWG #24 twisted, shielded pair wiring; that is, one positive wire and one

return wire, twisted and shielded.

l. From the monitor solar cell to the panel signal connector, the contractor shall provide one

AWG #20 twisted, shielded pair wiring, that is, one positive wire and one return wire,

twisted and shielded.






400-FORM-0002 (4/16/2014)

Figure 4-2. Panel 3 to Panel 2 Electrical Interface

Substrate Ground

Segment 17 Pos.

Segment 18 Pos.

Segment 19 Pos.

Segment 20 Pos.

Segment 21 Pos.

Segment 22 Pos.

Segment 23 Pos.

Segment 24 Pos.

Segment 17 Rtn.

Segment 18 Rtn.

Segment 19 Rtn.

Segment 20 Rtn.

Segment 21 Rtn.

Segment 22 Rtn.

Segment 23 Rtn.

Segment 24 Rtn.

Substrate Ground

Segment 17 Pos.

Segment 18 Pos.

Segment 19 Pos.

Segment 20 Pos.

Segment 21 Pos.

Segment 22 Pos.

Segment 23 Pos.

Segment 24 Pos.

Segment 17 Rtn.

Segment 18 Rtn.

Segment 19 Rtn.

Segment 20 Rtn.

Segment 21 Rtn.

Segment 22 Rtn.

Segment 23 Rtn.

Segment 24 Rtn.

Panel 3 Panel 2

Notes1. All wires except substrate ground and shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.6. The contractor shall furnish the outboard side of Panel 2 with a pigtail per this sepcification. The pigtail will be made into an inter-panel harness at GSFC and inserted into connectors on the inboard side of Panel 3 per the pin assignments defined by the contractor.

CSS 1 Pos.

CSS 2 Pos.

CSS 3 Pos.

PRT 5 Pos.

PRT 6 Pos.

CSS 1 Rtn.

CSS 2 Rtn.

CSS 3 Rtn.

PRT 5 Rtn.

PRT 6 Rtn.CSS 1 Shield

CSS 2 Shield

CSS 3 Shield

PRT 5 Shield

PRT 6 Shield






400-FORM-0002 (4/16/2014)

Figure 4-3. Panel 2 to Panel 1 Electrical Interface

Substrate Ground

Segment 9 Pos.

Segment 10 Pos.

Segment 11 Pos.

Segment 12 Pos.

Segment 13 Pos.

Segment 14 Pos.

Segment 15 Pos.

Segment 16 Pos.

Segment 9 Rtn.

Segment 10 Rtn.

Segment 11 Rtn.

Segment 12 Rtn.

Segment 13 Rtn.

Segment 14 Rtn.

Segment 15 Rtn.

Segment 16 Rtn.

Panel 2 Panel 1

Notes1. All wires except substrate ground and shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.6. The contractor shall furnish the outboard side of Panel 1 with a pigtail per this sepcification. The pigtail will be made into an inter-panel harness at GSFC and inserted into connectors on the inboard side of Panel 2 per the pin assignments defined by the contractor.

PRT 3 Pos.

PRT 4 Pos.

PRT 3 Rtn.

PRT 4 Rtn.

PRT 3 Shield

PRT 4 Shield

Substrate Ground

Segment 9 Pos.

Segment 10 Pos.

Segment 11 Pos.

Segment 12 Pos.

Segment 13 Pos.

Segment 14 Pos.

Segment 15 Pos.

Segment 16 Pos.

Segment 9 Rtn.

Segment 10 Rtn.

Segment 11 Rtn.

Segment 12 Rtn.

Segment 13 Rtn.

Segment 14 Rtn.

Segment 15 Rtn.

Segment 16 Rtn.

Hinge Damper Heater 2 Pos.

Hinge Damper Heater 2 Rtn.

Hinge Damper 2 Thermistor Shield

Hinge Damper 2 Thermistor Rtn.

Hinge Damper 2 Thermistor Pos.

Power and Signal Wires from Panel 3






400-FORM-0002 (4/16/2014)

Figure 4-4. Panel 1 to SADA Electrical Interface

Substrate Ground

Segment 1 Pos.

Segment 2 Pos.

Segment 3 Pos.

Segment 4 Pos.

Segment 5 Pos.

Segment 6 Pos.

Segment 7 Pos.

Segment 8 Pos.

Segment 1 Rtn.

Segment 2 Rtn.

Segment 3 Rtn.

Segment 4 Rtn.

Segment 5 Rtn.

Segment 6 Rtn.

Segment 7 Rtn.

Segment 8 Rtn.

Panel 1 SADA

Notes1. All wires except shields go through to the Panel 1-SADA interface. 2. Shields are grounded to connector shells. 3. Connector shells are grounded to the substrate honeycomb.4. Substrate ground wires are imbedded in honeycomb cells on all three panels.5. The contractor shall define the connector pinouts per the requirements of this specification.

PRT 1 Pos.

PRT 2 Pos.

PRT 1 Rtn.PRT 2 Rtn.

PRT 1 Shield

PRT 2 Shield

Hinge Damper Heater 1 Pos.

Hinge Damper Heater 1 Rtn.

Hinge Damper 1 Thermistor Shield

Hinge Damper 1 Thermistor Rtn.

Hinge Damper 1 Thermistor Pos.

Power and Signal Wires from Panels 2 and 3

Substrate Ground

Segment 1 Pos.

Segment 2 Pos.

Segment 3 Pos.

Segment 4 Pos.

Segment 5 Pos.

Segment 6 Pos.

Segment 7 Pos.

Segment 8 Pos.

Segment 1 Rtn.

Segment 2 Rtn.

Segment 3 Rtn.

Segment 4 Rtn.

Segment 5 Rtn.

Segment 6 Rtn.

Segment 7 Rtn.

Segment 8 Rtn.

Monitor Solar Cell Pos.

Monitor Solar Cell Rtn.

Monitor Solar Cell Shield






400-FORM-0002 (4/16/2014)

4.1.10 Platinum Resistor Thermometers (PRTs)

a. The contractor shall use Goodrich Platinum Resistor Thermometer type

118MM2000AFAAAC.

b. The contractor shall mount two PRTs on each panel; one on the back of the front

facesheet, and the other on the back facesheet.

c. The contractor shall propose the location of the PRTs, except that on Panel 1, the PRT on

the back of the front facesheet shall be directly under the monitor solar cell.

d. The contractor shall run the PRT lead wires along the back facesheet of each panel to the

signal connector using twisted, shielded wire.

e. The contractor shall ground the shield to the connector shell.

4.1.11 Insulated Substrate

NASA/GSFC will provide insulated substrates for each flight solar array panel and each

qualification coupon. The substrates will be made of composite facesheet with perforated

aluminum honeycomb core. The solar-cell-side facesheet will include co-cured Kapton

insulation.

NASA/GSFC will equip the flight substrates with fittings suitable for attaching handling fixtures.

The handling points will be provided by GSFC along with the GFE substrates.

The qualification substrate dimensions and keep-out zones are described in Figures 4-11 and 4-

12.

Prior to delivery to the contractor, the backs of the flight and qualification substrates will be

coated by NASA/GSFC with white paint.

The contractor shall provide drawings that show areas on the back of each flight and

qualification substrate that are not to be coated.






400-FORM-0002 (4/16/2014)

Figure 4-5. Panel 1 Front and Side View






400-FORM-0002 (4/16/2014)

Figure 4-6. Panel 1 Back View (KOZ = “Keep-Out Zone”)






400-FORM-0002 (4/16/2014)

Figure 4-7. Panel 2 Front and Side View






400-FORM-0002 (4/16/2014)







400-FORM-0002 (4/16/2014)

Figure 4-9 Panel 3 Front and Side View






400-FORM-0002 (4/16/2014)







400-FORM-0002 (4/16/2014)

Figure 4-11. Qualification Coupon Front and Side View






400-FORM-0002 (4/16/2014)

Figure 4-12. Qualification Coupon Back View (KOZ = “Keep-Out Zone”)

4.1.11.1 Substrate Insulation Resistance

The resistance between the substrate and the solar cell circuits shall be greater than 10 megohms

for the flight panels and qualification coupon(s).

4.1.11.2 Substrate Grounding

a. The contractor shall run two AWG #20 ground wires from each substrate to ground

contacts on the panel connectors.

b. The contractor shall ground both the facesheets and the aluminum honeycomb core.

c. The resistance between the substrate core and ground shall be less than 2 ohms.






400-FORM-0002 (4/16/2014)

d. The substrate ground leads from the three panels shall be combined into two leads at the

inboard panel connector at the spacecraft interface.

4.1.12 Panel Performance in Thermal Vacuum Environment

No flight panel shall degrade in peak power by more than 2 percent nor incur damage that may

question its reliability to meet the requirements of this document after exposure to the flight

thermal cycles in the vacuum of space as specified in Error! Reference source not found..

4.1.13 Panel Performance in Depressurization Environment

The flight panels and qualification coupons shall meet the requirements of this document after

depressurization from 1 atmosphere to 1E-05 Torr in thirty seconds.

4.1.14 Allowable Degradation Due to Charged Particle Radiation

The contractor shall consider hard particle radiation in its computation of end of life power, see

Sections 4.1.4 and 6.12.

4.1.15 Allowable Degradation Due to Humidity

The qualification and flight solar panels shall meet the requirements of this document during and

after exposure of 20% to 70% relative humidity after 2 years on the ground before launch.

4.2 RESOURCE ALLOCATIONS

4.2.1 Mass Allocation

The total add-on mass (total assembled panel mass minus the substrate mass) of the three Solar

Array Panels shall be less than 18.6 kg.

4.3 POWER

4.3.1 Solar Array Panels Power Wire Redundancy

The Solar Array Panels shall provide redundant contacts or connections for the segment output

power and return lines.

4.4 ELECTRICAL GROUNDING

4.4.1 Primary Power DC Isolation

The Solar Array Panels power and power returns shall be isolated from signal grounds by a DC

resistance of greater than or equal to 1 MΩ.






400-FORM-0002 (4/16/2014)

4.4.2 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal

to the component, and at the spacecraft interface) shall be less than or equal to 2.5 mΩ DC

resistance.

4.4.3 Connector DC Resistance

Component connectors shall be electrically connected to the insulated substrate with a DC

resistance less than or equal to 10 mΩ.

4.5 SIGNAL AND DATA INTERFACES

4.5.1 NA

4.5.2 Passive Analog Telemetry

4.5.2.1 Platinum Resistor Thermometer (PRT)

The contractor shall use Goodrich Platinum Resistor Thermometer type 118MM2000AFAAAC.

4.5.2.2 PRT Performance

The PRT shall meet its manufacturer’s specifications for resistance versus temperature.






400-FORM-0002 (4/16/2014)

5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE DOCUMENTATION

The contractor shall use metric units when interfacing with NASA GSFC including any

drawings, documents, models, except for the following cases:

Heritage Hardware: Hardware that has been previously qualified, or of similar design

heritage, may be specified in English units where use of metric equivalents would

lead to additional cost to the program.

Fasteners: Although bolt patterns will be defined using metric dimensioning, use of

English fasteners (with hole dimensioning and tolerancing) is permitted.

The dimensions of the GFE insulated substrates including stay-out zones are defined in Figures

4-1 through 4-6.

The electrical interface is defined in this Specification.

5.2 MASS PROPERTIES

5.2.1 NA

5.3 PHYSICAL ENVELOPE

No Solar Array Panel shall exceed the thermal and mechanical volume envelope described in

Figures 4-1 through 4-6.






400-FORM-0002 (4/16/2014)

6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the spacecraft components are specified in this section.

The SA Panels shall meet its performance requirements in Section 4.0 during and after exposure

to the environments specified in this section.

6.1 NA

6.2 NA

6.3 NA

6.4 NA

6.5 NA

6.6 NA

6.7 NA

6.8 PRESSURE

6.8.1 Operating Pressure Range

The Solar Array Panels shall be designed to meet all performance requirements while operating

over a pressure range of 1.08 x 105 Pa (813 Torr) to 1.3 x 10-12 Pa (1 x 10-14 Torr).

6.8.2 NA

6.9 NA

6.10 GROUND ENVIRONMENTS

The Solar Array Panels shall meet all of their performance requirements during exposure to air

temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

6.11 THERMAL REQUIREMENTS

6.11.1 Flight Interface Design Temperature Limits

The solar array panels shall meet the requirements of this document after exposure to the

temperature extremes and number of eclipse cycles in Error! Reference source not found..






400-FORM-0002 (4/16/2014)

Table 6-1. Solar Array Panel Temperature Environment

Number of Cycles Temperature Limits (ºC)

Operational: 17,057 -90°C to +115°C

Survival: 1 -100°C to +135°C

6.12 CHARGED PARTICLE RADIATION REQUIREMENTS

Components containing electronic parts will be exposed to a natural space radiation environment

that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2)

particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray

particles.

a. For solar cell degradation, the contractor shall use the surface-incident trapped electron,

trapped proton, and solar proton spectra in Tables 6-6 through 6-8 to determine the 1-

MeV electron equivalent fluences for Isc, Voc, and Pmax at EOL.

b. The contractor shall assume 750 µm backshielding for the purpose of this calculation.

6.12.1 Definitions

Total Ionizing Dose (TID) - the mean energy deposited by ionizing radiation in a device region

divided by the mass of the region. This is often given in units of rad(Si), where 1 rad(Si) = 100

erg deposited per gram of silicon.

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts

The top-level total ionizing dose requirement is shown in Error! Reference source not found.

and Error! Reference source not found.. The dose values are calculated as a function of

aluminum shield thickness in units of krad in silicon. For a nominal 2.54 mm (100 mils) of

equivalent aluminum shielding and a 3-year mission life, the expected dose is 6.67 krad-Si. This

includes a factor of 2 margin. EEE parts and materials shall be selected according to the level of

shielding shown in Error! Reference source not found. and Error! Reference source not

found..






400-FORM-0002 (4/16/2014)

Figure 6-1. Total Ionizing Dose-Depth Curve (includes x2 margin)






400-FORM-0002 (4/16/2014)

Table 6-2. Dose (including x2 margin) as a Function of Shielding

Dose

(mm): (mils): (g/cm2): (krad-Si):

0.002935 0.115532 0.000792 7.98E+03

0.004039 0.159003 0.00109 5.56E+03

0.005521 0.217353 0.00149 3.96E+03

0.007596 0.299042 0.00205 2.94E+03

0.010412 0.409907 0.00281 2.28E+03

0.014302 0.563075 0.00386 1.74E+03

0.019601 0.771675 0.00529 1.34E+03

0.0269 1.059047 0.00726 1.05E+03

0.036941 1.454366 0.00997 8.06E+02

0.050761 1.998477 0.0137 6.12E+02

0.069658 2.742436 0.0188 4.54E+02

0.095595 3.763556 0.0258 3.38E+02

0.131165 5.163948 0.0354 2.58E+02

0.180073 7.089488 0.0486 1.91E+02

0.246767 9.715225 0.0666 1.38E+02

0.339027 13.34749 0.0915 9.40E+01

0.466857 18.38015 0.126 6.32E+01

0.637297 25.09037 0.172 4.16E+01

0.87443 34.42632 0.236 2.66E+01

1.204194 47.40913 0.325 1.70E+01

1.64882 64.91404 0.445 1.10E+01

2.263885 89.12916 0.611 7.28E+00

3.108674 122.3885 0.839 5.40E+00

4.260995 167.7554 1.15 4.14E+00

5.854237 230.4813 1.58 3.22E+00

8.040313 316.5471 2.17 2.54E+00

11.04154 434.7053 2.98 2.06E+00

15.15432 596.6257 4.09 1.65E+00

20.78625 818.3545 5.61 1.32E+00

Aluminum Shield Thickness






400-FORM-0002 (4/16/2014)

Table 6-3. Trapped Electron Spectrum

Energy

(>MeV)

Integral

Fluence

(e/cm2)

0.001 4.416E+14

0.0013 4.178E+14

0.0017 3.935E+14

0.0021 3.757E+14

0.0028 3.520E+14

0.0036 3.314E+14

0.0046 3.110E+14

0.0059 2.893E+14

0.0077 2.644E+14

0.01 2.391E+14

0.013 2.130E+14

0.016 1.936E+14

0.021 1.733E+14

0.027 1.603E+14

0.035 1.516E+14

0.04 1.483E+14

0.07 9.981E+13

0.1 7.332E+13

0.25 2.767E+13

0.5 8.377E+12

0.75 3.175E+12

1 1.461E+12

1.5 3.884E+11

2 1.108E+11

2.5 3.907E+10

3 1.711E+10

3.5 7.712E+09

4 3.375E+09

4.5 1.307E+09

5 3.786E+08

5.5 7.789E+07

6 1.022E+07

6.5 0.000E+00

7 0.000E+00

7.5 0.000E+00

8 0.000E+00






400-FORM-0002 (4/16/2014)

Table 6-4. Trapped Proton Spectrum

Energy

(>MeV)

Integral

Fluence

(p/cm2)

Energy

(>MeV)

Integral

Fluence

(p/cm2)

0.00115 8.302E+13 6 1.636E+11

0.0021 7.756E+13 8 1.199E+11

0.0037 7.037E+13 10 9.614E+10

0.0065 6.194E+13 15 6.790E+10

0.01155 5.266E+13 20 5.530E+10

0.0204 4.268E+13 30 4.442E+10

0.036 3.107E+13 50 3.347E+10

0.06375 1.705E+13 60 3.016E+10

0.085 9.268E+12 80 2.533E+10

0.1 4.917E+12 100 2.153E+10

0.2 3.626E+12 150 1.447E+10

0.4 1.992E+12 200 9.695E+09

0.6 1.232E+12 300 4.056E+09

0.8 9.084E+11 400 1.637E+09

1 7.678E+11 700 1.912E+08

2 4.749E+11 1200 1.101E+07

4 2.523E+11 2000 2.000E+00






400-FORM-0002 (4/16/2014)

Table 6-5. Solar Proton Spectrum

Energy

(>MeV)

Integral

Fluence

(p/cm2)

Energy

(>MeV)

Integral

Fluence

(p/cm2)

Energy

(>MeV)

Integral

Fluence

(p/cm2)

0.1 5.045E+11 2 5.693E+10 55 9.945E+08

0.11 4.711E+11 2.2 5.314E+10 63 7.684E+08

0.12 4.423E+11 2.5 4.842E+10 71 6.068E+08

0.14 3.951E+11 2.8 4.453E+10 80 4.752E+08

0.16 3.582E+11 3.2 4.004E+10 90 3.701E+08

0.18 3.287E+11 3.5 3.708E+10 100 2.948E+08

0.2 3.045E+11 4 3.295E+10 110 2.406E+08

0.22 2.841E+11 4.5 2.970E+10 120 2.002E+08

0.25 2.588E+11 5 2.700E+10 140 1.421E+08

0.28 2.383E+11 5.5 2.464E+10 160 1.043E+08

0.32 2.162E+11 6.3 2.149E+10 180 7.911E+07

0.35 2.026E+11 7.1 1.900E+10 200 6.159E+07

0.4 1.838E+11 8 1.670E+10 220 4.889E+07

0.45 1.686E+11 9 1.465E+10 250 3.540E+07

0.5 1.562E+11 10 1.300E+10 280 2.629E+07

0.55 1.458E+11 11 1.160E+10 320 1.828E+07

0.63 1.320E+11 12 1.041E+10 45 1.427E+09

0.71 1.210E+11 14 8.554E+09 50 1.183E+09

0.8 1.109E+11 16 7.171E+09 55 9.945E+08

0.9 1.018E+11 18 6.103E+09 320 1.828E+07

1 9.430E+10 25 3.778E+09 350 1.435E+07

1.1 8.801E+10 28 3.165E+09 400 9.895E+06

1.2 8.264E+10 32 2.549E+09 450 7.068E+06

1.4 7.383E+10 35 2.201E+09 500 5.295E+06

1.6 6.696E+10 45 1.427E+09






400-FORM-0002 (4/16/2014)

1.8 6.145E+10 50 1.183E+09

6.12.2.2 NA

6.12.3 NA

6.12.4 NA

6.12.4.1 NA

6.12.4.2 NA

6.12.4.3 NA

6.12.5 Charging Environment

The Solar Array Panels shall be designed to withstand the degradation of surface materials and

associated surface charging effects due to the radiation environment for the PACE mission orbit.

6.13 ATOMIC OXYGEN FLUENCE

When a component has surfaces that will be exposed to the external space environment, all external

materials shall survive an atomic oxygen fluence (Observatory orbit velocity direction) of 2.0E+20

atoms/cm2 without loss of structural integrity or loss of critical performance criteria.

6.13.1 Atomic Oxygen Analysis

An analysis shall be performed for all external materials and finishes to verify the compatibility with

the AO environment. The analysis should show that any AO degradation does not pose a

contamination hazard for other components on the spacecraft (source of particles, molecular films,

debris, etc.).

6.13.2 Atomic Oxygen Testing

If no data exists for the proposed material and finishes, a test shall be performed, exposing a

representative sample to AO and verifying no loss of structural integrity or loss of critical

performance criteria.






400-FORM-0002 (4/16/2014)

7.0 CLEANLINESS

The requirements in this section ensure the cleanliness of the Solar Array Panels at delivery, so

as not to adversely affect its own performance, as well as not be a source of contamination to

other items, including not generating contaminants following delivery in excess of that permitted

below by virtue of its design, materials of construction, or operation.

7.1 SURFACE CONTAMINATION

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

The Solar Array Panels shall meet IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent when

inspected with both UV and white light in a darkened room.

7.1.1.2 Molecular Contamination

The Solar Array Panels shall meet a molecular surface cleanliness level of IEST-STD-CC1246E

R3.3E-1 on all external and critical surfaces.

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

The Solar Array Panels contractor shall not employ any of the following particle generating

materials or processes into the Solar Array Panels design or construction without prior approval

by the NASA/GSFC COR:

Paints prone to shedding due to large paint pigment molecules, overspray, poor adhesion,

etc.

Dry lubricants (e.g. molybdenum disulfide).

Surfaces prone to corrosion or oxides because of a lack of corrosion protection or

dissimilar metals in close contact.

Fabrics with brittle constituents (e.g., composites, graphite or glass).

Perforated materials when material is highly susceptible to tear propagation (e.g., MLI).

Metal oxides (bare [untreated] aluminum and magnesium, iron, non- corrosion resistant

steel, etc.).

Braided metallic or synthetic wires, ropes, slings, etc. unless measures have been taken to

contain any broken filaments or fibers (sheathing, sealing with polymers, covering, etc.).

Woven materials especially cut or unfinished ends (metal braid, EMI shielding, lacing

cord, expando sleeving), unless measures have been taken to prevent fraying or

generation of particles (cut with a hot knife, seal with polymer, bag, etc.).

Materials with thin films known to erode or crack or flake when subjected to normal

handling (e.g., indium tin oxide [ITO] or other rigid or brittle semiconductor or ceramic

coating on flexible substrates, Teflon, multi-layered insulation [MLI], etc.).






400-FORM-0002 (4/16/2014)

Foams, highly textured materials.

Trapped debris in holes.

7.1.2.2 Molecular Generation

7.1.2.2.1 Material Selection

The Solar Array Panels materials shall have a total mass loss (TML) less than 1.00% and a

collected volatile condensable mass (CVCM) less than 0.10%, when measured in accordance

with ASTM E-595 unless a materials usage agreement has been generated and approved by the

NASA/GSFC COR.

7.1.2.2.2 NA

7.1.2.2.3 Assembly Outgassing

a. The Solar Array Panels outgassing shall not exceed 1E-11 g/cm2-s that is condensable on

a Quartz Crystal Monitor (QCM) that is operated at -20 degrees C when the panels are

held at the hot operational temperature (115°C). The measurement will be made in a

chamber that has been certified clean (back ground outgassing rate and free of silicones

and other high molecular weight contaminants) and has been modeled by the GSFC

Contamination Analyst to account for mass sinks (cold fingers, pumps, cold surfaces,

etc.) that could influence the source outgassing rate.

b. A cold finger and/or a scavenger plate shall be used in tests for component that will be

mounted externally unless approved otherwise by the NASA/GSFC COR.

7.2 ELECTROSTATIC CLEANLINESS

The following paragraphs provide requirements and guidelines for minimizing the magnitude

and variations in the radiated electric field from the external surfaces of the Solar Array Panels

when exposed to the space plasma. All external Observatory surfaces that are exposed to the

space plasma will be sufficiently conductive and be connected to spacecraft ground through low

impedance paths.

7.2.1 Conductive Surface Ground Path

All Solar Array Panel external conductive surfaces except for the solar cell coverglasses shall be

connected to the spacecraft interface with a resistance less than 5 ohms, either through the use of

ground wire(s) or through metal-to-metal mounting contact. All external conductive surfaces

should have a minimum of two (2) connections to the nearest grounded spacecraft surface.






400-FORM-0002 (4/16/2014)

7.2.2 Conductive Surface Resistivity

All Solar Array Panel external conductive surfaces shall have a resistivity of less than 10^9

ohms/square.

7.2.3 Exposed Harness Specific Requirements

Harnesses that are exposed to sunlight or the ambient plasma shall be bundle shielded from

source to destination.

7.3 MAGNETIC CLEANLINESS

Stray fields are due to uncompensated current loops or stray currents that result in permanent or

variable magnetic moments. The stray magnetic field associated with the maximum total

beginning of life solar panel short-circuit current output shall not exceed 1 gauss at a point 2

meters from any point on the panel. The dipole moment of each solar array panel shall not

exceed 3 A-m2.






400-FORM-0002 (4/16/2014)

8.0 DESIGN & CONSTRUCTION REQUIREMENTS

8.1 PARTS, MATERIALS & PROCESSES (PMP)

8.1.1 EEE Parts

See the requirements in the SOW, PACE-PWR-SOW-0025.

8.1.2 Materials

See the requirements in the SOW, PACE-PWR-SOW-0025.

8.1.2.1 Material Conductivity

All parts should be passivated and mounting surfaces on Solar Array Panels shall be conductive

as defined in Section 4.4.

8.1.2.2 Material Limitations for Debris Casualty Area

a. Aluminum parts shall be finished with iridite per MIL-DTL-5541, Class 3.

b. Titanium surfaces shall be finished per AMS 2488.

a. The contractor shall report the mass, dimensions, and use of materials shown in Error!

Reference source not found.. Materials should be assessed using the total quantity

within that assembly. Isolated pieces of material may weigh less than 50 grams; however,

if attached to other parts of the same material it will stay together until demise or until

reaching ground.

Table 8-1. Limited materials for debris casualty area

Material Quantity

Titanium Mass > 50g

Steel Mass > 50g

Inconel Mass > 50g

Invar Mass > 50g

Materials with melting

points above 1200° C

(1473 K)

Mass > 50g






400-FORM-0002 (4/16/2014)

8.2 ELECTRICAL

8.2.1 Test Sensors

With the exception of thermocouples used in thermal ambient and thermal vacuum cycling, and

in high temperature electrical measurements, test sensors shall be designed for flight. Unless

specified to be removed before flight, test sensors will not be removed prior to flight.

8.2.2 Interface Requirements

8.2.2.1 Connector Selection

8.2.2.1.1 Connector Specifications

a. Selected connector types shall meet the Connector and Contact Requirements defined in

Section C2 of EEE-INST-002.

b. Environmental seals shall not be used in connectors especially if made from silicone

without the explicit approval by the NASA/GSFC COR.

8.2.2.1.2 NA

8.2.2.1.3 Contact Derating

The current carrying capacity of the contacts shall be derated for continuous operation at the

required current levels in a vacuum, as defined in Section C2 of EEE-INST-002.

8.2.2.1.4 Redundant Contact Derating

When redundant contacts are used for a single power source, each contact shall meet the

required derating criteria.

8.2.2.2 Signal Segregation

a. Wherever possible, different classes of signals (power, digital, analog, etc.) shall be

separated by using separate connectors and separate harness bundles.

b. If separate connectors are not feasible, classes of signals within a common connector

shall be isolated from one another. Connector pin assignments should be such that

sensitive circuits are separated from potential interference sources.

8.2.2.3 Test and Flight Signal Isolation

Test signals and flight signals shall not be located in the same connector.






400-FORM-0002 (4/16/2014)

8.2.2.4 Test Interfaces

Component test signals that require access during observatory-level testing will be handled as

follows:

8.2.2.4.1 Facility-Induced Noise

All test signals should be protected or isolated from facility-induced noise.

8.2.2.4.2 Facility-Induced ESD GSE Malfunction

All test signals shall be protected or isolated from facility-induced ESD GSE malfunction.

8.2.2.4.3 Facility-Induced GSE Malfunction

All test signals shall be protected or isolated from facility-induced GSE malfunction.

8.2.3 Mitigation of Internal Charging

Internal charging refers to the physical effect where high energy electrons deposit charge in a

dielectric, if the charging rate is higher than the leakage rate eventually a point is reached where

the dielectric discharges to the nearby structure.

8.2.3.1 NA

8.2.3.2 Floating Conductors

Floating conductors, if present, shall have a bleed path of less than 10 MΩ to the component

structure. This requirement is not applicable to small floating conductors (1 inch2 or less) or short

(1 inch or less) unterminated traces or wires that are inside of the components.

8.2.3.3 Dielectric Structures

8.2.3.3.1 Bulk Resistivity

Dielectric structures shall have a bulk resistivity less than 1012 ohm-cm.

8.2.3.3.2 Charge Bleed-Off

All dielectric structures shall have a charge bleed path to the spacecraft interface, designed to

route the discharge into the spacecraft structure in a controlled fashion.

8.3 SAFETY

All flight components shall meet the applicable sections of NASA-STD-8719.24 ELV Safety

Requirements.






400-FORM-0002 (4/16/2014)

a. All Ground Support Equipment being used at NASA/GSFC shall meet NASA

requirements specified in Section 8.8.

b. All Ground Support Equipment being used at the Launch site shall meet the applicable

sections of NASA-STD-8719.24 ELV Safety Requirements.

8.4 NA

8.5 IDENTIFICATION AND MARKING

Each unit shall be permanently marked with the part number and a unique sequential serial

number in the area designated on the interface control drawing in a manner to be approved by the

NASA/GSFC COR.

All markings shall use alcohol proof ink, engraving, or laser etching.

8.6 WORKMANSHIP

8.6.1 Workmanship Standards

See the workmanship standards and processes outlined in the SOW.

8.6.2 Connector

8.6.2.1 GSE Cable Connectors

GSE cable connectors that mate with flight test connectors shall be flight-approved connectors.

8.6.2.2 Prevention of Connector Mismating

Connector mismating prevention requirements are identified in this section.

8.6.2.2.1 Connector Uniqueness

Physically adjacent connectors shall be of different sizes or of different sexes or uniquely keyed

to facilitate proper mating.

8.6.2.2.2 NA

8.6.2.2.3 NA

8.6.2.2.4 Accessibility

The Solar Array Panels Spacecraft interface connectors shall be spaced far enough apart to allow

the mate and demate operations to be performed without a special tool.






400-FORM-0002 (4/16/2014)

8.6.2.2.5 Connector Gender

The connector half that sources power to another Solar Array Panel shall be female (socketed) to

protect against inadvertent grounding prior to mating.

8.6.2.3 NA

8.6.2.4 Connector Identification

Each connector shall be labeled and clearly visible to facilitate proper mating.

8.6.2.5 Protection of Unused Test Connectors

Test connectors shall be capped with flight-approved RF and static control covers when not in

use.

8.6.2.6 Connector Savers

Connector savers shall be used during integration and test to minimize wear on connector

contacts and must meet the same requirements as the flight connectors.

8.7 RELIABILITY AND MISSION LIFETIME

8.7.1 Mission Life

The Solar Array Panels shall meet all performance specifications through two (2) years of

ground testing and three (3) years of operation in space, following a two (2) month

commissioning period.

8.8 GROUND HANDLING

8.8.1 Ground Support Equipment (GSE) Design

All electrical EGSE or support equipment shall be in compliance with the National Electric Code

(NFPA 70) or equivalent standard.

Do not lift items heavier than you can comfortably lift. The National Institute for Occupational

Safety and Health (NIOSH) lifting guidelines state that a person could lift 37.65 lb (17.1 kg)

under ideal conditions in front of the body not involving trunk twisting. Ideal conditions include:

smooth lifting (no jerking), the hands spread 30 inches (76 cm) or less, lifting posture

unrestricted, and object held close to the body. Additionally, there must be good couplings

(handles, shoes, floor surfaces) and a favorable environment. If all of these conditions are not

met, the maximum weight must be decreased (or alternative lifting methods should be

considered)






400-FORM-0002 (4/16/2014)

8.8.2 NA

8.8.3 NA

8.8.4 NA

8.8.5 NA

8.8.6 NA

8.8.7 GSE Cleanliness

All Ground and Test support equipment shall be compatible with the flight component and the

environment where the flight component or test component will reside (cleanroom, thermal

vacuum chamber, vibration cell, etc.)

8.8.8 GSE Bakeout

Thermal Vacuum GSE shall be baked out and the outgassing rate certified prior to the test.

8.8.9 Test Harness

Test harnesses that will be used in vacuum during ground operations shall be vacuum

compatible.






400-FORM-0002 (4/16/2014)

9.0 MECHANICAL DESIGN REQUIREMENTS

For additional guidance on the design and analysis of threaded fastening systems in NASA

spaceflight hardware, consult NASA-STD-5020, Requirements for Threaded Fastening Systems

In Spaceflight Hardware.






400-FORM-0002 (4/16/2014)

10.0 LOGISTICS

10.1 NA

10.2 GROUND SUPPORT EQUIPMENT

a. Any GSE that will be used in a Thermal Vacuum test (harnesses, fixtures, stimulators,

etc.) shall be fabricated of vacuum compatible materials that meet the requirements of

Section 7.1.

b. Any GSE that will be used in a Thermal Vacuum test shall be capable of being baked out

at 115 degrees C. Special care will be taken to preserve the surface cleanliness of

thermal vacuum GSE items (especially harnesses) during integration and test activities in

non-cleanroom areas. Minimize contact with contaminating surfaces. Bagging is

recommended whenever possible.

c. GSE that will be used in contractor or NASA/GSFC cleanrooms shall be cleanroom

compatible.

d. Materials shall not outgas or generate particles that violate the requirements of Section

7.1.

e. Aluminum, for example, shall have protective finishes to prevent oxidation (anodize,

chromate conversion (irridite), or cleanroom compatible paint).

f. Any GSE that will be in contact or close proximity (up to approx. 6 feet) of flight

hardware shall meet the same surface cleanliness requirements as the flight hardware

unless the flight hardware can be thoroughly cleaned to remove any transferred

contaminants.

g. Materials that contact the flight hardware shall meet the vacuum compatibility

requirements in Section 7.1.

10.3 TRANSPORTATION EQUIPMENT

a. Transportation equipment shall assure that the solar array panels meet the requirements

of this specification after shipment.

b. Materials that are exposed to the same air that surrounds the flight panels, shall not

outgas hydrocarbons or other contaminants that can contaminate the panels. Bagging

materials for such large assemblies seldom provide 100 percent vapor barriers.

c. Unless the vendor can demonstrate 100 percent vapor barrier construction, the vendor

shall assure that all materials that are interior to a shipping container will not outgassing

during transport.

d. The containers shall maintain the cleanliness of the panels during long term storage – up

to 2 years.






400-FORM-0002 (4/16/2014)

e. Materials of construction as well as equipment in the transportation container shall not

degrade the surface or outgassing cleanliness of the solar panels.

Materials of concern include elastomeric seals, paints, surface finishes, substrates, elastomeric

dampeners that can outgas, lubricants, wire spring dampeners with lubricants to prevent

corrosion), monitoring equipment, surface finishes, etc. Where possible materials should meet

flight outgassing requirements. Containers should be compatible with a cleanroom and capable

of being cleaned with Isopropyl alcohol. Other materials may be used but tests should be run on

them to determine their outgassing potential in air over at the high end of the shipping

temperature range. Silicones should be avoided. Wood should not be used with prior approval of

the NASA/GSFC COR and review of contamination risk mitigation strategies.






400-FORM-0002 (4/16/2014)

11.0 VERIFICATION REQUIREMENTS

The contractor shall conduct a verification program that demonstrates the component design is

qualified. Per the SOW (PACE-PWR-SOW-0025), the contractor will provide a verification

matrix defining the method of verification for each specific requirement of this document.

11.1 VERIFICATION METHODS

Verification methods include inspection, analysis, as well as environmental, functional, and

performance testing, or a combination of these techniques.

11.1.1 Inspection

Verification by inspection includes (but is not limited to) visual inspection, simple physical

manipulation, gauging, measurement, and documentation examination.

11.1.2 Analysis

Verification by analysis will be used to show design margins. Also, when the particular tests

required for verification are impractical, risky, unacceptably long, or prohibitively expensive,

analysis may be used instead of testing, as noted in the verification matrices.

Analysis, including simulations where applicable, will also be used to guarantee that the Solar

Array Panels and their components will perform as expected under worst-case conditions.

11.1.3 Test

Verification by test includes, but is not limited to, the evaluation of performance by use of

special equipment or instrumentation, simulation techniques, and the application of established

principles and procedures to determine compliance with requirements.

11.2 INSPECTION REQUIREMENTS

Verification by inspection shall be by one of these three methods: 1) visual inspection of the

physical component; 2) a physical measurement of a property of the component, or; 3) a

documentation search demonstrating hardware of an identical design has demonstrated

fulfillment of a requirement.

11.2.1 Visual Inspection

Visual inspection of the physical component shall be performed by a customer appointed

qualified inspector to certify that the component has the properties/configuration specified in the

requirement.






400-FORM-0002 (4/16/2014)

11.2.2 Physical Measurement

Physical measurement of component property (mass, dimensions, etc.) shall be performed by a

customer appointed qualified inspector to demonstrate the component meets specific

requirement.

11.2.3 Documentation Search

Verification of requirements based on similarity shall include supporting rationale and

documentation and be approved by the NASA/GSFC COR.

11.3 ANALYSIS REQUIREMENTS

Verification of performance or function through detailed analysis, using all applicable tools and

techniques, is acceptable with NASA/GSFC COR approval. Detailed descriptions of the

minimum required analyses, as well as analysis requirements, are provided in the SOW.

11.4 TEST REQUIREMENTS

a. This section provides general test requirements on how testing is to be performed in the

process of verifying that the deliverable item meets its requirements. Performance

parameter measurements shall be taken to establish a baseline that can be used to assure

that there are no data trends established in successive tests that indicate a degradation of

performance trend within specification limits that could result in unacceptable

performance in flight.

b. Any requirement that exceeds previous qualification test data shall be presented to the

NASA/GSFC COR as part of the verification planning process described in the SOW, for

evaluation and a possible delta qualification test.

11.4.1 Definitions

The hardware definitions are reproduced here from Section 1.8 of GEVS (GSFC-STD-7000A).

Acceptance Tests: The verification process that demonstrates that hardware is acceptable for

flight. It also serves as a quality control screen to detect deficiencies and, normally, to provide

the basis for delivery of an item under terms of a contract.

Design Qualification Tests: Tests intended to demonstrate that the test item will function within

performance specifications under simulated conditions more severe than those expected from

ground handling, launch, and orbital operations. Their purpose is to uncover deficiencies in

design and method of manufacture. They are not intended to exceed design safety margins or to

introduce unrealistic modes of failure. The design qualification tests may be to either “prototype”

or “protoflight” test levels.






400-FORM-0002 (4/16/2014)

Protoflight Hardware: “Flight hardware of a new design; it is subject to a qualification test

program that combines elements of prototype and flight acceptance verification; that is, the

application of design qualification test levels and flight acceptance test durations.” The purpose

of the test on this hardware is to prove that a new design meets one or more of its design

requirements. Protoflight testing is performed at maximum expected flight levels plus a margin.

Test durations are typically the same as for acceptance tests.

Follow-On (Acceptance) Hardware: “Flight hardware built in accordance with a design that

has been qualified either as prototype or as protoflight hardware; follow-on hardware is subject

to a flight acceptance test program.” The purpose of the test on this hardware is to prove that a

particular flight unit has been manufactured properly. The design has already been proven

during a qualification or protoflight test program. Acceptance testing is performed at maximum

expected flight levels.

11.4.2 NA

11.4.3 Test Tolerances

Tolerances for the various mechanical test parameters are given in Error! Reference source not

found..

Table 11-1. Test Tolerances

Test Test Parameter Tolerance

Temperature 2 C

Humidity 5% RH

Mass Properties Weight: 25g

Center of Gravity: 6 mm

Moments of Inertia 10 %

Products of Inertia 10 %

Pressure >1.3 x 104 Pa (> 100 mm Hg): 5%

1.3 x 104 to 1.3 x 102 Pa (100 mm Hg to 1 mm Hg): 10%

1.3 x 102 to 1.3 x 101 Pa (1 mm Hg to 1 micron): 25%

< 1.3 x 101 Pa (< 1 micron): 80%

11.4.4 Test Restrictions

11.4.4.1 Failure During Tests

a. When a failure (non-conformance or trend indicating that an out-of-spec condition will

result) occurs, determination will be made as to the feasibility and value of continuing the

test to its specified conclusion. The test shall be stopped if equipment fails during testing

in cases where this failure will result in damage to the equipment.

b. Otherwise, the test shall be completed to obtain as much information as possible. If

corrective action is taken, the test will be repeated to the extent necessary to demonstrate






400-FORM-0002 (4/16/2014)

that the test item’s performance is satisfactory. If corrective action taken as a result of

failure affects the validity of previously completed tests (e.g., redesign of a component),

prior tests will be repeated.

If during a test sequence, a test item is operated in excess of design life and wears out or

becomes unsuitable for further testing from causes other than deficiencies, a spare will be

substituted, and previously completed tests will be repeated to the extent necessary.

c. No replacement, adjustment, maintenance, or repairs are authorized during testing. This

requirement does not prevent the replacement or adjustment of equipment that has

exceeded its design operating life during tests, provided that after such replacement, the

equipment is tested as necessary to assure its proper operation. A complete record of any

exceptions taken to this requirement shall be included in the test report.

11.4.4.2 Modification of Hardware

Once the formal acceptance test has started, cleaning, adjustment, or modification of test

hardware shall not be permitted.

11.4.4.3 External Adjustment

The Solar Array Panels shall be designed so that no external adjustments are required after start

of acceptance or qualification testing.

11.4.4.4 Re-Test Requirements

If any event, including test failure, requires that a Solar Array Panel be disassembled and

reassembled, then all tests performed prior to the event shall be considered for repeat. If the unit

has multiple copies of the same build, then all units must be examined to determine if the

problem is common. If all copies require disassembly for repair, then each must receive the

same test sequence.

11.5 REQUIRED TESTS

The orders of tests on the insulated substrates, qualification coupons, and flight panels are

specified in the applicable sections below.

The Intermediate Sequence tests consist of the following functional tests and shall be performed

as specified during qualification coupon and flight panel testing:

Visual Inspection of Add-on Hardware (Everything the contractor has added to the

substrate) per Section 11.5.22 of this specification

Visual Inspection of Insulated Substrate

Adhesion Check

Test Condition Power Verification per Section 11.5.19 of this specification






400-FORM-0002 (4/16/2014)

Substrate Insulation Resistance

Wiring Resistance

Resistance between each solar cell coverglass and the solar cell interconnect for the

applicable cell per Section 11.5.24 of this specification

Bypass Diode Functionality per Section 11.5.20 of this specification

11.5.1 NA

11.5.2 NA

11.5.3 NA

11.5.4 NA

11.5.5 NA

11.5.6 NA

11.5.7 NA

11.5.8 NA

11.5.9 NA

11.5.10 Thermal Vacuum Bake-out

The contractor shall bake out the Solar Array Panels prior to thermal vacuum cycling until the

outgassing requirements of Section 7.1.2.2 are met. The rate measured at the TQCM shall be

adjusted to account for chamber geometry, presence of cold sinks, chamber pumping speed, view

factors of the TQCM, and any other factors necessary to assure an accurate measurement of the

total outgassing per unit time per Kg mass of the unit under test. The chamber and GSE

containing organic materials shall be baked out and certified prior to installing the flight

hardware into the chamber per paragraph 11.5.10.3 below. NASA/GSFC shall perform the

calculations necessary to convert the outgassing rate to the test specific TQCM rate unless the

contractor elects to perform its own calculations and submits their assumptions, calculations and

results to NASA/GSFC for review. The cleanliness of the panels shall be preserved post bakeout

as described in paragraph 11.5.10.4 below.

The outgassing rate shall be considered acceptable when the flight hardware temperature does

not vary by more than 2 degrees C and the rate meets the requirements of Section 7.1.2.2.3 for 4

continuous hours.






400-FORM-0002 (4/16/2014)

11.5.10.1 Test Conditions for Thermal Vacuum Bakeout

The contractor shall bake the panels at a temperature of 135+0/-5 °C in a vacuum of 1 x 10-5

Torr or less for at least 96 hours.

The contractor shall perform the out gassing verification at the end of the thermal vacuum

bakeout when flight hardware temperature is lowered to 97C, Panel Performance in Thermal

Vacuum Environment with no break in chamber vacuum.

The contractor shall maintain the TQCM at -20 C throughout the test to measure total volatile

out gassed condensables without the influence of water vapor. The TQCM must have a

representative view of the hardware.

11.5.10.2 Calculation of the Configuration Adjusted TQCM Rate

The contractor shall collect and deliver to the NASA/GSFC COR the information necessary to

model the test chamber. NASA/GSFC shall calculate the acceptance QCM rate that indicates

that the hardware has met the outgassing rate criteria. This rate shall be provided to the

contractor within 24 hours of the test starting providing that (1) any changes in the configuration

are minor and (2) they are provided to the GSFC contamination control analyst within 12 hours

of the test starting. When these conditions are not met the rates may take longer to supply. The

following information shall be provided to the NASA/GSFC COR at least one month prior to the

test and updated as necessary to account for unforeseen last minute changes in configuration:

Drawing or sketch showing chamber configuration and geometry

Chamber dimensions

Shrouds dimensions,

Temperature and physical dimensions of all major elements that will not be isothermal

(within 5 degrees) of the flight hardware (shrouds, heater plates, scavenger plates, LN2

lines, windows, etc.),

QCM location,

Cold finger, QCM heat sink, scavenger plate locations, sizes, control temperatures during

the test and general set-up

Pump location and dimensions

For the S-112 Qualification Coupon panel thermal vacuum setup, and with the QCM

temperatures for each bakeout in accordance with Section 11.5.10.1 of this specification, the

QCM goal for each bakeout shall be <635Hz/hr above background.

11.5.10.3 Pre-Test Chamber Vacuum Bakeout and Chamber Certification

Prior to the test with the flight hardware, the chamber shall be baked out in vacuum with any

GSE containing organic materials to establish the chamber background rate.






400-FORM-0002 (4/16/2014)

The chamber shall be baked until the delta delta QCM rate is less than 5 Hz per hour with the

chamber at 130 degrees C. Once this is achieved, the chamber background rate shall be

measured for both certification temperatures specified in Section 11.5.10.1 with the chamber

conditions as specified in 11.5.10.1. Data shall be taken for a minimum of 4 hours at each

temperature. The chamber background rate shall be low and steady enough so as to not impede

measurement of the array outgassing rate. The PACE contamination analyst may relax these

requirements if later analysis proves that the measurement may be made.

Once this has been achieved, the contractor shall proceed to measure the background of the

chamber. Use of lower temperatures will lengthen the time required to complete the bakeout.

The following data shall be recorded during the test: QCM readings (once every 2 minutes

minimum), hardware temperature, chamber and shroud temperature, scavenger plate and other

cryosurface temperature, QCM temperature, and pressure.

If requested by the NASA/GSFC COR and prior to loading the chamber, the contractor shall

allow a minimum of two hours for a NASA/GSFC representative to inspect the chamber, its

equipment, and its configuration.

11.5.10.4 Post-Bakeout Cleanliness

The cleanliness of the flight hardware shall be maintained post bakeout. It is recommended that

the arrays be kept in sealed (i.e., low permeability) container or bagging material of the same

cleanliness level as the panels when the panels are in storage or during periods when a

contamination hazard is present in the area where the panels are located (painting, roof sealing,

curing of large quantities of polymers (adhesives/paints)).

11.5.10.4.1 Witness Foils

A NASA/GSFC provided aluminum witness foil shall be kept in the same environment (air

stream) as each panel at all times. The number of foils may be reduced to one if multiple panels

are kept in the same area or container. Each foil shall be replaced with a new foil and sent back

to the NASA/GSFC COR for analysis monthly. NASA/GSFC will send instructions on how to

handle and ship the foils with the foils. The collected non-volatile residue shall remain less than

0.2 mg/sq-ft per month. If this number is exceeded, the NASA/GSFC COR and the contractor

shall jointly determine a course of action. Normally deposited surface contaminants – for

example finger oils or phalates – should be able to be removed via a solvent cleaning of the

arrays. If there should be a contaminant that cannot be removed, then NASA/GSFC will conduct

further analysis to determine the volatility of the contaminant to determine if the bakeout shall

be required to be repeated.

Should a panel or group of panels need to be placed into a vacuum chamber after thermal

vacuum bakeout has been completed, the chamber shall meet the requirements of paragraphs

11.5.10.1. A QCM or witness foil that is kept at the same temperature as the flight hardware






400-FORM-0002 (4/16/2014)

shall be installed into the chamber to verify that the cleanliness of the hardware has not been

compromised by the exposure to the chamber or test conditions.

NASA/GSFC shall supply a sufficient quantity of 18 inch by 12 inch aluminum witness foils to

the contractor to monitor the cleanliness of the air to which the panels are exposed. A handling

procedure shall be supplied with the foils. The foils shall be used to monitor cleanliness post

bake-out. If there is a concern for the facility cleanliness, they may also be used by the

contractor to monitor facility cleanliness. The foils shall be replaced monthly or if there is a

contamination event and sent back to NASA/GSFC for analysis.

11.5.10.5 Surface Cleanliness Required for Performance

The contractor shall supply the assumed surface cleanliness level used in calculations of EOL

solar array performance. It should be expressed in percent obscuration ratio per IEST-STD-

CC1246D.

11.5.11 Thermal Vacuum Test

11.5.11.1 Thermal Vacuum Test Parameters

The contractor shall thermal cycle each flight panel in a vacuum of 1 x 10-5 Torr or less. The

contractor shall perform the thermal vacuum cycling in a chamber with a shroud at

approximately liquid nitrogen temperatures.

The contractor shall cycle the flight panels from -95°C to +120°C for 9 cycles,

and from -100°C to +135°C for 3 cycles.

The contractor shall fix at least 6 calibrated temperature sensors over each flight panel and use

the panels’ platinum resistor thermometers to monitor temperature.

The contractor shall cycle to the temperature extremes based on the average reading of the

temperature sensors.

The temperature gradients across the panels during temperature transitions shall be limited to

±25C.

The period for one cycle shall be greater than 1.5 hour, excluding the dwell.

The dwell at the temperature extremes shall be greater than 1 hour.

The rate of temperature change between cold and hot limits shall not exceed 30°C per minute.






400-FORM-0002 (4/16/2014)

The contractor shall send current through each panel circuit and through the platinum resistor

thermometers using the panel connectors during these tests.

The contractor shall continuously monitor the currents with a channel dedicated to that circuit,

heater, or resistor.

The contractor shall pass current through the cells during the first three cycles, pass current

through the bypass diodes during the second three cycles, pass current through the cells during

the seventh through ninth cycle, and pass current through the bypass diodes during the last three

cycles.

The reverse current through the cells shall be conducted through a flight quality test connector,

which the contractor shall mount to the flight panels that make connection to the anode side of

each string’s blocking diode and the service connector contacts that make connection to each

string’s return.

The Contractor shall not employ mechanical connections, other than flight quality crimps, during

the thermal vacuum exposure.

During the pump-down, the contractor shall monitor power line voltages, to demonstrate the

absence of corona discharge and multipaction.

Transitions from cold to hot conditions increase contamination hazards because material that has

accreted on the chamber walls may evaporate and deposit on the relatively cool solar panels.

Transitions will be conducted at rates sufficiently slow to prevent this from occurring. The

thermal vacuum test shall start and end with the flight hardware at 40 degrees C minimum to

minimize this risk and the thermal driving surfaces (shroud/heater) shall not be more than 40

degrees warmer than the flight hardware unless a QCM demonstrates that there is no risk.

Pass/Fail Criteria: Any discontinuity showing in recorded data shall fail the verification.

Any defects or reduction of power output outside the limits of the requirements of this

specification shall fail the verification.

Any mechanical damage to the panel, including the substrate and contractor installed

components, that put into question their ability to perform adequately during launch or flight

shall fail the verification.






400-FORM-0002 (4/16/2014)

11.5.12 NA

11.5.13 NA

11.5.14 NA

11.5.15 NA

11.5.16 Solar Cell and Bypass Diode Qualification Tests

The Contractor shall use solar cells for PACE that have been qualified to either AIAA-S-111-

2005 or AIAA-S-111A-2014 and shall demonstrate previous flight heritage for the cells in a

space environment that envelopes that of the PACE mission.

11.5.17 Solar Array Panel Qualification Tests

a. The Contractor shall qualify the PACE solar array panels using AIAA S-112A-2013,

Qualification and Quality Requirements for Electrical Components on Space Solar Panels as

tailored by this Specification.

b. The contractor shall use one or more coupons for all of the following tests. Section

numbers below refer to S-112A unless otherwise indicated.

Life Cycles Coupon Test with Humidity Exposure, Section 7.1

Solar Absorptance Characterization, Section 8.4

Bypass Diode Characterization, Section 8.5

The following tests are waived:

Humidity Exposure, Section 7.1.3.1

Acoustic Exposure, Section 7.2.4.

ESD Test, Section 7.3

UV Radiation Effects Characterization, Section 8.1

Angle of Incidence Characterization, Section 8.2

Atomic Oxygen Characterization, Section 8.6

c. The following tests or requirements are modified:






400-FORM-0002 (4/16/2014)

Sample Count, Section 6.1: The minimum sample count for CICs on the life cycle coupons shall

be the maximum number that will fit on the front sides of the GFE qualification coupon

substrates.

Life-Cycle Coupons/Subcoupons, Section 7.1.2: add the requirement that three cells be broken

and repaired prior to the start of environmental exposure. The contractor shall break one cell at a

positive string termination, another cell at a negative string termination, and another in the center

of a row.

Combined Effects Exposure, Section 7.1.3.4, Subsection 4 of S-112A: Dark forward bias may be

substituted for illumination during the heated portion of the cycles.

The VCM test in Section 7.2 shall be performed on a Life Cycle Coupon.

11.5.17.1 Solar Array Panel Life Cycle Coupon Tests

Section 7.1 of S-112A requires life cycle coupons. The coupons used in the cycling test must

include at least two samples of all of the parts to be used on the solar array panels.

a. In S-112A, the bakeout temperature shall be 135°C+0/-3°C.

b. Except for the 135°C hot survival temperature and the -100° C cold survival temperature,

the temperature extremes for the life cycle coupon test shall be 10°C beyond the

extremes and for 1.5X the number of cycles in Error! Reference source not found. of

this specification. The cycles to the hottest temperature shall not exceed 135°C.

c. The order of tests on the life cycle coupon shall be as follows:

Intermediate Sequence as specified in Section 11.5:

Damage and Repair

Panel Mass Measurement

Intermediate Sequence

Thermal Vacuum Bakeout/VCM Test per Sections 11.5.17and 11.5.17.1 of this specification

Thermal Vacuum Cycles 1 through 8


Rapid Temperature Cycles 1 through 10


Rapid Temperature Cycles 11 through 100


Rapid Temperature Cycles 101 through 5,000


Rapid Temperature Cycles 5,001 through 10,000


Rapid Temperature Cycles 10,001 through 24,074 (1.5X number of mission cycles)






400-FORM-0002 (4/16/2014)


d. Pass/Fail Criteria: The coupons shall fail the thermal cycle test if the test condition power

degrades by more than 3% from the baseline measurement prior to the start of cycling.

11.5.18 Flight Solar Array Panel Tests

The order of tests on the flight panels shall be as follows:

Intermediate Sequence as defined in Section 11.5 of this specification

Thermal Vacuum Cycles per Section 11.5.11 of this specification.

Thermal Vacuum Bakeout per Section 11.5.10 of this specification.


Panel Mass Properties Measurement per PACE-PWR-SOW-0025, PACE Solar Array Panels

Statement of Work.

11.5.19 Test Condition Power Verification

a. The contractor shall use a solar simulator calibrated with a set of primary or secondary

standard solar cells to determine the current-voltage (I-V) characteristics of each panel

through the panel terminals at 23º ± 5ºC.

b. The contractor shall extrapolate the measured data to obtain the I-V curve for a panel

operating at 28ºC and AM0.

Type: Measurement.

Level: Panel.

Schedule: Per Sections 11.5.17, 11.5.18, and 11.5.19

c. Pass/Fail Criteria: The verification shall fail if the power at the specified load voltage of

a panel does not meet the requirements of Section 4.1.1.

11.5.20 Bypass Diode Functionality Verification

a. The contractor shall pass a minimum of 120% of the Isc of each string on each panel

through the bypass diode circuits. This test shall be conducted at 23º ± 5ºC with the

results extrapolated to 28ºC.

Type: Test.

Level: Panel.






400-FORM-0002 (4/16/2014)

Schedule: Per Sections 11.5.17 and 11.5.18.

b. Pass/Fail Criteria: The voltage dropped by the bypass diodes shall not vary more than

±3% from the first panel level measurement to the last at an extrapolated temperature of

28ºC.

11.5.21 Substrate Insulation Resistance Verification

a. The contractor shall make connection to the cell circuits through each panel’s terminals

and shall tie all positive and negative power leads together.

b. The contractor shall then measure the resistance between these tied together leads and the

panel’s substrate.

c. The contractor shall make this measurement at 500 volts direct current (Vdc) with the

current limited to 20 microamperes or less with the positive test voltage on the cell

circuits

d. Type: Measurement. The insulation resistance values shall be recorded.

Level: Panel.

Schedule: Per Sections 11.5.17 and 11.5.18

e. Pass/Fail Criteria: The panel shall fail verification if its substrate insulation resistance is

less than that required by Section 4.1.11.1.

11.5.22 Solar Cell Mechanical Verification

a. The contractor shall visually inspect each solar cell on the qualification and flight panels

for compliance with Section 4.1.5.1.

b. The contractor shall inspect with the unaided eye and under a minimum of seven power

magnification.

c. The contractor shall perform optional inspections, which it determines are

advisable, at its discretion.

Type: Inspection.

Level: Panel.







400-FORM-0002 (4/16/2014)

d. Pass/Fail Criteria: Each solar cell shall meet the requirements of Section 4.1.5.1 or the

contractor shall remove the cell from the panel and replace it.

e. If more than 4% of the cells on the cell qualification coupon shall crack as a result of test,

the panel shall fail qualification.

11.5.23 Cover Orientation Verification

a. The contractor shall visually inspect each solar cell cover for compliance with Section

4.1.6.2.

Type: Inspection.

Level: Panel.


b. Pass/Fail Criteria: No more than 0.1% of a panel’s covers shall fail this test.

11.5.24 Cover Grounding Verification

a. The contractor shall verify the electrical connection between the geometric center of each

cover within ± .5 cm, and its connection to ground or the array circuitry at 23º ± 5ºC

b. Pass/Fail Criteria: The electrical connection shall fail the test if the resistance exceeds 1 x

109 ohms.

11.5.25 Flight Connector Type Verification

The contractor shall propose the type, level, schedule and pass/fail criteria for verifying the

connector type, connector mounting adequacy, and wiring.

11.5.26 Platinum Resistance Thermometer Type Verification

The contractor shall propose the type level, schedule and pass/fail criteria for verifying the

resistor type.

11.5.27 Platinum Resistance Thermometer Performance Verification

Type: Measurement.

Level: Component.







400-FORM-0002 (4/16/2014)

a. Pass/Fail Criteria: The output from each platinum resistor thermometer shall meet the

requirements of 4.1.10 or this requirement is failed.

b. During the thermal vacuum test or thermal cycling test, the platinum resistor thermometer

shall additionally be shown to operate without discontinuity in resistance or this

requirement is failed.

11.5.28 Parts and Assembly Layout Verification

The contractor shall propose methods and schedules for verifying the parts and assembly layout.

11.5.29 Mission Life Verification

a. The contractor shall certify that it has conducted a test program to demonstrate that the

solar panels have the required mission life.

b. The contractor shall certify that it has not knowingly limited the mission life of the

panels to less than the required mission life.

11.5.30 Shelf Life Verification

The contractor shall certify that the solar panels will have a shelf life greater than 5 years when

packaged to its specifications.

11.5.31 Substrate Ground Verification

The contractor shall verify the substrate ground with an ohm meter.

11.5.32 Cleanliness Verification

a. Immediately prior to shipment, the contractor shall inspect the panels to the requirements

of JSC-SN-C-005 to the highly sensitive level using both white and black light

inspections.

b. Immediately prior to shipment, the contractor shall verify the cleanliness of the array by

blacklight inspection per IEST-STD-CC1246D.

Type: Inspection.

Level: Panel.






400-FORM-0002 (4/16/2014)

Schedule: Prior to delivery to NASA/GSFC.

c. Pass/Fail Criteria: The inspection shall meet the visibly clean highly sensitive

requirements of JSC-SN-C-005.

d. If not, the contractor shall clean the “dirty” areas until the requirement is met.






400-FORM-0002 (4/16/2014)

APPENDIX A ABBREVIATIONS AND ACRONYMS

Abbreviation/

Acronym Definition

Al Aluminum

AM0 Air Mass Zero

AR/ITO Antireflective/Indium Tin Oxide Coating

CCB Configuration Control Board

CCR Configuration Change Request

CG Center of Gravity

CM Configuration Management

CMO Configuration Management Office

COR Contracting Officer's Representative

CSS Coarse Sun Sensor

CVCM Collected Volatile Condensable Mass

DC Direct Current

DDD Displacement Damage Dose

DILS Deliverable Items List and Schedule

EEE Electrical, Electronic, and Electromechanical

EOL End of Life

ESD Electrostatic Discharge

FS Factor of Safety

GFE Government Furnished Equipment

GSE Ground Support Equipment

GSFC Goddard Space Flight Center

ICD Interface Control Document

I&T Integration and Test

ITO Indium Tin Oxide

KOZ Keep-Out Zone

MLI Multi-Layer Insulation

Mohms Megaohms

MS Margin of Safety

NASA National Aeronautics and Space Administration

NIEL Non-Ionizing Energy Loss

PACE Plankton, Aerosol, Cloud, ocean Ecosystem

QCM Quartz Crystal Monitor

S/C Spacecraft

SOW Statement of Work

TBD To Be Defined

TBR To Be Reviewed

TID Total Ionizing Dose






400-FORM-0002 (4/16/2014)

Abbreviation/

Acronym Definition

TML Total Mass Loss

UUT Unit Under Test

VDA Vapor Deposited Aluminum

VDC Voltage, Direct Current

VCM Volatile, Condensible Material


PACE Solar Array Panels Specification - Amazon AWS

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