ATTACHMENT_B-_SPEC.pdf

Effective Date: December 18, 2018

Expiration Date: December 18, 2023

Check https://ipdtdms.gsfc.nasa.gov to verify that this is the correct version prior to use

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 ii

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

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/ iii

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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.

iv

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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 v

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Table of TBDs/TBRs/TBSs

Action Item

No.

Location Summary Individual/

Organization

Actionee vi

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Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 Hardware Description

3.1.1 Flight Solar Array Panels

3.1.2 Qualification Coupons

3.1.3 Insulated Substrates

4.0 Functional/Performance Requirements

4.1 Solar Array Panels Flight Unit Functional/Performance Requirements

4.1.1 Test Condition Power

4.1.1.1 Flight Panels

4.1.1.2 Qualification Coupons

4.1.2 Power at Highest Predicted Operating Temperature

4.1.3 Power at Highest Predicted Temperature

4.1.4 End-of-Life Power

4.1.5 Solar Cell and Bypass Diode

4.1.5.1 Solar Cell Mechanical

4.1.5.2 Solar Cell Layout

4.1.5.3 Solar Cell Power

4.1.5.4 Limit to Solar Cell Shadowing

4.1.6 Solar Cell Cover

4.1.6.1 Cover Material and Thickness

4.1.6.2 Cover Orientation

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

4.1.7 Blocking Diodes and Terminal Boards

4.1.8 Wire and Wire Layout

4.1.8.1 Coarse Sun Sensor (CSS) Wiring

4.1.8.2 Hinge Damper Heater and Thermistor Wiring

4.1.8.3 Monitor Solar Cell

4.1.9 Connector Wiring and Connector Type

4.1.10 Platinum Resistor Thermometers (PRTs)

4.1.11 Insulated Substrate

4.1.11.1 Substrate Insulation Resistance

4.1.11.2 Substrate Grounding

4.1.12 Panel Performance in Thermal Vacuum Environment

4.1.13 Panel Performance in Depressurization Environment

4.1.14 Allowable Degradation Due to Charged Particle Radiation

4.1.15 Allowable Degradation Due to Humidity

4.2 Resource Allocations

4.2.1 Mass Allocation

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4.3 Power

4.3.1 Solar Array Panels Power Wire Redundancy

4.4 Electrical Grounding

4.4.1 Primary Power DC Isolation

4.4.2 Mechanical Contact Resistance

4.4.3 Connector DC Resistance

4.5 Signal And Data Interfaces

4.5.1 NA

4.5.2 Passive Analog Telemetry

4.5.2.1 Platinum Resistor Thermometer (PRT)

4.5.2.2 PRT Performance

5.0 Physical Requirements

5.1 Interface Documentation

5.2 Mass Properties

5.2.1 NA

5.3 Physical Envelope

6.0 Environmental Requirements

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

6.8.2 NA

6.9 NA

6.10 Ground Environments

6.11 Thermal Requirements

6.11.1 Flight Interface Design Temperature Limits

6.12 Charged Particle Radiation Requirements

6.12.1 Definitions

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts

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

6.13 Atomic Oxygen Fluence

6.13.1 Atomic Oxygen Analysis

viii

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6.13.2 Atomic Oxygen Testing

7.0 Cleanliness

7.1 Surface Contamination

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

7.1.1.2 Molecular Contamination

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

7.1.2.2 Molecular Generation

7.2 Electrostatic Cleanliness

7.2.1 Conductive Surface Ground Path

7.2.2 Conductive Surface Resistivity

7.2.3 Exposed Harness Specific Requirements

7.3 Magnetic Cleanliness

8.0 Design & Construction Requirements

8.1 Parts, Materials & Processes (PMP)

8.1.1 EEE Parts

8.1.2 Materials

8.1.2.1 Material Conductivity

8.1.2.2 Material Limitations for Debris Casualty Area

8.2 Electrical

8.2.1 Test Sensors

8.2.2 Interface Requirements

8.2.2.1 Connector Selection

8.2.2.2 Signal Segregation

8.2.2.3 Test and Flight Signal Isolation

8.2.2.4 Test Interfaces

8.2.3 Mitigation of Internal Charging

8.2.3.1 NA

8.2.3.2 Floating Conductors

8.2.3.3 Dielectric Structures

8.3 Safety

8.4 NA

8.5 Identification and Marking

8.6 Workmanship

8.6.1 Workmanship Standards

8.6.2 Connector

8.6.2.1 GSE Cable Connectors

8.6.2.2 Prevention of Connector Mismating

8.6.2.3 NA

8.6.2.4 Connector Identification

8.6.2.5 Protection of Unused Test Connectors

8.6.2.6 Connector Savers

8.7 Reliability and Mission Lifetime

ix

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8.7.1 Mission Life

8.8 Ground Handling

8.8.1 Ground Support Equipment (GSE) Design

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

8.8.8 GSE Bakeout

8.8.9 Test Harness

9.0 Mechanical Design Requirements

10.0 Logistics

10.1 NA

10.2 Ground Support Equipment

10.3 Transportation Equipment

11.0 Verification Requirements

11.1 Verification Methods

11.1.1 Inspection

11.1.2 Analysis

11.1.3 Test

11.2 Inspection Requirements

11.2.1 Visual Inspection

11.2.2 Physical Measurement

11.2.3 Documentation Search

11.3 Analysis Requirements

11.4 Test Requirements

11.4.1 Definitions

11.4.2 NA

11.4.3 Test Tolerances

11.4.4 Test Restrictions

11.4.4.1 Failure During Tests

11.4.4.2 Modification of Hardware

11.4.4.3 External Adjustment

11.4.4.4 Re-Test Requirements

11.5 Required Tests

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

x

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11.5.9 NA

11.5.10 Thermal Vacuum Bake-out

11.5.11 Thermal Vacuum Test

11.5.11.1 Thermal Vacuum Test Parameters

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

11.5.17 Solar Array Panel Qualification Tests

11.5.17.1 Solar Array Panel Life Cycle Coupon Tests

11.5.18 Flight Solar Array Panel Tests

11.5.19 Test Condition Power Verification

11.5.20 Bypass Diode Functionality Verification

11.5.21 Substrate Insulation Resistance Verification

11.5.22 Solar Cell Mechanical Verification

11.5.23 Cover Orientation Verification

11.5.24 Cover Grounding Verification

11.5.25 Flight Connector Type Verification

11.5.26 Platinum Resistance Thermometer Type Verification

11.5.27 Platinum Resistance Thermometer Performance Verification

11.5.28 Parts and Assembly Layout Verification

11.5.29 Mission Life Verification

11.5.30 Shelf Life Verification

11.5.31 Substrate Ground Verification

11.5.32 Cleanliness Verification

Appendix A Abbreviations and Acronyms xi

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List of Figures Figure Page

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

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

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

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

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

List of Tables Table Page

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

Table 6-3. Trapped Electron Spectrum

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

Table 11-1. Test Tolerances

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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.

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

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)

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

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

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

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.

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

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.

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

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.

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

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.

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

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

Substrate Ground

Substrate Ground

Substrate Ground

Substrate Ground

Substrate Ground

Substrate Ground

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

SA

D

A In te rface

Monitor Solar Cell

Twisted, Shielded Trio

Thermistor Thermistor

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

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.

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

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.

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

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

Notes

1. 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

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

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

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

Notes

1. 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

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

Figure 4-4. Panel 1 to SADA Electrical Interface

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

Notes

1. 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

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

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.

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

Figure 4-5. Panel 1 Front and Side View

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

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

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

Figure 4-7. Panel 2 Front and Side View

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

Figure 4-8. Panel 2 Back View (KOZ = “Keep-Out Zone”)

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

Figure 4-9 Panel 3 Front and Side View

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

Figure 4-10. Panel 3 Back View (KOZ = “Keep-Out Zone”)

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

Figure 4-11. Qualification Coupon Front and Side View

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

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.

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

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.

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

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.

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

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.

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

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..

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

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..

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

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

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

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

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

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

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

8 0.000E+00

Table 6-4. Trapped Proton Spectrum

(>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

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

Table 6-5. Solar Proton Spectrum

(>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

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

1.8 6.145E+10 50 1.183E+09

15 6.790E

+10

20 5.530E

+10

30 4.442E

+10

50 3.347E

+10

60 3.016E

+10

80 2.533E

+10

100 2.153E

+10

150 1.447E

+10

200 9.695E

+09

300 4.056E

+09

400 1.637E

+09

700 1.912E

+08

1200 1.101E

+07

2000 2.000E

+00

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

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.

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

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.).

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

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.

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

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.

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

8.0…