RST-EPS-SPEC-0173-_Draft2020.8.24.pdf

Effective Date: TBD

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

RST-EPS-SPEC-0173, Revision -

Roman Space Telescope (RST), Code 448

RST Solar Array Panels Specification

DRAFT

Roman Space Telescope Reviewed – Not Subject to Export Control

RST Solar Array Spec RST-EPS-SPEC-0173, Revision - Draft ii CHECK https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm

TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.

RST Solar Array Specification

Review/Signature/Approval Page

Prepared by:

David Kim

Approved by:

Electronic Approval available on-line at the RST CM Tool https://ipdtdms.gsfc.nasa.gov/frontmenu_dsp.cfm iii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Preface This document is a Roman Space Telescope (RST) Configuration Management (CM)-controlled document.

Note: Prior to May 20, 2020, the project name was Wide Field Infrared Survey Telescope

(WFIRST).

For the purposes of configuration management, the prefixes “WFIRST” and “RST” are completely interchangeable. For example, RST-MGMT-PROC-0024 is the same as WFIRST-

MGMT-PROC-0024.

Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the RST CM Office (CMO), along with supportive material justifying the proposed change.

In this document, a requirement is identified by “shall,” a good practice by “should,” permission by “may” or “can,” expectation by “will,” and descriptive material by “is.”

Questions or comments concerning this document should be addressed to:

RST Configuration Management Office Mail Stop 448 Goddard Space Flight Center Greenbelt, Maryland 20771 iv Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Change History Log

Revision Effective Date Description of Changes (Reference the CCR & CCB/ERB Approval Date)

Revision - TBD Initial Release per CCR-0275 v Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Table of TBDs/TBRs/TBSs [optional]

Item No. Location Summary Individual/ Organization

Actionee

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

Table of Contents

1 INTRODUCTION

1.1 General Information

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents and Drawings

2 CONTRACT DESCRIPTION

2.1 Solar Array Description

2.2 Qualification Coupons Description

2.3 Insulated Substrates Description

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 Solar Array Functional/Performance Requirements

3.1.1 Test Condition Power

3.1.1.1 Flight Panels

3.1.1.2 Qualification Coupons

3.1.2 Power at Highest Predicted Operating Temperature

3.1.3 End of Life Power

3.1.4 PWM Segment Current Relative to Digital Segment Current

3.1.5 Solar Cell and Bypass Diode

3.1.5.1 Solar Cell Mechanical

3.1.5.2 Solar Cell Layout

3.1.5.3 Limit to Solar Cell Shadowing

3.1.6 Solar Cell Cover

3.1.6.1 Cover Material and Thickness

3.1.6.2 Cover Orientation

3.1.6.3 Anti-Reflective Coating (AR)

3.1.7 Optical Solar Reflectors (OSR), Terminal Boards, and Blocking Diodes

3.1.8 Wire and Wire Layout

3.1.8.1 Coarse Sun Sensor (CSS) Wiring

3.1.8.2 Monitor Solar Cell

3.1.9 Connector Wiring and Connector Type

3.1.10 Platinum Resistor Thermometers (PRT)

3.1.11 Insulated Substrate

3.1.11.1 Substrate Insulation Resistance

3.1.11.2 Substrate Grounding

3.1.12 Panel Performance in Thermal Vacuum Environment

3.1.13 Panel Performance in Depressurization Environment

3.1.14 Allowable Degradation Due to Charged Particle Radiation

3.1.15 Allowable Degradation Due to Humidity

3.2 Resource Allocations

3.2.1 Mass Allocation

3.3 Power

3.3.1 Solar Array Panel Power Wire Redundancy

3.4 Electrical Grounding

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

3.4.1 Primary Power DC Isolation

3.4.2 Mechanical Contact Resistance

3.4.3 Connector DC Resistance

3.4.4 Passive Analog Telemetry

3.4.4.1 Platinum Resistor Thermometer (PRT)

3.4.4.2 PRT Performance

4 PHYSICAL REQUIREMENTS

4.1 Interface Documentation

4.2 Physical Envelope

5 ENVIRONMENTAL REQUIREMENTS

5.1 Acoustic Deflection

5.2 Pressure

5.2.1 Operating Pressure Range

5.3 Ground Environments

5.4 Thermal

5.4.1 Flight Interface Design Temperature Limits

5.5 Charged Particle Radiation Requirements

5.5.1 Definitions

5.5.2 Total Ionizing Dose

5.5.2.1 Minimum TID Tolerance for EEE Parts and Materials

5.5.3 Charging Environment

6 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 Particulate Contamination

6.1.1.2 Molecular Contamination

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

6.1.2.2 Molecular Generation

6.1.2.2.1 Material Selection

6.1.2.2.2 Assembly Outgassing

6.2 Electrostatic Cleanliness

6.2.1 Conductive Surface Ground Path

6.2.2 Conductive Surface Resistivity

6.3 Magnetic Cleanliness

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

7.1.2 Materials

7.1.2.1 Material Conductivity

7.1.2.2 Material Limitations for Debris Casualty Area .... Error! Bookmark not defined.

7.2 Electrical

7.2.1 Test Sensors

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

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

7.2.2.1.1 Connector Specifications

7.2.2.1.2 Contact Derating

7.2.2.1.3 Redundant Contact Derating

7.2.2.2 Signal Segregation

7.2.2.3 Test and Flight Signal Isolation

7.2.2.4 Test Interfaces

7.2.2.4.1 Facility-Induced Noise

7.2.2.4.2 Facility-Induced ESD GSE Malfunction

7.2.2.4.3 Facility-Induced GSE Malfunction

7.2.3 Mitigation of Internal Charging

7.2.3.1 Floating Conductors

7.2.3.2 Dielectric Structures

7.2.3.2.1 Bulk Resistivity

7.2.3.2.2 Charge Bleed-Off

7.3 Safety

7.4 Identification and Marking

7.5 Workmanship

7.5.1 Workmanship Standards

7.5.2 Connector

7.5.2.1 GSE Cable Connectors

7.5.2.2 Prevention of Connector Mismating

7.5.2.2.1 Connector Uniqueness

7.5.2.2.2 Accessibility

7.5.2.2.3 Connector Gender

7.5.2.3 Connector Identification

7.5.2.4 Protection of Unused Test Connectors

7.5.2.5 Connector Savers

7.6 Reliability and Mission Lifetime

7.6.1 Mission Life

7.7 Ground Handling

7.7.1 Ground Support Equipment (GSE) Design

7.7.2 GSE Cleanliness

7.7.3 GSE Bake-out

7.7.4 Test Harness

8 MECHANICAL DESIGN REQUIREMENTS

9 LOGISTICS

9.1 Ground Support Equipment (GSE)

9.2 Transportation Equipment

10 VERIFICATION REQUIREMENTS

10.1 Verification Methods

10.1.1 Inspection

10.1.2 Analysis

10.1.3 Test

10.2 Inspection Requirements

10.2.1 Visual Inspection

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

10.2.2 Physical Measurement

10.2.3 Documentation Search

10.3 Analysis Requirements

10.4 Test Requirements

10.4.1 Definitions

10.4.2 Test Tolerances

10.4.3 Test Restrictions

10.4.3.1 Failure During Tests

10.4.3.2 Modification of Hardware

10.4.3.3 External Adjustment

10.4.3.4 Re-Test Requirements

10.5 Required Tests

10.5.1 Mass Properties Measurement

10.5.2 Thermal Vacuum Bake-out

10.5.2.1 Test Conditions for Thermal Vacuum Bake-out

10.5.2.2 Calculation of the Configuration Adjusted QCM Rate

10.5.2.3 Pre-Test Chamber Vacuum Bake-out and Chamber Certification

10.5.2.4 Post Bake-out Cleanliness

10.5.2.4.1 Witness Foils

10.5.2.5 Surface Cleanliness Required for Performance

10.5.3 Thermal Vacuum Test

10.5.3.1 Thermal Vacuum Test Parameters

10.5.4 Solar Cell and Bypass Diode Qualification Tests

10.5.5 Solar Array Panel Qualification Tests

10.5.5.1 Solar Array Panel Life Cycle Coupon Tests

10.5.6 Flight Solar Array Panel Tests

10.5.7 Power Verification

10.5.7.1 Test Condition Power

10.5.7.2 End of Life Power

10.5.8 Bypass Diode Functionality Verification

10.5.9 Substrate Insulation Resistance Verification

10.5.10 Solar Cell Mechanical Verification

10.5.11 Coverglass Orientation Verification

10.5.12 Flight Connector Type Verification

10.5.13 PRT Type Verification

10.5.14 PRT Performance Verification

10.5.15 Parts and Assembly Layout Verification

10.5.16 Mission Life Verification

10.5.17 Shelf Life Verification

10.5.18 Substrate Ground Verification

10.5.19 Cleanliness Verification

APPENDIX A ABBREVIATIONS AND ACRONYMS

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

List of Figures Figure 2-1. Solar Array Panels in Stowed and Deployed Positions Figure 3-1. Suggested OSR locations Figure 3-2. Rear facesheet wiring block diagram

List of Tables

Table 5-1. Solar Array Panel Temperature Environment Table 5-2. Total Ionizing Dose in Silicon Table 5-3. Surface Incident Integral Solar Proton Fluences Table 5-4. Displacement Damage Dose in Gallium Arsenide Table 7-1. Limited Materials for Debris Casualty Area ............... Error! Bookmark not defined.

Table 10-1. Test Tolerances

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

1 INTRODUCTION

1.1 General Information

The Roman Space Telescope (RST) is a NASA observatory designed to settle essential questions in the areas of dark energy, exoplanets, and infrared astrophysics. The telescope has a primary mirror that is 2.4 meters in diameter and is the same size as the Hubble Space Telescope’s primary mirror. RST will have two instruments, the Wide Field Instrument and the Coronagraph Instrument.

1.2 Scope

This specification describes the electrical, mechanical, environmental, and verification testing requirements for space qualified solar array panels for the NASA Goddard Space Flight Center (GSFC) RST Mission.

1.3 Related Documentation

The latest versions of all documents below should be used. RST documents can be obtained from the RST CM Tool.

1.3.1 Applicable Documents and Drawings

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 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 RST Solar Array Statement of Work (RST-EPS-SOW-0055), in which case the Statement of Work takes precedence.

Document Number Title

AIAA S-111A-2014

AIAA S-111A-2005

Qualification and Quality Requirements for Space Solar Cells

AIAA S-112A-2013 Qualification and Quality Requirements for Electrical Components on Space Solar Panels

AMS 2488 Anodic Treatment – Titanium and Titanium Alloys Solution pH 13 or Higher

ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

DOD-HDBK-83575 General Handbook for Space Vehicle Wiring Harness Design and Testing

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

FAA AC 20-71 Federal Aviation Administration Advisory Circular (AC) 20- 71, “Dual Locking Devices on Fasteners”

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.

IEST-STD-CC1246E Product Cleanliness Levels and Contamination Control Program

MIL-A-8625F Anodic Coatings for Aluminum and Aluminum Alloys MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and Aluminum

Alloys

MIL-DTL-27500

(NEMA WC27500)

Cable, Power, Electrical, and Cable Special Purpose, Electrical Shielded and Unshielded, General Specification

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NASA-HDBK-7005 Dynamic Environment Criteria NASA-STD-5001B Structural Design and Test Factors of Safety for Spaceflight

Hardware NASA-STD-5017A Design and Development Requirements for Mechanisms NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware NASA-STD-5020 Requirements for Threaded Fastening Systems in Spaceflight

Hardware NASA-STD-6016 Standard Materials and Processes Requirements for Spacecraft NASA-STD-7001 Payload Vibroacoustic Test Criteria NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General

Practices for NFPA 70 National Fire Protection Association National Electric Code RST-EPS-LIST-0114 RST Solar Array Panels Deliverable Items List and Schedule RST-EPS-SOW-0055 RST Solar Array Panels Statement of Work WFIRST-SYS-SPEC-0033A RST Cleanliness Specification

Drawing Number Title 2231111 Panel 1 EPS ICD 2231112 Panel 2 EPS ICD 2231113 Panel 3 EPS ICD 2231114 Panel 4 EPS ICD 2231115 Panel 5 EPS ICD 2231116 Panel 6 EPS ICD

2231016 SUBSTRATE ASSEMBLY, TEST COUPON SASS

2231022 SOLAR PANEL QUAL COUPON, INSERT INSTALLATION SASS

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

2 CONTRACT DESCRIPTION

2.1 Solar Array Description

Two flight solar array panels (Panels 3-4) are part of a body-mounted solar array and four flight solar array panels (Panels 1-2, 5-6) are part of a deployed solar array for the RST Spacecraft.

Once deployed all solar array panels will be coplanar with each other. The RST solar array will convert solar energy to electrical power and also act as a sunshield 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.

Figure 2-1. Solar Array Panels in Stowed and Deployed Positions

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

2.2 Qualification Coupons Description

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 RST Solar Array panel design.

2.3 Insulated Substrates Description

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.

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the Solar Array described in Section 2.1.

3.1 Solar Array Functional/Performance Requirements

3.1.1 Test Condition Power

3.1.1.1 Flight Panels

Under simulated Air Mass Zero (AM0) illumination at 28°C, normal incidence, the sum of all the solar cell string power output, taken at the panel terminals, at Beginning of Life (BOL) shall be no less than 10100 watts at 53 volts or greater.

Two solar cell string circuits will be connected to the Spacecraft Solar Array Module (SAM) and will be referred to as the PWM segments. Twenty-seven solar cell string circuits will be connected to the Spacecraft Segment Module (SM) and will be referred to as the Digital segments. One solar cell circuit will be connected to the SM and will be referred to as the Unswitched segment.

3.1.1.2 Qualification Coupons

a) The qualification coupon(s) shall, to the extent practicable due to size limitations, have at least one example of the type of circuit used on the flight panels.

b) The qualification coupon output from each of these circuits shall be proportional to the panel power requirement in Section 3.1.1.1, scaled based on circuit size.

3.1.2 Power at Highest Predicted Operating Temperature

The Contractor shall extrapolate the current-voltage curves of the flight panels and qualification coupon(s) for 0.948 AM0, 136°C at BOL. The sum of all the solar cell circuit power output, taken at the panel terminals, shall be no less than 6600 watts at 36.5 volts or greater.

3.1.3 End of Life Power

The Contractor shall predict the combined power output from of all six flight panels under 0.948 AM0, 60°C, with 35 degrees roll and 15 degrees pitch solar incidence angles, and exposure to the space environment at Lagrangian point L2 for an interval of 5 years and 90 days. This End of Life (EOL) power output, taken at the panel terminals, shall be no less than 6100 watts at 45 volts or greater. Additionally, the circuit voltage at maximum power (Vmp) shall be no less than

34.5 volts when measured at the panel terminals at 136°C EOL.

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

3.1.4 PWM Segment Current Relative to Digital Segment Current The current produced by a PWM segment at any possible operating voltage, temperature, or intensity shall be at least twice the current produced by any Digital segment operating at the same voltage, temperature, and intensity.

3.1.5 Solar Cell and Bypass Diode

3.1.5.1 Solar Cell Mechanical

No cell on a panel shall have a crack visible at 3x or less magnification.

3.1.5.2 Solar Cell Layout

a) The Contractor shall propose the solar cell layout and comply with stay out zones, configuration, and other requirements specified in the substrate drawings listed in Section 1.3.1.

b) The solar cell layout of the six solar array panels shall be divided into 30 parallel circuits;

2 PWM segments, 27 Digital segments, and 1 Unswitched segment.

c) Each PWM segment and, to the extent possible, each Digital segment shall have an equal number of solar cell strings in parallel and sized in proportion relative to each other according to Section 3.1.4. The Unswitched segment shall have four solar cell strings in parallel.

d) At BOL, 1.05 AM0, 136°C, the short-circuit current for any Digital segment shall be no greater than 8 amperes and for any PWM segment shall be no greater than 16 amperes.

e) Panels 3 and 4 shall contain one PWM segment each located in the upper portion of each panel.

f) To the extent practicable, the Contractor shall not divide individual segments between panels and will group strings in each segment next to each other. Except for the Unswitched segment, where two strings shall be placed on Panels 3 and 4 each, and located around the Optical Solar Reflectors (OSR) described in Section 3.1.7.

3.1.5.3 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.

3.1.6 Solar Cell Cover

3.1.6.1 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.

3.1.6.2 Cover Orientation

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

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

3.1.6.3 Anti-Reflective Coating (AR)

The Contractor shall use covers that are coated with AR.

3.1.7 Optical Solar Reflectors (OSR), Terminal Boards, and Blocking Diodes

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

GSFC.

c) Each terminal board shall contain blocking diodes, which connect in series with the solar cell strings and redundantly paralleled as necessary to limit the short-circuit current (Isc) supported by each diode to be no greater than 0.65 amperes at BOL, 1.05 AM0, 136°C.

d) The Contractor shall conformally coat the terminal boards covering at minimum all exposed metal surfaces prior to delivery to GSFC.

e) The terminal boards shall be grouped on each panel as close to each other and near a panel edge and/or cell-less area as possible to minimize heat from surroundings.

f) The Contractor shall bond OSRs to the front facesheet of the panels directly opposite the terminal boards and as centered over the boards as possible.

g) The area of the OSRs on each panel shall be sufficient to maintain a blocking diode junction temperature of no greater than 125°C when carrying the maximum predicted current at BOL, 1.05 AM0, 136°C and with a surrounding substrate front facesheet temperature of 136°C. The back facesheet will be covered in Multi-Layer Insulation

(MLI).

h) The Contractor shall provide details of the terminal board thermal design so GSFC can provide and/or verify the OSR area necessary to not exceed the required temperature (initial estimate of 0.167 m2 per panel).

i) Each OSR shall be 100 µm (nominal) thick and have an absorptivity and emittance of no greater than 0.06 and no less than 0.83 (nominal) respectively.

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

Figure 3-1. Suggested OSR locations

3.1.8 Wire and Wire Layout

Rear facesheet wiring block diagram is shown in Figure 3-2.

a) The Contractor shall provide at least 688 and 297 cm pigtails for the power and signal harness from the clamp points on Panels 3 and 4 respectively. Power and signal, primary and redundant wires shall be split into separate bundles. These pigtails shall be secured near the outboard edge using Government Furnished Equipment (GFE) clamps. These will be terminated and mated to the spacecraft at GSFC in accordance with the pin assignments to be defined by the contractor.

b) The Contractor shall provide at least 60 cm flying leads for the power and signal harness at the inter-panel output connector locations provided by GSFC.

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

c) Any power wire that carries more than 1 ampere at BOL, 1.05 AM0, 136°C, shall be spaced, to the extent possible, at least 2 mm or the width of one AWG #20 wire from each other to spread local heating due to power dissipation in the wires.

d) For added insulation in critical areas, the Contractor shall place Kapton insulation underneath the power and return wires after the solar cell strings are paralleled as described in Section 3.1.9.

e) The Contractor shall protect wire wherever abrasion may be a problem.

f) The Contractor shall use stress relief between wire tie points to avoid strains, particularly on the solar cell string terminations.

g) The Contractor shall address how it will stake wire.

Figure 3-2. Rear facesheet wiring block diagram

3.1.8.1 Coarse Sun Sensor (CSS) Wiring

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

a) The Contractor shall route wiring for the CSSs to the panel pigtail harness described in Section 3.1.8 on the rear facesheet of the panels using GFE twisted shielded wire.

b) The Contractor shall ground the wire shield to a panel ground point.

c) The Contractor shall terminate the CSS wiring with at least a 60 cm pigtail at the CSS locations on top of Panel 3.

3.1.8.2 Monitor Solar Cell

The Contractor shall place a monitor cell on Panels 3 and 4 each. The Contractor shall wire the cells to the panel connector using twisted shielded wire.

3.1.9 Connector Wiring and Connector Type

The Contractor shall provide a pigtail with or without Ground Support Equipment (GSE) connectors for the power and signal inter-panel connectors.

a) GSFC will use Sommer DW-747 connectors for the power and signal inter-panel connectors. Power and signal wires shall be split between separate primary and secondary connectors. The Contractor shall define the connector pin assignments.

b) Test connectors shall be Glenair Series 79 Micro-Crimp or other GSFC approved connector, which the Contractor must fix to each panel to test each individual solar cell string as specified in Paragraph j) of this section.

c) For power, the Contractor shall parallel the strings at a terminal board or boards for each segment.

d) The Contractor shall use JAN TXV 1N5811 blocking diodes for each solar cell string on the terminal board or boards. Parallel-redundant blocking diodes shall be used for each solar cell string as necessary to support a current no greater than defined in Section 3.1.7.

e) The cathodes (output) of the diodes shall be paralleled and the circuit returns shall be paralleled to form each of the segments.

f) The Contractor shall use wire meeting MIL-W-22759/44 or MIL-W-22759/33. Shielded wires shall meet NEMA-WC27500.

g) From each PWM segment to the panel power connectors, the Contractor shall provide eight AWG #20 pair wiring; that is eight positive and eight return wires.

h) From each Digital segment to the panel power connectors, the Contractor shall provide four AWG #20 pair wiring; that is four positive and four return wires.

i) From the Unswitched segment to the panel power connectors, the Contractor shall provide four AWG #20 pair wiring; that is four positive and four return wires. Two of the four pair shall be on Panels 3 and 4 each as described in Section 3.1.5.2.

j) From the anode (input) of the blocking diodes, for each solar cell string to the panel test connectors, the Contractor shall provide one AWG #20 or #24 wire dependent on the test connector used. From the return side of each solar cell circuit to the panel test connectors, the Contractor shall provide one AWG #20 or #24 wire.

k) From each hinge damper heater to the panel power connectors, the Contractor shall provide two AWG #20 pair wiring; that is two positive and two return wires.

l) From each potentiometer to the panel signal connectors, the Contractor shall provide two AWG #24 pair wiring; that is two positive and two return wires.

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

m) From each PRT and CSS to the panel signal connectors, the Contractor shall provide one AWG #24 twisted shielded pair wiring; that is one positive and one return wire twisted and shielded.

n) From each thermostat to the panel signal connectors, the Contractor shall provide two AWG #24 twisted shielded pair wiring; that is two positive and two return wires twisted and shielded.

o) From each monitor solar cell to the panel signal connector, the contractor shall provide one AWG #24 twisted shielded pair wiring; that is one positive and one return wire twisted and shielded.

3.1.10 Platinum Resistor Thermometers (PRT)

a) The Contractor shall use Goodrich Platinum Resistor Thermometer type

0118MM2000AFAFBC.

b) The Contractor shall mount four PRTs on each panel; two on the back of the front facesheet and two on the rear facesheet next to the terminal boards.

c) The Contractor shall propose the location of the PRTs, except that on Panel 3 and 4, one PRT on the back of the front facesheet shall be directly under each monitor solar cell.

d) The Contractor shall run the PRT lead wires along the back facesheet of each panel to the signal connector or pigtail using AWG #24 twisted shielded wire.

e) The Contractor shall ground the wire shield to the connector shell or panel ground point.

3.1.11 Insulated Substrate

GSFC will provide insulated substrates for each flight solar array panel and each qualification coupon. The substrates will be made of composite facesheet with 3.1 and 6.1 pcf perforated aluminum honeycomb core. The front facesheet (solar cell side) will also include co-cured Kapton insulation.

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.

3.1.11.1 Substrate Insulation Resistance

The resistance between the substrate and solar cell circuits shall be greater than 10 megohms for the flight panels and qualification coupon(s).

3.1.11.2 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.

d) The substrate ground leads from Panels 1, 3, 5, and Panels 2, 4, 6 shall be combined into two leads for the Panel 3 and 4 pigtails respectively.

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

3.1.12 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 Section 5.4.1.

3.1.13 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-5 Torr in thirty seconds. Verified by analysis or shown by heritage.

3.1.14 Allowable Degradation Due to Charged Particle Radiation The Contractor shall consider hard particle radiation in its computation of EOL power based on Section 5.5.

3.1.15 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.

3.2 Resource Allocations

3.2.1 Mass Allocation

The total add-on mass (total assembled panel mass minus the substrate mass) of the six Solar Array panels shall be no greater than 82 kg, not including the pigtail lengths from Panels 3 and 4 described in Section 3.1.8.

3.3 Power

3.3.1 Solar Array Panel Power Wire Redundancy

The Solar Array panels shall provide redundant contacts or connections for the circuit output power and return lines.

3.4 Electrical Grounding

3.4.1 Primary Power DC Isolation

The Solar Array panel power and returns shall be isolated from signal grounds by a DC resistance of no less than 1 megohm.

3.4.2 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the component) shall be no greater than 2.5 milliohm.

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3.4.3 Connector DC Resistance

Component connectors shall be electrically connected to chassis with a DC resistance no greater than 10 milliohm.

3.4.4 Passive Analog Telemetry

3.4.4.1 Platinum Resistor Thermometer (PRT)

The Contractor shall use PRT type specified in Section 3.1.10.

3.4.4.2 PRT Performance

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

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4 PHYSICAL REQUIREMENTS

4.1 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 tolerances) is permitted.

The Electrical Interface is defined in this Specification.

The Mechanical Interface is defined by the Panel EPS ICDs listed in Section 1.3.1.

4.2 Physical Envelope

No Solar Array panel shall exceed the thermal and mechanical volume envelope described in the Panel EPS ICDs listed in Section 1.3.1.

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

5 ENVIRONMENTAL REQUIREMENTS

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

The Solar Array panels shall meet its performance requirements in Section 3 during and after exposure to the environments specified in this section.

5.1 Acoustic Deflection

The solar cells, wiring, and other hardware provided by the Contractor shall withstand deflections with a minimum radius of curvature of 30.64 meters. Verified by analysis or shown by heritage.

5.2 Pressure

5.2.1 Operating Pressure Range

The Solar Array panels shall be designed to meet all performance requirements while operating over a pressure range of 1.08E5 Pa (813 Torr) to 1.3E-12 Pa (1E-14 Torr).

5.3 Ground Environments

The Solar Array panels shall meet all of their performance requirements during exposure to air temperature between 5°C and 30°C and relative humidity between 30% and 70%.

5.4 Thermal

5.4.1 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 cycles listed below.

Table 5-1. Solar Array Panel Temperature Environment

Number of Cycles Temperature Limits Operational: 3000 50°C to 136°C

LEOP: 1 -25°C to 136°C MCC: 1 -80°C to 136°C

5.5 Charged Particle Radiation Requirements

Components containing electronic parts will be exposed to a natural Space radiation environment that consists of trapped particles which include electrons, protons, and heavier ions; particles from solar events (flares and coronal mass ejections); and galactic cosmic rays.

For solar cell degradation, the Contractor shall use the surface incident solar proton spectra and/or displacement damage dose in Table 5-3 and Table 5-4 to determine the solar cell loss factors due to radiation.

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5.5.1 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 silicon.

5.5.2 Total Ionizing Dose

5.5.2.1 Minimum TID Tolerance for EEE Parts and Materials

The top-level total ionizing dose requirement is shown in Table 5-2. 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 5.25-year mission life, the expected dose is 22.3 krad-Si at a 95% confidence level. EEE parts and materials shall be selected according to the level of shielding shown in Table 5-2.

Table 5-2. Total Ionizing Dose in Silicon

Aluminum Shield Thickness Dose (krad-Si)

(mm): (mils): (g/cm2): C = 95%

0.026 1.024 7.02E-03 1.03E+03

0.033 1.299 8.91E-03 8.46E+02

0.041 1.614 1.11E-02 7.24E+02

0.052 2.047 1.40E-02 6.06E+02

0.063 2.48 1.70E-02 5.25E+02

0.082 3.228 2.21E-02 4.43E+02

0.1 3.937 2.70E-02 3.75E+02

0.126 4.961 3.40E-02 3.13E+02

0.159 6.26 4.29E-02 2.63E+02

0.2 7.874 5.40E-02 2.24E+02

0.256 10.079 6.91E-02 1.82E+02

0.319 12.559 8.61E-02 1.53E+02

0.404 15.906 1.09E-01 1.26E+02

0.508 20 1.37E-01 1.02E+02

0.637 25.079 1.72E-01 8.58E+01

0.804 31.654 2.17E-01 6.83E+01

1.012 39.843 2.73E-01 5.47E+01

1.275 50.197 3.44E-01 4.51E+01

1.604 63.15 4.33E-01 3.57E+01

2.019 79.488 5.45E-01 2.84E+01

2.542 100.079 6.86E-01 2.23E+01

3.198 125.906 8.63E-01 1.75E+01

4.024 158.425 1.09E+00 1.34E+01

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5.069 199.567 1.37E+00 1.03E+01

6.38 251.181 1.72E+00 7.75E+00

8.033 316.26 2.17E+00 5.85E+00

10.112 398.11 2.73E+00 4.35E+00

12.731 501.22 3.44E+00 3.16E+00

16.025 630.906 4.33E+00 2.31E+00

20.175 794.291 5.45E+00 1.65E+00

25.399 999.961 6.86E+00 1.18E+00

Table 5-3. Surface Incident Integral Solar Proton Fluences

Energy, E (MeV): Fluence > E (p/cm2) C = 95%

0.1 2.49E+12

0.3 1.49E+12

0.5 1.16E+12

0.7 9.79E+11 1 8.12E+11 3 4.04E+11 5 2.79E+11 7 2.12E+11 10 1.53E+11 15 9.94E+10 20 7.03E+10 25 5.22E+10 30 4.02E+10 35 3.17E+10 40 2.56E+10 45 2.10E+10 50 1.74E+10 55 1.46E+10 60 1.24E+10 70 9.18E+09 80 6.96E+09 90 5.40E+09

100 4.27E+09 125 2.57E+09 150 1.67E+09 175 1.14E+09 200 8.16E+08 225 6.02E+08 250 4.55E+08 275 3.51E+08 300 2.76E+08 400 1.25E+08 500 6.73E+07

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Table 5-4. Displacement Damage Dose in Gallium Arsenide

Aluminum Shield Thickness DDD(GaAs) (MeV/g) C = 95% (mm): (mils): (g/cm2):

0.026 1.024 7.017E-03 9.28E+10

0.033 1.299 8.906E-03 6.37E+10

0.041 1.614 1.107E-02 4.96E+10

0.052 2.047 1.403E-02 3.54E+10

0.063 2.480 1.700E-02 2.75E+10

0.082 3.228 2.213E-02 2.05E+10

0.100 3.937 2.699E-02 1.72E+10

0.126 4.961 3.401E-02 1.26E+10

0.159 6.260 4.291E-02 9.36E+09

0.200 7.874 5.398E-02 7.96E+09

0.256 10.079 6.909E-02 5.84E+09

0.319 12.559 8.609E-02 4.57E+09

0.404 15.906 1.090E-01 3.58E+09

0.508 20.000 1.371E-01 2.76E+09

0.637 25.079 1.719E-01 2.21E+09

0.804 31.654 2.170E-01 1.70E+09

1.012 39.843 2.731E-01 1.25E+09

1.275 50.197 3.441E-01 9.98E+08

1.604 63.150 4.329E-01 8.06E+08

2.019 79.488 5.449E-01 6.07E+08

2.542 100.079 6.861E-01 4.81E+08

3.198 125.906 8.631E-01 3.71E+08

4.024 158.425 1.086E+00 2.84E+08

5.069 199.567 1.368E+00 2.21E+08

6.380 251.181 1.722E+00 1.66E+08

8.033 316.260 2.168E+00 1.29E+08

10.112 398.110 2.729E+00 9.89E+07

12.731 501.220 3.436E+00 7.31E+07

16.025 630.906 4.325E+00 5.55E+07

20.175 794.291 5.445E+00 4.10E+07

25.399 999.961 6.855E+00 3.07E+07

5.5.3 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 RST mission orbit.

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6 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.

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 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.

6.1.1.2 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.

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

The 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 GSFC:

• 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 the material is highly susceptible to tear propagation (e.g.

multi-layer insulation (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, crack, or flake when subjected to normal handling (e.g. indium tin oxide, other rigid or brittle semiconductors, or ceramic coating on flexible substrates such as Teflon, MLI, etc.)

• Foams or highly textured materials

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• Trapped debris in holes

6.1.2.2 Molecular Generation

6.1.2.2.1 Material Selection

The Solar Array materials shall have a total mass loss (TML) of less than 1.00% and a collected volatile condensable mass (CVCM) of less than 0.10%, when measured in accordance with ASTM E-595 unless a materials usage agreement has been generated and approved by GSFC.

6.1.2.2.2 Assembly Outgassing

a) The Solar Array outgassing shall be measured with a Quartz Crystal Microbalance

(QCM).

b) The measurement shall be made in a chamber that has been certified clean (background outgassing rate) and has been modeled by a GSFC Contamination Analyst to account for mass sinks (e.g. cold fingers, pumps, cold surfaces, etc.) that could influence the source outgassing rate.

c) Outgassing shall not exceed 1.06E-5 g/s that is condensable on a QCM operated at - 20°C when the Solar Array panels are held at 125°C.

d) A cold finger and/or scavenger plate shall be used in tests for components that will be mounted externally unless approved otherwise by GSFC.

6.2 Electrostatic Cleanliness

The following Sections 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.

6.2.1 Conductive Surface Ground Path

All Solar Array external conductive surfaces except for the solar cell coverglasses shall be connected to the Spacecraft interface with a resistance of 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 connections to the nearest grounded Spacecraft surface.

6.2.2 Conductive Surface Resistivity

All Solar Array external conductive surfaces shall have a resistivity of less than 109 ohms/square.

6.3 Magnetic Cleanliness

The Contactor shall use best practices when designing the solar cell layout and wiring to minimize magnetic fields.

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The Contractor shall calculate the dipole moment of each Solar Array panel using the highest predicted BOL short-circuit current for each string.

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

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 EEE Parts

See requirements in the SOW.

7.1.2 Materials

See requirements in the SOW.

7.1.2.1 Material Conductivity

All parts should be passivated and mounting surfaces on Solar Array panels shall be conductive as defined in Section 3.4.

7.2 Electrical

7.2.1 Test Sensors

With the exception of thermocouples used in thermal ambient, 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.

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

7.2.2.1.1 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 explicit approval by the NASA/GSFC COR.

7.2.2.1.2 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.

7.2.2.1.3 Redundant Contact Derating

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

7.2.2.2 Signal Segregation

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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.

7.2.2.3 Test and Flight Signal Isolation

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

7.2.2.4 Test Interfaces

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

7.2.2.4.1 Facility-Induced Noise

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

7.2.2.4.2 Facility-Induced ESD GSE Malfunction

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

7.2.2.4.3 Facility-Induced GSE Malfunction

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

7.2.3 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.

7.2.3.1 Floating Conductors

Floating conductors, if present, shall have a bleed path of less than 10 megohm to the component structure. This requirement is not applicable to small floating conductors ≤ 1 inch2 (≤ 6.45cm2) or short ≤ 1 inch (≤ 2.54cm) unterminated traces or wires that are inside of the components.

7.2.3.2 Dielectric Structures

7.2.3.2.1 Bulk Resistivity

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

7.2.3.2.2 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.

7.3 Safety

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

a) All GSE being used at GSFC shall meet NASA requirements specified in Section 7.7.

b) All GSE being used at the Launch site shall meet the applicable sections of NASA-STD-

8719.24 ELV Safety Requirements.

7.4 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 approved by GSFC.

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

7.5 Workmanship

7.5.1 Workmanship Standards

See the workmanship standards and processes outlined in the SOW.

7.5.2 Connector

7.5.2.1 GSE Cable Connectors

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

7.5.2.2 Prevention of Connector Mismating

7.5.2.2.1 Connector Uniqueness

Physically adjacent connectors shall be of different sizes, genders, or uniquely keyed to facilitate proper mating where possible.

7.5.2.2.2 Accessibility

The Solar Array Spacecraft interface connectors shall be spaced far enough apart to allow the mate and demate operations to be performed without a special tool.

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7.5.2.2.3 Connector Gender

The connector half that sources power to another Solar Array panel shall be socketed (female) to protect against inadvertent grounding prior to mating.

7.5.2.3 Connector Identification

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

7.5.2.4 Protection of Unused Test Connectors

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

7.5.2.5 Connector Savers

Connector savers shall be used during integration and test to minimize wear on flight connector contacts and must meet the same requirements as the flight connectors.

7.6 Reliability and Mission Lifetime

7.6.1 Mission Life

The Solar Array shall meet all performance specifications through 2 years of ground testing and

5.25 years of operation in space.

7.7 Ground Handling

7.7.1 Ground Support Equipment (GSE) Design

All electrical GSE or support equipment shall be in compliance with the National Electric Code (NFPA 70) or equivalent standard.

7.7.2 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.)

7.7.3 GSE Bake-out

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

7.7.4 Test Harness

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

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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.

8 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.

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9 LOGISTICS

9.1 Ground Support Equipment (GSE)

a) Any GSE that will be used in a thermal vacuum test (harness, fixtures, stimulators, etc.)

shall be fabricated of vacuum compatible materials that meet the requirements of Section 6.1.

b) Any GSE that will be used in a thermal vacuum test shall be capable of being baked out at 146°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.